Weight & Balance

Purpose- Aircraft weight and balance is safety and secondary to achieve efficiency. Improper loading reduces efficiency of aircraft maneuverability and increase fuel consumption.Aircraft have tendency to gain weight because of accumulation of dirt grease etc.

Theory- Influence of weight is directly dependent upon instance from the fulcrum. To Balance the lever weight must be distributed so that turning effect is same on one side of fulcrum as on other.The distance of any object from the fulcrum is called Lever Arm.Lever arm multiplied by weight of object is its turning effect about fulcrum is known as Moment.An aircraft is balanced if it remains level when suspended from an imaginary point. This point is the location of ideal C.G.

Weight and balance data can be obtained from 1. Aircraft specification 2. Flight manual 3. operating limitation 4. Weight and balance report

Datum –An imaginary vertical plane from which all horizontal measurements are taken for balance purposes with aircraft in level flight attitude. It is a plane at right angle to longitudinal axis of aircraft. In most cases it is located on nose of aircraft on same point on aircraft structure itself.

Arm-is horizontal distance that an item of equipment is located from datum.Arm distance is always given or measured in inches and except for location which might be exactly on datum, it is proceeded by plus or minus. The plus sign indicate a distance aft of datum and minus sign indicates distance forward of datum.

Moment- it is a product of weight multiplied arm. The plus and minus sign depend whether moment is added or removed.

Center of gravity – is a point about which nose heavy and tail heavy moment are exactly equal in magnitude.

Max. weight- is maximum authorized weight of an aircraft.

Empty weight- include all operating equipment that has fixed location and actually installed in aircraft. It includes weight of airframe, power plant, required equipment, optional or special equipment, fixed ballast, hydraulic fluid and residual fuel and oil.

Residual fuel and oil are fluids that will not be normally drain out because they are trapped in fuel , oil line or tanks. they must be included in aircraft empty weight.

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Useful load- is determined by subtracting empty weight from max. allowable gross weight. It consists of max. oil,fuel ,passenger,baggage,pilot,co-pilot and other crew member.Determinig the distribution of this weight is called Weight check.

Empty C.G.- it is a C.G. of aircraft in empty weight condition.EWCG range- it is allowable variation of travel within C.G. limits. When EWCG of aircraft falls within this range.

Operating C.G. range- is a distance between forward and rearward C,G, limit indicated in Type certificate data sheet.These limit are shown either in % of MAC or inches from datum of aircraft. The loaded aircraft CG location must remain within this limit in all condition.Calculation of CG location in% of MAC- 1. find diff. between distance to empty weight CG location from datum and distance to leading edge of MAC from datum.2) divide by length of MAC and multiply by 100Weighing point- Point on scale at which weight is concentrated.Zero fuel weight- is max. weight of loaded aircraft without fuel.Minimum fuel- is amount of fuel must shown weight & balance report when aircraft is loaded for an extreme condition check.Full oil- when weighing an aircraft , oil tank must either contain no.of gallons of oil specified or drained. When an aircraft with full oil is weighed, weight of oil must be subtracted from recorded reading to arrive at actual reading.Tare weight- include weight of all extra items, such as jack, block & chocks on weighing scale platform, except the item being weighed.Prepare aircraft for weighing-

1) drain fuel system completely only residual fuel is left in tank, considered part of empty weight. In special cases, aircraft must be weighed with full fuel tanks, provided a mean of determining exact amount of fuel is available.

2) Drain all engine oil and entrapped oil will be considered as residual oil. Hydraullic reservoir and system should be filled , drinking and water reservoir and lavatory tank should be drained, and CSD oil tank should be filled.

Some aircraft are not weighed with wheel on scales, but weighed with scales placed either at jacking point on special weighing point. When weighing an aircraft with wheel placed on scales release brake to reduce incorrect reading caused by side load on scales.

Empty weight is determined by adding net weight on each weighing point. The net weight is actual reading less tare weight.

Fuels and Fuel System

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Fuel may be classified as 1)Solid 2) Liquid 3) GaseousLiquid fuel are classified as non volatile or volatile. The nonvolatile fuel are heavy oil used in diesel engine. Volatile fuel used in fuel metering devices that easily vaporizes.Volatility- it is a measure of the tendency of a liquid substance to vaporize under gjven conditions.Reid vapor pressure test- measures tendency of gasoline to vaporize. A sample of fuel is stored in bomb and it is submerged in a bath at constant temperature and indicated pressure is noted. The higher the pressure more tendency of gasoline to vaporize.

Jet fuels are composed of hydro carbons with a little more carbon and usually a higher sulphur content then gasoline. Inhibitors may be added to reduce corrosion and oxidation. Anti icing additives also blended to prevent fuel icing.Two type of jet fuel- 1)Kerosene grade turbine fuel, named JET A 2)A blend of gasoline and kerosene fractions designated JET B. There is third type JET A-1, which is made for operation at extremely low temperatures.

Jet A was developed as heavy kerosene having higher flash point and lower frezzing point than most of kerosene. It has very low vapor pressure. It contains more heat nergy per gallon then does Jet BJet B is similer to Jet A. It is a blend of kerosene and gasoline.these fuels are not interchangeable.A high volatile fuel is required for starting in cold weather and to make arieal restart easier. Low volatile is desired to reduce the possibility of vapor lock and to reduce fuel losses by evaporation.Jet fuels range in color from colorless liquid to straw colored (amber) liquid, depending on age or the crude petroleum sources.Fuel contamination- The higher the viscosity of fuel the greater is its ability to hold contaminants. The contaminants are-Water, Rust or Scale & Dirt.

Water- can be present in fuel in two forms-1)dissolved in fuel 2) suspended in fuel. Suspended water can be detected by naked eyes. The fineally droplets reflect light and in high concentrations give the fuel a dull, hazy or cloudy appearance. Particles of suspended water may unite to form droplets of free water. If cloud disappear at bottom, air is present. If the cloud disappear at top, water is present. Free water can cause icing of the aircraft fuel system.

Foreign particles- most common types are rust, sand, aluminium or magnesium products, brass shavings and rubber. Rust is found in two forms 1) Red rust, which is non magnetic 2) Black rust, which is magnetic

MettalaurgyAGE HARDENING

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The term as applied to soft or low carbon steels, relates to slow, gradual changes that take place in properties of steels after the final treatment. These changes, which bring about a condition of increased hardness, elastic limit, and tensile strength with a consequent loss in ductility, occur during the period in which the steel is at normal temperatures.

AGINGSpontaneous change in the physical properties of some metals, which occurs on standing, at atmospheric temperatures after final cold working or after a final heat treatment. Frequently synonymous with the term " Age-Hardening

ALCLADThe common name for a type of clad wrought aluminum products, such as sheet and wire, with coatings of high-purity aluminum or an aluminum alloy different from the core alloy in composition. The coatings are anodic to the core so they protect exposed areas on the core electrolytic ally during exposure to corrosive environments

ALLOY STEELSteel containing substantial quantities of elements other than carbon and the commonly-accepted limited amounts of manganese, sulfur, silicon, and phosphorous. Addition of such alloying elements is usually for the purpose of increased hardness, strength or chemical resistance. The metals most commonly used for forming alloy steels are: nickel, chromium, silicon, manganese, tungsten, molybdenum and vanadium. "Low Alloy" steels are usually considered to be those containing a total of less than 5% of such added constituents.

AGINGANNEALING

A heating and cooling operation implying usually a relatively slow cooling. Annealing is a comprehensive term. The process of such a heat treatment may be: to remove stresses; to induce softness; to alter ductility; toughness; electrical magnetic, or other physical properties; to refine the crystalline structure; to remove gases; to produce a definite micro-structure. In annealing, the temperature of the operation and the rate of cooling depend upon the material being heat treated and the purpose of the treatment.

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ANODIZING (Aluminum Anodic Oxide Coating)A process of coating aluminum by anodic treatment resulting in a thin film of aluminum oxide of extreme hardness. A Wide variety of dye colored coatings are possible by impregnation in process

ARTIFICIAL AGINGAn aging treatment above room temperature. (See Precipitation Heat Treatment and compare with natural aging

Objectives of Heat Treatments

Heat Treatment is the controlled heating and cooling of metals to alter their physical and mechanical properties without changing the product shape. Heat treatment is sometimes done inadvertently due to manufacturing processes that either heat or cool the metal such as welding or forming.

Heat Treatment is often associated with increasing the strength of material, but it can also be used to alter certain manufacturability objectives such as improve machining, improve formability, restore ductility after a cold working operation. Thus it is a very enabling manufacturing process that can not only help other manufacturing process, but can also improve product performance by increasing strength or other desirable characteristics.

Steels are particularly suitable for heat treatment, since they respond well to heat treatment and the commercial use of steels exceeds that of any other material. Steels are heat treated for one of the following reasons:

1.

Softening

2.

Hardening

3.

Material Modification

Common Heat Treatments

Softening: Softening is done to reduce strength or hardness, remove

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residual stresses, improve toughnesss, restore ductility, refine grain size or change the electromagnetic properties of the steel. Restoring ductility or removing residual stresses is a necessary operation when a large amount of cold working is to be performed, such as in a cold-rolling operation or wiredrawing. Annealing — full Process, spheroidizing, normalizing and tempering — austempering, martempering are the principal ways by which steel is softened.

Hardening: Hardening of steels is done to increase the strength and wear properties. One of the pre-requisites for hardening is sufficient carbon and alloy content. If there is sufficient Carbon content then the steel can be directly hardened. Otherwise the surface of the part has to be Carbon enriched using some diffusion treatment hardening techniques.

Full Annealing

Full annealing is the process of slowly raising the temperature about 50 ºC (90 ºF) above the Austenitic temperature line A3 or line ACM in the case of Hypoeutectoid steels (steels with < 0.77% Carbon) and 50 ºC (90 ºF) into the Austenite-Cementite region in the case of Hypereutectoid steels (steels with > 0.77% Carbon).

It is held at this temperature for sufficient time for all the material to transform into Austenite or Austenite-Cementite as the case may be. It is then slowly cooled at the rate of about 20 ºC/hr (36 ºF/hr) in a furnace to about 50 ºC (90 ºF) into the Ferrite-Cementite range. At this point, it can be cooled in room temperature air with natural convection.

The grain structure has coarse Pearlite with ferrite or Cementite (depending on whether hypo or hyper eutectoid). The steel becomes soft and ductile.

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Normalizing

Normalizing is the process of raising the temperature to over 60 º C (108 ºF), above line A3 or line ACM fully into the Austenite range. It is held at this temperature to fully convert the structure into Austenite, and then removed form the furnace and cooled at room temperature under natural convection. This results in a grain structure of fine Pearlite with excess of Ferrite or Cementite. The resulting material is soft; the degree of softness depends on the actual ambient conditions of cooling. This process is considerably cheaper than full annealing since there is not the added cost of controlled furnace cooling.

The main difference between full annealing and normalizing is that fully annealed parts are uniform in softness (and machinablilty) throughout the entire part; since the entire part is exposed to the controlled furnace cooling. In the case of the normalized part, depending on the part geometry, the cooling is non-uniform resulting in non-uniform material properties across the part. This may not be desirable if further machining is desired, since it makes the machining job somewhat unpredictable. In such a case it is better to do full annealing. Process Annealing

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Process Annealing is used to treat work-hardened parts made out of low-Carbon steels (< 0.25% Carbon). This allows the parts to be soft enough to undergo further cold working without fracturing. Process annealing is done by raising the temperature to just below the Ferrite-Austenite region, line A1on the diagram. This temperature is about 727 ºC (1341 ºF) so heating it to about 700 ºC (1292 ºF) should suffice. This is held long enough to allow recrystallization of the ferrite phase, and then cooled in still air. Since the material stays in the same phase through out the process, the only change that occurs is the size, shape and distribution of the grain structure. This process is cheaper than either full annealing or normalizing since the material is not heated to a very high temperature or cooled in a furnace.

Stress Relief Annealing

Stress Relief Anneal is used to reduce residual stresses in large castings, welded parts and cold-formed parts. Such parts tend to have stresses due to thermal cycling or work hardening. Parts are heated to temperatures of up to 600 - 650 ºC (1112 - 1202 ºF), and held for an extended time (about 1 hour or more) and then slowly cooled in still air. Introduction

Tempering is a process done subsequent to quench hardening. Quench-hardened parts are often too brittle. This brittleness is caused by a predominance of Martensite. This brittleness is removed by tempering. Tempering results in a desired combination of hardness, ductility, toughness, strength, and structural stability. Tempering is not to be confused with tempers on rolled stock-these tempers are an indication of the degree of cold work performed.

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The mechanism of tempering depends on the steel and the tempering temperature. The prevalent Martensite is a somewhat unstable structure. When heated, the Carbon atoms diffuse from Martensite to form a carbide precipitate and the concurrent formation of Ferrite and Cementite, which is the stable form. Even though a little strength is sacrificed, toughness (as measured by impact strength) is increased substantially. Springs and such parts need to be much tougher — these are tempered to a much lower hardness.

Tempering is done immediately after quench hardening. When the steel cools to about 40 ºC (104 ºF) after quenching, it is ready to be tempered. The part is reheated to a temperature of 150 to 400 ºC (302 to 752 ºF). In this region a softer and tougher structure Troostite is formed. Alternatively, the steel can be heated to a temperature of 400 to 700 ºC (752 to 1292 ºF) that results in a softer structure known as Sorbite. This has less strength than Troostite but more ductility and toughness.

The heating for tempering is best done by immersing the parts in oil, for tempering upto 350 ºC (662 ºF) and then heating the oil with the parts to the appropriate temperature. Heating in a bath also ensures that the entire part has the same temperature and will undergo the same tempering. For temperatures above 350 ºC (662 ºF) it is best to use a bath of nitrate salts. The salt baths can be heated upto 625 ºC (1157 ºF). Regardless of the bath, gradual heating is important to avoid cracking the steel. After reaching the desired temperature, the parts are held at that temperature for about 2 hours, then removed from the bath and cooled in still air.

Austempering

Austempering is a quenching technique. The part is not quenched through the Martensite transformation. Instead the material is quenched above the temperature when Martensite forms MS, around 315 ºC (600 ºF). It is held till at this temperature till the entire part reaches this temperature. As the part is held longer at this temperature, the Austenite transforms into Bainite. Bainite is

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tough enough so that further tempering is not necessary, and the tendency to crack is severely reduced.

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Martempering

Martempering is similar to Austempering except that the part is slowly cooled through the martensite transformation. The structure is martensite, which needs to tempered just as much as martensite

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that is formed through rapid quenching. The biggest advantage of Austempering over rapid quenching is that there is less distortion and tendency to crack. Hardness is a function of the Carbon content of the steel. Hardening of a steel requires a change in structure from the body-centered cubic structure found at room temperature to the face-centered cubic structure found in the Austenitic region. The steel is heated to Autenitic region. When suddenly quenched, the Martensite is formed. This is a very strong and brittle structure. When slowly quenched it would form Austenite and Pearlite which is a partly hard and partly soft structure. When the cooling rate is extremely slow then it would be mostly Pearlite which is extremely softHardenability, which is a measure of the depth of full hardness achieved, is related to the type and amount of alloying elements. Different alloys, which have the same amount of Carbon content, will achieve the same amount of maximum hardness; however, the depth of full hardness will vary with the different alloys. The reason to alloy steels is not to increase their strength, but increase their hardenability — the ease with which full hardness can be achieved throughout the material.

Usually when hot steel is quenched, most of the cooling happens at the surface, as does the hardening. This propagates into the depth of the material. Alloying helps in the hardening and by determining the right alloy one can achieve the desired properties for the particular application.

Such alloying also helps in reducing the need for a rapid quench cooling — thereby eliminate distortions and potential cracking. In addition, thick sections can be hardened fully.

Quench Media

Quenching is the act of rapidly cooling the hot steel to harden the steel.

Water: Quenching can be done by plunging the hot steel in water. The water adjacent to the hot steel vaporizes, and there is no direct contact of the water with the steel. This slows down cooling until the bubbles break and allow water contact with the hot steel. As the water contacts and boils, a great amount of heat is removed from the steel. With good agitation, bubbles can be prevented from sticking to the steel, and thereby prevent soft spots.

Water is a good rapid quenching medium, provided good agitation is done. However, water is corrosive with steel, and the rapid cooling

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can sometimes cause distortion or cracking.

Salt Water: Salt water is a more rapid quench medium than plain water because the bubbles are broken easily and allow for rapid cooling of the part. However, salt water is even more corrosive than plain water, and hence must be rinsed off immediately.

Brine

A faster quench than water is brine. In some steels that have a low hardenability it may be necessary to go to a brine quench. Brine solution is made by adding salt, sodium chloride, to water. The effect of brine on the quench is to make the water more efficient by precipitating on the steel and then blowing off very rapidly creating rapid agitation and disrupting the vapor jacket.

Oil: Oil is used when a slower cooling rate is desired. Since oil has a very high boiling point, the transition from start of Martensite formation to the finish is slow and this reduces the likelihood of cracking. Oil quenching results in fumes, spills, and sometimes a fire hazard.

Transformations on Cooling

Annealing, normalizing, sphereodizing

The structure and hardness of the steel is established by the rate of cooling from the austenitic condition. If brought down slowly the steel will be annealed and soft. The structure will be mostly ferrite and cementite, carbides. This can be done in a temperature controlled furnace by dropping the temperature through a known rate over a set period of time dependent on the type of steel. Another method is to preheat a heavy bar of low carbon to the same temperature as critical for the steel and bury both of them together in vermiculite. The vermiculite, obtained in bags from garden supply, is made from chipped mica and is an excellent insulating material. It will slow the cooling rate down so that the blade will still be hot to the touch the next day. For most of the carbon steels this will be enough to anneal the piece.

If allowed to air cool it will be normalized, a tougher condition comprised of fine pearlite and carbides. Blades can be ground and prepared for heat treatment in either normalized or annealed states. Another treatment that is particularly effective for workability and for dimensional stability is called sphereodizing. With the steel in a normalized condition you reheat, usually in salt to inhibit oxidization, to a temperature just below lower critical, 1300F and hold for at least an hour. What occurs is that the carbides will begin

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to aglomulate or pool into larger more evenly spaced particles in a ferrite matrix. It makes handfinishing much easier.

Quenchants

The method of controlling the speed of cooling is the quenchant.  The quench rate is determined by how quickly the quenchant can remove the heat from the steel. When a piece of hot steel enters the quenchant the area surrounding the blade absorbs heat from the blade until it is heated itself.

Heat treatment of precipitation hardening alloysAluminium alloys are strengthened in a number of ways including: solid solution hardening, cold working, dispersion hardening and precipitation hardening.

Precipitation hardening (otherwise known as age hardening) is a process whereby a fine precipitate structure is formed in the alloy matrix following a heat treatment process.

The precipitation hardening process follows three main steps:

1. Solution treatment. The alloy is heated above the solvus temperature to dissolve any precipitates and ensure the alloying elements are in solid solution.

2. Quench, The alloy is quenched. The alloying elements in solution do not have time to diffuse and form precipitates. Thus, the alloying elements remain in solution forming what is known as a supersaturated solid solution.

3. Ageing. The alloy is heated to an intermediate temperature below the solvus temperature. The alloying elements are able to diffuse to form coherent precipitate clusters (known as GP zones).

Example age hardening 2XXX series aluminium alloy system

The coherent precipitates increase the strength of the alloy by distorting the crystal lattice and creating resistance to dislocation motion. The number of precipitates increases with increasing time

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thus increasing the strength of the alloy. However, with excessive time the precipitates become large and incoherent and their strengthening effect decreases. Thus, during precipitation hardening there are four main stages:

1. solid solution strengthening in the supersaturated solid solution

2. coherency stress hardening from the coherent precipitates 3. precipitation hardening by resistance to dislocation cutting 4. hardening through resistance to dislocation between

precipitates.

Steel Alloys

Steel Alloys can be divided into five groups

        Carbon Steels

        High Strength Low Alloy Steels

        Quenched and Tempered Steels

        Heat Treatable Low Alloy Steels

        Chromium-Molybdenum Steels

Steels are readily available in various product forms.   The American Iron and Steel Institute defines carbon steel as follows:

Steel is considered to be carbon steel when no minimum content is specified or required for chromium, cobalt, columbium [niobium], molybdenum, nickel, titanium, tungsten, vanadium or zirconium, or any other element to be added to obtain a desired alloying effect; when the specified minimum for copper does not exceed 0.40 per cent; or when the maximum content specified for any of the following elements does not exceed the percentages noted: manganese 1.65, silicon 0.60, copper 0.60.  Carbon steels are normally classified as shown below.

Low-carbon steels contain up to 0.30 weight percent C. The largest category of this class of steel is flat-rolled products (sheet or strip) usually in the cold-rolled and annealed condition.

Medium-carbon steels are similar to low-carbon steels except that the carbon ranges from 0.30 to 0.60 weight percent and the manganese from 0.60 to 1.65 weight percent. Increasing the carbon content to approximately 0.5 weight percent with an accompanying increase in manganese allows medium-carbon steels to be used in the quenched and tempered condition.

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High-carbon steels contain from 0.60 to 1.00 weight percent C with manganese contents ranging from 0.30 to 0.90weight percent.

Types, Characteristics, and Uses of Alloyed Steels

While the plain carbon type of steel remains the principal product of the steel mills, so-called alloy or special steels are being turned out in ever increasing tonnage. Let us now consider those alloyed steels and their uses in aircraft.

CARBON STEELS. -Steel containing carbon in percentages ranging from 0.10 to 0.30 percent are classed as low-carbon steel. The equivalent SAE numbers range from 1010 to 1030. Steels of this grade are used for making such items as safety wire, certain nuts, cable bushings, and threaded rod ends. Low-carbon steel in sheet form is used for secondary structural parts and clamps, and in tubular form for moderately stressed structural parts.

Steels containing carbon in percentages ranging from 0.30 to 0.50 percent are classed as medium-carbon steel. This steel is especially adaptable for machining or forging and where surface hardness is desirable. Certain rod ends and light forgings are made from SAE 1035 steel.

Steel containing carbon in percentages ranging from 0.50 to 1.05 percent are classed as high-carbon steel. The addition of other elements in varying quantities adds to the hardness of this steel. In the fully heat-treated condition, it is very hard and will withstand high shear and wear and have little deformation. It has limited use in aircraft. SAE 1095 in sheet form is used for making flat springs, and in wire form for making coil springs.

NICKEL STEELS. -The various nickel steels are produced by combining nickel with carbon steel. Steels containing from 3 to 3.75 percent nickel are commonly used. Nickel increases the hardness, tensile strength, and elastic limit of steel without appreciably decreasing the ductility. It also intensifies the hardening effect of heat treatment. SAE 2330 steel is used extensively for aircraft parts such as bolts, terminals, keys, clevises, and pins.

CHROMIUM STEELS. -Chromium steels are high in hardness, strength, and corrosion-resistant properties. SAE 51335 is particularly adaptable for heat-treated forgings that require greater toughness and strength than may be obtained in plain carbon steel. It is used for such articles as the balls and rollers of antifriction bearings.

CHROMIUM-NICKEL OR STAINLESS STEELS. anticorrosive degree is determined by the surface condition of the metal as well as by the

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composition, temperature, and concentration of the corrosive agent. The principal part of stainless steel is chromium, to which nickel may or may not be added. The corrosion-resisting steel most often used in aircraft construction is known as 18-8 steel because of its content of 18 percent chromium and 8 percent nickel. One of the distinctive features of 18-8 steel is that its strength maybe increased by cold-working.

Stainless steel may be rolled, drawn, bent, or formed to any shape. Because these steels expand about 50 percent more than mild steel and conduct heat only about 40 percent as rapidly, they are more difficult to weld. Stainless steel, with but a slight variation in its chemical composition, can be used for almost any part of an aircraft. Some of its more common applications are in the fabrication of exhaust collectors, stacks and manifolds, structural and machined parts, springs, castings, and tie rods and cables.

CHROME-VANADIUM STEELS. -These are made of approximately 0.18 percent vanadium and about 1.00 percent chromium. When heat treated, they have strength, toughness, and resistance to wear and fatigue. A special grade of this steel in sheet form can be cold-formed into intricate shapes. It can be folded and flattened without signs of breaking or failure. SAE 6150 is used for making springs; and chrome-vanadium with high-carbon content, SAE 6195, is used for ball and roller bearings.

CHROME-MOLYBDENUM STEELS. -Molyb-denum in small percentages is used in combination with chromium to form chrome- molybdenum steel, which has various uses in aircraft. Molybdenum is a strong alloying element, only 0.15 to 0.25 percent being used in the chrome-molybdenum steels; the chromium content varies from 0.80 to 1.10 percent. Molybdenum raises the ultimate strength of steel without affecting ductility or workability. Molybdenum steels are tough, wear resistant, and harden throughout from heat treatment. They are especially adaptable for welding, and for this reason are used principally for welded structural parts and assemblies. SAE 4130 is used for parts such as engine mounts, nuts, bolts, gear structures, support brackets for accessories, and other structural parts.

Effects of Elements on Steel

Carbon has a major effect on steel properties.  Carbon is the primary hardening element in steel.  Hardness and tensile strength increases as carbon content increases up to about 0.85% C as shown in the figure above.  Ductility and weldability decrease with increasing carbon. 

Manganese is generally beneficial to surface quality especially in resulfurized steels. Manganese contributes to strength and

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hardness, but less than carbon.  The increase in strength is dependent upon the carbon content.  Increasing the manganese content decreases ductility and weldability, but less than carbon. Manganese has a significant effect on the hardenability of steel.

Phosphorus  increases strength and hardness and decreases ductility and notch impact toughness of steel. Phosphorous levels are normally controlled to low levels.  Higher phosphorus is specified in low-carbon free-machining steels to improve machinability.

Sulfur decreases ductility and notch impact toughness especially in the transverse direction.  Weldability decreases with increasing sulfur content.The only exception is free-machining steels, where sulfur is added to improve machinability.

Silicon  is one of the principal deoxidizers used in steelmaking.  Silicon is less effective than manganese in increasing as-rolled strength and hardness.

Copper  in significant amounts is detrimental to hot-working steels.  Copper negatively affects forge welding, but does not seriously affect arc or oxyacetylene welding.Copper is beneficial to atmospheric corrosion resistance when present in amounts exceeding 0.20%.

Lead is virtually insoluble in liquid or solid steel.  However, lead is sometimes added to carbon and alloy steels by means of mechanical dispersion during pouring to improve the mach inability.

Boron is added to fully killed steel to improve harden ability. Boron-treated steels are produced to a range of 0.0005 to 0.003%.A very small amount of boron (about 0.001%) has a strong effect on hardenability.Boron is most effective in lower carbon steels.   

Chromium  is commonly added to steel to increase corrosion resistance and oxidation resistance, to increase hardenability, or to improve high-temperature strength.  As a hardening element, Chromium is frequently used with a toughening element such as nickel to produce superior mechanical properties. At higher temperatures, chromium contributes increased strength.

Nickel is a ferrite strengthener.  Nickel does not form carbides in steel.  It remains in solution in ferrite,  strengthening and toughening the ferrite phase.  Nickel increases the hardenability and impact strength of steels. 

Molybdenum increases the hardenability of steel.  Molybdenum may produce secondary hardening during the tempering of quenched

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steels. It enhances the creep strength of low-alloy steels at elevated temperatures. 

Aluminum  is widely used as a deoxidizer.  Aluminum can control austenite grain growth in reheated steels and is therefore added to control grain size. Titanium, zirconium, and vanadium are also valuable grain growth inhibitors, but there carbides are difficult to dissolve into solution in austenite.

Titanium is used to retard grain growth and thus improve toughness. Titanium is also used to achieve improvements in inclusion characteristics.  Titanium causes sulfide inclusions to be globular rather than elongated thus improving toughness and ductility in transverse bending.

Vanadium increases the yield strength and the tensile strength of carbon steel. The addition of small amounts of Niobium can significantly increase the strength of steels.  Vanadium is one of the primary contributors to precipitation strengthening in microalloyed steels.   When thermomechanical processing is properly controlled the ferrite grain size is refined and there is a corresponding increase in toughness.  The impact transition temperature also increases when vanadium is added.

All microalloy steels contain small concentrations of one or more strong carbide and nitride forming elements.  Vanadium, niobium, and titanium combine preferentially with carbon and/or nitrogen to form a fine dispersion of precipitated particles in the steel matrix.

FERROUS AIRCRAFT METALS.

Table 1-1.-SAE Numerical Index

Type of steel Classification

Carbon 

Nickel 

Nickel-chromium 

Molybdenum 

Chromium 

Chromium-vanadium 

1xxx

2xxx

3xxx

4xxx

5xxx

6xxx

7xxx

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Tungsten 

Silicon-manganese 

9xxx

 

Among the common materials used are ferrous metals. The term ferrous applies to the group of metals having iron as their principal constituent.

Identification

If carbon is added to iron, in percentages ranging up to approximately 1.00 percent, the product will be vastly superior to iron alone and is classified as carbon steel. Carbon steel forms the base of those alloy steels produced by combining carbon with other elements known to improve the properties of steel. A base metal (such as iron) to which small quantities of other metals have been added is called an alloy. The addition of other metals is to change or improve the chemical or physical properties of the base metal.

SAE NUMERICAL INDEX. -The steel classification of the Society of Automotive Engineers (SAE) is used in specifications for all high-grade steels used in automotive and aircraft construction. A numerical index system identifies the composition of SAE steels.

Each SAE number consists of a group of digits, the first of which represents the type of steel; the second, the percentage of the principal alloying element; and usually the last two or three digits, the percentage, in hundredths of 1 percent, of carbon in the alloy. For example, the SAE number 4150 indicates a molybdenum steel containing 1 percent molybdenum and 50 hundredths of 1 percent of carbon.

Stainless Steels

Stainless Steels are iron-base alloys containing Chromium.  Stainless steels usually contain less than 30% Cr and more than 50% Fe. They attain their stainless characteristics because of the formation of an invisible and adherent chromium-rich oxide surface film. This oxide establishes on the surface and heals itself in the presence of oxygen.  Some other alloying elements added to enhance specific characteristics include nickel, molybdenum, copper, titanium, aluminum, silicon, niobium, and nitrogen.  Carbon is usually present in amounts ranging from less than 0.03% to over 1.0% in certain martensitic grades.  Corrosion resistance and mechanical properties

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are commonly the principal factors in selecting a grade of stainless steel for a given application.

Selecting a Stainless Steel

There are a large number of stainless steels produced.  Corrosion resistance, physical properties, and mechanical properties are generally among the properties considered when selecting stainless steel for an application.  A more detailed list of selection criteria is listed below:

←   Corrosion resistance ←   Resistance to oxidation

and sulfidation ←   Toughness ←   Cryogenic strength ←   Resistance to abrasion

and erosion ←   Resistance to galling and

seizing ←   Surface finish ←   Magnetic properties

←   Retention of cutting edge

←   Ambient strength ←   Ductility ←   Elevated temperature

strength ←   Suitability for intended

cleaning procedures ←   Stability of properties in

service ←   Thermal conductivity ←   Electrical resistivity

←   Suitability for intended fabrication  techniques

Aluminum Alloys

TYPES, CHARACTERISTICS, AND USES. -Aluminum is one of the most widely used metals in modern aircraft construction. It is vital to the aviation industry because of its high strength/weight ratio, its corrosion-resisting qualities, and its comparative ease of fabrication. The outstanding characteristic of aluminum is its light weight. In color, aluminum resembles silver, although it possesses a characteristic bluish tinge of its own. Commercially pure aluminum melts at the comparatively low temperature of 1,216°F. It is nonmagnetic, and is an excellent conductor of electricity.

Commercially pure aluminum has a tensile strength of about 13,000 psi, but by rolling or other cold-working processes, its strength may be approximately doubled. By alloying with other metals, together with the use of heat-treating processes, the tensile strength may be raised to as high as 96,000 psi, or to well within the strength range of structural steel.

Aluminum alloy material, although strong, is easily worked, for it is very malleable and ductile. It may be rolled into sheets as thin as 0.0017 inch or drawn into wire 0.004 inch in diameter. Most aluminum alloy sheet stock used in aircraft construction ranges from

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0.016 to 0.096 inch in thickness; however, some of the larger aircraft use sheet stock that may be as thick as 0.0356 inch.

One disadvantage of aluminum alloy is the difficulty of making reliable soldered joints. Oxidation of the surface of the heated metal prevents soft solder from adhering to the material; therefore, to produce good joints of aluminum alloy, a riveting process is used. Some aluminum alloys are also successfully welded.

The various types of aluminum maybe divided into two classes-casing alloys (those suitable for casting in sand, permanent mold, and die castings) and the wrought alloys (those that may be shaped by rolling, drawing, or forging). Of the two, the wrought alloys are the most widely used in aircraft construction, being used for stringers, bulkheads, skin, rivets, and extruded sections. Casting alloys are not extensively used in aircraft.

WROUGHT ALLOYS. -Wrought alloys are divided into two classes-nonheat treatable and heat treatable. In the nonheat-treatable class, strain hardening (cold-working) is the only means of increasing the tensile strength. Heat-treatable alloys may be hardened by heat treatment, by cold-working, or by the application of both processes.

Aluminum products are identified by a universally used designation system. Under this arrangement, wrought aluminum and wrought aluminum alloys are designated by a four-digit index system.

The first digit of the designation indicates the major alloying element or alloy group, as shown in table 1-2. The lxxx indicates aluminum of 99.00 percent or greater; 2xxx indicates an aluminum alloy in which copper is the major alloying element; 3xxx indicates an aluminum alloy with manganese as the major alloying element; etc. Although most aluminum alloys contain several alloying elements, only one group (6xxx) designates more than one alloying element.

In the 1xxx group, the second digit in the designation indicates modifications in impurity limits. If the second digit is zero, it indicates that there is no special control on individual impurities. The last two of the four digits indicate the minimum aluminum percentage. Thus, alloy 1030 indicates 99.30 percent aluminum without special control on impurities. Alloys 1130, 1230, 1330, etc., indicate the same aluminum purity with special control on one or more impurities. Likewise, 1075, 1175, 1275, etc., indicate 99.75 percent aluminum.

Aluminum Alloys can be divided into nine groups.

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Designation Major Alloying Element

1xxxUnalloyed (pure) >99% Al

2xxxCopper is the principal alloying element, though other elements (Magnesium) may be specified

3xxxManganese is the principal alloying element

4xxx Silicon is the principal alloying element

5xxx Magnesium is the principal alloying element

6xxx Magnesium and Silicon are principal alloying elements

7xxxZinc is the principal alloying element, but other elements such as Copper, Magnesium, Chromium, and Zirconium may be specified

8xxxOther elements  (including Tin and some Lithium compositions)

9xxx Reserved for future use

1xxx Series.  These grades of aluminum are characterized by excellent corrosion resistance, high thermal and electrical conductivities, low mechanical properties, and excellent workability. Moderate increases in strength may be obtained by strain hardening. Iron and silicon are the major impurities.

2xxx Series.  These alloys require solution heat treatment to obtain optimum properties; in the solution heat-treated condition, mechanical properties are similar to, and sometimes exceed, those of low-carbon steel. In some instances, precipitation heat treatment (aging) is employed to further increase mechanical properties. This treatment increases yield strength, with attendant loss in elongation; its effect on tensile strength is not as great.The alloys in the 2xxx series do not have as good corrosion resistance as most other aluminum alloys, and under certain conditions they may be subject to intergranular corrosion.

3xxx Series.  These alloys generally are non-heat treatable but have about 20% more strength than 1xxx series alloys. Because only a limited percentage of manganese (up to about 1.5%) can be effectively added to aluminum, manganese is used as major element in only a few alloys.

4xxx Series. The major alloying element in 4xxx series alloys is silicon, which can be added in sufficient quantities (up to 12%) to cause substantial lowering of the melting range.  For this reason,

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aluminum-silicon alloys are used in welding wire and as brazing alloys for joining aluminum, where a lower melting range than that of the base metal is required.

5xxx Series. The major alloying element is Magnesium an when it is used as a major alloying element or with manganese, the result is a moderate-to-high-strength work-hardenable alloy.Alloys in this series possess good welding characteristics and relatively good resistance to corrosion in marine atmospheres. However, limitations should be placed on the amount of cold work and the operating temperatures (150 degrees F) permissible for the higher-magnesium alloys to avoid susceptibility to stress-corrosion cracking.

6xxx Series. Alloys in the 6xxx series contain silicon and magnesium approximately in the proportions required for formation of magnesium silicide (Mg2Si), thus making them heat treatable. Although not as strong as most 2xxx and 7xxx alloys, 6xxx series alloys have good formability, weldability, machinability, and relatively good corrosion resistance, with medium strength. Alloys in this heat-treatable group may be formed in the T4 temper (solution heat treated but not precipitation heat treated) and strengthened after forming to full T6 properties by precipitation heat treatment.

7xxx Series. Zinc, in amounts of 1 to 8% is the major alloying element in 7xxx series alloys, and when coupled with a smaller percentage of magnesium results in heat-treatable alloys of moderate to very high strength. Usually other elements, such as copper and chromium, are also added in small quantities. 7xxx series alloys are used in airframe structures, mobile equipment, and other highly stressed parts.  Higher strength 7xxx alloys exhibit reduced resistance to stress corrosion cracking and are often utilized in a slightly overaged temper to provide better combinations of strength, corrosion resistance, and fracture toughness.

 The temper designation follows the alloy designation and shows the actual condition of the metal. It is always separated from the alloy designation by a dash.

The letter F following the alloy designation indicates the "as fabricated condition, in which no effort has been made to control the mechanical properties of the metal,

The letter O indicates dead soft, or annealed, condition.

The letter W indicates solution heat treated. Solution heat treatment consists of heating the metal to a high temperature followed by a rapid quench in cold water,

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This in an unstable temper, applicable only to those alloys that spontaneously age at room temperature, Alloy 7075 may be ordered in the W condition.

The letter H indicates strain hardened, cold-worked, hand-drawn, or rolled. Additional digits are added to the H to indicate the degree of strain hardening. Alloys in this group cannot be strengthened by heat treatment, hence the term nonheat-treatable.

The letter T indicates fully heat treated. Digits are added to the T to indicate certain variations in treatment.

W -solution heat treated , unstable temper

T2- Annealed cast product only

T3-Solution heat treated and then cold worked

T4- solution heat treated

T5- Artifficialy aged only

T6- solution heat treated and then Artifficialy aged

T7- Solution heat treated and then stabilized

T8- Soluton heat treated, cold worked and then artificially aged

T9- Soluton heat treated, artificially aged, and then cold worked

T10 Artificially aged and then cold worked

Greater strength is obtainable in the heat-treatable alloys. They are often used in aircraft in preference to the nonheat-treatable alloys. Heat-treatable alloys commonly used in aircraft construction (in order of increasing strength) are 6061, 6062, 6063, 2017, 2024, 2014,7075, and 7178.

Alloys 6061, 6062, and 6063 are sometimes used for oxygen and hydraulic lines and in some applications as extrusions and sheet metal.

Alloy 2017 is used for rivets, stressed-skin covering, and other structural members. Alloy 2024 is used for airfoil covering and fittings. It may be used wherever 2017 is specified, since it is stronger.

Alloy 2014 is used for extruded shapes and forgings. This alloy is similar to 2017 and 2024 in that it contains a high percentage of

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copper. It is used where more strength is required than that obtainable from 2017 or 2024.

Alloy 7178 is used where highest strength is necessary, Alloy 7178 contains a small amount of chromium as a stabilizing agent, as does alloy 7075.

Nonheat-treatable alloys used in aircraft construction are 1100, 3003, and 5052. These alloys do not respond to any heat treatment other than a softening, annealing effect. They may be hardened only by cold- working.

Alloy 1100 is used where strength is not an important factor, but where weight, economy, and corrosion resistance are desirable. This alloy is used for fuel tanks, fairings, oil tanks, and for the repair of wing tips and tanks.

Alloy 3003 is similar to 1100 and is generally used for the same purposes. It contains a small percentage of manganese and is stronger and harder than 1100, but retains enough work ability that it is usually preferred over 1100 in most applications.

Alloy 5052 is used for fuel lines, hydraulic lines, fuel tanks, and wing tips. Substantially higher strength without too much sacrifice of workability can be obtained in 5052. It is preferred over 1100 and 3003 in many applications. 

Alclad is the name given to standard aluminum alloys that have been coated on both sides with a thin layer of pure aluminum. Alclad has very good corrosion-resisting qualities and is used exclusively for exterior surfaces of aircraft. Alclad sheets are available in all tempers of 2014, 2017, 7075, and 7178.

CASTING ALLOYS. -Aluminum casting alloys, like wrought alloys, are divided into two groups. In one group, the physical properties of the alloys are determined by the elements added and cannot be changed after the metal is cast. In the other group, the elements added make it possible to heat-treat the casting to produce desired physical properties.

The casting alloys are identified by a letter preceding the alloy number. This is exactly opposite from the case of wrought alloys, in which the letters follow the number. When a letter precedes a number, it indicates a slight variation in the composition of the original alloy. This variation in composition is made simply to impart some desirable quality. In casting alloy 214, for example, the addition of zinc, to increase its pouring qualities, is designated by the letter A in front of the number, thus creating the designation A214. When castings have been treated, the heat treatment and the

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composition of the casting are indicated by the letter T and an alloying number. An example of this is the sand casting alloy 355, which has several different compositions and tempers and is designated by 355-T6, 355-T51, and A355-T51.

Aluminum alloy castings are produced by one of three basic methods-sand mold, permanent mold, and die cast. In casting aluminum, in most cases, different types of alloys must be used for different types of castings. Sand castings and die castings require different types of alloys than those used in permanent molds

Copper Alloys

Copper alloys are commonly used for their electrical and thermal conductivities, corrosion resistance, ease of fabrication, surface appearance, strength and fatigue resistance.  Copper alloys can be readily soldered and brazed, and a number of copper alloys can be welded by arc, and resistance methods. Color of copper alloys is a significant reason for using them for decorative purposes.  For decorative parts, conventional copper alloys having specific colors are readily available

Along with ease of fabrication, some of the principal selection criteria for copper alloys are:

←   Corrosion resistance ←   Electrical conductivity ←   Thermal conductivity ←   Color and surface appearance

Corrosion resistance of copper alloys is good in many environments; however copper alloys may be attacked by some common reagents and environments.  Pure copper resists attack under some corrosive conditions.  Some copper alloys, on the other hand, sometimes have inadequate performance in certain environments.

Stress corrosion cracking most commonly occurs in brass.  Brasses containing more than 15% Zn are the most susceptible.

Dealloying is another form of corrosion that affects zinc containing copper alloys.  During dezincification of brass, selective removal of zinc results in gradual replacement of sound brass by weak, porous copper.  Unless stopped the metal is weakened and liquids or gases may be capable of leaking through the porous structure.

Electrical and thermal conductivity of copper and its alloys are relatively good.  This is why copper is the most commonly used electrical conductor.  Alloying decreases electrical conductivity to a greater extent than thermal conductivity.  This is why copper and

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high-copper alloys are preferred over other copper alloys when high electrical or thermal conductivity is required

Monel

Monel, the leading high-nickel alloy, combines the properties of high strength and excellent corrosion resistance. This metal consists of 67 percent nickel, 30 percent copper, 1.4 percent iron, 1 percent manganese, and 0.15 percent carbon. It cannot be hardened by heat treatment; it responds only to cold-working. Monel, adaptable to castings and hot- or cold-working, can be successfully welded and has working properties similar to those of steel. It has a tensile strength of 65,000 psi that, by means of cold-working, may be increased to 160,000 psi, thus entitling this metal to classification among the tough alloys. Monel has been successfully used for gears and chains, for operating retractable landing gears, and for structural parts subject to corrosion. In aircraft, Monel has long been used for parts demanding both strength and high resistance to corrosion, such as exhaust manifolds and carburetor needle valves and sleeves.

K-Monel

K-Monel is a nonferrous alloy containing mainly nickel, copper, and aluminum. It is produced by adding a small amount of aluminum to the Monel formula. It is corrosion resistant and capable of hardening by heat treatment. K-Monel has been successfully used for gears, chains, and structural members in aircraft that are subjected to corrosive attacks. This alloy is nonmagnetic at all temperatures. K-Monel can be successfully welded.

Titanium Alloys

The density of Titanium is roughly 55% that of steel.  Titanium alloys are extensively utilized for significantly loaded aerospace components.  Titanium is used in applications requiring somewhat elevated temperatures.  The good corrosion resistance experienced in many environments is based on titanium’s ability to form a stable oxide protective layer.  This makes titanium useful in surgical implants and some chemical plant equipment applications.

TYPES, CHARACTERISTICS, AND USES. -Titanium alloys are being used in quantity for jet engine compressor wheels, compressor blades, spacer rings, housing compartments, and airframe parts such as engine pads, ducting, wing surfaces, fire walls, fuselage skin adjacent to the engine outlet, and armor plate. In view of titanium’s high melting temperature, approximately 3,300°F, its high-temperature properties are disappointing. The ultimate and yield strengths of titanium drop fast above 800°F. In applications where

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the declines might be tolerated, the absorption of oxygen and nitrogen from the air at temperatures above 1,000°F makes the metal so brittle on long exposure that it soon becomes worthless. Titanium has some merit for short-time exposure up to 2,000°F where strength is not important, as in aircraft fire walls.

IDENTIFICATION OF TITANIUM. -Titanium metal, pure or alloyed, is easily identified. When touched with a grinding wheel, it makes white spark traces that end in brilliant white bursts. When rubbed with a piece of glass, moistened titanium will leave a dark line similar in appearance to a pencil mark.

Corrosion Failures Analysis

←  Uniform corrosion  ←  Pitting corrosion ←  Intergranular corrosion ←  Crevice corrosion ←  Galvanic corrosion ←  Stress corrosion cracking

Uniform Corrosion

Uniform or general corrosion is typified by the rusting of steel.  Other examples of uniform corrosion are the tarnishing of silver or the green patina associated with the corrosion of copper

Some common methods used to prevent or reduce general corrosion are listed below:

←   Coatings ←   Inhibitors ←   Cathodic protection ←   Proper materials selection

  Pitting Corrosion

Pitting is a localized form of corrosive attack.  Pitting corrosion is typified by the formation of holes or pits on the metal surface.  Pitting can cause failure due to perforation while the total corrosion, as measured by weight loss, might be rather minimal.  The rate of penetration may be 10 to 100 times that by general corrosion.

Pits may be rather small and difficult to detect.  In some cases pits may be masked due to general corrosion.  Pitting may take some time to initiate and develop to an easily viewable size.

Pitting occurs more readily in a stagnant environment.  The aggressiveness of the corrodent will affect the rate of pitting.  Some

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methods for reducing the effects of pitting corrosion are listed below:

←   Reduce the aggressiveness of the environment ←   Use more pitting resistant materials

←   Improve the design of the system

Crevice Corrosion

Crevice corrosion is a localized form of corrosive attack.  Crevice corrosion occurs at narrow openings or spaces between two metal surfaces or between metals and nonmetal surfaces.  A concentration cell forms with the crevice being depleted of oxygen.  This differential aeration between the crevice (microenvironment) and the external surface (bulk environment) gives the the crevice an anodic character.  This can contribute to a highly corrosive condition in the crevice.  Some examples of crevices are listed below:

←   Flanges ←   Deposits ←   Washers ←   Rolled tube ends ←   Threaded joints ←   O-rings ←   Gaskets ←   Lap joints  ←   Sediment

Some methods for reducing the effects of crevice corrosion are listed below:

←   Eliminate the crevice from the design ←   Select materials more resistant to crevice corrosion ←   Reduce the aggressiveness of the environment

PROPERTIES OF METALS

Hardness

Hardness refers to the ability of a metal to resist abrasion, penetration, cutting action, or permanent distortion. Hardness may be increased by working the metal and, in the case of steel and certain titanium and aluminum alloys, by heat treatment and cold-working (discussed later). Structural parts are often formed from metals in their soft state and then heat treated to harden them so that the finished shape will be retained. Hardness and strength are closely associated properties of all metals.

Brittleness

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Brittleness is the property of a metal that allows little bending or deformation without shattering. In other words, a brittle metal is apt to break or crack without change of shape. Because structural metals are often subjected to shock loads, brittleness is not a very desirable property. Cast iron, cast aluminum, and very hard steel are brittle metals.

Malleability

A metal that can be hammered, rolled, or pressed into various shapes without cracking or breaking or other detrimental effects is said to be malleable. This property is necessary in sheet metal that is to be worked into curved shapes such as cowlings, fairings, and wing tips. Copper is one example of a malleable metal.

Ductility

Ductility is the property of a metal that permits it to be permanently drawn, bent, or twisted into various shapes without breaking. This property is essential for metals used in making wire and tubing. Ductile metals are greatly preferred for aircraft use because of their ease of forming and resistance to failure under shock loads. For this reason, aluminum alloys are used for cowl rings, fuselage and wing skin, and formed or extruded parts, such as ribs, spars, and bulkheads. Chrome-molybdenum steel is also easily formed into desired shapes. Ductility is similar to malleability.

Elasticity

Elasticity is that property that enables a metal to return to its original shape when the force that causes the change of shape is removed. This property is extremely valuable, because it would be highly undesirable to have a part permanently distorted after an applied load was removed. Each metal has a point known as the elastic limit, beyond which it cannot be loaded without causing permanent distortion. When metal is loaded beyond its elastic limit and permanent distortion does result, it is referred to as strained. In aircraft construction, members and parts are so designed that the maximum loads to which they are subjected will never stress them beyond their elastic limit.

NOTE: Stress is the internal resistance of any metal to distortion.

Toughness

A material that possesses toughness will withstand tearing or shearing and may be stretched or otherwise deformed without breaking. Toughness is a desirable property in aircraft metals.

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Density

Density is the weight of a unit volume of a material. In aircraft work, the actual weight of a material per cubic inch is preferred, since this figure can be used in determining the weight of a part before actual manufacture. Density is an important consideration when choosing a material to be used in the design of a part and still maintain the proper weight and balance of the aircraft.

Fusibility

Fusibility is defined as the ability of a metal to become liquid by the application of heat. Metals are fused in welding. Steels fuse at approximately 2,500°F, and aluminum alloys at approximately 1, 110°F.

Conductivity

Conductivity is the property that enables a metal to carry heat or electricity. The heat conductivity of a metal is especially important in welding, because it governs the amount of heat that will be required for proper fusion. Conductivity of the metal, to a certain extent, determines the type of jig to be used to control expansion and contraction. In aircraft, electrical conductivity must also be considered in conjunction with bonding, which is used to eliminate radio interference. Metals vary in their capacity to conduct heat. Copper, for instance, has a relatively high rate of heat conductivity and is a good electrical conductor

Contraction and Expansion

Contraction and expansion are reactions produced in metals as the result of heating or cooling. A high degree of heat applied to a metal will cause it to expand or become larger. Cooling hot metal will shrink or contract it. Contraction and expansion affect the design of welding jigs, castings, and tolerances necessary for hot-rolled material

METAL WORKING PROCESSES

When metal is not cast in a desired manner, it is formed into special shapes by mechanical working processes. Several factors must be considered when determining whether a desired shape is to be cast or formed by mechanical working. If the shape is very complicated, casting will be necessary to avoid expensive machining of mechanically formed parts. On the other hand, if strength and quality of material are the prime factors in a given part, a cast will be unsatisfactory. For this reason, steel castings are seldom used in aircraft work.

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There are three basic methods of metal working. They are hot-working, cold-working, and extruding. The process chosen for a particular application depends upon the metal involved and the part required, although in some instances you might employ both hot- and cold-working methods in making a single part.

Hot-Working

Almost all steel is hot-worked from the ingot into some form from which it is either hot- or cold-worked to the finished shape. When an ingot is stripped from its mold, its surface is solid, but the interior is still molten. The ingot is then placed in a soaking pit, which retards loss of heat, and the molten interior gradually solidifies. After soaking, the temperature is equalized throughout the ingot, which is then reduced to intermediate size by rolling, making it more readily handled.

The rolled shape is called a bloom when its sectional dimensions are 6 x 6 inches or larger and approximately square. The section is called a billet when it is approximately square and less than 6 x 6 inches. Rectangular sections that have width greater than twice the thickness are called "slabs." The slab is the intermediate shape from which sheets are rolled.

HOT-ROLLING. -Blooms, billets, or slabs are heated above the critical range and rolled into a variety of shapes of uniform cross section. The more common of these rolled shapes are sheets, bars, channels, angles, I-beams, and the like. In aircraft work, sheets, bars, and rods are the most commonly used items that are rolled from steel. As discussed later in this chapter, hot-rolled materials are frequently finished by cold-rolling or drawing to obtain accurate finish dimensions and a bright, smooth surface.

FORGING. -Complicated sections that cannot be rolled, or sections of which only a small quantity is required, are usually forged. Forging of steel is a mechanical working of the metal above the critical range to shape the metal as desired. Forging is done either by pressing or hammering the heated steel until the desired shape is obtained.

Pressing is used when the parts to be forged are large and heavy, and this process also replaces hammering where high-grade steel is required. Since a press is slow acting, its force is uniformly transmitted to the center of the section, thus affecting the interior grain structure as well as the exterior to give the best possible structure throughout.

Hammering can be used only on relatively small pieces. Since hammering transmits its force almost instantly, its effect is limited to a small depth. Thus, it is necessary to use a very heavy hammer

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or to subject the part to repeated blows to ensure complete working of the section. If the force applied is too weak to reach the center, the finished forging surface will be concave. If the center is properly worked, the surface will be convex or bulged. The advantage of hammering is that the operator has control over the amount of pressure applied and the finishing temperature, and is able to produce parts of the highest grade.

This type of forging is usually referred to as smith forging, and it is used extensively where only a small number of parts are needed. Considerable machining and material are saved when a part is smith forged to approximately the finished shape.

Cold-Working

Cold-working applies to mechanical working performed at temperatures below the critical range, and results in a strain hardening of the metal. It becomes so hard that it is difficult to continue the forming process without softening the metal by annealing.

Since the errors attending shrinkage are eliminated in cold-working, a much more compact and better metal is obtained. The strength and hardness as well as the elastic limit are increased, but the ductility decreases. Since this makes the metal more brittle, it must be heated from time to time during certain operations to remove the undesirable effects of the working.

While there are several cold-working processes, the two with which you are principally concerned are cold-rolling and cold-drawing. These processes give the metals desirable qualities that cannot be obtained by hot-working.

COLD-ROLLING. -Cold-rolling usually refers to the working of metal at room temperature. In this operation, the materials that have been hot-rolled to approximate sizes are pickled to remove any scale, after which they are passed through chilled finished rolls. This action gives a smooth surface and also brings the pieces to accurate dimensions. The principal forms of cold-rolled stocks are sheets, bars, and rods.

COLD-DRAWING. -Cold-drawing is used in making seamless tubing, wire, streamline tie rods, and other forms of stock. Wire is made from hot-rolled rods of various diameters. These rods are pickled in acid to remove scale, dipped in lime water, and then dried in a steam room, where they remain until ready for drawing. The lime coating adhering to the metal serves as a lubricant during the drawing operation. Figure 1-23 shows the drawing of rod, tubing, and wire.

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The size of the rod used for drawing depends upon the diameter wanted in the finished wire. To reduce the rod to the desired wire size, it is drawn cold through a die. One end of the rod is filed or hammered to a point and slipped through the die opening, where it is gripped by the jaws of the draw, then pulled through the die. This series of operations is done by a mechanism known as the drawbench, as shown in figure 1-23.

To reduce the rod gradually to the desired size, it is necessary to draw the wire through successively smaller dies. Because each of these drawings reduces the ductility of the wire, it must be annealed from time to time before further drawings can be accomplished. Although cold-working reduces the ductility, it increases the tensile strength of the wire enormously. In making seamless steel aircraft tubing, the tubing is cold-drawn through a ring-shaped die with a mandrel or metal bar inside the tubing to support it while the drawing operations are being performed. This forces the metal to flow between the die and the mandrel and affords a means of controlling the wall thickness and the inside and outside diameters.

Extruding

The extrusion process involves the forcing of metal through an opening in a die, thus causing the metal to take the shape of the die opening. Some metals such as lead, tin, and aluminum may be extruded cold; but generally, metals are heated before the operation is begun.

The principal advantage of the extrusion process is in its flexibility. Aluminum, because of its workability and other favorable properties, can be economically extruded to more intricate shapes and larger sizes than is practicable with many other metals. Extruded shapes are produced in very simple as well as extremely complex sections.

A cylinder of aluminum, for instance, is heated to 750°F to 850°F, and is then forced through the opening of a die by a hydraulic ram. Many structural parts, such as stringers, are formed by the extrusion process

ALLOYING OF METALS

A substance that possesses metallic properties and is composed of two or more chemical elements, of which at least one is a metal, is called an "alloy." The metal present in the alloy in the largest proportion is called the "base metal." All other metals and/or elements added to the alloy are called "alloying elements." The metals are dissolved in each other while molten, and they do not separate into layers when the solution solidifies. Practically all the metals used in aircraft are made up of a number of alloying

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elements. In addition to increasing the strength, alloying may change the heat-resistant qualities of a metal, its corrosion resistance, electrical conductivity, or magnetic properties. It may cause an increase or decrease in the degree to which hardening occurs after cold-working. Alloying may also make possible an increase or decrease in strength and hardness by heat treatment. Alloys are of great importance to the aircraft industry in providing materials with properties that pure metals alone do not possess.

HARDNESS TESTING METHODS. -Hardness testing is a factor in the determination of the results of heat treatment as well as the condition of the metal before heat treatment. There are two commonly used methods of hardness testing, the Brinell and the Rockwell tests. These tests require the use of specific machines and are covered later in this chapter. An additional, and somewhat indirect, method known as spark testing is used in identifying ferrous metals. This type of identification gives an indication of the hardness of the metal.

Spark testing is a common means of identifying ferrous metals that have become mixed. In this test, the piece of iron or steel is held against a revolving stone, and the metal is identified by the sparks thrown off. Each ferrous metal has its own peculiar spark characteristics. The spark streams vary from a few tiny shafts to a shower of sparks several feet in length. Few nonferrous metals give off sparks when touched to a grinding stone. Therefore, these metals cannot be successfully identified by the spark test.

Wrought iron produces long shafts that are a duIl red color as they leave the stone, and they end up a white color. Cast iron sparks are red as they leave the stone, but turn to a straw color. Low-carbon steels give off long, straight shafts that have a few white sprigs. As the carbon content of the steel increases, the number of sprigs along each shaft increases, and the stream becomes whiter in color. Nickel steel causes the spark stream to contain small white blocks of light within the main burst

.HARDNESS TESTING

Learning Objective: Recognize hardness testing methods, related equipments, and their operation

Hardness testing is a method of determining the results of heat treatment as well as the state of a metal prior to heat treatment. Since hardness values can be tied in with tensile strength values and, in part, with wear resistance, hardness tests are an invaluable check of heat-treatment control and of material properties. Practically all hardness testing equipments now in service use the resistance to penetration as a measure of hardness. Included among the better known bench-type hardness testers are the Brinell and

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the Rockwell, both of which are described and illustrated in this section. Also included are three portable type hardness testers now being used by maintenance activities

BRINELL TESTER

The Brinell hardness tester, shown in figure 1-25, uses a hardened spherical ball, which is forced into the surface of the metal. The ball is 10 millimeters (0.3937inch) in diameter. A pressure of 3,000 kilograms (6,600 pounds) is used for ferrous metals and 500 kilograms for nonferrous metals. Normally, the load should be applied for 30 seconds. In order to produce equilibrium, this period may be increased to 1 minute for extremely hard steels. The load is applied by means of hydraulic pressure. The hydraulic pressure is built up by a hand pump or an electric motor, depending on the model of tester. A pressure gauge indicates the amount of pressure. There is a release mechanism for relieving the pressure after the test has been made, and a calibrated microscope is provided for measuring the diameter of the impression in millimeters. The machine has various shaped anvils for supporting the specimen and an elevating screw for bringing the specimen in contact with the ball penetrator. There are attachments for special tests.

To determine the Brinell hardness number for a metal, the diameter of the impression is first measured, using the calibrated microscope furnished with the tester. Figure 1-26 shows an impression as seen through he microscope. After measuring the diameter of the impression, the measurement is converted into the Brinell hardness number on the conversion table furnished with the tester

ROCKWELL TESTER

The Rockwell hardness tester, shown in figure 1-27, measures the resistance to penetration as does the Brinell tester, but instead of measuring the diameter of the impression, the Rockwell tester measures the depth, and the hardness is indicated directly on a dial attached to the machine. The more shallow the penetration, the higher the hardness number.

Two types of penetrators are used with the Rockwell tester–a diamond cone and a hardened steel ball. The load that forces the penetrator into the metal is called the "major load," and is measured in kilograms. The results of each penetrator and load combination are reported on separate scales, designated by letters. The penetrator, the major load, and the scale vary with the kind of metal being tested.

For hardened steels, the diamond penetrator is used, the major load is 150 kilograms, and the hardness is read on the C scale. When this

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reading is recorded, the letter C must precede the number indicated by the pointer. The C-scale setup is used for testing metals ranging in hardness from C-20 to the hardest steel (usually about C-70). If the metal is softer than C-20, the B-scale setup is used. With this setup, the 1/16-inch ball is used as a penetrator, the major load is 100 kilograms, and the hardness is read on the B scale. numbers in the outer circle are black, and the inner numbers are red.

The Rockwell tester is equipped with a weight pan, and two weights are supplied with the machine. One weight is marked in red. The other weight is marked in black. With no weight in the weight pan, the machine applies a major load of 60 kilograms. If the scale setup calls for a 100-kilogram load, the red weight is placed in the pan. For a 150-kilogram load, the black weight is added to the red weight. The black weight is always used in conjunction with the red weight; it is never used alone. Practically all testing is done with either the B-scale setup or the C-scale setup. For these scales, the colors may be used as a guide in selecting the weight (or weights) and in reading the dial. For the B-scale test, use the red weight and read the red numbers. For a C-scale test, add the black weight to the red weight and read the black numbers.

In setting up the Rockwell machine, use the diamond penetrator for testing materials that are known to be hard. If in doubt, try the diamond, since the steel ball may be deformed if used for testing hard materials. If the metal tests below C-22, then change to the steel ball.

Use the steel ball for all soft materials-those testing less than B-100. Should an overlap occur at the top of  the B scale and the bottom of the C scale, use the C-scale setup.

Before the major load is applied, the test specimen must be securely locked in place to prevent slipping and to properly seat the anvil and penetrator. To do this, a load of 10 kilograms is applied before the lever is tripped. This preliminary load is called the "minor load." The minor load is 10 kilograms regardless of the scale setup. When the machine is set up properly, it auto-matically applies the 10-kilogram load.

The metal to be tested in the Rockwell tester must be ground smooth on two opposite sides and be free of scratches and foreign matter. The surface should be perpendicular to the axis of penetration, and the two opposite ground surfaces should be parallel. If the specimen is tapered, the amount of error will depend on the taper. A curved surface will also cause a slight error in the hardness test. The amount of error depends on the curvature–the smaller the radius of curvature, the greater the error. To eliminate such error, a small flat should be ground on the curved surface if possible.

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NONMETALLIC MATERIALS

Transparent plastics, reinforced plastics, and composite materials are common materials used in aircraft construction. Sandwich construction is used for radomes as well as for structural areas where strength and rigidity are important.

TRANSPARENT PLASTICS

Transparent plastic materials used in aircraft canopies, windshields, and other transparent enclosures may be divided into two major classes, or groups, depending on their reaction to heat. They are the thermoplastic materials and the thermosetting materials. Thermoplastic materials will soften when heated and harden when cooled. These materials can be heated until soft, formed into the desired shape, and when cooled, will retain this shape. The same piece of plastic can be reheated and reshaped any number of times without changing the chemical composition of the material.

Thermosetting plastics harden upon heating, and reheating has no softening effect. They cannot be reshaped after once being fully cured by the application of heat. These materials are rapidly being phased out in favor of stretched acrylic, a thermoplastic material. Transparent plastics are manufactured in two forms of material-solid (monolithic) and laminated. Laminated plastic consists of two sheets of solid plastic bonded to a rubbery inner layer of material similar to the sandwich materials used in plate glass.

Laminated transparent plastics are well suited to pressurized applications in aircraft because of theirshatter resistance, which is much higher than that of the stretched solid plastics.

Stretched acrylic is a thermoplastic conforming to Military Specification MIL-P-25690. This specification covers transparent, solid, modified acrylic sheet material having superior crack propagation resistance (shatter resistance, craze resistance, fatigue resistance) as a result of proper hot stretching.

Stretched acrylic is prepared from modified acrylic sheets, using a processing technique in which the sheet is heated to its forming temperature and then mechanically stretched so as to increase its area approximately three or four times with a resultant decrease in its thickness. Most of the Navy’s high-speed aircraft are equipped with canopies made from stretched acrylic plastic.

COMPOSITE MATERIAL

Composites are materials consisting of a com-bination of high-strength stiff fibers embedded in a common matrix (binder) material;

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for example, graphite fibers and epoxy resin. Composite structures are made of a number of fiber and epoxy resin laminates. These laminates can number from 2 to greater than 50, and are generally bonded to a substructure such as aluminum or nonmetallic honeycomb. The much stiffer fibers of graphite, boron, and Kevlar® epoxies have given com-posite materials structural properties superior to the metal alloys they have replaced.

There are numerous combinations of composite materials being studied in laboratories and a number of types currently used in the production of aircraft components. Examples of composite materials are as follows: graphite/epoxy, Kevlar®/epoxy, boron poly-amide, graphite polyamide, boron-coated boron aluminum, coated boron titanium, boron graphite epoxy hybrid, and boron/epoxy. The trend is toward minimum use of boron/epoxy because of the cost when compared to current generation of graphite/epoxy composites. Composites are attractive structural materials because they provide a high strength/weight ratio and offer design flexibility. In contrast to traditional materials of construction, the properties of these materials can be adjusted to more efficiently match the

 

Figure 1-32.-Sandwich construction.

requirements of specific applications. However, these materials are highly susceptible to impact damage, and the extent of the damage is difficult to determine visually. Nondestructive inspection (NDI) is required to analyze the extent of damage and the effectiveness of repairs.

SANDWICH CONSTRUCTION

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From the standpoint of function, sandwich parts in naval aircraft can be divided into two broad classes: (1) radomes and (2) structural. The first class, radomes, is a reinforced plastic sandwich construction designed primarily to permit accurate and dependable functioning of the radar equipment. This type of construction was discussed in the preceding section under "Reinforced Plastics."

The second class, referred to as structural sandwich, normally has either metal or reinforced plastic facings on cores of aluminum or balsa wood. This material is found in a variety of places such as wing surfaces, decks, bulkheads, stabilizer surfaces, ailerons, trim tabs, access doors, and bomb bay doors. Figure 1-32 shows one type of sandwich construction using a honeycomb-like aluminum alloy core, sandwiched between aluminum alloy sheets, called "facings." The facings are bonded to the lightweight aluminum core with a suitable adhesive so as to develop a strength far greater than that of the components themselves when used alone.

Another type of structural sandwich construction consists of a low-density balsa wood core combined with high-strength aluminum alloy facings bonded to each side of the core. The grain in the balsa core runs perpendicular to the aluminum alloy facings, and the core and aluminum facings are firmly bonded together under controlled temperatures and pressures.

The facings in this type of construction carry the major bending loads, and the cores serve to support the facings and carry the shear loads. The outstanding characteristics of sandwich construction are strength, rigidity, lightness, and surface smoothness.

AIRCRAFT HARDWARE AND SEALS

Aircraft hardware is usually identified by its specification number or trade name. Threaded fasteners and rivets are usually identified by AN (Air Force-Navy), NAS (national aircraft standard), and MS (military standard) numbers. Quick-release fasteners are usually identified by factory trade names and size designations. 

AIRCRAFT STRUCTURAL HARDWARE.

The term aircraft structural hardware refers to many items used in aircraft construction. You should be concerned with such hardware as rivets, fasteners, bolts, nuts, screws, washers, cables, guides, and you should be familiar with common electrical system hardware

Solid Rivets

Solid rivets are classified by their head shape, by the material from which they are manufactured, and by their size. Rivet head shapes

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and their identifying code numbers are shown in figure 2-1. The prefix MS identifies hardware that conforms to written military standards. The prefix AN identifies specifications that are developed and issued under the joint authority of the Air Force and the Navy.

Rivet Identification Code 

The rivet codes shown in figure 2-1 are sufficient to identify rivets only by head shape. To be meaningful and precisely identify a rivet, certain other information is encoded and added to the basic code.

Figure 2-1.—Rivet head shapes and code numbers.

A letter or letters following the head-shaped code identify the material or alloy from which the rivet was made. Table 2-1 includes a listing of the most common of these codes. The alloy code is followed by two numbers separated by a dash. The first number is the numerator of a fraction, which specifies the shank diameter in thirty-seconds of an inch. The second number is the numerator of a fraction in sixteenths of an inch, and identifies the length of the rivet. The rivet code is shown in figure 2-2.

Rivet Composition 

Most of the rivets used in aircraft construction are made of aluminum alloy. A few special-purpose rivets are made of mild steel, Monel, titanium, and copper. Those aluminum alloy rivets made of 1100, 2117, 2017,2024, and 5056 are considered standard.

ALLOY 1100 RIVETS.— Alloy 1100 rivets are supplied as fabricated (F) temper, and are driven in this condition. No further treatment of the rivet is required before use, and the rivet’s properties do not change with prolonged periods of storage. They are relatively soft and easy to drive. The cold work resulting from driving increases their strength slightly. The 1100-F rivets are used only for riveting nonstructural parts. These rivets are identified by their plain head, as shown in table 2-1.

ALLOY 2117 RIVETS.— Like the 1100-F rivets, these rivets need no further treatment before use and can be stored indefinitely. They are furnished in the solution-heat-treated (T4) temper, but change to the

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Figure 2-2.—Rivet coding example.

solution-heat-treated and cold-worked (T3) temper after driving. The 2117-T4 rivet is in general use throughout aircraft structures, and is by far the most widely used rivet, especially in repair work. In most cases the 2117-T4 rivet may be substituted for 2017-T4 and 2024-T4 rivets for repair work by using a rivet with the next larger diameter. This is desirable since both the 2017-T4 and 2024-T4 rivets must be heat treated before they are used or kept in cold storage. The 2117-T4 rivets are identified by a dimple in the head.

ALLOY 2017 AND 2024 RIVETS.— As mentioned in the preceding paragraph, both these rivets are supplied in the T4 temper and must be heat treated. These rivets must be driven within 20 minutes after quenching or refrigerated at or below 32°F to delay the aging time 24 hours. If either time is exceeded, reheat treatment is required. These rivets may be reheated as many times as desired, provided the proper solution heat-treatment temperature is not exceeded. The 2024-T4 rivets are stronger than the 2017-T4 and are, therefore, harder to drive. The

Table 2-1.-Rivet Material Identification

 

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Figure 2-3.-Self-plugging rivet (mechanical lock).

2017-T4 rivet is identified by the raised teat on the head, while the 2024-T4 has two raised dashes on the head.

Letter A – Aluminium alloy,1100 or 3003 composition

AD- Aluminium alloy, 2117T composition

D- Aluminium alloy, 2017T

DD-A luminium alloy, 2024T

B – Aluminium alloy, 5056 composition

C- Copper alloy

M- monel

Absence of letter shows rivet manufactured from mild steel

ALLOY 5056 RIVETS.— These rivets are used primarily for joining magnesium alloy structures because of their corrosion-resistant qualities. They are supplied in the H32 temper (strain-hardened and then stabilized). These rivets are identified by a raised cross on the head. The 5056-H32 rivet may be stored indefinitely with no change in its driving characteristics.

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Round head rivet- used in interior of aircraft, except where clearance required for adjacent members.it offers resistance to tension.

Flat head rivet- like round head rivet,is used on interior structures. It is used where maximum strength is needed and where there is not sufficient clearance to use a round head rivet.

Brazier head rivets- has head large diameter, which makes it particularly adaptable for riveting thin sheet stock. Tahe brazier head rivet offers slight resistance to airflow, used on exterior surfaces. A modified brazier head having a reduced diameter

Universal head rivet is combination of round head, flat head, brazier head rivets. It is used in both interior and exterior surfaces.

Countersunk rivet are used to fasten sheet over other sheet must fix. They are also used on exterior surfaces it offer only small resistance to airflow

Blind Rivets 

In places accessible from only one side or where space on one side is too restricted to properly use a bucking bar, blind rivets are usually used. Blind rivets may also be used to secure nonstructural parts to the airframe.

Figure 2-3 shows a blind rivet that uses a mechanical lock between the head of the rivet and the pull stem. Note in view B that the collar that is attached to the head has been driven into the head and has assumed a wedge or cone shape around the groove in the pin. This holds the shank firmly in place from the head side.

The self-plugging rivet is made of 5056-H14 aluminum alloy and includes the conical recess and locking collar in the rivet head. The stem is made of 2024-T36 aluminum alloy. Pull grooves that fit into the jaws of the rivet gun are provided on the stem end that protrudes above the rivet head. The blind end portion of the stem incorporates a head and a land (the raised portion of the grooved surface) with an extruding angle that expands the rivet shank.

Applied loads for self-plugging rivets are comparable to those for solid shank rivets of the same shear strength, regardless of sheet thickness. The composite shear strength of the 5056-H14 shank and the 2024-T36 pin exceeds 38,000 psi. Their tensile strength is in excess of 28,000 psi; Pin retention characteristics are excellent in these rivets. The possibility of the pin working out is minimized by the lock formed in the rivet head.

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Hi-Shear Rivets(special rivet)

Hi-shear (pin) rivets are essentially thread less bolts. The pin is headed at one end and is grooved about the circumference at the other. A metal collar is swaged onto the grooved end. They are available in two head styles—the flat protruding head and the flush 100-degree countersunk head. Hi-shear rivets are made in a variety of materials, and are used only in shear applications. Because the shear strength of the rivet is greater than either the shear or bearing strength of sheet aluminum alloys, they are used primarily to rivet thick gauge sheets together. They are never used where the grip length is less than the shank diameter. Hi-shear rivets are shown in figure 2-4.

 

Figure 2-4.—Hi-shear rivet.

 

Figure 2-5.—Sectional view of rivnut showing head and end designs.

Hi-shear rivets are identified by code numbers similar to the solid rivets. The size of the rivet is measured in increments of thirty-seconds of an inch for the diameter and sixteenths of an inch for the grip length. For example, an NAS 1055-5-7 rivet would be a hi-shear rivet with a countersunk head. Its diameter would be 5/32 of an inch and its maximum grip length would be 7/16 of an inch.

The collars are identified by a basic code number and a dash number that correspond to the diameter of the rivet. An A before the dash number indicates an aluminum alloy collar. The NAS528-A5 collar would be used on a 5/32-inch-diameter rivet pin. Repair procedures

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involving the installation or replacement of hi-shear rivets generally specify the collar to be used.

Rivnuts

The rivnut is a hollow rivet made of 6063 aluminum alloy, counterbored and threaded on the inside. They are manufactured in two head styles, flat and countersunk, and in two shank designs, open and closed ends. See figure 2-5. Each of these rivets is available in three sizes: 6-32, 8-32, and 10-32. These numbers indicate the nominal diameter and the actual number of threads per inch of the machine screw that fits into the rivnut.

Open-end rivnuts are the most widely used, and are recommended in preference to the closed-end type. However, in sealed flotation or pressurized compartments, the closed-end rivnut must be used.

FASTENERS (SPECIAL)

Fasteners on aircraft are designed for many different functions. Some are made for high-strength requirements, while others are designed for easy installation and removal.

Lock-Bolt Fasteners

Lock-bolt fasteners are designed to meet high-strength requirements. Used in many structural applications, their shear and tensile strengths equal or exceed the requirements of AN and NAS bolts. The lock-bolt pin, shown in view A of figure 2-6, consists of a pin and collar. It is available in two head styles: protruding and countersunk. Pin retention is accomplished by swaging the collar into the locking grooves on the pin.

The blind lock bolt, shown in view B of figure 2-6, is similar to the self-plugging rivet shown in figure 2-3. It features a positive mechanical leek for pin retention.

Hi-Lok Fasteners 

The hi-lok fastener, shown in figure 2-7, com-bines the features of a rivet and a bolt and is used for high-strength, interference-free fit of primary structures. The hi-lok fastener consists of a threaded pin and threaded locking collar. The pins are made of

 

 

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Figure 2-6.—Lock bolts.

 

 

Figure 2-7.—Hi-lok fastener.

cadmium-plated alloy steel with protruding or 100-degree flush heads. Collars for the pins are made of anodized 2024-T6 aluminum

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or stainless steel. The threaded end of the pin is recessed with a hexagon socket to allow installation from one side. The major diameter of the threaded part of the pin has been truncated (cut undersize) to accommodate a 0.004-inch maximum interference-free fit. One end of the collar is internally recessed with a 1/16-inch, built-in variation that automatically provides for variable material thickness without the use of washers and without fastener preload changes. The other end of the collar has a torque-off wrenching device that controls a predetermined residual tension of preload (10%) in the fastener.

Jo-Bolt Fasteners 

The jo-bolt, shown in figure 2-8, is a high-strength, blind structural fastener that is used on difficult riveting jobs when access to one side of the work is impossible. The jo-bolt consists of three factory-assembled parts: an aluminum alloy or alloy steel nut, a threaded alloy steel bolt, and a corrosion-resistant steel sleeve. The head styles available for jo-bolts are the 100-degree flush head, the hexagon protruding head, and the 100-degree flush millable head.

 Turlock Fasteners

Turn lock fasteners are used to secure panels that require frequent removal. These fasteners are available in several different styles and are usually referred to by the manufacturer’s trade name.

CAMLOC FASTENERS.— The 4002 series Camloc fastener consists of four principal parts: the receptacle, the grommet, the retaining ring, and the

 

Figure 2-8.—Jo-bolt.

stud assembly. See figure 2-9. The receptacle is an aluminum alloy forging mounted in a stamped sheet metal base. The receptacle assembly is riveted to the access door frame, which is attached to the structure of the aircraft. The grommet is a sheet metal ring held in the access panel with the retaining ring. Grommets are furnished

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in two types: the flush type and the protruding type. Besides serving as a grommet for the hole in the access panel, it also holds the stud assembly. The stud assembly consists of a stud, a cross pin, a spring, and a spring cup. The assembly is designed so it can be quickly inserted into the grommet by compressing the spring. Once installed in the grommet, the stud assembly cannot be removed unless the spring is again compressed.

The Camloc high-stress panel fastener, shown in figure 2-10, is a high-strength, quick-release rotary fastener, and may be used on flat or curved inside or outside panels. The fastener may have either a flush or protruding stud. The studs are held in the panel with flat or cone-shaped washers—the latter being used with flush fasteners in dimpled holes. This fastener may be distinguished from screws by the deep No. 2 Phillips recess in the stud head and by the bushing in which the stud is installed. 

A threaded insert in the receptacle provides an adjustable locking device. As the stud is inserted and turned counterclockwise one-half turn or more, it screws out the insert to permit the stud key to engage the insert cam when turned clockwise. Rotating the

 

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Figure 2-9.-Camloc 4002 series fastener.

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Figure 2-10.-Camloc high-stress panel fastener.

stud clockwise one-fourth turn engages the insert. Continued rotation screws the insert in and tightens the fastener. Turning the stud one-fourth turn counterclockwise will release the stud, but will not screw the insert out far enough to permit re-engagement. The stud should be turned at least one-half turn counterclockwise to reset the insert.

AIRLOC FASTENERS.— Figure 2-11 shows the parts that make up an Airloc fastener. The Airloc fastener also consists of a receptacle, a stud, and a cross pin. The stud is attached to the access panel and is held in place by the cross pin. The receptacle is riveted to the access panel frame.Two types of Airloc receptacles are available: the fixed (view A) and the floating (view B). The floating receptacle makes for easier alignment of the stud in the receptacle. Several types of studs are also available, but in each instance the stud and cross pin come as separate units so the stud may be easily installed in the access panel.The Airloc receptacle is fastened to the inner surface of the access panel frame by two rivets. The rivet heads must be flush with the outer surface of the panel frame. When you are replacing receptacles, drill out the two old rivets and attach the new receptacle by flush riveting. Be careful not to mar the sheet. When you are inserting the stud and cross pin, insert the stud through the access panel and, by using a special hand tool, insert the cross pin in the stud. Cross pins can be removed by means of special ejector pliers.

 DZUS FASTENERS.— DZUS fasteners are available in two types. A light-duty type is used on box covers, access hole covers, and lightweight fairings. The heavy-duty type is used on cowling and heavy fairings. The main difference between the two Dzus fasteners is a grommet, which is only used on

 

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Figure 2-11.—Airloc fastener.

the heavy-duty fasteners. Otherwise, their construction features are about the same. Figure 2-12 shows the parts of a light-duty Dzus fastener. Notice that they include a spring and a stud. The spring is made of cadmium-plated steel music wire, and is usually riveted to an aircraft structural member. The stud comes in a number of designs (as shown in views A, B, and C) and mounts in a dimpled hole in the cover assembly.

When the panel is being positioned on an aircraft, the spring riveted to the structural member enters the hollow center of the stud. Then, when the stud is turned about one-fourth turn, the curved jaws of the stud slip over the spring and compress it. The resulting tension locks the stud in place and secures the panel.

FLAT-HEAD PINS.— The flat-head pin is used with tie rod terminals or secondary controls, which do not operate continuously. The flat-head pin should be secured with a cotter pin. The pin is normally installed with the head up. See figure 2-17. This precaution is taken

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to maintain the flat-head pin in the installed position in case of cotter pin failure.

SNAP RINGS.— A snap ring is a ring of metal, either round or flat in cross section, that is tempered to have springlike action. This springlike action will hold the snap ring firmly seated in a groove. The external types are designed to fit in a groove around the outside of a shaft or cylinder. The internal types fit in a groove inside a cylinder. Special pliers are designed to install each type of snap ring. Snap rings can be reused as long as they retain their shape and springlike action. External snap rings may be safety wired, but internal types are never safetied.

STUDS.— There are four types of studs used in aircraft structural applications. They are the coarse thread, fine thread, stepped and lockring studs. Studs may be drilled or undrilled on the nut end. Coarse (NAS183) and fine (NAS184) thread studs are manufactured from alloy steel and are heat treated. They have identical threads on both ends. The stepped stud has a different thread on each end of the stud. The lockring stud may be substituted for undersize or oversize studs. The lockring on this stud prevents it from backing out due to vibration, stress, or temperature variations. Refer to the Hardware Manual, detailed information on studs.

HELI-COIL INSERTS.— Heli-coil thread inserts are primarily designed to be used in materials that arc not suitable for threading because of their softness. The inserts are made of a diamond cross-sectioned stainless steel wire that is helically coiled and, in its finished form, is similar to a small, fully compressed spring. There are two types of heli-coil inserts. See figure 2-18. One is the plain insert, made with a tang that forms a portion of the bottom coil offset, and is used to drive the insert. This tang is left on the insert after installation, except when its removal is necessary to provide clearance for the end of the bolt. The tang is notched to break off from the body of the insert, thereby providing full penetration for the fastener. 

Bolts

Many types of bolts are used on aircraft. However, before discussing some of these types, it might be helpful to list and explain some commonly used bolt terms. You should know the names of bolt parts and be aware of the bolt dimensions that must be considered in selecting a bolt. Figure 2-19 shows both types of information.

 

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Figure 2-19.—Bolt terms and dimensions.

The three principal parts of a bolt are the head, thread, and grip. The head is the larger diameter of the bolt and may be one of many shapes or designs. The head keeps the bolt in place in one direction, and the nut used on the threads keeps it in place in the other direction.

To choose the correct replacement, several bolt dimensions must be considered. One is the length of the bolt. Note in figure 2-19 that the bolt length is the distance from the tip of the threaded end to the head of the bolt. Correct length selection is indicated when the chosen bolt extends through the nut at least two full threads. In the case of flat-end bolts or chamfered (rounded) end bolts, at least the full chamfer plus one full thread should extend through the nut. See figure 2-19. If the bolt is too short, it may not extend out of the bolt hole far enough for the nut to be securely fastened. If it is too long, it may extend so far that it interferes with the movement of nearby parts.

Unnecessarily long bolts can affect weight and balance and reduce the aircraft payload capacity. In addition, if a bolt is too long or too short, its grip is usually the wrong length. As shown in figure 2-20, grip length should be approximately the same as the thickness of the material to be fastened. If the grip is too short, the threads of the bolt will extend into the bolt hole and may act like a reamer when the material is vibrating. To prevent this, make certain that no more than two threads extend into the bolt hole. Also make certain that any threads that enter the bolt hole extend only into the thicker member that is being fastened. If the grip is too long, the nut will run out of threads before it can be tightened. In this event, a bolt with a shorter grip should be used, or if the bolt grip extends only a short distance through the hole, a washer maybe used. A second bolt dimension that must be considered is diameter. Figure 2-19 shows that the diameter of the bolt is the thickness of its shaft. If this thickness is 1/4 of an inch or more, the bolt diameter is usually given in fractions of an inch; for example, 1/4, 5/16, 7/16, and 1/2.

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However, if the bolt is less than 1/4 of an inch thick, the diameter is usually expressed as a whole number. For instance, a bolt that is 0.190 inch in diameter is called a No. 10 bolt, while a bolt that is 0.164 inch in diameter is called a No. 8.

Class 1 is loose fit, Class 2 is free fit, C lass 3 is medium fit, Class 4 is close fit. Generally aircraft bolt are manufactured in class 3 fit. Aircraft bolt are manufactured from Cadmium plated or zinc plated corrosion resistant steel , unplated corrosion resistant steel and anodized aluminium alloy.

AN bolt are manufactured in 3 head styles- hex head, clevis, eyebolt,used in both shear and tension load. NAS bolt are manufactured in hex , countersunk and internal wrenching head, MS bolt are in hex head and internal wrenching head type.

The results of using a bolt of the wrong diameter should be obvious. If the bolt is too big, it cannot enter the bolt hole. If the diameter is too small, the bolt has too much play in the bolt hole, and the chances are that it is not as strong as the correct bolt. The third and fourth bolt dimensions that should be considered when choosing a bolt replacement are head thickness and width. If the head is too thin or too narrow, it may not be strong enough to bear the load imposed on it. If the head is too thick or too wide, it may extend so far that it interferes with the movement of adjacent parts.

 BOLT HEADS.— The most common type of head is the hex head. See figure 2-20. This type of head may be thick for greater strength or relatively thin in order to fit in places having limited clearances. In addition, the head may be common or drilled to lockwire the bolt. A hex-head bolt may have a single hole drilled through it between two of the sides of the hexagon and still be classed as common. The drilled head-hex bolt has three holes drilled in the head, connecting opposite sides of the hex. Figure 2-20.-Correct and incorrect grip lengths.

 

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Figure 2-21.—Bolt heads

Seven additional types of bolt heads are shown in figure 2-21. Notice that view A shows an eyebolt, often used in flight control systems. View B shows a countersunk-head, close-tolerance bolt. View C shows an internal-wrenching bolt. Both the countersunk-head bolt and the internal-wrenching bolt have hexagonal recesses (six-sided holes) in their heads. They are tightened and loosened by use of appropriate sized Allen wrenches. View D shows a clevis bolt with its characteristic round head. This head may be slotted, as shown, to receive a common screwdriver or recessed to receive a Reed-and-Prince or a Phillips screwdriver.

View E shows a torque-set wrenching recess that has four driving wings, each one offset from the one opposite it. There is no taper in the walls of the recess. This permits higher torque to be applied with less tendency for the driver to slip or cam out of the slots.

View F shows an external-wrenching head that has a washer face under the head to provide an increased bearing surface. The 12-point head gives a greater wrench gripping surface.

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View G shows a hi-torque style driving slot. This single slot is narrower at the center than at the outer portions. This and the center dimple provide the slot with a bow tie appearance. The recess is also undercut in a taper from the center to the outer ends, producing an inverted keystone shape. These bolts must be installed with a special hi-torque driver adapter. They must also be driven with some type of torque-limiting or torque-measuring device. Each diameter of bolt requires the proper size of driver for that particular bolt. The bolts are available in standard and reduced 100-degree flush heads. The reduced head requires a driver one size smaller than the standard headBOLT THREADS.— Another structural feature in which bolts may differ is threads. These usually come in one of two types: coarse and fine. The two are not interchangeable. For any given size of bolt

 

 

Figure 2-22.—Bolt head markings. 

there is a different number of coarse and fine threads per inch. For instance, consider the 1/4-inch bolts. Some are called 1/4-28 bolts because they have 28 fine threads per inch. Others have only 20 coarse threads per inch and are called 1/4-20 bolts. To force one size of threads into another size, even though both are 1/4 of an inch, can strip the finer threads or softer metal. The same thing is true concerning the other sizes of bolts; therefore, make certain that bolts you select have the correct type of threads.

BOLT MATERIAL.— The type of metal used in an aircraft bolt helps to determine its strength and its resistance to corrosion. Therefore, make certain that material is considered in the selection of replacement bolts. Like solid shank rivets, bolts have distinctive head markings that help to identify the material from which they are manufactured. Figure 2-22 shows the tops of several hex-head bolts, each marked to indicate the type of bolt material.

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BOLT IDENTIFICATION.— Unless current directives specify otherwise, every unserviceable bolt should be replaced with a bolt of the same type. Of course, substitute and interchangeable items are sometimes available, but the ideal fix is a bolt-for-bolt replacement. The part number of a needed bolt may be obtained by referring to the illustrated parts breakdown (IPB) for the aircraft concerned. Exactly what this part number means depends upon whether the bolt is AN (Air Force-Navy), NAS (National Aircraft Standard), or MS (Military Standard). 

AN Part Number.— There are several classes of AN bolts, and in some instances their part numbers reveal slightly different types of information. However, most AN numbers contain the same type of information.

Figure 2-23 shows a breakdown of a typical AN bolt part number. Like the AN rivets discussed earlier, it starts with the letters AN. Next, notice that a number follows the letters. This number usually consists of two digits. The first digit (or absence of it) shows the class of the bolt. For instance, in figure 2-23, the series number has only one digit, and the absence of one digit shows that this part number represents a general-purpose hex-head bolt. However, the part numbers for some bolts of this class have two digits. In fact, general-purpose hex-head bolts include all part numbers beginning with AN3, AN4, and so on, through AN20. Other series numbers and the classes of bolts that they represent are as follows:

AN21 through AN36—clevis bolts

AN42 through AN49—eyebolts

The series number shows another type of information other than bolt class. With a few exceptions, it indicates bolt diameter in sixteenths of

 

 

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Figure 2-23.—AN bolt part number breakdown.

an inch. For instance, in figure 2-23, the last digit of the series number is 4; therefore, this bolt is 4/16 of an inch (1/4 of an inch) in diameter. In the case of a series number ending in 0, for instance AN30, the 0 stands for 10, and the bolt has a diameter of 10/16 of an inch (5/8 of an inch).

Refer again to figure 2-23, and observe that a dash follows the series number. When used in the part numbers for general-purpose AN bolts, clevis bolts, and eyebolts, this dash indicates that the bolt is made of carbon steel. With these types of bolts, the letter D means 2017 aluminum alloy. The letters DD stand for 2024 aluminum alloy. For some bolts of this type, a letter H is used with these letters or with the dash. If it is so used, the letter H shows that the bolt has been drilled for safetying. C indicates corrosion resitant steel. A indicates that shank is undrilled.

Next, observe the number 20 that follows the dash. This is called the dash number. It represents the bolt’s grip (as taken from special tables). In this instance the number 20 stands for a bolt that is 2 1/32 inches long. 

The last character in the AN number shown in figure 2-23 is the letter A. This signifies that the bolt is not drilled for cotter pin safetying. If no letter were used after the dash number, the bolt shank would be drilled for safetying. 

NAS Part Number.— Another series of bolts used in aircraft construction is the NAS. See figure 2-24. In considering the NAS 144-25 bolt (special internal-wrenching type), observe that the bolt identification code starts with the letters next number (4) indicates the bolt diameter in sixteenths of an inch. The dash number (25) indicates bolt grip in sixteenths of an inch.

 

 

Figure 2-24.-NM bolt part number breakdown.

 

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Figure 2-25.—MS bolt part number breakdown.

MS Part Number.— MS is another series of bolts used in aircraft construction. In the part number shown in figure 2-25, the MS indicates that the bolt is a Military Standard bolt. The series number (20004) indicates the bolt class and diameter in sixteenths of an inch (internal-wrenching, 1/4-inch diameter). The letter H before the dash number indicates that the bolt has a drilled head for safetying. The dash number (9) indicates the bolt grip in sixteenths of an inch.

Clevis Bolt- is round and slotted to rcieve a commom screw driver or recessed to receive cross point screw driver. This type of bolt is used only in Shear application.

Eye bolt- is used in tension load

Nuts

Aircraft nuts differ in design and material, just as bolts do, because they are designed to do a specific job with the bolt. For instance, some of the nuts are made of cadmium-plated carbon steel, stainless steel, brass, or aluminum alloy. The type of metal used is not identified by markings on the nuts themselves. Instead, the material must be recognized from the luster of the metal.

Nuts also differ greatly in size and shape. In spite of these many and varied differences, they all fall under one of two general groups: self-locking and nonself-locking. Nuts are further divided into types such as plain nuts, castle nuts, check nuts, plate nuts, channel nuts, barrel nuts, internal-wrenching nuts, external-wrenching nuts, shear nuts, sheet spring nuts, wing nuts, and Klincher locknuts

NONSELF-LOCKING NUTS.— Nonself-locking nuts require the use of a separate locking device for security of installation. There are several types of these locking devices mentioned in the following paragraphs in connection with the nuts on which they are used. Since no single locking device can be used with all types of nonself-locking nuts, you must select one suitable for the type of nut being used.

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SELF-LOCKING NUTS.— Self-locking nuts provide tight connections that will not loosen under vibrations. Self-locking nuts approved for use on aircraft meet critical strength, corrosion-resistance, and temperature specifications. The two major types of self-locking nuts are prevailing torque and free spinning. The two general types of prevailing torque nuts are the all-metal nuts and the nonmetallic insert nuts. New self-locking nuts must be used each time components are installed in critical areas throughout the entire aircraft, including all flight, engine, and fuel control linkage and attachments. The flexloc nut is an example of the all-metal type. The elastic stop nut is an example of the nonmetallic insert type. All-metal self-locking nuts are constructed with the threads in the load-carrying portion of the nut out of phase with the threads in the locking portion, or with a saw cut top portion with a pinched-in thread. The locking action of these types depends upon the resiliency of the metal when the locking section and load-carrying section are forced into alignment when engaged by the bolt or screw threads. 

PLAIN HEX NUTS.— These nuts are available in self-locking or nonself-lotting styles. When the nonself-locking nuts are used, they should be locked with an auxiliary locking device such as a check nut or lock washer. Used in large tension loads.

CASTLE NUTS.— These nuts are used with drilled shank bolts, hex-head bolts, clevis bolts, eyebolts, and drilled-head studs. These nuts are designed to be secured with cotter pins or safety wire. Used in tension loads.

CASTELLATED SHEAR NUTS.— Like the castle nuts, these nuts are castellated for safetying. They are not as strong or cut as deep as the castle nuts. Used in shear load only.

CHECK NUTS.— These nuts are used in locking devices for nonself-locking plain hex nuts, setscrews, and threaded rod ends.

PLATE NUTS.— These nuts are used for blind mounting in inaccessible locations and for easier maintenance. They are available in a wide range of sizes and shapes. One-lug, two-lug, and right-angle shapes are available to accommodate the specific physical requirements of nut locations. Floating nuts provide a controlled amount of nut movement to compensate for subassembly misalignment. They can be either self-locking or nonself-locking. See figure 2-27.

CHANNEL NUTS.— These nuts are used in applications requiring anchored nuts equally spaced around openings such as access and inspection doors and removable leading edges. Straight or curved channel nut strips offer a wide range of nut spacings and provide a

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multinut unit that has all the advantages of floating nuts. They are usually self-locking.

BARREL NUTS.— These nuts are installed in drilled holes. The round portion of the nut fits in the drilled hole and provides a self-wrenching effect. They are usually self-locking.

INTERNAL-WRENCHING NUTS.— These nuts are generally used where a nut with a high tensile strength is required or where space is limited and the use of external-wrenching nuts would not permit the use of conventional wrenches for installation and removal. This is usually where the bearing surface is counterbored. These nuts have a nonmetallic insert that provides the locking action.

 

 

 

Figure 2-26.–Nuts.

 

 

Figure 2-27.–Self-locking nuts.

 

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Figure 2-28.-Sheet spring nut.

Screws

The most common threaded fastener used in aircraft construction is the screw. The three most used types are the structural screw, machine screw, and the self-tapping screw.

STRUCTURAL SCREWS.— Structura.l screws are used for assembling structural parts. They are made of alloy steel and are heat treated. Structural screws have a definite grip length and the same shear and tensile strengths as the equivalent size bolt. They differ from structural bolts only in the type of head. These screws are available in round-head, countersunk-head, and brazier-head types, either

 

Figure 2-29.—Typical installations of the Wincher locknut.

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Figure 2-30.-Structural screws. 

slotted or recessed for the various types of screwdrivers. See figure 2-30.

MACHINE SCREWS.— The commonly used machine screws are the flush-head, round-head, fillister-head, socket-head, pan-head and truss-head types. 

Flush-Head.— Flush-head machine screws are used in countersunk holes where a flush finish is desired. These screws are available in 82 and 100 degrees of head angle, and have various types of recesses and slots for driving.

Round-Head.— Round-head machine screws are frequently used in assembling highly stressed aircraft components.

Fillister-Head.— Fillister-head machine screws are used as general-purpose screws. They may also be used as cap screws in light applications such as the attachment of cast aluminum gearbox cover plates.

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Socket-Head.— Socket-head machine screws are designed to be screwed into tapped holes by internal wrenching. They are used in applications that require high-strength precision products, compactness of the assembled parts, or sinking of the head into holes.

Pan- and Truss-Head.— Pan-head and truss-head screws are general-purpose screws used where head height is unimportant. These screws are available with cross-recessed heads only.

SELF-TAPPING SCREWS.— A self-tapping screw is one that cuts its own internal threads as it is turned into the hole. Self-tapping screws can be used only in comparatively soft metals and materials. Self-tapping screws may be further divided into two classes or groups: machine self-tapping screws and sheet metal self-tapping screws 

Machine self-tapping screws are usually used for attaching removable parts, such as nameplates, to castings. The threads of the screw cut mating threads in the casting after the hole has been predrilled. Sheet metal self-tapping screws are used for such purposes as temporarily attaching sheet metal in place for riveting. They may also be used for permanent assembly of nonstructural parts, where it is necessary to insert screws in blind applications.

CAUTION

Self-tapping screws should never be used to replace standard screws, nuts, or rivets in the original structure. Over a period of time, vibration and stress will loosen this type of fastener, causing it to lose its holding ability.

WASHERS

Washers such as ball socket and seat washers, taper pin washers, and washers for internal-wrenching nuts and bolts have been designed for special applications. See figure 2-31.

Ball socket and seat washers are used where a bolt is installed at an angle to the surface, or where perfect alignment with the surface is required at all times.

These washers are used together.

Taper pin washers are used in conjunction with threaded taper pins. They are installed under the nut to effect adjustment where a plain washer would distort. 

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Washers for internal-wrenching nuts and bolts are used in conjunction with NAS internal-wrenching bolts. The washer used under the head is countersunk to seat the bolt head or shank radius. A plain washer is used under the nut.

 

 

 

Figure 2-31.—Various types of special washers.

CABLES

CABLE-A cable is a group of wires or a group of strands of wires twisted together into a strong wire rope. The wires or strands may be twisted in various ways. The relationship of the direction of twist of each strand to each other and to the cable as a whole is called the opposite to the twist of the strands around the center strand or core, the cable will not stretch (or set) as much as one in which they are all twisted in the same direction. This direction of twist (in opposite direction) is most commonly adopted, and it is called a regular or an ordinary lay. Cables may have a right regular lay or a left regular lay. If the strands are twisted in the direction of twist around the center strand or core, the lay is called a lang lay. There is a right and left lang lay. The only other twist arrangement-twisting the strands alternately right and left, then twisting them all either to the right or to the left about the core—is called a reverse lay. Most aircraft cables have a right regular lay.

When aircraft cables are manufactured, each strand is first formed to the spiral or helical shape to fit the position it is to occupy in the finished cable. The process of such forming is called preforming, and cables made by such a process are said to be preformed. The process of preforming is adopted to ensure flexibility in the finished cable and to relieve bending and twisting stresses in the strands as they are woven into the cable. It also keeps the strands from spreading when the cable is cut. All aircraft cables are internally lubricated during construction.

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Aircraft control cables are fabricated either from flexible, preformed carbon steel wire or from flexible, preformed, corrosion-resistant steel wire. The small corrosion-resistant steel cables are made of steel containing not less than 17 percent chromium and 8 percent nickel, while the larger ones (those of the 5/16-, 3/8-, and 7/16-inch diameters) are made of steel that, in addition to the amounts of chromium and nickel just mentioned, also contains not less than 1.75 percent molybdenum. 

Cables may be designated 7 x 7, 7 x 19, or 6 x 19 according to their construction. A 7 x 7 cable consists of six strands of seven wires each, laid around a center strand of seven wires. A 7 x 19 cable consists of six strands of 19 wires, laid around a 19-wire central strand, A 6 x 19 IWRC cable consists of six strands of 19 wires each, laid around an independent wire rope center.

The size of cable is given in terms of diameter measurement. A 1/8-inch cable or a 5/16-inch cable means that the cable measures 1/8 inch or 5/16 inch in diameter, as shown in figure 2-32. Note that the cable diameter is that of the smallest circle that would enclose the entire cross section of the cable. Aircraft

 

Figure 2-32.-Cable cross section.

control cables vary in diameters, ranging from 1/16 of an inch to 3/8 of an inch.

Fittings

Cable ends may be equipped with several different types of fittings such as terminals, thimbles, bushings, and shackles. Terminal fittings are generally of the swaged type. (The swaging process is described in detail in chapter 9 of this manual.) Terminal fittings are

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available with threaded ends, fork ends, eye ends, and single-shank and double-shank ball ends. 

Threaded-end, fork-end, and eye-end terminals are used to connect the cable to turnbuckles, bell cranks, and other linkage in the system. The ball terminals are used for attaching cable to quadrants and special connections where space is limited. The

 

Figure 2-33.—Types of cable terminal fittings. 

single-shank ball end is usually used on the ends of cables, and the double-shank ball end may be used at either the ends or in the center of a cable run. Figure 2-33 shows the various types of terminal fittings. 

Thimble, bushing, and shackle fittings may be used in place of some types of terminal fittings when facilities and supplies are limited and immediate replacement of the cable is necessary. Figure 2-34 shows these fittings

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Turnbuckles

A turnbuckle is a mechanical screw device consisting of two threaded terminals and a threaded barrel. Figure 2-35 shows a typical turnbuckle assembly. Turnbuckles are fitted in the cable assembly for the purpose of making minor adjustments in cable length and for adjusting cable tension. One of the terminals has right-hand threads

 

Figure 2-34.-Thimble, bashing, and shackle fittings.

 

 

Figure 2-35.—Typical turnbuckle assembly.

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and the other has left-hand threads, The barrel has matching right- and left-hand threads internally. The end of the barrel, with left-hand threads inside, can usually be identified by either a groove or knurl around the end of the barrel. Barrels and terminals are available in both long and short lengths. When you install a turnbuckle in a control system, it is necessary to screw both of the terminals an equal number of turns into the turnbuckle barrel. It is also essential that all turnbuckle terminals be screwed into the barrel at least until not more than three threads are exposed. On initial installation, the turnbuckle terminals should not be screwed inside the turnbuckle barrel more than four threads. Figure 2-36 shows turnbuckle thread tolerances. 

After a turnbuckle is properly adjusted, it must be safetied. There are several methods of safetying turnbuckles. However, only two methods have been adopted as standard procedures by the services. These methods are discussed later in this chapter.

GUIDES

Fairleads (rubstrips), grommets, pressure seals, and pulleys are all types of cable guides. They are used to protect control cables by preventing the cables

 

 

Figure 2-36.-Turnbuckle tolerances.

 

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Figure 2-37.—Typical cable guides.

from rubbing against nearby metal parts. They are also used as supports to reduce cable vibration in long stretches (runs) of cable. Figure 2-37 shows some typical cable guides.

Fairleads

Fairleads maybe made of a solid piece of material to completely encircle cables when they pass through holes in bulkheads or other metal parts. Fairleads may be used to reduce cable whipping and vibration in long runs of cable. Split fairleads are made for easy installation around single cables to protect them from rubbing on the edges of holes.

Grommets

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Grommets are made of rubber, and they are used on small openings where single cables pass through the walls of unpressurized compartments.

Pressure Seals

Pressure seals are used on cables or rods that must move through pressurized bulkheads. They fit tightly enough to prevent air pressure loss, but not so tightly as to hinder movement of the unit.

Pulleys

Pulleys (or sheaves) are grooved wheels used to change cable direction and to allow the cable to move with a minimum of friction. Most pulleys used on aircraft are made from layers of cloth impregnated with phenolic resin and fused together under high temperatures and pressures. Aircraft pulleys are extremely strong and durable, and cause minimum wear on the cable passing over them. Pulleys are provided with grease-sealed bearings, and usually do not require further lubrication. However, pulley bearings may be pressed out, cleaned, and relubricated with special equipment. This is usually done only by depot-level maintenance activities. Pulley brackets made of sheet or cast aluminum are required with each pulley installed in the aircraft. See figure 2-38. Besides holding the pulley in the correct position and at the correct angle, the brackets prevent the cable from slipping out of the groove on the pulley wheel.

SECTORS AND QUADRANTS

These units are generally constructed in the form of an arc or in a complete circular form. They are grooved around the outer circumference to receive the cable, as shown in figure 2-38. The names sector and quadrant are used interchangeably. Sectors and quadrants are similar to bell cranks and walking

 

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Figure 2-38.-Control system components.

 

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Figure 2-39.-Series 145 and 155 quick-disconnect couplings.

beams, which are used for the same purpose in rigid control systems.

AIRCRAFT HYDRAULIC HARDWARE AND SEALS

Learning Objective: Identify the various hydraulic hardware and seals used in naval aircraft. 

Hardware, such as the quick-disconnect coupling, and seals and packings are used throughout the aircraft. They are essential for safe and proper operation of aircraft systems. You must be familiar with the various types used on naval aircraft.

STATIC DISCHARGERS

Static dischargers are commonly known as static wicks static discharge wicks. They are used on aircraft to allow the continuous satisfactory operation of onboard navigation and radio communication systems. During adverse charging conditions, they limit the potential static buildup on the aircraft and control interference generated by static charge. Static dischargers are not

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lighting arrestors and do not reduce or increase the likelihood of an aircraft being struck by lightning. Static dischargers are subject to damage or significant changes in resistance characteristics as a result of lightning strike to the aircraft, and should be inspected after a lightning strike to ensure proper static discharge operation. Static dischargers are fabricated with a wick of wire or a conductive element on one end, which provides a high resistance discharge path between the aircraft and the air. See figure 2-56. They are attached on some aircraft to the ailerons, elevators, rudder, wing, horizontal and vertical stabilizer tips, etc. Refer to your applicable aircraft’s MIM for maintenance procedures.

AIRCRAFT SAFETYING METHODS

Learning Objective: Identify the various safety methods used on aircraft hardware. 

You will come in contact with many different types of safetying materials. These materials are used to stop rotation and other movement of fasteners. They are also used to secure other equipment that may come loose due to vibration in the aircraft.

COTTER PINS

Cotter pins are used to secure bolts, screws, nuts, and pins. Some cotter pins are made of low-carbon steel, while others consist of stainless steel and are more resistant to corrosion. Also, stainless steel cotter pins may be used in locations where nonmagnetic material is required. Regardless of shape or material, all cotter pins are used for the same general purpose—safetying. Figure 2-57 shows three types of cotter pins and how their size is determined.

NOTE: Whenever uneven prong cotter pins are used, the length measurement is to the end of the shortest prong.

SAFETY WIRE

Safety wire comes in many types and sizes. You must first select the correct type and size of wire for the job. Annealed corrosion-resistant wire is used in high-temperature, electrical equipment, and aircraft instrument applications. All nuts except the self-locking types must be safetied; the method used depends upon the particular installation.

TORQUING OF FASTENERS

Learning Objective: Recognize the importance of the proper torquing of fasteners and the required torquing procedures.

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Fastener fatigue failure accounts for the majority of all fastener problems. Fatigue breaks are caused by insufficient tightening and the lack of proper preload or clamping force. This results in movement between the parts of the assembly and bending back and forth or cyclic stressing of the fastener. Eventually, cracks will progress to the point where the fastener can no longer support its designed load. At this point the fastener fails with varying consequences.

TORQUING PROCEDURES

For the nut to properly load the bolt and prevent premature failure, a designated amount of torque must be applied. Proper torque reduces the possibility of the fastener loosening while in service. The correct torque to apply when you are tightening an assembly is based on many variables. The fastener is subjected to two stresses when it is tightened. These stresses are torsion and tension. Tension is the desired stress, while torsion is the undesirable stress caused by friction. A large percentage of applied torque is used to overcome this friction, so that only tension remains after tightening. Proper tension reduces the possibility of fluid leaks.

Table 3-4.—Recommended Torque Values (Inch-Pounds)

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The recommended torque values provided in table 3-4 have been established for average dry, cadmium- plated nuts for both the fine and coarse thread series of nuts. Thread surface variations such as paint, lubrication, hardening, plating, and thread distortion may alter these values considerably. The torque values must be followed unless the MIM or structual repair manual for the specific aircraft requires a specific torque for a given nut. Torque values vary slightly with manufacturers. When the torque values are included in a

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Figure 3-12.—Torque wrenches.

technical manual, these values take precedence over the standard torque values provided in the of nuts, bolts, and screws used in aircraft construction. You should use this manual when specific torque values are not provided as a part of the removal/replacement instructions.

To obtain values in foot-pounds, divide inch-pound values by 12. Do not lubricate nuts or bolts except for corrosion-resistant instructed to do so, steel parts or where specifically Always tighten by rotating

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the nut first if possible. When space considerations make it necessary to tighten the fastener by rotating the bolt head, approach the high side of the indicated torque range. Do not exceed the maximum allowable torque value. Maximum torque ranges should be used only when materials and surfaces being joined are of sufficient thickness, area, and strength to resist breaking, warping, or other damage. 

For corrosion-resisting steel nuts, use the torque values given for shear-type nuts. The use of any type of drive-end extension on a torque wrench changes the dial reading required to obtain the actual values indicated in the torque range tables. See figure 3-12.

TORQUING COMPUTATION

When you are using a drive-end extension, the torque wrench reading must be computed using the following formula:

where:

S = handle setting or reading

T = torque applied at end of adapter

La = length of handle in inches

Ea = length of extension in inches

If you desire to exert 100 inch-pounds at the end of the wrench and extension, when La equals 12 inches and Ea equals 6 inches, it is possible to determine the handle setting by making the following calculation:

S = 66.7 inch–pounds

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Whenever possible, attach the extension in line with the torque wrench. When it is necessary to attach the extension at any angle to the torque wrench, the effective length of the assembly will be La + Ea, as shown in figure 3-12. In this instance, length Eb must be substituted for length Ea in the formula.

NOTE: It is not advisable to use a handle extension on a flexible beam-type torque wrench at any time. The use of a drive-end extension on any type of torque wrench makes use of the formula necessary. When the formula has been used, force must be applied to the handle of the torque wrench at the point from which the measurements were taken. If this is not done, the torque obtained will be in error.

HOSE FABRICATION AND MAINTENANCE

Hose assemblies are used to connect moving parts with stationary parts and in locations subject to severe vibration. Hose assemblies are heavier than aluminum-alloy tubing and deteriorate more rapidly. They are used only when absolutely necessary. Hose assemblies are made up of hose and hose fittings. A hose consists of multiple layers of various materials. An example of the hose most often used in medium-pressure applications is shown in figure 5-1.

TYPES OF HOSE

There are two basic types of hose used in military aircraft and related equipment. They are synthetic rubber and polytetrafluoroethylene, commonly known as Teflon@ or PTFE. 

Bulk hose identification will vary with the materials from which the hose is constructed. It is important that you are able to clearly identify the proper hose to be used by recognizing the various hose markings.

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Figure 5-1.—Medium pressure synthetic rubber hose, MIL-H-8794.

Synthetic Rubber Hose 

Synthetic rubber hose has a seamless synthetic rubber inner tube covered with layers of cotton and wire braid, and an outer layer of rubber impregnated cotton braid. The hose is provided in low-, medium-, and high-pressure types.

Synthetic rubber hose (if rubber-covered) is identified by the indicator stripe and markings that are stencilled along the length of the hose. The indicator stripe (also called the lay line because of its use in determining the straightness or lie of a hose) is a series of dots or dashes. The markings (letters and numerals) contain the military specification, the hose size, the cure date, and the manufacturer’s federal supply code number. This information is repeated at intervals of 9 inches. Refer to figure 5-2.

Size is indicated by a dash followed by a number (referred to as a dash number). The dash number does not denote the inside or outside diameter of the hose. It refers to the equivalent outside diameter of rigid tube size in sixteenths (1/16) of an inch. A dash 8 (-8) mates to a number 8 rigid tube, which has an outside diameter of one-half inch (8/16). The inside of the hose will not be one-half inch, but slightly smaller to allow for tube thickness.

The cure date is provided for age control. It is indicated by the quarter of the year and year. The year is divided into four quarters.

1st quarter — January, February, March

2d quarter — April, May, June

3d quarter — July, August, September

4th quarter — October, November, December

The cure date is also marked on bulk hose containers in accordance with Military Standard 129 (MIL-STD-129). 

Synthetic rubber hose (if wire-braid covered) is identified by bands wrapped around the hose at the ends and at intervals along the length of the hose. Each band is marked with the same information (fig.5-2).

Teflon® Hose The Teflon® hose is made up of a tetrafluoro- ethylene resin, which is processed and extruded into tube shape to a desired

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size. It is covered with stainless steel wire, which is braided over the tube for strength and protection. The advantages of this hose are its operating temperature range, its chemical inertness to all fluids normally used in hydraulic and engine lubrication systems, and its long life. At this time, only medium-pressure and high-pressure types are available. These are complete assemblies with factory-installed end fittings. The fittings may be either the detachable type or the swaged type. When failures occur, replacement must be made on a complete assembly basis.

Teflon® hose is identified by metal bands or pliable plastic bands at the ends and at 3-foot intervals. These bands contain the hose military specification number, size indicated by a dash (-) and a number, operating pressure, and the manufacturer’s federal supply code number. Refer to figure 5-2.

Flared Fitting

There are two types of flared tubing joints—the single-flared joint and the double-flared joint. The single-flared tube joint is used on all sizes of steel tubing and 5052 aluminum alloy tubing that conforms to Federal Specification WW-T-700/6 with 1/2 inch or larger outside diameter, Use the tube flaring tool (fig. 6-12) to prepare tube ends for flaring. Check tube ends for roundness, square cut, cleanliness, and no draw marks or scratches. Draw marks can spread and split the tube when it is flared. Use a deburring tool to remove burrs from the inside and outside of the tubing. Remove filings, chips, and grit from inside the tube. Clean the tube. Slip the fitting nut and sleeve onto the tube. Place the tube into the proper size hole in the grip die. Make sure the end of the tube extends 1/64 inch above the surface of the grip die. Center the plunger over the end of the tube and

Figure 6-12.—Tube flaring toot (single-flare).

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tighten the yoke setscrew to secure the tube in the grip die and hold the yoke in place. Strike the top of the plunger several light blows with a hammer or mallet, turning the plunger a half turn after each blow. Loosen the setscrew and remove the tube from the grip die. Check to make sure that no cracks are evident and that the flared end of the tube is no larger than the largest diameter of the sleeve being used. The double-flare tube joint is used on all 5052 aluminum alloy tubes with less than 1/2-inch outside diameter, except when used with NAS 590 series tube fittings and NAS 591 connectors or NAS 593 con-nectors. Aluminum alloy tubing used in low-pressure oxygen systems or corrosion-resistant steel used in brake systems must be double flared. Double flare reduces the chance of cutting the flare by overtightening. When fabricating oxygen lines, make sure that all tube material and tools are kept free of oil and grease. Use the tube flaring tool (fig. 6-13) to prepare tube ends. Check tube end for roundness, square cut, cleanliness, and make sure there are no draw marks or scratches. Draw marks can split the tubing when it is flared.

Figure 6-13.—Tube flaring tool (double-flare).

Use a deburring tool to remove burrs from the inside and outside of tube. Remove filings, chips, and grit from inside the tube. Clean the tube. Select the proper size die blocks, and place one-half of the die block into the flaring tool body with the countersunk end towards the ram guide. Install the nut and sleeve, and lay the tube in the die block with 1/2 inch protruding beyond countersunk end. Place the other half of the die block into the tool body, close latch plate, and tighten the clamp nuts fingertight. Insert the upset flare punch in the tool body with the gauge end toward the die blocks. The upset flare punch has one end counterbored or recessed to gauge the amount of tubing needed to form a double lap flare. Insert the ram and tap lightly with a hammer or mallet until the upset flare punch contacts the die blocks, and the die blocks are set against the stop plate on the bottom. Use a wrench to tighten the latch plate nuts alternately, beginning with the closed side, to prevent distortion of the tool. Reverse the upset flare punch; insert the upset flare punch and ram into the tool body. Tap lightly with a hammer or mallet until

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the upset flare punch contacts the die blocks. Remove the upset flare punch and ram. Insert the finishing flare punch and ram. Tap the ram lightly until a good seat is formed (fig. 6-14). Check the seat at intervals during the finishing operation to avoid overseating.

Flareless Fitting

Preparing tube ends for flareless fitting requires a presetting operation whereby the sleeve is set onto the tubing. Presetting is necessary to form the seal between the sleeve and the tube without damaging the connector. Presetting should always be accomplished with a presetting tool, such as the one shown in figure 6-15. These tools are machined from tool steel and hardened so that they may be used with a minimum of distortion and wear.

Figure 6-14.—Tube position and resulting flare.

NOTE: A flareless-tube connector may be used as a presetting tool in case of an emergency. However, when connectors are used as presetting tools, aluminum connectors should be used only once, and steel connectors should not be used more than five times. 

Special procedures are used in the presetting operation. Select the correct size presetting tool or a flareless fitting body. Clamp the presetting tool or flareless fitting body in a vise. Slide a nut and then a sleeve onto the tube, and make sure the pilot and cutting edge of the sleeve points toward the end of tube. Select the lubricant from table 6-4, and lubricate fitting threads, tool seat, and shoulder sleeve. Place the tube end firmly against the bottom of the presetting tool seat, while slowly screwing the nut onto the tool threads with a wrench until the tube

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BASIC HYDRAULIC/PNEUMATIC AND EMERGENCY POWER SYSTEMS

Open Center

An open center system is one having fluid flow, but no pressure in the system when the actuating mechanisms are idle. The pump circulates the fluid from the reservoir, through the selector valves, and back to the reservoir. Figure 7-1 shows a basic open center system. The open center system may employ any number of subsystems, with a selector valve for each subsystem. Unlike the closed center system, the selector valves of the open center system are always connected in series with each other. In this arrangement, the system pressure line goes through each selector valve, Fluid is always allowed free passage through each selector valve and back to the reservoir until one of the selector valves is positioned to operate a mechanism.

When one of the selector valves is positioned to operate an actuating device, fluid is directed from the pump through one of the working lines to the actuator. See view B of figure 7-1. With the

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selector valve in this position, the flow of fluid through the valve to the reservoir is blocked. The pressure builds up in the system to overcome the resistance and moves the piston of the actuating cylinder, The fluid from the opposite end of the actuator returns to the selector valve and flows back to the reservoir. Operation of the system following actuation of the component depends on the type of selector valve being used. Several types of selector valves are used in conjunction with the open center system. One type is both manually engaged and manually disengaged. First the valve is manually moved to an operating position. Then, the actuating mechanism reaches the end of its operating cycle, and the pump output continues until the system relief valve relieves the pressure. The relief valve unseats and allows the fluid to flow back to the reservoir. The system pressure remains at the relief valve set pressure until the selector valve is manually returned to the neutral position. This action reopens the open center flow and allows the system pressure to drop to line resistance pressure.

The manually engaged and pressure disengaged type of selector valve is similar to the valve pre-viously discussed. When the actuating mechanism reaches the end of its cycle, the pressure continues to rise to a predetermined pressure. The valve auto-matically returns to the neutral position and to open center flow.

Closed Center

In the closed center system, the fluid is under pressure whenever the power pump is operating. Figure 7-2 shows a complex closed center system.

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Figure 7-1.—Basic open center hydraulic system.

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The power pump may be one used with a separate pressure regulator control. The power pump may be used with an integral pressure control valve that eliminates the need for a pressure regulator. This system differs from the open center system in that the selector or directional control valves are arranged in parallel and not in series. The means of controlling pump pressure will vary in the closed center system. If a constant delivery pump is used, the system pressure will be regulated by a pressure regulator. A relief valve acts as a backup safety device in case the regulator fails. If a variable displacement pump is used, system pressure is controlled by the pump’s integral pressure mechanism compensator. The compensator automatically varies the volume output. When pressure approaches normal system pressure, the compensator begins to reduce the flow output of the pump. The pump is fully compensated (near zero flow) when normal system pressure is attained. When the pump is in this fully compensated condition, its internal bypass

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mechanism provides fluid circulation through the pump for cooling and lubrication. A relief valve is installed in the system as a safety backup. An advantage of the open center system over the closed center system is that the continuous pressurization of the system is eliminated. Since the pressure is built up gradually after the selector valve is moved to an operating position, there is very little shock from pressure surges. This action provides a smoother operation of the actuating mechanisms. The operation is slower than the closed center system, in which the pressure is available the moment the selector valve is positioned. Since most aircraft applications require instantaneous operation, closed center systems are the most widely used.

Power systems are designed to produce and maintain a given pressure. The pressure output of most of the Navy’s high-performance aircraft is 3,000 psi. The hydraulic system, shown in figure 7-2, is an example of a representative 3,000 psi hydraulic power system. The aircraft has three independent hydraulic power systems. The two primary systems are the flight hydraulic power system and the combined hydraulic power system. These systems are pressurized by two independent engine-driven hydraulic pumps on each engine. The auxiliary power system also operates on 3,000 psi pressure. It is pressurized by the hydraulic hand pump and/or the electric motor-driven hydraulic pump. The auxiliary power system is similar to the combined hydraulic power system. The primary difference is that the combined system supplies hydraulic pressure to utility hydraulic circuits and the flight controls.

The hydraulic control valves and actuators that operate the primary flight controls are of the tandem construction type. This design permits operation from either or both of the two power systems. With this arrangement, either engine can fail or be shut down without complete loss of hydraulic power to either system. The flight system reservoir supplies fluid to the two engine-driven flight system pumps. The combined system reservoir supplies fluid to the two engine-driven combined system pumps and to the auxiliary hydraulic power system. Both reservoirs are of the pressurized piston type. They are pressurized by engine bleed air during engine operations and by an external air (nitrogen) source during maintenance operations.

Hydraulic system pressure is indicated on the integrated hydraulic pressure indicator. This indicator displays the output pressure of the flight and combined hydraulic power systems. The flight hydraulic power system provides power for the operation of the rudder, stabilizer, and flaperons. It also provides power for operation of the automatic flight control system actuators, which are an integral part of the rudder and stabilizer control surface actuators. The flight hydraulic system also controls the automatic operation of the

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isolation valve. This valve is a part of the combined hydraulic system.

The combined hydraulic power system consists of two parallel circuits—one to power the primary systems and the other to power the secondary systems. The primary system consists of spin recovery, rudder, stabilizer flaperon, speed brakes, and electric ram air turbine systems. The secondary system consists of wing slats, wing flaps, wing fold, landing gear, arresting gear, wheel brakes, nosewheel steering, and the nose strut locking systems. 

The isolation valve shuts off flow to the secondary systems during flight and limits the combined system’s pressure requirements to operation of the primary circuit. Operation of the isolation valve is both automatic and manual. The reservoir pressurization system provides the reservoir with a differential pressure of 40 psi to prevent engine-driven pump cavitation. The pressure is maintained at 40 psi by the air regulator. In the event of regulator failure, the relief valve installed between the regulator and the reservoir prevents overpressurization. The relief valve opens at 50 psi. The chemical air drier removes excessive moisture from the bleed air. Dry, clean air is sent to the reservoir through the check valve, air regulator, and relief valve.

TWO bleeder valves are installed in the flight and combined system reservoirs. One is found on the air side of the reservoir and the other on the fluid side. The air side valve permits the bleeding of air pressure during system maintenance. It allows the bleeding of any hydraulic fluid seepage past seals to the air side. The fluid side bleeder reduces excessive fluid level and bleeds air from the fluid side.

Quick-disconnect fittings in the hydraulic power systems permit easy pump or engine removal without loss of fluid to the system. The fittings connect ground hydraulic test stands for maintenance purposes. The pump disconnects should not be forced together against the back pressure of a pressurized reservoir or system. Forcing disconnects together may result in damaged seals in the male ends of the disconnects. When the disconnects do not slide in smoothly, they should be removed and checked for proper seating of the O-rings. Replace seals if they are damaged. The seal goes on top of the O-ring. When the disconnects are uncoupled, the ends not being used should be properly protected from dirt and other contamination. Use only approved metal closures.

RESERVOIRS

The reservoir is a tank in which an adequate supply of fluid for the system is stored. Fluid flows from the reservoir to the pump, where

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it is forced through the system and eventually returned to the reservoir.

The reservoir not only supplies the operating needs of the system, but it also replenishes fluid lost through leakage. Furthermore, the reservoir serves as an overflow basin for excess fluid forced out of the system by thermal expansion (the increase of fluid volume caused by temperature changes), the accumulators, and by piston and rod displacement. The reservoir also furnishes a place for the fluid to purge itself of air bubbles that may enter the system. Foreign matter picked up in the system may also be separated from the fluid in the reservoir, or as it flows through line filters.

Most nonpressurized reservoirs contain filters to maintain the hydraulic fluid in a clean state, free from foreign matter. They are usually located in filler necks and internally within the reservoir. The mesh-type filter (finger strainer), usually installed in the filler neck, removes foreign particles from fluid that is added to the reservoir. Internally installed filters clean the fluid as it returns to the reservoir from the system. This type of installation may have a bypass valve incorporated to allow fluid to bypass the filter if it becomes clogged. Some modern aircraft hydraulic reservoirs do not incorporate this feature. All reservoirs containing filters are designed to permit easy removal of the filter element for cleaning or replacement.

A reservoir instruction plate is usually attached to the reservoir, or it may be attached to the aircraft structure adjacent to the filler opening. Navy specifications designate the minimum information that must be contained on this plate. Figure 7-3 shows the reservoir instruction plate. Information on an instruction plate must include the following:

1. Simple and complete instructions for tilling

2. Reservoir fluid capacity at full level

3. Full level indication

4. Refill level indication

5. Specification number and color of fluid

6. Position of operating cylinders during filling

7. System pressure (accumulator charged or discharged)

8.Instructions regarding air bleeding

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Additional information may be added, when required, such as the following:

1. Additional full and refill levels under various conditions of system pressure

2. Safety precautions

3. Filter element servicing information

4. Total fluid capacity of the system

There are two classes of hydraulic reservoirs— class I and class II. Class I reservoirs are constructed in such a manner that the air and hydraulic fluid are not separated. Class II reservoirs are constructed in such a manner that the pressurizing agent and fluid chambers are separated. This is accomplished by installing a piston between the chambers.

Nonpressurized reservoirs are vented to the atmosphere so the reservoir can "breathe." This is done to prevent a vacuum from being formed as the fluid level in the reservoir is lowered. The vent also makes it possible for air that has entered the system to find a means of escape.

The reservoir on aircraft designed for high-altitude flying is usually pressurized. Pressurizing assures a positive flow of fluid to the pump at high altitudes when low atmospheric pressures are encountered.

On some aircraft, the reservoir is pressurized by bleed air taken from the compressor section of the engine. On others, the reservoir may be pressurized by hydraulic system pressure.

Nonpressurized Reservoirs

Nonpressurized reservoirs are used in several transport, patrol, and utility aircraft. These aircraft are not designed for violent maneuvers; in some cases, they do not fly at high altitudes. Those aircraft that incorporate nonpressurized reservoirs and fly at high altitudes have the reservoirs installed within a pressurized area. High altitude in this situation means an altitude where atmospheric pressure is inadequate to maintain sufficient flow of fluid to the hydraulic pumps. Most nonpressurized reservoirs are constructed in a cylindrical shape. The outer housing is manufactured from a strong corrosion-resistant metal. 

Filter elements are normally installed internality within the reservoir to clean returning system hydraulic fluid. In some of the older aircraft, a filter bypass valve is incorporated to allow fluid to bypass

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the filter in the event the filter becomes clogged. Reservoirs serviced by pouring fluid directly into the reservoir have a filler strainer (finger strainer) assembly incorporated within the filler well to strain out impurities as the fluid enters the reservoir.

Generally, reservoirs described in the above paragraph use a visual gauge to indicate the fluid quantity. Gauges incorporated on or in the reservoir may be either a glass tube, a direct reading gauge, or a float-type rod, which is visible through a transparent dome. In some cases, the fluid quantity may also be read in the cockpit through the use of quantity transmitters.

A typical nonpressurized reservoir is shown in figure 7-4, This reservoir consists of a welded body and cover assembly clamped together. Gaskets are incorporated to seal against leakage between assemblies.

QUANTITY INDICATING GAUGE. —The reservoir fluid quantity is indicated through a mechanically operated float and arm (liquidometer) type of unit. The quantity gauge is mounted directly on the side of the reservoir. As shown in figure 7-4, the float and arm unit extends into the reservoir. The shell of the liquidometer provides a glass window over a pointer and dial, with the pointer mechanically linked to the float arm. As the float arm moves to correspond to the fluid level, the pointer, through mechanical linkage, moves to indicate the quantity available. This provides a direct reading sight gauge at the reservoir.

This same float movement actuates the potentiometer wiper arm of an integral transmitter potentiometer. The remote indicating circuit is energized, and a duplicate indication of the reservoir

AIR PRESSURE REGULATORS. —Air pressure used in pressurizing hydraulic reservoirs must be controlled within safe limits. Specific pressure requirements vary between aircraft. In some aircraft, the air pressure is controlled by an air pressure regulator (fig. 7-9). This regulator normally maintains 40 psi pressure in the reservoir. It also incorporates a relief valve to relieve excessive pressure and a differential valve to allow equalization

AIR RELIEF VALVE. —An air relief valve is normally incorporated in the air portion of the hydraulic power system to relieve excessive air pressure entering the reservoir due to a mal-functioning air pressure regulator. The relief valve shown in figure 7-10 is cylindrical in shape and consists of a housing, poppet, spring, and adjusting screw. This valve may be mounted directly to the reservoir or in a line leading from the reservoir, depending on the aircraft system design. During operation, air pressure enters the inlet port and contacts the poppet surface. When system air pressure increases to 50 psi, the poppet is forced off its seat, which allows excessive air pressure to be

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exhausted to the atmosphere. When system pressure is lowered to 49 psi, the poppet spring tension overcomes system pressure and reseats the poppet, thus closing the valve.

Maintenance of the valve usually includes the replacement of all seals and the adjustment of its controlling pressures. This valve is designed to relieve at a cracking (just open) pressure of 50 psi; the reseating pressure is 49 psi. The valve will operate at full flow when the pressure reaches 60 psi. All pressure adjustments of relief valves must be performed on a test bench. You can control valve pressures by adjusting the adjusting screw on the valve until the proper settings are obtained.

PUMPS

All aircraft hydraulic systems have one or more power-driven pumps and may have a hand pump as an additional source of power. Power-driven pumps are the primary source of energy, and may be either engine-driven or electric-motor driven. As a general rule, motor-driven pumps are installed for use in emergencies; that is, for operation of actuating units when the engine-driven pump is inoperative. Hand pumps are generally installed for testing purposes as well as for use in emergencies.

In this section, the various types of pumps used in naval aircraft, both hand- and power-driven, are described and illustrated.

Hand Pumps

Hand pumps are used in hydraulic systems to supply fluid under pressure to subsystems, such as the landing gear, flaps, canopy, and bomb-bay doors, and to charge brake accumulators. Systems using hand pumps are classified as emergency systems. Most of these systems may be used effectively during preventive maintenance.

Double-action type of hand pumps are used in hydraulic systems. Double action means that a flow of fluid is created on each stroke of the pump handle instead of every other stroke, as in the single-action type. There are several versions of the double-action hand pump, but all use the reciprocating piston principle, and operation is similar to the one shown in figure 7-13.

This pump consists of a cylinder, a piston containing a built-in check valve (A), a piston rod, an operating handle, and a check valve (B) at the inlet port. When the piston is moved to the left in the illustration, check valve (A) closes and check valve (B) opens.

Fluid from the reservoir then flows into the cylinder through inlet port (C). When the piston is moved to the right, check valve (B)

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closes. The pressure created in the fluid then opens check valve (A), and fluid is admitted behind the piston. Because of the space occupied by the piston rod, there is room for only part of the fluid; therefore, the remainder is forced out port (D) into the pressure line. If the piston is again moved to the left, check valve (A) again closes. The fluid behind the piston is then forced through outlet port (D). At the same time, fluid from the reservoir flows into the cylinder through check

Power-Driven Pumps

As previously mentioned, power pumps are generally driven by the aircraft engine, but may also be electric-motor driven. Power pumps are classified according to the type of pumping action used, and may be either the gear type or piston type. Power pumps may be further classified as constant displacement or variable displacement.

A constant displacement pump is one that displaces or delivers a constant fluid output for any rotational speed. For example, a pump might be designed to deliver 3 gallons of fluid per minute at a speed of 2,800 revolutions per minute. As long as it runs at that speed, it will continue to deliver at that rate, regardless of the pressure in the system. For this reason, when the constant displacement pump is used in a system, a pressure regulator or unloading valve must also be incorporated. The pressure regulator valve will maintain a set pressure in the system by diverting excess pump flow back to the reservoir. The unloading valve will divert all pump flow back to the reservoir when the preset system pressure is reached. This condition remains in effect until further demand is placed on the system. 

A variable displacement pump has a fluid output that varies to meet the demand of the system. For example, a pump might be designed to maintain system pressure at 3,000 psi by varying its fluid output from 0 to 7 gallons per minute. When this type of pump is used, no external pressure regulator or unloading valve is needed. This function is incorporated in the pump and controls the pumping action by maintaining a variable volume, at near constant pressure, to meet the hydraulic system demands.

GEAR-TYPE PUMP. —A gear-type pump consists of two meshed gears that revolve in a housing (fig. 7-14). The drive gear in the installation is turned by a drive shaft that engages an electric motor. The clearance between the gear teeth as they mesh and between the teeth and pump housing is very small. The inlet port is connected to the reservoir line, and the outlet port is connected to the pressure line. In the illustration, the drive gear is turning in a counterclockwise direction, and the driven (idle) gear is turning in a clockwise direction. As the teeth pass the inlet port, fluid is trapped between the teeth and the housing. This fluid is carried around the housing to the outlet port. As the teeth mesh again, the

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fluidbetween the teeth is displaced into the outlet port. This action produces a positive flow of fluid under pressure into the pressure line. A shear pin or shear section that will break under excessive loads is incorporated in the drive shaft. This is to protect the engine accessory drive if pump failure is caused by excessive load or jamming of parts.All gear-type pumps are constant displacement pumps. These pumps are usually driven by a dc wound electric motor. For those aircraft using batteries, the pump may be used to build up hydraulic pressure for the brake system during towing operation.

Maintenance of a pump at the organizational level consists of replacement of the complete assembly. The motor and pump may be ordered separately; however, this is normally done by intermediate- and depot-level maintenance only.

RELIEF VALVES

Relief valves are not new to most people; different types of relief valves are used in our homes and automobiles, as well as many other places. Relief valves are pressure limiting or safety devices commonly used to prevent pressure from building up to a point where it might blow seals or burst or damage the container in which it is installed, etc. In aircraft, relief valves are installed within hydraulic systems to relieve excessive pressurized fluid caused from thermal expansion, pressure surges, and the failure of a hydraulic pump’s compensator or other regulating devices.

Main System Relief Valves

Main system relief valves are designed to operate within certain specific pressure limits and to relieve complete pump output when in the open position. Relief valves are set to open and close at pressures determined by the system in which they are installed. In systems designed to operate at 3,000 psi normal pressure, the relief valve might be set to be completely open at 3,650 psi and reseat at 3,190 psi. These pressure ranges may vary from one aircraft to another. When the relief valve is in the open position, it directs excessive pressurized fluid to the reservoir return line.

Thermal Relief Valves 

Thermal relief valves are usually smaller as compared to system relief valves. They are used in systems where a check valve or selector valve prevents pressure from being relieved through the main system relief valve.

Figure 7-25 shows a typical thermal relief valve. As pressurized fluid in the line in which it is installed builds up to an excessive amount,

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the valve poppet is forced off its seat; this allows excessive pressurized fluid to flow through the relief valve to the reservoir return line, as shown in view B of figure 7-25. When system pressure decreases to a predetermined pressure, spring tension overcomes system pressure and forces the valve poppet to the closed position, as shown in view A.

Relief valve maintenance is limited to adjusting the valve for proper relieving pressure and checking the valve for leakage. If you think a relief valve is leaking internally, a flexible hose maybe connected to the return port of the valve and the drippings, if any, caught in a container. The opening and closing pressure of the valve may also be checked in this manner provided an external source of rower is used.

SHUTOFF VALVES

All hydraulic systems do not have shutoff valves incorporated; however, in some systems a shutoff valve is installed in the fluid supply line between the reservoir and the engine-driven pumps, and other places where shutting off the fluid is desirable. These valves, like other valves, may be electrically or manually controlled, depending upon the design of the valve.

The purpose of shutoff valves differ according to their installation. All shutoff valves control the flow of fluid; however, they may isolate troubles by shutting off a complete system or subsystem, or they may control the speed a component moves by partially closing the valve (manual type).

ACCUMULATORS

The purpose of the accumulator in a hydraulic system is to store a volume of fluid under pressure. There are several reasons why it is advantageous to store a volume of fluid under pressure. Some of these are listed below:

1. An accumulator acts as a cushion against pressure surges that may be caused by the pulsating fluid delivery from the pump or from system operations.

2. The accumulator supplements the pump’s output when the pump is under a peak load by storing energy in the form of fluid under pressure.

3. The energy stored in the accumulator may be used to actuate a unit in the event of normal hydraulic system failure. For example, sufficient energy can be stored in the accumulator for several applications of the wheel brakes.

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There are two general types of accumulators in use on naval aircraft. They are the spherical type and the cylindrical type. Until a few years ago, the spherical type was the more commonly used accumulator; however, the cylindrical type has proved more satisfactory for high-pressure hydraulic systems, and is now more commonly used than the spherical type. Examples of both types are shown in figure 7-34.

Spherical Type

The spherical type accumulator is constructed in two halves that are screwed together. A synthetic rubber diaphragm is installed between both halves, making two chambers. Two threaded openings exist in the assembled component. The opening at the top, as shown in figure 7-34, contains a screen disc that prevents the diaphragm from extruding through the threaded opening when system pressure is depleted, thus rupturing the diaphragm. On some designs the screen is replaced by a button protector fastened to the center of the diaphragm. The top threaded opening provides a means for connection of the fluid chamber of the accumulator to the hydraulic system. The bottom threaded opening provides a means for installation of an air filler valve. This valve (when open) allows an air/nitrogen source to be connected to and enter the accumulator; moreover, when the valve is closed, it traps the air/nitrogen within the accumulator.

Cylindrical Type

Cylindrical accumulators consist of a cylinder and piston assembly. End caps are attached to both ends of the cylinder. The internal piston separates the fluid and air/nitrogen chambers. Both the end caps and piston are sealed with gaskets and packings to prevent external leakage around the end caps and internal leakage between the chambers. In one end cap, a hydraulic fitting is used to attach the fluid chamber to the hydraulic system. In the other end cap, an air filler valve is installed to perform the same function as the filler valve installed in the spherical accumulator.

Operation

In operation, the compressed-air chamber is charged to a predetermined pressure, which is somewhat lower than the system operating pressure. This initial charge is referred to as the accumulator preload.

As an example of accumulator operation, let us assume that the cylindrical accumulator in figure 7-34 is designed for a preload of 1,300 psi in a 3,000 psi system. When the initial charge of 1,300 psi is introduced into the unit, hydraulic system pressure is zero. As air

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pressure is applied through the air pressure port, it moves the piston toward the opposite end until it bottoms. If the air behind the piston has a pressure of 1,300 psi, the hydraulic system pump will have to create a pressure within the system greater than 1,300 psi before the hydraulic fluid can actuate the piston. Thus, at 1,301 psi the piston will start to move within the cylinder, compressing the air as it moves. At 2,000 psi it will have backed up several inches. At 3,000 psi the piston will have backed up to its normal operating position, compressing the air until it occupies a space less than one-half the length of the cylinder.

When actuation of hydraulic units lowers the system pressure, the compressed air will expand against the piston, forcing fluid from the accumulator. This supplies an instantaneous supply of fluid to the hydraulic system.

Many aircraft have several accumulators in the hydraulic system. There may be a main system accumulator and an emergency system accumulator. There may also be auxiliary accumulators located in various unit systems. Regardless of the number and their location within the system, all accumulators perform the same function-that of storing an extra volume of hydraulic fluid under pressure.

Maintenance

Accumulators should be visually examined for indications of external hydraulic fluid leaks. They should then be examined for external air leaks by brushing the exterior with soapy water, which will form bubbles where the air leaks occur.

The air valve assembly should be loosened to examine the accumulator for internal leaks. If hydraulic fluid comes out of the air valve, the accumulator should be removed and replaced. The overhaul or repair of the accumulator is not a line maintenance function, but it is the responsibility of an intermediate-level activity.

The air preload pressure should be checked after relieving the hydraulic system pressure by operating the wing flaps or other hydraulically actuated unit. The majority of the accumulators installed in naval aircraft are equipped with air pressure gauges for this purpose. When the accumulator is not equipped with a high-pressure air gauge, you may install one at the air preload fitting for this purpose. The required pressure can be found in the MIM for each aircraft. The preload pressure may be checked by another method in case the accumulator is not equipped with an air pressure gauge. With the system pressure (as indicated by the cockpit gauge) at the normal operating value, relieve system pressure by operating the wing flaps or another unit slowly. The pressure gauge reading must be watched carefully. The last reading before the indicator

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needle drops suddenly to zero is accepted as the accumulator preload air pressure.

Before disassembly of any accumulator, ensure that the air preload has been completely exhausted. This may be accomplished by loosening the swivel nut on the air filler valve until all air is out; then remove the valve.

Servicing

The purpose of the hydraulic system accumulator is to store an extra volume of fluid under pressure. The energy stored in an accumulator is used for various purposes, such as the actuation of a unit in the event of normal hydraulic system failure. For example, sufficient energy can be stored in an accumulator for several applications of the wheel brakes.

Most accumulators are installed with an air gauge and a high-pressure air valve mounted on a panel of the structure near the accumulator. Figure 7-35 shows the brake system accumulator installation used on one type of aircraft. The air valve used in the accumulator installations is usually the same type as that used on shock struts.

To service an accumulator, the hydraulic pressure that is trapped in the accumulator must be relieved. This is accomplished by actuating the units involved. For example, the hydraulic pressure in a brake accumulator may be relieved by applying the emergency brake several times. When the hydraulic pressure is relieved, the accumulator gauge should indicate the air or nitrogen pressure specified for the particular accumulator installation. If the pressure indicated is below the specified pressure, the accumulator must be recharged with dry compressed air or nitrogen

PRESSURE INDICATORS

Pressure gauges installed in hydraulic and pneumatic systems are used to indicate existing hydraulic and pneumatic pressures, and are calibrated in pounds per square inch. Naval aircraft use both the direct reading gauges and the synchro (electric) type.

Direct Reading Type

Direct reading gauges are used in installations such as accumulators, emergency air bottles, arresting gear snubbers, and brake systems. The gauge is connected directly into units or lines leading from units and become part of the container or system. At these points the gauge is able to sample existing pressure.

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The main part of the direct reading gauge is the Bourdon tube. The Bourdon tube is a curved metal tube that is oval in cross-sectional shape (fig. 7-36).

 One end of the Bourdon tube is closed, while the other end has a fitting for connecting it to a pressure source. The fitting end is fastened to the gauge frame, while the other end is free to move so it can operate the mechanical linkage.

Assume that fluid pressure enters the Bourdon tube. Since fluid pressure will be transmitted equally in all directions and the area on the outside radius of the tube is greater than that of its inside, the force will also be greater on the outside radius, which tends to straighten the tube. As the movable end of the tube tries to turn outward, it turns the pivot segment gear. This gear meshes with a smaller rotary gear to which a pointer is attached, and its movement causes a reading on the pressure gauge. The gauge dial is calibrated so that the needle points to a number that corresponds to the exact pressure that is applied. When the pressure is removed, the Bourdon tube acts as a spring, and returns to its normal position.

Synchro Type

On most newer aircraft, an electrically operated (synchro) pressure indicator is used. Figure 7-37 shows the pressure indicator of a typical naval aircraft. This aircraft is equipped with three hydraulic systems—No. 1 flight control system, No. 2 flight control system, and utility system. One indicator provides pressure indication for all three systems. This type of arrangement is desirable because it saves instrument panel space.

The indicator system consists of three pressure transmitters, one located in each of the system lines, The transmitters operate on the Bourdon tube principle. Expansion and contraction of the Bourdon tube is transmitted by mechanical linkage to the rotor of a transmitter synchro. The synchro transmits an electrical signal through wiring to the pressure indicator. The indicator contains two synchros mechanically attached to the two separate pointers. When the HYD PRESS SELECTOR switch (fig. 7-37) is in the No. 1 and No. 2 FLT CONT position, the pointers (marked "1" and "2") indicate the pressure in their respective systems, independent of each other. When the HYD PRESS SELECTOR switch is in the UTILITY position, the synchros are connected in electrical parallel, and the pointers align with each other and act as one.

Although the Aviation Electrician’s Mate is responsible for inspecting and maintaining all the aircraft gauges and other instruments, you must know how to read the hydraulic pressure gauge to inspect and maintain the hydraulic system.

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Pressure gauges on some naval aircraft are calibrated to register from 0 to 2,000 psi; on others they register from 0 to 4,000 psi. The gauge in figure 7-37 is an example of the latter type. As shown in figure 7-37, on gauges designed for a range of 0 to 4,000 psi, the dial is calibrated with four major markings with the numerals 1,2,3, and 4. One major intermediate graduation between each numeral and four minor intermediate markings between the major markings are for reading to the nearest 100 psi. On these gauges, the numeral reading must be multiplied by 1,000 to obtain the actual pressure in psi. 

On gauges designed for a range of 0 to 2,000 psi, the dial is calibrated with two major markings, the numerals 1,000 and 2,000, and four intermediate graduations for reading to the nearest 200 psi. A gauge of this type is shown in figure 7-38.

Pneumatic System

Two types of pneumatically operated emergency systems are currently used in naval aircraft. One type consists merely of one or more storage cylinders, a control in the cockpit for releasing the contents of the cylinders, a ground charge valve, and the connecting lines and fittings. This type of system must be serviced with compressed air or nitrogen. The other type of system in current use has its own compressor and other equipment necessary for maintaining an adequate supply of compressed air during flight. Provision for ground charging this type of system is also provided. In addition to a compressor, the components in this type of system usually include a filter, a pressure regulator, a moisture separator, a relief valve, a chemical drier, and storage cylinder(s).

AIR COMPRESSORS. —A typical air com-pressor is shown in figure 7-43. An installation of this type receives its supply of air from the compressor section of the aircraft engine. This air is then compressed further to the required pressure for operating the system. Compressors of this type are capable of maintaining up to and above 3,000 psi pressure during flight.

On some aircraft, the compressor is operated by an electric motor. On others, a hydraulic motor is used to drive the compressor. Compressors must be serviced with oil periodically, as outlined in the aircraft MIM. An oil level sight gauge is provided on the compressor (fig. 7-43).

AIR FILTERS. —An air filter is usually located in the line leading into the system compressor. Additional filters may be located at various points in the system lines to remove any foreign matter that may enter the system

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Like hydraulic filters, air filters have a removable element and a built-in relief valve. The relief valve is designed to open and bypass the air supply around the filter element should the element become clogged. Some air filters are equipped with the micronic-type element, which must be replaced periodically. Others have the screen mesh type, which requires periodic cleaning. The latter type may be reinstalled after cleaning and drying.

AIR PRESSURE REGULATORS. —A pressure regulator is generally located in the line between the engine compressor and the pneumatic system compressor; however, it may be incorporated within the system moisture separator. Its purpose is to regulate the pressure of the supply air before it enters the system compressor. The pressure regulator maintains a stable outlet pressure regardless of the inlet pressure.

MOISTURE SEPARATORS. —The moisture separator in a pneumatic system is always located downstream of the compressor. Its purpose is to remove any moisture caused by the compressor. A complete moisture separator consists of a reservoir, a pressure switch, a dump valve, and a check valve, and it may also include a regulator and a relief valve. The dump valve is energized and de-energized by the pressure switch. When de-energized, it completely purges the separator reservoir and lines up to the compressor. The check valve protects the system against pressure loss during the dumping cycle and prevents reverse flow through the separator.

RELIEF VALVES. —A relief valve is incorporated in a pneumatic system to protect the system from overpressurization. Overpressurization is generally caused by thermal expansion (heat). Relief valves are generally adjusted to open and close at pressures slightly above normal system operating pressure. For example, in a system designed to operate at 3,000 psi, the relief valve might be set to open at 3,750 psi and reseat at 3,250 psi.

CHEMICAL DRIERS. —Chemical driers are incorporated at various locations in a pneumatic system. Their purpose is to absorb any moisture that may collect in the lines and other parts of the system. Each drier contains a cartridge, which should be blue in color. If otherwise noted, the cartridge is to be considered contaminated with moisture and should be replaced.

STORAGE CYLINDERS. —Pneumatic storage cylinders (bottles) are made of steel and maybe either cylindrical or spherical in shape. Both types of cylinders are made up of two main parts—the container itself and a manifold assembly. The container serves as a trap for moisture, as well as an air storage space. The manifold assembly is made up of the "in" and "outlet" ports and a moisture drain fitting. See figure 7-44.

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Cooling of the high-pressure air in the storage cylinders will cause some condensation to collect in them. To ensure positive operation of systems, storage cylinders must be purged of moisture periodically. This is accomplished by slightly cracking the moisture drain fitting, located on the cylinder manifold.

Some aircraft have a pneumatic system that will maintain the required pressure in these bottles in flight. However, most of these pneumatic systems require servicing on the ground with an external source of high-pressure air or nitrogen prior to each flight.

Air storage bottles are serviced in the same manner as accumulators. Most air bottles have an air filler valve and a pressure gauge. These systems generally require higher servicing pressure than accumulators.

Since gases expand with heat and contract when cooled, air storage bottles are usually filled to a given pressure at ambient temperature. A graph similar to that shown in figure 7-45 is usually mounted on a plate or decal on or near the bottle or air filler valve. If the instruction plate is missing or not readable, the information may be found in the General Information and Servicing section of the applicable MIM.

Pressure should be added to air storage bottles slowly in order not to build up heat from rapid transfer. You should take care to ensure that air storage bottles are not overinflated

Single-Acting Actuating Cylinder

The single-acting, piston-type cylinder uses fluid pressure to apply force in only one direction. In some designs of this type, the force of gravity moves the piston in the opposite direction. However, most cylinders of this type apply force in both directions. Fluid pressure provides the force in one direction, and spring tension provides the force in the opposite direction, In some single-acting cylinders, com-pressed air or nitrogen is used instead of a spring for movement in the direction opposite that achieved with fluid pressure. A three-way directional control valve is normally used to control the operation of this type of cylinder. To extend the piston rod, fluid under pressure is directed through the port and into the cylinder. See figure 8-1. This pressure acts on the surface area of the blank side of the piston, and forces the piston to the right. This action, of course, extends the rod to the right, through the end of the cylinder. The actuated unit is moved in one direction. During this action, the spring is compressed between the rod side of the piston and the end of the cylinder. Within limits of the cylinder, the length of the stroke depends upon the desired movement of the actuated unit.

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Double-Acting Actuating Cylinder

Most piston-type actuating cylinders are double-acting, which means that fluid under pressure can be applied to either side of the piston to provide movement and apply force in the corresponding direction. One design of the double-acting, piston-type actuating cylinder is shown in view A of figure 8-2. This cylinder contains one piston and piston rod assembly. The stroke of the piston and piston rod assembly in either direction is produced by fluid pressure. The two fluid ports, one near each end of the cylinder, alternate as inlet and outlet, depending upon the "direction of flow from the directional control valve.

This is referred to as an unbalanced actuating cylinder; that is, there is a difference in the effective working areas on the two sides of the piston. Refer to view A of figure 8-2. Assume that the cross-sectional area of the piston is 3 square inches and the cross-sectional area of the rod is 1 square inch. In a 2,000 psi system, pressure acting against the blank side of the piston creates a force of 6,000 pounds (2,000 x 3). When the pressure is applied to the rod side of the piston, the 2,000 psi pressure acts on 2 square inches (the cross-sectional area of the piston less the cross-sectional area of the rod) and creates a force of 4,000 pounds (2,000 x 2). For this reason, this type of cylinder is normally installed in such a manner that the blank side of the piston carries the greater load; that is, the cylinder carries the greater load during the piston rod extension stroke.

A four-way directional control valve is normally used to control the operation of this type of cylinder. The valve can be positioned to direct fluid under pressure to either end of the cylinder, and to allow the displaced fluid to flow from the opposite end of the cylinder through the control valve to return/exhaust. The piston of the cylinder shown in view A of figure 8-2 is equipped with an O-ring seal and backup rings to prevent internal leakage of fluid from one side of the piston to the other. Suitable seals and backup rings are also used between the hole in the end cap and the piston rod to prevent external leakage. In addition, some cylinders of this type have a felt wiper ring attached to the inside of the end cap and fitted around the piston rod to guard against the entrance of dirt and other foreign matter into the cylinder.

SELECTOR VALVES

Selector valves are used in a hydraulic system to direct the flow of fluid. A selector valve directs fluid under system pressure to the desired working port of an actuating unit (double-acting), and, at the same time, directs return fluid from the opposite working port of the actuating unit to the reservoir.

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Some aircraft maintenance instruction manuals (MIMs) refer to selector valves as control valves. It is true that selector valves may be placed in this classification, but you should understand that all control valves are not selector valves. In the strict sense of the term, a selector valve is one that is engaged at the will of the pilot or copilot for the purpose of directing fluid to the desired actuating unit. This is not true of all control valves. 

Selector valves may be located in the pilot’s compartment and be directly engaged manually through mechanical linkage, or they maybe located in some part of the aircraft and be engaged by remote control. Remote-controlled selector valves are generally solenoid operated.

The typical four-way selector valve has four ports—a pressure port, a return port, and two cylinder (or working) ports. The pressure port is connected to the main pressure line from the power pump, the return port is connected to the reservoir return line, and the two cylinder ports are connected to opposite working ports of the actuating unit. 

Three general types of selector valves are discussed in this chapter. They are the poppet, slide, and solenoid-operated valves. Practically all selector valves currently in use come under one of these three general types.

Poppet-Type Selector Valve

Poppet-type selector valves are manufactured in both the balanced and unbalanced design. An unbalanced poppet selector valve offers unequal working areas on the poppets. The larger area of the poppet is in contact with the working lines of the system; consequently, when excessive pressure exists within the working lines due to thermal expansion, the poppet will open. This action allows the excessive pressurized fluid to flow into the pressure line, where it is relieved by the main system relief valve.

The balanced poppet selector valve has equal poppet areas. The poppets will remain in the selected position during thermal expansion of working line fluid. For this reason, thermal relief valves are installed in working lines that incorporate balanced poppet selector valves.

Figure 8-7 shows a typical four-port poppet selector valve. This is a manually operated valve, and consists of a group of conventional spring-loaded poppets. The poppets are enclosed in a common housing and interconnected by passageways to direct the flow of fluid in the desired direction. The poppets are actuated by cams on a camshaft, as shown in figure 8-8. They are arranged so that rotation

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of the shaft by its controlling lever will open the proper combination of poppets to direct the flow of hydraulic fluid to the desired port of the actuating unit. At the same time, fluid will be directed from the opposite port of the actuating unit, through the selector valve, and back to the reservoir.

All poppet-type selector valves are provided with a stop for the camshaft. The stop is an integral part of the shaft, and strikes against a stop pin in the body to prevent overrunning. A poppet selector valve housing usually contains poppets, poppet seats, poppet springs, and a camshaft.

When the camshaft is rotated, either clockwise or counterclockwise from neutral, the cam lobes unseat the desired poppets and allow a fluid flow. One cam lobe operates the two pressure poppets, and the other lobe operates the two return poppets. To stop the rotation of the camshaft at an exact position, a stop pin is secured to the body, and extends through a cutout section of the camshaft flange. This stop pin prevents overtravel by ensuring that the cam lobes stop rotating when the poppets have been unseated as high as they can go, where any further rotation would allow them to return to their seats.

The poppet-type selector valve has three positions-neutral and two working positions. In the neutral position, the camshaft lobes are not contacting any of the poppets. This position assures that the poppet springs will hold all four poppets firmly seated. With all poppets seated, there is no fluid flow through the valve. This action also blocks the two cylinder ports, so when this valve is in neutral, the fluid in the unit system is trapped. To allow for thermal expansion buildup, thermal relief valves must be installed in both working lines.

CHECK VALVES

The purpose of a check valve is to allow the fluid to flow in only one direction. In some installations, such as brake systems, the check valve confines fluid under pressure within the desired section of the hydraulic system. The valve prevents the fluid from reversing its normal direction of flow. The valve prevents pressure from escaping into adjacent sections of the system.

Automatic Check Valves

Automatic check valves contain a seat on which a movable body (ball, cone, or poppet) seats by means of spring tension. See figure 8-12. The valve opens when pressure in the direction of flow (indicated by an arrow on the body of the valve) is strong enough to unseat the movable body. Flow in the reverse direction, along with

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spring tension, tends to seal the movable body against the valve seat.

When the pressure on the downstream side of the valve exceeds that on the upstream side, the resultant unbalanced force seals the valve closed, as shown in view A of figure 8-12. When the pressure is reversed, the valve is forced open against the tension of the spring, and the fluid flows freely through the valve, as shown in view B of figure 8-12. The tension of the spring is relatively weak, and is intended to be barely sufficient to support the ball in its proper position.

Bypass Check Valves

Bypass check valves serve the same purpose as automatic check valves, but are so constructed that they may be opened manually to allow the flow of fluid in both directions. An example of the possible use of a bypass check valve is in the line between the hand pump and the accumulator. bypass check valve in this line would allow hand pump pressure to be directed to either the accumulator or the selector valve.

SEQUENCE VALVES

Sequence valves are used to control a sequence of operations; they ensure that actuating units operate at the proper time and in the proper sequence. Sequence valves may be mechanically operated or pressure-operated valves. An example of the use of a sequence valve is in a landing gear actuating system. In a landing gear actuating system, the landing gear doors must open before the landing gear starts to extend. Conversely, the landing gear must be retracted before the doors close. A sequence valve installed in each landing gear actuating line performs this function.

Sequence valves may be installed in one or both cylinder lines of an actuating system, depending upon the type of action desired. A direct line will go to the first unit to be operated, and a branch line goes from the sequence valve to the second unit.

SHUTTLE VALVES

All aircraft incorporate emergency systems that provide alternate methods of operating essential systems required to land the aircraft safely. These emergency systems usually provide pneumatic or hydraulic operation of the essential systems; however, in some cases due to the design, they maybe operated satisfactorily through mechanical linkage. When using the pneumatic or hydraulic emergency system, that pressure must be directed to the unit concerned; emergency pressure must not enter the normal system,

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especially if the pneumatic type system is used. To allow operating pressure to reach the actuating unit and still not enter the other system, a shuttle valve is installed in the working line to the actuating unit. The main purpose of the shuttle valve is to isolate the normal system from the emergency system.

Shuttle valves are located close to the actuating unit concerned. This location reduces to a minimum the units to be bled and isolates as much of the normal system from the emergency system as possible. In some installations, the shuttle valve is an integral part of the actuating unit.

A typical shuttle valve is shown in figure 8-15. The body contains three ports-the normal system inlet port, the emergency system inlet port, and the unit outlet port. A shuttle valve used to operate more than one actuating cylinder may contain additional unit outlet ports.

Enclosed in the body is a sliding part called the shuttle. It is used to seal one of the two inlet ports. A shuttle seat is installed at each inlet port. During operation, the shuttle is held against one of these seats, sealing off that port. These parts are held in the body by end caps. External leakage is prevented by an O-ring gasket at each end cap.

RESTRICTORS

Restrictors are used in hydraulic systems to limit the flow of hydraulic fluid to or from actuators where speed control of the cylinders is necessary to provide specific actions. If control in one direction only is desired, a one-way restrictor is used. If restricted fluid flow both to and from an actuating cylinder is necessary, a two-way restrictor is installed.

One-Way Restrictor

One-way restrictors provide reduced hydraulic flow in one direction only, to limit actuating speed of hydraulic cylinders for the purpose of proper timing or sequence of operation. Also, they provide free flow of fluid in the opposite direction to permit the actuating cylinder to actuate at a faster rate of speed during the reverse action of the cylinder.

One-way restrictors are used in some landing gear systems to regulate the speed and sequence of landing gear retraction or extension. If sequenced action (that is, one cylinder to be actuated before other cylinders on the same line) is desired, one-way restrictors are placed in the line upstream of all cylinders except one. Figure 8-16 shows both the one-way and two-way restrictors.

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The main parts of a one- way restrictor are the cylindrical body and cap, which contain a spring-loaded poppet, a cage, and a stainless steel filter element.

The one-way restrictor allows free flow in one direction and restricted flow in the opposite direction. Both directions of flow are indicated by arrows found on the body of the valve.

In a restricted direction, pressurized fluid entering port R (fig. 8-16) flows through the filter assembly and enters the cage through drilled passages. Fluid from the interior of the cage is forced through the poppet’s orifice, thus causing the required metering action. In the free flow direction, pressurized fluid entering port F overcomes poppet spring tension and allows fluid to flow past the poppet’s seat, through drilled passages within the larger flange of the cage, and out through port R.

Two-Way Restrictor

Two-way restrictors are used to limit the flow of hydraulic fluid where it is desirable to retard the action of a hydraulic cylinder in both directions. Figure 8-16 shows two types of two-way restrictors, one of which has a machined orifice with two integral stainless steel filters. The other type shown contains an orifice plate between two stainless steel filters. The filters contained within the restrictors are identical in construction and provide protection in both directions of flow. The filter size specification for the two-way restrictor is identical to those found within one-way restrictors.

Two-way restrictors, regardless of whether they are of the machined orifice type or of the plate orifice type, operate identically. Fluid entering either port is filtered prior to flowing through the orifice, thus protecting the orifice from possible stoppage. As the fluid is metered through the orifice, the prescribed rate flow is directed out the opposite port of the restrictor and to the actuating unit.

PRESSURE-REDUCING VALVES

Pressure-reducing valves are used in hydraulic systems where it is necessary to lower the normal system operating pressure a specified amount. Figure 8-17 shows the operation of a pressure-reducing valve. View A of figure 8-17 shows system pressure being ported to a subsystem through the shuttle and sleeve assembly. Subsystem pressurized fluid works on the large flange area of the shuttle, which causes the shuttle to move to the left after reaching a specified pressure, thus closing off the normal system. The valve will stay in this position until the subsystem pressure is lowered, at which time the shuttle will move to its prior position and allow the required amount of pressurized fluid to enter the subsystem. During normal

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operation of the subsystem, the pressure-reducing valve continuously meters fluid to the subsystem.

HYDRAULIC FUSES 

A hydraulic fuse is a safety device. Fuses may be installed at strategic locations throughout a hydraulic system. They are designed to detect line or gauge rupture, fitting failure, or other leak-producing failure or damage.

One type of fuse, referred to as the automatic resetting type, is designed to allow a certain volume of fluid per minute to pass through it. If the volume passing through the fuse becomes excessive, the fuse will close and shut off the flow. When the pressure is removed from the pressure supply side of the fuse, it will automatically reset itself to the open position. Fuses are usually cylindrical in shape, with an inlet and outlet port at opposite ends, as shown in figure 8-18. A stationary sleeve assembly is con-tained within the body. Other parts contained within the body, starting at the inlet port, are a control head, piston and piston subassembly stop rod, a lock spring, and a lock piston and return spring.

Fluid entering the fuse is divided into two flow paths by the control head. The main flow is between the sleeve and body, and a secondary flow is to the piston. Fluid flowing through the main path exerts a force on the lock piston, causing it to move away from the direction of flow, This movement uncovers ports, allowing fluid to flow through the fuse.

The movement of the locking piston also causes a lock spring to release the piston subassembly stop rod, thus allowing the piston to be displaced by fluid from the secondary flow. If the flow through the fuse exceeds a specified amount, the piston, moving in the direction of flow, will block the ports originally covered by the locking piston, thus blocking the flow of fluid.

Any interruption of the flow of fluid through the fuse removes the operating force from the lock piston. This allows the lock piston spring to return the piston to the original position, which resets the fuse.

Proof-Testing Cables

All newly fabricated cables should be tested for proper strength before they are installed in aircraft The test consists of applying a specified tension load on the cable for a specified number of minutes. The proof loads for testing various size cables are given in

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tables contained in NAVAIR 01-1A-8. Proof loading will result in a certain amount of permanent stretch being imparted to the cable. This stretch must be taken into account when You fabricate cable assemblies. Cables that are made up long slightly may be entirely too long after proof loading

AIRCRAFT WHEELS, TIRES, AND TUBES

Aircraft wheels are made from either aluminum or magnesium alloys. These materials provide a strong, lightweight wheel that requires very little maintenance. The wheels used on naval aircraft are of two general types—divided and remountable flange. Both of these designs make wheel buildup a fairly simple operation.

The wheels used with tires and tubes have knurled flanges to prevent the tire from slipping on the wheel. Wheels used with tubeless tires have the wheel sections sealed by an O-ring, and they use special valves that are a part of the wheel.

DIVIDED (SPLIT) WHEEL

Figure 11-1 shows a typical divided (split) wheel. This type of wheel is divided into two halves. The two halves are sealed by an O-ring and held together with nuts and bolts. Each wheel half is statically balanced. This procedure allows any two opposite halves of the same size and type to be joined together to form one wheel assembly. If the outboard half of a wheel is beyond repair, a new outboard half may be drawn from supply. The new outboard half is then matched to the old inboard half. This type of wheel is used on nose, main, and tail landing gears.

REMOUNTABLE FLANGE WHEEL

The remountable flange wheel is made so one flange of the wheel can be removed to change the tire. The flange is held in place by a lockring. The wheel is balanced with the flange mounted on the wheel. Then, both the wheel and flange are marked. To ensure proper balance of the wheel during assembly, the two marks should be lined up. Figure 11-2 shows a typical remountable flange wheel. This type of wheel is commonly used on the main landing gear.

The similarity of one wheel to another in size and shape is not proof that the wheels can be inter-changed. One wheel may be designed for heavy duty while the other may be designed to carry a lighter load. Also, the wheels may be designed for use with different types of brake assemblies.

TYPICAL WHEEL ASSEMBLY

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A complete wheel assembly is shown in figure 11-3. The wheel casting is the basic unit of the wheel assembly. It is to this part that all other com-ponents are assembled and upon which the tire is mounted.

Figure 11-1.—Typical divided (split) wheel assembly.

The demountable flange is attached to the wheel to simplify tire removal and installation. The remountable flange lockring secures the flange to the wheel. The flange is fitted into a groove in the wheel casting.

The bearing cups are shrink-fitted into the hub of the wheel casting, and are the parts on which the bearings ride. The bearings are tapered roller bearings. Each bearing is made of a cone and rollers. This type of bearing absorbs side thrust as well as radial loads and landing shocks. These bearings must be cleaned and lubricated in accordance with the NAVAIR 04-10-1 manual.

A three-piece grease retainer keeps the grease in the inboard bearing and keeps out dirt and moisture. It is composed of a felt seal

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and inner and outer closure rings. A lockring secures the assembly inside the wheel hub.

The hubcap seals the outboard side of the hub. It is secured with a lockring. On some aircraft, the hubcap is secured with screws.

All wheels designed to be used on the main landing gear are equipped with braking components. These components are attached to the wheel casting. They may consist of either a brake drum or brake drive keys. The wheel shown in  figure 11-3 is

TIRE CONSTRUCTION

Figure 11-10 shows the construction details of a tube-type aircraft tire. Tubeless tires are similar to tube tires except they have a rubber inner liner that is mated to the inside surface of the tire. The rubber liner helps retain air in the tire. The beaded area of a tubeless tire is designed to form a seal with the wheel flange. Wear indicators have been built into some tires as an aid in measuring tread wear. These indicators are holes in the tread area or lands in the bottom of the tread grooves.

The cord body consists of multiple layers of nylon with individual cords arranged parallel to each other and completely encased in rubber. The cord fabric has its strength in only one direction. Each layer of coated fabric constitutes one ply of the cord body. Adjacent cord plies in the body are assembled with the cords crossing at nearly right angles to each other. This arrangement provides a strong and flexible tire that distributes impact shocks over a wide area. The functions of the cord body are to give the tire tensile strength, to resist internal pressures, and to maintain tire shape.

The tread is a layer of rubber on the outer surface of the tire. It protects the cord body from abrasion, cuts, bruises, and moisture. It is the surface that contacts the ground.

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Figure 11-10.–Sectional view of aircraft tire showing construction details.

The sidewall is an outer layer of rubber adjoining the tread and extending to the beads. Like the tread, it protects the cord body from abrasion, cuts, bruises, and moisture.

The beads are multiple strands of high-tensile strength steel wire imbedded in robber and wrapped in strips of open weave fabric. The beads hold the tire firmly on the rims and serve as an anchor for the fabric plies that are turned up around the bead wires. The chafing strips are one or more plies of rubber-impregnated woven fabric wrapped around the outside of the beads. They provide additional rigidity to the bead and prevent the metal wheel rim from chafing the tire. Tubeless tires have an additional ply of rubber over the chafing strips to function as an air seal.

The breakers are one or more plies of cord or woven fabric impregnated with rubber. They are used between the tread rubber and the cord body to provide extra reinforcement to prevent bruise damage to the tire. Breakers are not part of the cord body.

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Tread Patterns

There are three tread patterns or tread designs used on naval aircraft. They are plain, ribbed, and nonskid. A plain tread has a smooth, uninterrupted surface. A ribbed tread has three or more continuous circumferential ribs separated by grooves. A nonskid tread is any grooved or ribbed tread. Other tread designs may be provided under specific circum-stances or as required by applicable MS standards or drawing. The most common design used on naval aircraft is the ribbed pattern.

Tread Construction

The tread construction will usually be one of four types. Other tread types may be necessary for specific circumstances or as required by military standards, such as ice and snow treads.

NOTE: Additional safety precautions are required in handling ice and snow treads.

• Rubber tread. A rubber tread is constructed from 100-percent new (no reclaim) rubber. It maybe new natural rubber, new synthetic material, or a blend of new material and new synthetic materials.

• Cut-resistant tread. A cut-resistant tread has improved cut-resistant properties that are imparted to the tire by incorporating a barrier into the undertread that resists penetration of cutting objects.

• Reinforced tread. A reinforced tread is constructed with fabric cord or other reinforcing materials as an integral part of the tread. See figure 11-11.

• Reinforced cut-resistant tread. A reinforced cut-resistant tread combines the features of both the cut-resistant and reinforced-tread designs.

Ply Rating

Reference to the number of cord fabric plies in a tire has been superseded by the term ply rating. This term is used to identify a tire’s maximum recom-mended load for specific types of service. It does not necessarily represent the number of cord fabric plies in a tire. Most nylon cord tires have ply ratings greater than the actual number of fabric plies in the cord body.

Size Designation

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Figure 11-12 shows the points of measurement used to designate the size of a tire. For example, a tire with a size designation of 26 X 6.6 would have an outside diameter (measurement A) of 26 inches and a cross-sectional width (measurement B) of 6.6 inches. The letter X merely separates the two measurements. If the tire’s size designation were 26 X 6.6-10, then the tire would have a rim diameter (measurement C) of 10 inches. If only one numerical designation is used for a tire, you should assume that it is the outside diameter (measurement A).

Standard Identification Markings

You should be familiar with the markings on the sidewall of a tire. You will need this information to complete a VIDS/MAF for a tire change. The

Figure 11-12.—Size designation of tires.

markings engraved or embossed on a sidewall are shown in figure 11-13. 

Most of the markings are self-explanatory, Item 10 has a maximum of 10 characters. The first four positions show the date of manufacture in the form of a Julian date (last digit of the year followed by the day of the year, or 17 Oct 1985 = 5290). The next

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positions are completed by the manufacturer and are either numbers or letters. They are used to create a unique serial number for a particular tire. The cut limit (11 ) is expressed in thirty-seconds of an inch and

Figure 11-13.—New tire identification markings.

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is used to evaluate the depth of cuts in the thread area. Tires are marked with a red dot (14) on the sidewall to indicate the lightweight (balance) point of the tire.

TIRE STORAGE

The life of a tire, whether mounted or unmounted, is directly affected by storage conditions. Tires should always be stored indoors in a dark, cool, dry room. It is necessary to protect them from light, especially sunlight. Light causes ultraviolet (UV) damage by breaking down the rubber compounds. The elements, such as wind, rain, and temperature changes, also break down the rubber compounds. Damage from the elements is visible in the form of surface cracking or weather checking. UV damage may not be visible. Tires can be protected from light by painting the storeroom windows. Tires must not be allowed to come in contact with oils, greases, solvents, or other petroleum products that cause rubber to soften or deteriorate. The storeroom should not contain fluorescent lights or sparking electrical equipment that could produce ozone.

Tires should be stored vertically in racks and according to size. See figure 11-15. The edges of the racks must be smooth so the tire tread does not rest on a sharp edge. Tires must never be stacked in horizontal piles. The issue of tires from the storeroom should be based on age from the date of manufacture so the older tires will be used first. This procedure helps to prevent the chance of deterioration of the older tires in stock.

TIRE INSPECTION

There are two types of inspections conducted on tires. One is conducted with the tire mounted on the wheel. The other inspection is conducted with the tire dismounted.

Mounted Inspection

During each daily or special inspection, tires must be inspected for correct pressure, tire slippage on the wheel (tube tires), cuts, wear, and general condition. Tires must also be inspected before each flight for obvious damage that may have been caused during or after the previous flight.

Maintaining the correct inflation pressure in an aircraft tire is essential to safety and to obtain its maximum service life. Military aircraft inner tubes and tubeless tire liners are made of natural rubber to satisfy extreme low-temperature performance requirements. Natural rubber is a relatively poor air retainer. This accounts for the daily inflation pressure loss and the need for frequent pressure checks. If this check discloses more than a normal

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loss of pressure, you should check the valve core for leakage byputting a small amount of suitable leak detection solution (Leaktec) or soapy water on the end of the valve and watch for bubbles. Replace the valve core if it is leaking. If no bubbles appear, it is an indication that the inner tube (or tire) has a leak. When the tire and wheel assembly shows repeated pressure loss exceeding 5 percent of the correct operating inflation pressure, it should be removed from the aircraft and sent to the AIMD or IMA.

WARNING

Overinflation or underinflation can cause catastrophic failure of aircraft tire and wheel assemblies. This could result in injury, death, and/or damage to aircraft or other equipment.

After making a pressure check, you should always replace the valve cap. Be sure that it is screwed on fingertight. The cap prevents moisture, salt, oil, and dirt from entering the valve stem and damaging the valve core. It also acts as a secondary seal if a leak develops in the valve core.

Tires that are equipped with inner tubes, and operate with less than 150 psi, and all helicopter tube tires must use tire slippage marks. The slippage mark is a red paint strip 1 inch wide and 2 inches long. It extends equally across the tire sidewall and the wheel rim, as shown in figure 11-16. Tires should be inspected for slippage on the rim after each flight. If the markings do not align within one-fourth of an inch, the wheel assembly should be replaced and the defective assembly forwarded to the AIMD or IMA for repair. Failure to correct tire slippage may cause the valve stem to be ripped from the tube.

Tire treads should be inspected to determine the extent of wear. The maximum allowable thread wear for tires without wear depth indicators is when the tread pattern is worn to the bottom of the tread groove at any spot on the tire. The maximum allowable tread wear for tires with tread wear indicators is when the tread pattern is worn either to the bottom of the wear depth indicator or the bottom of the tread groove. These limits apply regardless of whether the wear is the result of skidding or normal use.

The tread and sidewall should be examined for cuts and embedded foreign objects. Figure 11-17 shows the method for measuring the depth of cuts, cracks, and holes. Glass, stones, metal, and other materials embedded in the tread should be removed to prevent cut growth and eventual carcass damage. A blunt awl or screwdriver maybe used for this purpose. You should be careful to avoid enlarging the hole or damaging the cord body fabric.

WARNING

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When you are probing for foreign objects, be sure you keep the probe from penetrating deeper into the tire. Objects being pried from the tire frequently are ejected suddenly and with considerable force. To avoid eye injury, safety glasses or a face shield should be worn. A gloved hand over the object may be used to deflect it.

Aircraft should not be parked in areas where the tires may stand in spilled hydraulic fluids, lubricating oils, fuel, or organic solvents. If any of these materials is accidentally spilled on a tire, it should be immediately wiped with a clean, absorbent cloth. The tires should then be washed with soap and thoroughly rinsed with water.

Extra care should be taken when you inspect mounted helicopter tires. Because of the long intervals between tire changes, helicopter tires are subject to weather and UV damage.

Nylon Flat Spotting

If the aircraft stands in one place under a heavy static load for several days, local stretching may cause an out-of-round condition with a resultant thumping during takeoff and landing.

Dual Installations

On dual-wheel installations, tires should be matched according to the dimensions indicated in table 11-1. Tires vary somewhat in size between manufacturers and can vary a great deal after being used. When two tires are not matched, the larger one supports most or all of the load. Since one tire is not designed to carry this increase in load, a failure may result.

CHAPTER 12

LANDING GEAR, BRAKES, AND HYDRAULIC UTILITY SYSTEMS

Main Landing Gear

The typical aircraft landing gear assembly consists of two main landing gears and one steerable nose landing gear. As you can see in figure 12-1, a main gearis installed under each wing. Because aircraft are different in size, shape, and construction, every landing gear is specially designed. Although main landing gears are designed differently, all main gear struts are attached to strong members of the wings or fuselage so that the landing shock is distributed throughout the main body of the structure. The main gears are also equipped with brakes that are used to shorten the landing roll of the aircraft and to guide the aircraft during taxiing.

Nose Landing Gear

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On aircraft with tricycle landing gear, the nose gear is retracted either rearward or forward into the aircraft fuselage. Generally, the nose gear consists of a single shock strut with one or two wheels attached. On most aircraft the nose gear has a steering mechanism for taxiing the aircraft. The mechanism also acts as a shimmy damper to prevent oscillation or shimmy of the nosewheel. Since the nosewheel must be centered before it can be retracted into the wheel well, a centering device aligns the strut and wheel when the weight of the aircraft is off the gear. Damping, steering, and centering devices are discussed later in this chapter.

SHOCK STRUTS

Shock struts are self-contained hydraulic units. They carry the burden of supporting the aircraft on the ground and protecting the aircraft structure by absorbing and dissipating the tremendous shock of landing. Shock struts must be inspected and serviced regularly for them to function efficiently. This is one of your important responsibilities.

Each landing gear is equipped with a shock strut. In addition to the landing gear shock struts, carrier aircraft are equipped with a shock strut on the arresting gear. The shock strut’s primary purpose is to reduce arresting hook bounce during carrier landings.

Because of the many different designs of shock struts, only information of a general nature will be included in this chapter. For specific information on a particular installation, you should refer to the applicable aircraft MIM or accessories manual.

A typical pneumatic/hydraulic shock strut (metering pin type) is shown in figure 12-8. It uses compressed air or nitrogen combined with hydraulic fluid to absorb and dissipate shock, and it is often

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Figure 12-8.–Landing gear shock strut (metering pin type).

referred to as the "air-oil" type strut. This particular strut is designed for use on the main landing gear. As shown in the illustration, the shock strut is essentially two telescoping cylinders or tubes, with externally closed ends. When assembled, the two cylinders, known as cylinder and piston, form an upper and lower chamber for movement of the fluid. The lower chamber is always filled with fluid, while the upper chamber contains compressed air or nitrogen. An orifice (small opening) is placed between the two chambers. The fluid passes through this orifice into the upper chamber during compression, and returns during extension of the strut.

Most shock struts employ a metering pin similar to that shown in figure 12-8 to control the rate of fluid flow from the lower chamber into the upper chamber. During the compression stroke, the rate of

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fluid flow is not constant, but is controlled automatically by the variable shape of the metering pin as it passes through the orifice.

 

Figure 12-9.–Landing gear shock strut (metering tube type).

On some types of shock struts now in service, a metering tube replaces the metering pin, but shock strut operation is the same. An example of this type of shock strut is shown in figure 12-9.

Some shock struts are equipped with a dampening or snubbing device, which consists of a recoil valve on the piston or recoil tube. The purpose of the snubbing device is to reduce the rebound during the extension stroke and to prevent a too rapid extension of the shock strut, which would result in a sharp impact at the end of the stroke.

The majority of shock struts are equipped with an axle that is attached to the lower cylinder to provide for tire and wheel installation. Shock struts not equipped with axles have provisions on the end of the lower cylinder for ready installation of the axle

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assembly. Suitable connections are also provided on all shock struts to permit attachment to the aircraft.

A fitting, which consists of a fluid filler inlet and a high-pressure air valve, is located near the upper end of

 

Figure 12-10.–Nose gear shock strut.

each shock strut to provide a means of filling the strut with hydraulic fluid and inflating it with air or nitrogen.

A packing gland designed to seal the sliding joint between the upper and lower telescoping cylinders is installed in the open end of the outer cylinder. A packing gland wiper ring is also installed in a groove in the lower bearing or gland nut on most shock struts to keep the sliding surface of the piston or inner cylinder free from dirt, mud, ice, and snow. Entry of foreign matter into the packing gland will result in leaks. The majority of shock struts are equipped with torque arms attached to the upper and lower cylinders to maintain correct alignment of the wheel.

Nose gear shock struts are provided with an upper centering cam that is attached to the upper cylinder and a mating lower centering cam that is attached to the lower cylinder. See figure 12-10. These

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cams serve to line up the wheel and axle assembly in the straight-ahead position when the shock strut is fully extended. This prevents the nosewheel from being cocked to one side when the nose gear is retracted, preventing possible structural damage to the aircraft. These mating cams

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Figure 12-11.–Shock strut operation.

also keep the nosewheel in a straight-ahead position prior to landing when the strut is fully extended. Some nose gear shock struts have

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the attachments for installation of an external shimmy damper, which is discussed later in this chapter.

Nose and main gear shock struts are usually provided with jacking points and towing lugs. Jacks should always be placed under the prescribed points. When towing lugs are provided, the towing bar should be attached only to these lugs. 

All shock struts are provided with an instruction plate that gives, in a condensed form, instructions relative to the filling of the strut with fluid and inflation of the strut. The instruction plate also specifies the correct type of hydraulic fluid to use in the strut. The plate is attached near the high-pressure air valve. It is of the utmost importance that you always consult the applicable aircraft MIMs and familiarize yourself with the instructions on the plate prior to servicing a shock strut with hydraulic fluid and nitrogen or air.

Figure 12-11 shows the inner construction of a shock strut and the movement of the fluid during compression and extension of the strut. The com-pression stroke of the shock strut begins as the aircraft hits the ground. The center of mass of the aircraft con-tinues to move downward, compressing the strut and sliding the inner cylinder into the outer cylinder. The metering pin is forced through the orifice, and by its variable shape, controls the rate of fluid flow at all points of the compression stoke. In this manner, the greatest possible amount of heat is dissipated through the walls of the shock strut. At the end of the downward stroke, the compressed air or nitrogen is further compressed, limiting the compression stroke of the strut. If there is an insufficient amount of fluid and/or air or nitrogen in the strut, the compression stroke will not be limited, and the strut will "bottom" out, resulting in severe shock and possible damage to the aircraft.

The extension stroke occurs at the end of the compression stroke, as the energy stored in the compressed air or nitrogen causes the aircraft to start moving upward in relation to the ground and wheels. At this instant, the compressed air or nitrogen acts as a spring to return the strut to normal. At this point, a snubbing or dampening effect is produced by forcing the fluid to return through the restrictions of the snubbing device (recoil valve). If this extension were not snubbed, the aircraft would rebound rapidly and tend to oscillate up and down because of the action of the compressed air. A sleeve, spacer, or bumper ring incorporated in the strut limits the extension stroke.

MECHANICAL LINKAGE

The landing gear drag brace (fig. 12-12) consists of an upper and lower brace that is hinged at the center to

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Figure 12-12.–Landing gear drag brace adjustment. 

permit the brace to jackknife during retraction of the gear. The upper brace pivots on a trunnion attached to the wheel well overhead. The lower brace is connected to the lower portion of the shock strut outer cylinder. On the drag brace shown in figure 12-12, a locking mechanism is used where the lower and upper drag braces meet. Usually in this type of installation, the locking mechanism is adjusted so that it is allowed to be positioned slightly overcentered. You must be able to inspect and adjust landing gear braces and lccking mechanisms as specified in the applicable MIM.

To adjust the drag brace shown in figure 12-12, you would first remove the cotter pin and nut (not shown) from the lock arm shaft. With the drag brace in the full extended position, rotate the eccentric bushings that are located on each end of the lock arm shaft.

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Both bushings must be rotated together to ensure that the high point of the eccentricity is the same on both bushings. Failure to do this may result in damage to the equipment or sluggish operation. The bushings maybe rotated in either direction until the end of the leek arm shaft, shown as point "A" in figure 12-12, is a distance of 0.003 inch to 0.015 inch from the striker. This clearance is checked with a feeler gauge.

Other portions of the drag brace are nonadjustable, except for the length of its down leek cylinder. Figure 12-12 indicates the cylinder should be adjusted to a length of 12 3/8 inches.

In the design of drag braces, the tendency has been directed toward lessening the adjustment requirements. In some installations, drag braces are manufactured to exact dimensions and do not require adjustments.

Bleeding the System

Bleed the system every time you replace a part or disconnect a line. Clear the nose gear from the deck with the hydraulic and electrical power connected. Depress the nose gear steering switch and operate the rudder pedals. As the nose gear steering cylinder moves, open and close the extend and retract bleed ports. Do the same with the relief valve bleed port at the steering cylinder until the hydraulic fluid is free of air. Cycle the steering system five complete cycles. Secure the bleed ports and lockwire. Disconnect electrical and hydraulic power and remove the jack

SERVICING, BLEEDING, AND INSPECTING SHOCK STRUTS

For efficient operation of shock struts, the proper fluid level and pneumatic pressure must be maintained. Before you check the fluid level, you should consult the aircraft MIM. Deflating a strut can be a dangerous operation unless the servicing personnel are thoroughly familiar with high-pressure air valves and observe all the necessary safety precautions.

Servicing

The high-pressure air valve shown in figure 12-18 is used on most naval aircraft. This air valve is used on struts, accumulators, and various other components that must be serviced with high-pressure air or nitrogen. The following procedures for deflating a typical shock strut, servicing with hydraulic fluid, and reinflating is for instructional purposes only. See figure 12-19. For specific aircraft consult the appropriate aircraft MIM.

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1. Position the aircraft so that the shock struts are in the normal ground operating position. Ensure that personnel workstands, and other obstacles are clear of the aircraft.

NOTE: Some aircraft must be placed on jacks with their struts completely extended for servicing.

2. Remove the cap from the air valve, as shown in view A of figure 12-19.

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Figure 12-19.-Servicing a landing gear strut.

 

 

Figure 12-20.-Landing gear strut servicing instruction plate.

3. Release the air pressure in the strut by slowly turning the air valve swivel nut counterclockwise approximately 2 turns. This action can normally be accomplished with the use of a combination wrench.

WARNING

When loosening the swivel nut ensure that the 3/4-inch hex body nut is either lockwired in place or held tightly with a wrench. If the swivel nut is loosened before the air pressure has been released, serious injury may result to personnel.

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4. Ensure that the shock strut compresses as the air or nitrogen pressure is released. In some cases, it may be necessary to rock the aircraft after deflating to ensure complete compressing of the strut.

5. When the strut is fully compressed, the air valve assembly may be removed by breaking the safety wire and turning the 3/4-inch body nut counter-clockwise.

6. Use the type of hydraulic fluid specified on the shock strut inspection plate to fill the strut to the level of the air valve opening. Figure 12-20 shows the instruction plate found on one type of aircraft main landing gear strut.

NOTE: The instruction plate may be found on the strut or on the wheel door near the strut. Improper oil level in the strut chamber will decrease the shock absorbing capabilities of the strut and could cause the strut to bottom out during landing. This would damage the strut and/or wing structure.

7. Reinstall the air valve assembly, using a new O-ring packing. Torque the air valve body hex nut from 100 inch-pounds to 110 inch-pounds, as shown in view B of figure 12-19.

8. Lockwire the air valve assembly to the strut, using the holes provided in the body nut.

9. Inflate the strut, using a regulated high-pressure source of nitrogen or dry air. Under no circumstances should any type of bottle gas other than nitrogen or compressed air be used to inflate shock struts. The amount a strut is inflated depends upon the specific aircraft strut being serviced. One manufacturer may use a strut inflation chart, such as the one shown in view D of figure 12-19. The strut is measured as indicated at dimension "A." This measurement, in inches, is then located on the bottom of the inflation chart. For example, locate the measurement of 1.75 inches on the chart. From this point, vertically trace an imaginary line until it intersects the curved line. At this point of intersection, horizontally trace a second imaginary line to the left edge of the chart. The figure indicated at this point (550 psi) is the required pressure for that particular extension of the strut.

All aircraft struts are not measured from the same points. View E of figure 12-19 shows another location where strut extension is measured. The proper procedure to use will always be found on the instruction plate attached to the shock strut. If these instructions are not legible, consult the applicable MIM.

If the strut’s chamber is underpressurized, the strut may not overcome normal O-ring friction during extension on takeoff. This

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condition could prevent the strut from fully extending, thus the torque scissors limit switch would not actuate to close the electrical circuit to retract the gear. It would also cause the strut to bottom during taxiing and landing operations. 

If the strut’s chamber is overpressurized, the additional pressure will tend to keep the strut pressurized after takeoff. On those aircraft that use shrink mechanisms, the shrink mechanisms may be overloaded or stall the strut actuator as the gear retracts. If the gear retracts in the wing without shrinking, due to the failure of the shrink mechanism, damage to both the wing and landing gear may result.

10. Tighten the air valve swivel hex nut to a recommended torque of 50-70 inch-pounds.

11. Remove the high-pressure air-line chuck and install the valve cap fingertight

Because some aircraft struts require special servicing procedures, the General Information and Servicing section of the applicable MIM should always be checked before servicing the shock struts of any aircraft.

Bleeding

If the fluid level of a shock strut has become extremely low or, if for any other reason, air is trapped in the strut cylinder, it may be necessary to bleed the strut during the servicing operation. Bleeding is performed with the aircraft placed on jacks. In this position, the shock struts can be extended and compressed during the filling operation, expelling all of the entrapped air. As mentioned earlier, certain aircraft must be placed on jacks for routine servicing of the shock struts. The following is a typical bleeding procedure.

1. Construct a bleed hose that contains a fitting suitable for making an airtight connection to the shock strut filler opening. The hose should be long enough to reach from the shock strut tiller opening to the deck when the aircraft is on jacks.

2. Jack the entire aircraft until all shock struts are fully extended.

3. Release the air or nitrogen pressure in the strut to be bled, as previously described in this chapter.

4. Remove the air tiller valve assembly.

5. Fill the strut to the level of the filler port with hydraulic fluid.

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6. Attach the bleed hose to the filler port, and insert the opposite end of the hose into a quantity of clean hydraulic fluid.

7. Place an exerciser jack or other suitable single-base jack under the shock strut jacking point. See view C of figure 12-19. Compress and extend the strut fully (by raising and lowering the jack) until the flow of air bubbles from the strut has completely stopped.

NOTE: Compress the strut slowly and allow it to extend by its own weight.

8. Remove the exerciser jack, and then lower and remove all other jacks.

9. Remove (he bleed hose from the shock strut.

10. Install the air tiller valve and inflate the strut, as previously described.

Inspection

Shock struts should be inspected regularly for leakage of fluid and for proper extension. Exposed portions of the strut pistons should be cleaned in the same manner as actuating cylinder pistons during preflight and postflight inspections. Exposed pistons should be inspected closely for scoring and corrosion. Excessive leakage of fluid can usually be stopped by deflating the strut and tightening the packing gland nut. If leakage still persists after tightening the packing gland

 

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Figure 12-21.-Landing gear shock strut tools.

nut and reinflating the strut, the strut must be dis-assembled and the packings replaced.

The tools shown in figure 12-21 are typical of the tools used during disassembly and assembly of landing gear shock struts. Normally, each tool is designed for, and should be used only on, one type of installation. When using wrenches, you must take care to maintain the lugs of the wrenches in their respective positions.

Slippage of the wrench, when under torquing conditions, may cause damage to aircraft parts, the tool, or even injury to personnel. NEVER place extension handles of any type on these tools to increase the applied force.

These tools, like other special tools, should be kept where they will not be subjected to rough handling, which could cause mushroomed or deformed surfaces, making them useless for aircraft repair. Shock strut disassembly and replacement of packings is a requirement for advancement to first class; therefore, it is not covered in this training manual

Viscosity,

More viscous a fluid more difficult for it to flow.We understand viscosity as a property that tends to retard fluid motion It has the

dimensions and units of in the SI system. Viscosity of a fluid is strongly dependent on temperature and is a weak function of pressure. For example, when the pressure of air is increased from 1 atm to 50 atm, its viscosity increases only by about 10 percent allowing one to ignore its dependence on pressure. It is seen that the viscosity of liquids deceases with temperature while that for the gases increases with temperature. This difference in behaviour is explained by the cohesive and intermolecular forces within the fluid. Liquids are characterized by strong cohesive forces and close packing of molecules. When temperature increases cohesive forces are weakened and there is less resistance to motion. Hence viscosity decreases. With gases, the cohesive forces are very weak and the molecules are spaced apart. Viscosity is due to the exchange of momentum between molecules as a result of random motion. As the temperature increases the molecular activity increases giving rise to an increased resistance to motion or in other words viscosity increases.

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Kinematic Viscosity,

In fluid flow problems viscosity often appears in combination with density in the form

(1.12)

One of the common examples is Reynolds Number, defined as VL/ being one of the very important parameters in Fluid Dynamics.This term is referred to as Kinematic Viscosity and has the dimensions

of Viscosity (dynamic viscosity)

The SI physical unit of dynamic viscosity is the pascal-second (Pa·s), which is identical to 1 N·s/m2 or 1 kg/(m·s).

The cgs physical unit for dynamic viscosity is the poise (P). It is more commonly expressed, particularly in ASTM standards, as centipoise (cP). The centipoise is commonly used because water has a viscosity of 1.0020 cP (at 20 °C; the closeness to one is a convenient coincidence).

1 poise = 100 centipoise = 1 g/(cm·s) = 0.1 Pa·s.

1 centipoise = 1 mPa·s.

Kinematic viscosity

The SI physical unit of kinematic viscosity is the (m2/s). The cgs physical unit for kinematic viscosity is the stokes (abbreviated S or St). It is sometimes expressed in terms of centistokes (cS or cSt).

1 stokes = 100 centistokes = 1 cm2/s = 0.0001 m2/s.

Gases

Viscosity in gases arises principally from the molecular diffusion that transports momentum between layers of flow. The kinetic theory of gases allows accurate prediction of the behaviour of gaseous viscosity, in particular that, within the regime where the theory is applicable:Viscosity is independent of pressure; and Viscosity increases as temperature increases.

Liquids

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In liquids, the additional forces between molecules become important. This leads to an additional contribution to the shear stress though the exact mechanics of this are still controversial. Thus, in liquids:Viscosity is independent of pressure (except at very high pressure); and Viscosity tends to fall as temperature increases

The dynamic viscosities of liquids are typically several orders of magnitude higher than dynamic viscosities of gases

Density,

Density is defined as mass per unit volume of the substance. Unit of

density in the SI system is Under ordinary conditions the

density of water is while that for air at C and

atmospheric pressure is .

Density of liquids is somewhat insensitive to the changes in pressure and temperature. For gases there is a strong dependence of density on these quantities and is given by the equation of state of the particular gas.

Specific Volume, v

Specific Volume, v of a fluid is defined as the volume per unit mass and its numerical value is given by the reciprocal of density.

Specific Weight,

Specific Weight, of a fluid is defined as the weight per unit volume and is related to density.

(1.15)

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where g is acceleration due to gravity. Its units (in SI units) are

Specific Gravity, SG

Specific Gravity, SG of a fluid is the ratio of its density to that of

water under reference conditions, usually at (i.e., .)

Pressure, p

Pressure is the normal force per unit area exerted on the plate. Dimensions of pressure are F/L2 which is also called a Pascal. Pressure values read by measuring devices such as a manometer are the pressure levels above the atmospheric pressure and are called gauge pressure. When pressure is referred to a vacuum it becomes an Absolute Pressure being the sum of the gauge pressure and atmospheric pressure Pressure is the application of force to a surface, and the concentration of that force in a given area.

More formally, pressure (symbol: p) is defined as the magnitude of the normal force per unit area.

p = F / A

where p is the pressure, F is the normal force, and A is the area. Pressure is transmitted to solid boundaries or across arbitrary sections of fluid normal to these boundaries or sections at every point. Unlike stress, pressure is defined as a scalar quantity.

The gradient of pressure is force density.

Pressure is sometimes measured not as an absolute pressure, but relative to atmospheric pressure; such measurements are sometimes called gauge pressure.

Atmospheric pressure is the pressure above any area in the Earth's atmosphere caused by the weight of air. As elevation increases, fewer air molecules are above. Therefore, atmospheric pressure decreases with increasing altitude.

Hydrostatic pressure

Hydrostatic pressure is the pressure due to the weight of a fluid.

p = ρgh

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where ρ (rho) is density of the fluid, g is acceleration due to gravity, and h is height of the fluid above the point being measured. See also Pascal's Law.

Stagnation pressure

Stagnation pressure is the pressure a fluid exerts when it is motionless. Consequently, although a fluid moving at higher speed will have a lower static pressure, it may have a higher stagnation pressure. Static pressure and stagnation pressure are related by the Mach number of the fluid. In addition, there can be differences in pressure due to differences in the elevation (height) of the fluid. See Bernoulli's equation.

The pressure of a moving fluid can be measured using a Pitot probe, or one of its variations such as a Kiel probe or Cobra probe, connected to a manometer. Depending on where the inlet holes are located on the probe, it can measure static pressure or stagnation pressure.

Units

The SI unit for pressure is the pascal (Pa), equal to one newton per square metre (N·m-2 or kg·m-1·s-2). This special name for the unit was added in 1971; before that, pressure in SI was expressed in units such as N/m².

Non-SI measures (still in use in some parts of the world) include the pound-force per square inch (psi) and the bar.

Some meteorologists prefer the hectopascal (hPa) for atmospheric air pressure, which is equivalent to the older unit millibar (mbar). Similar pressures are given in kilopascals (kPa) in practically all other fields, where the hecto prefix is hardly ever used. In Canadian weather reports, the normal unit is kPa. The obsolete unit inch of mercury (inHg) is still sometimes used in the United States.

The standard atmosphere (atm) is an established constant. It is approximately equal to typical air pressure at earth mean sea level and is defined as follows.

standard atmosphere = 101325 Pa = 101.325 kPa = 1013.25 hPa.

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Pascal's law

In a body of equally dense fluid at rest, the pressure is the same for all points in the fluid so long as those points are at the same depth below the fluid's surface Formula

The intuitive formulation is that the pressure at the base of a column of water is due to the weight of the column.

The difference of pressure between two differents heights h1 and h2 is given by :

where ρ (rho) is the density or volumic mass of the fluid, g the acceleration due to gravity, and h1, h2 are elevations

Work (abbreviated W) is the energy transferred by a force to a moving object. Work is a scalar quantity, but it can be positive or negative. It is associated with a change in energy, but not all changes in energy can be readily analysed in terms of work.In physics, work is defined as the integral of dot product of force times infinitesimal translation:

The SI derived unit of work is the joule (J), which is defined as the work done by a force of one newton acting over a distance of one metre. The dimensionally equivalent newton-meter (N·m) is sometimes used instead; however, it is also sometimes reserved for torque to distinguish its units from work or energy.

Temperature, T

Temperature is a measure of the random molecular motion of the fluid at a point. The hotter the fluid the more energy is stored in random motion of molecules.

The unit of temperature is Kelvin (K), as an absolute measure of thermal energy or Centigrade (oC), as a relative measure with 0 degrees at the freezing point of water.

Velocity, V

Velocity (symbol: v) is a vector measurement of the rate and direction of motion. The units of Velocity are m/s in the metric

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system The average speed v of an object moving a distance d during a time interval t is described by the formula:

Acceleration is the rate of change of an object's velocity over time. The average acceleration of a of an object whose speed changes from vi to vf during a time interval t is given by:

Where vi = an object's initial velocity and vf = the object's final velocity over a period of time t

Speed (symbol: v) is the rate of motion, or equivalently the rate of change of position, expressed as distance d moved per unit of time t.

Speed is a scalar quantity with dimensions distance/time; the equivalent vector quantity to speed is known as velocity. Speed is measured in the same physical units of measurement as velocity, but does not contain the element of direction that velocity has. Speed is thus the magnitude component of velocity.

Units of speed include:metres per second, (symbol m/s), the SI derived unit kilometres per hour, (symbol km/h) miles per hour, (symbol mph)

Kinetic energy is energy that a body has as a result of its speed.

Potential energy is stored energy. The energy is stored by doing work against a force such as gravity or the spring in a clockwork motor

Energy is a fundamental quantity that every physical system possesses. Energy of physical system in a certain given state is defined as the amount of work W needed to change the state of the system from some initial state (called reference state or reference level) to the given state. The SI unit for both energy and work is the joule (J),

Ideal Gas Law

Pressure, density and temperature of a gas are related through an equation of state. Under ordinary conditions for air,

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(1.17)

where p is the absolute pressure the, density, T the absolute temperature and R is a gas constant. The above equation is called the Ideal Gas Law or the Perfect Gas Equation. The gases obeying this equation are called Ideal Gases.

Bernoulli's Principle

The pressure of a fluid (liquid or gas) decreases at points where the speed of the fluid increases. In other words, Bernoulli found that within the same fluid, in this case air, high speed flow is associated with low pressure, and low speed flow with high pressure. This principle was first used to explain changes in the pressure of fluid flowing within a pipe whose cross-sectional area varied. In the wide section of the gradually narrowing pipe, the fluid moves at low speed, producing high pressure. As the pipe narrows it must contain the same amount of fluid. In this narrow section, the fluid moves at high speed, producing low pressure. In fluid dynamics, Bernoulli's equation, derived by Daniel Bernoulli, describes the behavior of a fluid moving along a streamline.

v = fluid velocity along the streamline g = acceleration due to gravity on Earth h = height from an arbitrary point in the direction of gravity p = pressure along the streamline ρ = fluid density

Zeroth-law definition of temperature

While most people have a basic understanding of the concept of temperature, its formal definition is rather complicated. Before jumping to a formal definition, let us consider the concept of thermal equilibrium. If two closed systems with fixed volumes are brought together, so that they are in thermal contact, changes may take place in the properties of both systems. These changes are due to the transfer of heat between the systems. When a state is reached in which no further changes occur, the systems are in thermal equilibrium.

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Now a basis for the definition of temperature can be obtained from the so-called zeroth law of thermodynamics which states that if two systems, A and B, are in thermal equilibrium and a third system C is in thermal equilibrium with system A then systems B and C will also be in thermal equilibrium (being in thermal equilibrium is a transitive relation; moreover, it is an equivalence relation).

Therefore, it is useful to establish a temperature scale based on the properties of some reference system. Then, a measuring device can be calibrated based on the properties of the reference system and used to measure the temperature of other systems. One such reference system is a fixed quantity of gas. The ideal gas law indicates that the product of the pressure and volume (P · V) of a gas is directly proportional to the temperature:

(1)

where 'T is temperature, n is the number of moles of gas and R is the gas constant. Thus, one can define a scale for temperature based on the corresponding pressure and volume of the gas: the temperature in kelvins is the pressure in pascals of one mole of gas in a container of one cubic metre, divided by 8.31... In practice, such a gas thermometer is not very convenient, but other measuring instruments can be calibrated to this scale.

Equation 1 indicates that for a fixed volume of gas, the pressure increases with increasing temperature. Pressure is just a measure of the force applied by the gas on the walls of the container and is related to the energy of the system. Thus, we can see that an increase in temperature corresponds to an increase in the thermal energy of the system. When two systems of differing temperature are placed in thermal contact, the temperature of the hotter system decreases, indicating that heat is leaving that system, while the cooler system is gaining heat and increasing in temperature. Thus heat always moves from a region of high temperature to a region of lower temperature and it is the temperature difference that drives the heat transfer between the two systems.

Temperature in gases

The second law of thermodynamics states that any two given systems when interacting with each other will later reach the same average energy. Temperature is a measure of the average kinetic energy of a system. The formula for the kinetic energy of an atom is:

Thus, particles of greater mass (say a neon atom relative to a hydrogen molecule) will move slower than lighter

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counterparts, but will have the same average energy. This average energy is independent of the mass because of the nature of a gas, all particles are in random motion with collisions with other gas molecules, solid objects that may be in the area and the container itself (if there is one). A visual illustration of this from Oklahoma State University makes the point more clear. Not all the particles in the container have different velocities, regardless of whether there are particles of more than one mass in the container, but the average kinetic energy is the same because of the ideal gas law. In a gas the distribution of energy (and thus speeds) of the particles corresponds to the Boltzmann distribution.

An electronvolt is a very small unit of energy, approximately 1.602×10-19 joule.

Second-law definition of temperature

In the previous section temperature was defined in terms of the Zeroth Law of thermodynamics. It is also possible to define temperature in terms of the second law of thermodynamics, which deals with entropy. Entropy is a measure of the disorder in a system. The second law states that any process will result in either no change or a net increase in the entropy of the universe. This can be understood in terms of probability. Consider a series of coin tosses. A perfectly ordered system would be one in which every coin toss would come up either heads or tails. For any number of coin tosses, there is only one combination of outcomes corresponding to this situation. On the other hand, there are multiple combinations that can result in disordered or mixed systems, where some fraction are heads and the rest tails. As the number of coin tosses increases, the number of combinations corresponding to imperfectly ordered systems increases. For a very large number of coin tosses, the number of combinations corresponding to ~50% heads and ~50% tails dominates and obtaining an outcome significantly different from 50/50 becomes extremely unlikely. Thus the system naturally progresses to a state of maximum disorder or entropy.

Entropy

The thermodynamic entropy S, often simply called the entropy in the context of thermodynamics, is a measure of the disorder present in a physical system. Equivalently, it can be understood as a measure of the amount of energy in a system that cannot be used to do work (the precise meaning of this will be explained in the following article). The SI unit of entropy is J·K-1 (joules per kelvin), which is the unit of heat capacity.

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Thermodynamic definition of entropy

In this section, we discuss the original definition of entropy, as introduced by Clausius in the context of classical thermodynamics. Clausius defined the change in entropy of a thermodynamic system, during a reversible process in which an amount of heat dQ is introduced at constant absolute temperature T, as

This definition makes sense when absolute temperature has been defined.

Clausius gave the quantity S the name "entropy"

Measuring entropy

In real experiments, it is quite difficult to measure the entropy of a system. The techniques for doing so are based on the thermodynamic definition of the entropy, and require extremely careful calorimetry

Now, we have stated previously that temperature controls the flow of heat between two systems and we have just shown that the universe, and we would expect any natural system, tends to progress so as to maximize entropy. Thus, we would expect there to be some relationship between temperature and entropy. In order to find this relationship let's first consider the relationship between heat, work and temperature. A heat engine is a device for converting heat into mechanical work and analysis of the Carnot heat engine provides the necessary relationships we seek. The work from a heat engine corresponds to the difference between the heat put into the system at the high temperature, qH and the heat ejected at the low temperature, qC. The efficiency is the work divided by the heat put into the system or:

(2)

where wcy is the work done per cycle. We see that the efficiency depends only on qC/qH. Because qC and qH correspond to heat transfer at the temperatures TC and TH, respectively, qC/qH should be some function of these temperatures:

(3)

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Carnot's theorem states that all reversible engines operating between the same heat reservoirs are equally efficient. Thus, a heat engine operating between T1 and T3 must have the same efficiency as one consisting of two cycles, one between T1 and T2, and the second between T2 and T3. This can only be the case if:

which implies:

q13 = f(T1,T3) = f(T1,T2)f(T2,T3)

Since the first function is independent of T2, this temperature must cancel on the right side, meaning f(T1,T3) is of the form g(T1)/g(T3) (i.e. f(T1,T3) = f(T1,T2)f(T2,T3) = g(T1)/g(T2)· g(T2)/g(T3) = g(T1)/g(T3)), where g is a function of a single temperature. We can now choose a temperature scale with the property that:

(4)

Substituting Equation 4 back into Equation 2 gives a relationship for the efficiency in terms of temperature:

(5)

Notice that for TC = 0 K the efficiency is 100% and that efficiency becomes greater than 100% below 0 K. Since an efficiency greater than 100% violates the first law of thermodynamics, this implies that 0 K is the minimum possible temperature. In fact the lowest temperature ever obtained in a macroscopic system was 20 nK, which was achieved in 1995 at NIST. Subtracting the right hand side of Equation 5 from the middle portion and rearranging gives:

where the negative sign indicates heat ejected from the system. This relationship suggests the existence of a state function, S, defined by:

(6)

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where the subscript indicates a reversible process. The change of this state function around any cycle is zero, as is necessary for any state function. This function corresponds to the entropy of the system, which we described previously. We can rearranging Equation 6 to get a new definition for temperature in terms of entropy and heat:

(7)

For a system, where entropy S may be a function S(E) of its energy E, the temperature T is given by:

(8)

The reciprocal of the temperature is the rate of increase of entropy with energy

In thermodynamics, a thermodynamic system is in thermodynamic equilibrium if its energy distribution equals a Maxwell-Boltzmann distribution. This allows a single temperature to be attributed to the system. The key idea is that the macroscopic parameters are unchanging. The term "thermal equilibrium" is also used to describe this situation.

The process that leads to a thermodynamic equilibrium is called thermalisation. An example of this is a system of interacting particles that is left undisturbed by outside influences. By interacting, they will share energy/momentum among themselves and reach a state where the global statistics are unchanging in time

The ideal gas law or equation is the equation of state of an ideal gas. It combines the three primitive gas laws formulated by early physics researchers. Although roughly accurate for gases at low pressures and high temperatures, it becomes increasingly inaccurate at higher pressures and lower temperatures. The equation has the form:

PV = nRT

where P is the pressure of an ideal gas, V is the volume, n is the number of moles, R is the gas constant[0.08206, in (atm*liters)/(moles*degrees Kelvin)], and T is its temperature

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Some isotherms of an ideal gas (i.e. the relation between pressure P and volume V at fixed temperature T; plotted for a set of temperatures, with increasing T from lower to upper curve)

Using statistical mechanics, the ideal gas law can be derived by assuming that a gas is composed of a large number of small molecules, with no attractive or repulsive forces. In reality, gas molecules do interact with attractive and repulsive forces. In fact it is these forces that result in the formation of liquids.

Volume, also called capacity, is a quantification of how much space an object occupies. The SI unit for volume is the cubic metre. American spelling is cubic meter.

The volume of a solid object is a numerical value given to describe the three-dimensional concept of how much space it occupies. One-dimensional objects (such as lines) and two-dimensional objects (such as squares) are assigned zero volume in the three-dimensional space.

Mathematically, volumes are defined by means of integral calculus, by approximating the given body with a large amount of small cubes, and adding the volumes of those cubes. The generalization of volume to arbitrarily many dimensions is called content. In differential geometry, volume is expressed by means of the volume form.

Volume and capacity are sometimes distinguished, with capacity being used for how much a container can hold (with contents measured commonly in litres or its derived units), and volume being how much space an object displaces (commonly measured in cubic mVolume measures: other SI units

A commonly used SI unit for volume is the litre (American spelling liter), and one thousand litres is the volume of a cubic metre (American spelling cubic meter), which was formerly termed a stere and often called a "cube" in engineering slang. A cubic centimetre (American spelling cubic centimeter) is the same volume as a millilitre.

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etres or its derived units).

Relationship to density

The volume of an object is equal to its mass divided by its average density. This is a rearrangement of the calculation of density as mass per unit volume.

The term specific volume is used for volume divided by mass. This is the reciprocal of the mass density, expressed in units such as cubic meters per kilogram (m³/kg).

The internal energy of a system (abbreviated E or U) is the total kinetic energy due to the motion of molecules (translational, rotational, vibrational) and the total potential energy associated with the vibrational and electric energy of atoms within molecules or crystals. Internal energy is a quantifiable state function of a system. The SI unit of energy is the joule.

For systems consisting of molecules, the internal energy is partitioned among all of these types of motion. In systems consisting of monatomic particles, such as helium gas and other noble gases, the internal energy consists only of the translational kinetic energy of the individual atoms. Monatomic particles, of course, do not rotate or vibrate, and are not excited to higher electrical energies, except at very high temperatures.

From the standpoint of statistical mechanics, the internal energy is shown to be equivalent to the ensemble average of the total energy of the system.

Measurement

Internal energy can not be measured directly; it is only measured as a change (ΔU). The equation for change in internal energy is

where

Q is heat input to or output of the system, measured in joules W is work done on or by the system, measured in joules

A positive value of W represents work done on the system, a negative value representing work done by the system. Generally physicists use negative "U" indicating work done by a system; chemists use positive "U" indicating work done on a system. A positive value of Q represents heat flow into the system, a negative value representing heat flow out of the system.

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physics, a force is an external cause responsible for any change of a physical system. For instance, a person holding a dog by a rope is experiencing the force applied by the rope on his hand, and the cause for its pulling forward is the force exercised by the rope. The kinetic expression of this change is, according to Newton's second law, acceleration, non kinetic expressions such as deformation can also occur. The SI unit for force is the newton.

Enthalpy (symbolized H, also called heat content) is the sum of the internal energy of matter and the product of its volume multiplied by the pressure. Enthalpy is a quantifiable state function, and the total enthalpy of a system cannot be measured directly; the enthalpy change of a system is measured instead. Enthalpy is a thermodynamic potential, and is useful particularly for nearly-constant pressure process, where any energy input to the system must go into internal energy or the mechanical work of expanding the system.

Equations

Enthalpy is defined by the following equation:

where

H is the enthalpy, measured in joules U is the internal energy, measured in joules P is the pressure of the system, measured in pascals V is the volume, measured in cubic metres

The total enthalpy of a system cannot be measured directly; the enthalpy change of a system is measured instead. Enthalpy change is defined by the following equation:

where

ΔH is the enthalpy change, measured in joules Hfinal is the final enthalpy of the system, measured in joules. In a chemical reaction, Hfinal is the enthalpy of the products. Hinitial is the initial enthalpy of the system, measured in joules. In a chemical reaction, Hinitial is the enthalpy of the reactants.

Enthalpy is most useful when pressure is held constant through exposure to the surroundings, to analyse reactions that increase the volume of the system, causing it to do mechanical work on the surroundings and lose energy. Conversely, reactions that cause a

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decrease in volume cause the surroundings to do work on the system, and an increase in the energy of the system. In this case, enthalpy change may be expressed as:

DH = DU + P DV

where

D may indicate an infinitesimal change (often denoted "d") or a finite difference (often denoted "Δ").

Regardless of whether the external pressure is constant, infinitesimal enthalpy change obeys:

dH = T dS + V dP

(where S is the entropy) so long as the only work done is through volume change. Since the expression T dS always represents transfer of heat, it makes sense to treat the enthalpy as a measure of the total heat in the system, so long as the pressure is held constant; this explains the term heat content.

For an exothermic reaction at constant pressure, the system's change in enthalpy is equal to the energy released in the reaction, including the energy retained in the system and lost through expansion against its surroundings. Similarly, for an endothermic reaction, the system's change in enthalpy is equal to the energy absorbed in the reaction, including the energy lost by the system and gained from compression from its surroundings.

Standard enthalpy

The standard enthalpy change of reaction (denoted Ho or HO)is the enthalpy change that occurs in a system when 1 equivalent of matter is transformed by a chemical reaction under standard conditions.

A common standard enthalpy change is the standard enthalpy change of formation, which has been determined for a vast number of substances. The enthalpy change of any reaction under any conditions can be computed, given the standard enthalpy change of formation of all of the reactants and products. Other reactions with standard enthalpy change values include combustion (standard enthalpy change of combustion) and neutralisation (standard enthalpy change of neutralisation).

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Mechanical power

In physics, power (symbol: P) is the amount of work W done per unit of time t. This can be modeled as an energy flow, equivalent to the rate of change of the energy in a system, or the time rate of doing work, as defined by

whereP is power E is energy or work W t is time

The units of power are therefore energy divided by time.

SI unit

The SI unit of power is the watt, which is equal to one joule per second.

Non-SI units

Non-SI units of power include horsepower (HP), Pferdestärke (PS), cheval vapeur (CV) and foot-pounds per minute. One unit of horsepower is equivalent to 33,000 foot-pounds per minute, or the power required to lift 550 pounds one foot in one second, and is equivalent to about 746 watts. Other units include: dBm, logarithmic measure with 1 milliwatt as reference; kilocalorie per hour (often referred to as Calories per hour

The horsepower (hp) is the name of several non-metric units of power. In scientific discourse the term "horsepower" is rarely used due to the various definitions and the existence of an SI unit for power, the watt (W). However, the idea of horsepower persists as a legacy term in many languages, particularly in the automotive industry for listing the maximum power of internal-combustion engines.

The various types of horsepower (metric) are:

Horsepower (hp)

According to the most common definition of horsepower, one horsepower is defined as exactly:

1 hp = 745.69987158227022 W

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A common memory aid is based on the fact that Christopher Columbus first sailed to the Americas in 1492. The memory aid states that 1 hp = 1/2 Columbus or 746 W.

In fourteen hundred and ninety-two Columbus sailed the ocean blue. Divide that son-of-a-gun by two And that's how many watts there are in a horsepower.

The horsepower was first used by James Watt during a business venture where his steam engines replaced horses. It was defined that a horse can lift 33,000 pounds force with a speed of 1 foot per minute: 33,000 ft·lbf·min−1. This is sometimes called a mechanical horsepower to distinguish it from the other definitions of horsepower below.

Put into perspective, a healthy human can sustain about 0.1 horsepower. Most observers familiar with horses and their capabilities estimate that Watt was a bit optimistic; few horses could maintain that effort for long.

Simpler formulae

In the simplest case, that of a body moving in a steady direction, and acted on by a force parallel to that direction, the work is given by the formula

where

F is the force and s is the distance traveled by the object.

The work is taken to be negative when the force opposes the motion. More generally, the force and distance are taken to be vector quantities, and combined using the dot product:

where φ is the angle between the force and the displacement vector.

This formula holds true even when the force acts at an angle to the direction of travel. To further generalize the formula to situations in which the force and the object's direction of motion changes over time, it is necessary to use differentials, d, to express the infinitesimal work done by the force over an infinitesimal displacement, thus:

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The integration of both sides of this equation yields the most general formula, as given above.

The centripetal force is the force pulling an object toward the center of a circular path as the object goes around the circle. An object can travel in a circle only if there is a centripetal force on it.

In the case of an orbiting satellite the centripetal force is its weight and acts toward the satellite's primary; in the case of an object at the end of a rope, the centripetal force is the tension of the rope and acts towards whatever the rope is anchored to.

Centripetal force must not be confused with centrifugal force. In an inertial reference frame (not rotating or accelerating), the centripetal force accelerates a particle in such a way that it moves along a circular path. In a corotating reference frame, a particle in circular motion has zero velocity. In this case, the centripetal force appears to be exactly cancelled by a pseudo-force, the centrifugal force. Centripetal forces are true forces, appearing in inertial reference frames; centrifugal forces appear only in rotating frames.

Centripetal force must not be confused with central force either.

Objects moving in a straight line with constant speed also have constant velocity. However an object moving in an arc with constant speed has a changing direction of motion. As velocity is a vector of speed and direction, a changing direction implies a changing velocity. The rate of this change in velocity is the centripetal acceleration. Differentiating the velocity vector gives the direction of this acceleration towards the center of the circle.

By Newton's second law of motion, as there is an acceleration there has to be a force in the direction of the acceleration. This is the centripetal force, and is equal to:

(where m is mass, v is velocity, r is radius of the circle, and the minus sign denotes that the vector points to the center of the circle and ω = v / r is the angular velocity)

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Centripetal means towards the center.

There are two different definitions for the term Centrifugal force. One is that centrifugal force is one of the fictitious forces that appears to act on an object when its motion is viewed from a rotating frame of reference. Another, less popular definition is that Centrifugal force is the reaction force exerted by an object moving in a circular path upon the object that is causing its circular motion, according to Newton's Third Law.

The force that maintains circular motion is called centripetal force. If no force is exerted on an object, it moves in a straight line at a constant speed. To make the object deviate from that straight path into in a circular one, a centripetal ("center seeking") force must be exerted at right angles to the object's velocity, directed toward the center of the circle. Since this causes a change in the direction of the object's velocity, the centripetal force causes a corresponding centripetal acceleration, also toward the center.

When viewed from an Inertial Frame of Reference, what is really happening is that the passenger's inertia resists any change of motion and keeps the passenger moving along the initial straight line of motion. From this point of view, the only reason that the passenger is pushed to the outside of the car is that the person is still travelling in a straight line, and the car has accelerated. Once the passenger hits the door of the car, the car is then able to apply the centripetal force on the passenger to accelerate him or her around the turn with the car. Friction between the seat of the car and the seat of the passenger's pants is also a component of the centripetal force, and at lower speeds, where passengers do not slide, friction accounts for all of it. In turn, the passenger also exerts a reaction force upon the door: according to the alternative definition, this would also be called a centrifugal force.

Confusion has emerged over the term centrifugal force because of these two quite different definitions. According to one definition, centrifugal force acts on the object and is a fictitious force, that only exists in rotating frames of reference. The other force that has been referred to as centrifugal force is the real reaction force exerted by the object.

In physics, a force is an external cause responsible for any change of a physical system. For instance, a person holding a dog by a rope is experiencing the force applied by the rope on his hand, and the cause for its pulling forward is the force exercised by the rope. The kinetic expression of this change is, according to Newton's second law, acceleration, non kinetic expressions such as deformation can also occur. The SI unit for force is the newton.

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The rate of change of speed with respect to time is termed acceleration.

In physics, momentum is a physical quantity related to the velocity and mass of an object.

Momentum can be defined as "mass in motion." All objects have mass, so if an object is moving, then it has momentum. The amount of momentum that an object has depends two variables: the mass of the moving object and its velocity. This can be written as:

Momentum = mass × velocity

In physics, the symbol for momentum is a small p; so the above equation can be rewritten as:

p = m × v

where m is the mass and v the velocity. The SI unit of momentum is kilogram metres per second (kg m/s). The equation shows that an object with twice the mass of another object travelling at the same velocity would have twice the momentum. An object travelling twice as fast as another in the same direction with the same mass would have twice the momentum.

The velocity of an object is given by its speed and its direction. This means that if the direction of an object changes, its velocity changes. Likewise if the speed changes, the velocity also changes. Because momentum depends on velocity, it too has a magnitude and a direction: it is a vector quantity. For example the momentum of a 5 kg bowling ball would have to be described by the statement that it was moving westward at 2 m/s. It is insufficient to say that the ball has 10 kg m/s of momentum; the momentum of the ball is not fully described until information about its direction is given.

A change in an object's momentum is known as an impulse:

The impulse (mass × change in velocity) = force applied × the time over which the force was applied.

Kinetic energy is energy that a body has as a result of its speed.

It is formally defined as work needed to accelerate a body from rest to a velocity v. Having gained this energy during its acceleration, the body maintains this kinetic energy unless its speed changes. The same amount of work would also be required to return the body to a state of rest from that velocity.

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Potential energy is stored energy. The energy is stored by doing work against a force such as gravity or the spring in a clockwork motor

Newton's first law: law of inertia

When no force acts on an object (or when the forces acting on it cancel), it moves in a straight line at constant speed.

This law is also called the Law of Inertia or Galileo's Principle.

Newton's second law: fundamental law of dynamics

The acceleration of an object equals the total force acting on it,

divided by its (constant) mass, .

Where m is the mass of the object in question, is the total force acting on the object is the object's acceleration, i.e., the rate of change of its velocity with respect to time

When the forces on the object all act along the same line, they can be added as positive and negative numbers, depending on their direction. When they do not all act along the same line, the total must be found by vector addition.

The quantity m, or mass, is a characteristic of the object. The greater the total force acting on an object, the greater the change in its acceleration will be. This equation, therefore, indirectly defines the concept of mass.

In the equation, F = ma, a is directly measurable but F is not. The second law only has meaning if we are able to assert, in advance, the value of F. Rules for calculating force include Newton's law of universal gravitation.

Newton's second law can also be stated in terms of momentum as

. The physical meaning of this equation is that objects interact by exchanging momentum, and they do this via a force. When the mass of an object is varying, this form is valid, and the

form is not. The statement in terms of momentum is also valid in special relativity if we express the momentum as ,

where γ is .

Newton's first law appears to be a special case of the second law, and Newton may have stated the first law separately simply in order to throw down the gantlet to the Aristotelians. However, many

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modern physicists prefer to think of the First Law as defining the reference frames in which the other two laws are valid. These reference frames are called inertial reference frames or Galilean reference frames, and are moving at constant velocity, that is to say, without acceleration. (Note that an object may have a constant speed and yet have a non-zero acceleration, as in the case of uniform circular motion. This means that the surface of the Earth is not an inertial reference frame, since the Earth is rotating on its axis and orbits around the Sun. However, for many experiments, the Earth's surface can safely be assumed to be inertial. The error introduced by the acceleration of the Earth's surface is minute.)

An illustration of Newton's third law. The skaters' forces on each other are equal in magnitude, and in opposite directions. Although the forces are equal, the accelerations are not: the less massive skater will have a greater acceleration. In terms of conservation of momentum, the total momentum is zero before they push off, because they aren't yet moving, and zero after they push off, because the momenta have opposite signs, and the momentum of the less massive skater (small m, large v) cancels the momentum of the more massive skater (large m, small v).

Newton's third law: law of reciprocal actions

Whenever one body exerts force upon a second body, the second body exerts an equal and opposite force upon the first body.

or:

Momentum is conserved, i.e., momentum cannot be created or destroyed, but only transferred from one object to another.

The two forces in Newton's third law are of the same type, e.g., if the road exerts a forward frictional force on an accelerating car's tires, then it is also a frictional force that Newton's third law predicts for the tires pushing backward on the road.

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Newton's third law should not be interpreted as a prediction that forces always cancel, or that equilibrium always exists. When objects A and B interact, the forces referred to are acting on different objects: A's force on B, and B's on A. We add forces acting on the same object, not on different objects, so it doesn't make sense physically to say that these two forces add up to zero.

The forces acting between particles A and B lie along parallel lines, but need not lie along the line connecting the particles. One example of this is a force on an electric dipole due to a point charge, when the dipole points in a direction perpendicular to the line connecting the point charge and the dipole. The force on the dipole due to the point charge is perpendicular to the line connecting them, so there is a reaction force on the point charge in the opposite direction, but these two force vectors are parallel and, even when extended to a line, they never cross each other in space.

The gas laws are a set of laws that describe the relationship between thermodynamic temperature (T), pressure (P) and volume (V) of gases. Three of these laws, Boyle's law, Charles's law, and Gay-Lussac's law come together to form the combined gas law, which with the addition of Avogadro's law later gave way to the ideal gas law. Other important gas laws include Dalton's law of partial pressures. The kinetic theory of gases, Graham's law of effusion, and root mean square velocity all explain how individual molecules in a gas act and their relation to pressure, volume, and temperature.

A gas which obeys these gas laws exactly is hypothetical and is known as an ideal gas (or perfect gas). An ideal gas does not exist, however, some gases follow the laws more closely than others given standard conditions.

The most important gas law is the ideal gas law which states that:

-where: P is pressure in atmospheres (atm) or kilopascals (kPa) V is volume in liters n is the number of moles of gas R is the ideal gas constant in L atm/mol K or kPa/mol K T is temperature in kelvins.

Other gas laws, such as van der Waals equation, seek to eliminate the value differences between the ideal gas laws and the actual gases. The van der Waals equation alters the ideal gas law to more truely reflect how actual gases function using a series of calculated values called van der Waals constants.

Boyle's law (sometimes called the Boyle Mariotte law) is one of the gas laws

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Boyle's Law states that the product of the volume and pressure of an ideal gas is constant, given constant temperature. Expressed mathematically, the formula for Boyle's law is:

-where: V is volume of the gas. P is the pressure of the gas. k is a constant.

To maintain the constant during an increase in pressure of a gas, at fixed temperature, requires that the volume decrease. Conversely, reducing the pressure of the gas increases the volume.

The exact value of the constant need not be known to make use of the law in comparison between two volumes of the same amount of gas at equal temperature:

Boyle's law, Charles's Law, and Gay-Lussac's Law form the combined gas law. The three gas laws in combination with Avogadro's Law can be generalized by the ideal gas law.

Charles's law (sometimes called the Law of Charles and Gay-Lussac) is one of the gas laws

Charles law states that, at constant pressure, the volume of a given mass of a gas at 0 degrees Celsius increases or decreases by 1/273 times its volume for every degree Celsius rise or fall in temperature. The formula for this law is:

-where:V is the volume. T is the temperature (measured in kelvins). k is a constant.

To maintain the constant, k, during heating of a gas at fixed pressure, the volume must increase. Conversely, cooling the gas decreases the volume. The exact value of the constant need not be known to make use of the law in comparison between two volumes of gas at equal pressure:

.

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The Reynolds number is the most important dimensionless number in fluid dynamics and provides a criterion for determining dynamic similitude. Where two similar objects in perhaps different fluids with possibly different flowrates have similar fluid flow around them, they are said to be dynamically similar.

It is named after Osborne Reynolds (1842-1912), who proposed it in 1883. Typically it is given as follows:

or

With: vs - mean fluid velocity, L - characteristic length (equal to diameter 2r if a cross-section is circular), μ - (absolute) dynamic fluid viscosity, ν - kinematic fluid viscosity: ν = μ / ρ, ρ - fluid density. The Reynolds number is the ratio of inertial forces (vsρ) to viscous forces (μ/L) and is used for determining whether a flow will be laminar or turbulent. Laminar flow occurs at low Reynolds numbers, where viscous forces are dominant, and is characterized by smooth, constant fluid motion, while turbulent flow, on the other hand, occurs at high Reynolds numbers and is dominated by inertial forces, producing random eddies, vortices and other flow fluctuations.

The transition between laminar and turbulent flow is often indicated by a critical Reynolds number (Recrit), which depends on the exact flow configuration and must be determined experimentally. Within a certain range around this point there is a region of gradual transition where the flow is neither fully laminar nor fully turbulent, and predictions of fluid behaviour can be difficult. For example, within circular pipes the critical Reynolds number is generally accepted to be 2300, where the Reynolds number is based on the pipe diameter and the mean velocity vs within the pipe, but engineers will avoid any pipe configuration that falls within the range of Reynolds numbers from about 2000 to 4000 to ensure that the flow is either laminar or turbulent.

Laminar flow is when a fluid flows in parallel layers, with no disruption between the layers. In fluid dynamics, laminar flow is a flow regime characterized by high momentum diffusion, low momentum convection, and pressure and velocity independence from time. It is the opposite of turbulent flow. The (dimensionless) Reynolds number characterizes whether flow conditions lead to laminar or turbulent flow.For example, consider the flow of air over

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an airplane wing. The boundary layer is a very thin sheet of air lying over the surface of the wing (and, for that matter, all other surfaces of the airplane). Because air has viscosity, this layer of air tends to adhere to the wing. As the wing moves forward through the air, the boundary layer at first flows smoothly over the streamlined shape of the airfoil. Here the flow is called the laminar layer.

As the boundary layer approaches the centre of the wing, it begins to lose speed due to skin friction, and it becomes thicker and turbulent. Here it is called the turbulent layer. The process of a laminar boundary layer becoming turbulent is known as boundary layer transition.The point at which the boundary layer changes from laminar to turbulent is called the transition point. Where the boundary layer becomes turbulent, drag due to skin friction is relatively high. As speed increases, the transition point tends to move forward. As the angle of attack increases, the transition point also tends to move forward. One way to limit the size and effect of the turbulent region is to use swept-back "delta" wings. This is particularly important in supersonic aircraft.

In fluid dynamics, turbulence or turbulent flow is a flow regime characterized by low momentum diffusion, high momentum convection, and rapid variation of pressure and velocity in space and time. Flow that is not turbulent is called laminar flow. The (dimensionless) Reynolds number characterizes whether flow conditions lead to laminar or turbulent flow.

Consider the flow of water over a simple smooth object, such as a sphere. At very low speeds the flow is laminar; i.e., the flow is smooth (though it may involve vortices on a large scale). As the speed increases, at some point the transition is made to turbulent ("chaotic") flow. In turbulent flow, unsteady vortices appear on many scales and interact with each other. Drag due to boundary layer skin friction increases. The structure and location of boundary layer separation often changes, sometimes resulting in a reduction of overall drag. Because laminar-turbulent transition is governed by Reynolds number, the same transition occurs if the size of the object is gradually increased, or the viscosity of the fluid is decreased, or if the density of the fluid is increased.

In physics and fluid mechanics, the boundary layer is that layer of fluid in the immediate vicinity of a bounding surface. In the atmosphere the boundary layer is the air layer near the ground affected by diurnal heat, moisture or momentum transfer to or from the surface. On an aircraft wing the boundary layer is the part of the flow close to the wing. The Boundary layer effect occurs at the field region in which all changes occur in the flow pattern. The boundary layer distorts surrounding nonviscous flow. It is a phenonomen of viscous forces. This effect is related to the Leidenfrost effect and the Reynolds number.

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Aerodynamics

The aerodynamic boundary layer was discovered by Ludwig Prandtl at the beginning of the twentieth century and represents one of the greatest discoveries in the history of aerodynamics. Two effects must to be considered. First, the boundary layer adds to the effective thickness of the body, through the displacement thickness hence increasing the pressure drag. Secondly, the shear forces at the surface of the wing create skin friction drag.

At high Reynolds numbers, typical of full-sized aircraft, it is desirable to have a laminar boundary layer. This results in a lower skin friction due to the characteristic velocity profile of laminar flow. However, the boundary layer inevitably thickens and becomes less stable as the flow develops along the body, and eventually becomes turbulent, the process known as boundary layer transition. One way of dealing with this problem is to suck the boundary layer away through a porous surface. This can result in a reduction in drag, but is usually impractical due to the mechanical complexity involved.

At lower Reynolds numbers, such as those seen with model aircraft, it is relatively easy to maintain laminar flow. This gives low skin-friction, which is desirable. However, the same velocity profile which gives the laminar boundary layer its low skin friction also causes it to be badly affected by adverse pressure gradients. As the pressure begins to recover over the rear part of the wing chord, a laminar boundary layer tends to separate from the surface. Such separation causes a large increase in the pressure drag, since it greatly increases the effective size of the wing section. In these cases, it can be advantageous to deliberately trip the boundary layer into turbulence at a point prior to the location of laminar separation. The fuller velocity profile of the turbulent boundary layer allows it to sustain the adverse pressure gradient without separating. Thus, although the skin friction is increased, overall the drag is decreased. Special wing sections have also been designed which tailor the pressure recovery so that laminar separation is reduced or even eliminated. This represents an optimum compromise between the pressure drag from flow separation and skin friction the induced turbulence

Types of drag are generally divided into three categories: parasitic drag, lift-induced drag and wave drag. Parasitic drag includes form drag, skin friction and interference drag. Lift-induced drag is only relevant when wings or a lifting body are present, and is therefore usually discussed only in the aviation perspective of drag. Beyond these two kinds of drag there is a third kind of drag, called wave drag, that occurs when the solid object is moving through the fluid at or near the speed of sound in that fluid. The overall drag of an object is characterized by a dimensionless number called the drag coefficient, and is calculated using the drag equation. Assuming a

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constant drag coefficient, drag will vary as the square of velocity. Thus, the resultant power needed to overcome this drag will vary as the cube of velocity

Parasitic drag is drag caused by moving a solid object through a fluid. Parasitic drag is made up of many components, the most prominent being form drag. Skin friction and interference drag are also major components of parasitic drag.

In aviation, induced drag tends to be greater at lower speeds because a high angle of attack is required to maintain lift. However, as speed increases the induced becomes much less, but parasitic drag necessarily increases because the fluid is flowing faster. At even higher speeds in the transonic, wave drag enters the picture. Each of these forms of drag changes in proportion to the others based on speed. The combined overall drag curve therefore shows a minimum at some airspeed - an aircraft flying at this speed will be at or close to its optimal efficiency. Pilots will use this speed to maximise endurance (minimum fuel consumption), or maximise gliding range in the event of an engine failure.

In aerodynamics, form drag, profile drag, or pressure drag, is a component of parasitic drag. The general size and shape of the body is the most important factor in form drag; bodies with a larger apparent cross-section will have a higher drag than thinner bodies. "Clean" designs, or designs that are streamlined and change cross-sectional area gradually are also critical for achieving minimum form drag. Form drag follows the drag equation, meaning that it rises with the square of speed, and thus becomes more important for high speed aircrat

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In physics, the drag equation gives the drag experienced by an object moving through a fluid.

where D is the force of drag, Cd is the drag coefficient (a dimensionless constant, e.g. 0.25 to 0.45 for a car), ρ is the density of the fluid*, v is the velocity of the object relative to the fluid, and A is the reference area.

In aerodynamics, skin friction is the component of parasitic drag arising from the friction of the fluid against the "skin" of the object that is moving through it. Skin friction is a function of the interaction between the fluid and the skin of the body, as well as the wetted surface, or the area of the surface of the body that would become wet if sprayed with water flowing in the wind. As with other components of parasitic drag, skin friction follows the drag equation and rises with the square of the velocity.

Interference drag

In aerodynamics, interference drag is a component of parasitic drag which is caused by vortices. Whenever two surfaces meet at a sharp angle on an airplane, the airflow has a tendency to form a vortex. Accelerating the air into this vortex causes drag on the plane, and the resulting low pressure area behind the plane also contributes. Thus, the primary method of reducing interference drag is eliminating sharp angles by adding fairings which smooth out any sharp angles on the aircraft. As with other components of parasitic drag, interference drag follows the drag equation and rises with the square of the velocity.

Wave drag is an aerodynamics term that refers to a sudden and very powerful form of drag that appears on aircraft flying at high-subsonic speeds. Wave drag is caused by the formation of shock waves around the aircraft. Shock waves radiate away a considerable amount of energy, energy that is "seen" by the aircraft as drag. Although shock waves are typically associated with supersonic flow, they can actually form at much lower speeds at areas on the aircraft where the Bernoulli effect accelerates local airflow to supersonic speeds over curved areas. The effect is typically seen at speeds of about Mach 0.8, but it is possible to notice the problem at any speed over that of the critical mach of that aircraft's wing. The magnitude of the rise in drag is impressive, typically peaking at about four times the normal subsonic drag.

These research projects were quickly put to use by aircraft designers. One common solution to the problem of wave drag due to

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the wings was to use a swept-wing, which had actually been developed before WWII and used on some German wartime designs (none of which saw service). Sweeping the wing to the rear makes it appear thinner and longer in the direction of the airflow, making a "normal" wing shape closer to that of the von Kármán ogive, while still remaining useful at lower speeds where curvature and thickness are important.

One does not have to sweep the wing, it is possible to build a wing that is simply extremely thin.

Lift-induced drag

In aerodynamics, lift-induced drag, or more simply, induced drag, is a drag force arising from the generation of lift by wings or a lifting body during flight.Induced drag will be present whenever the wings are producing lift. To that extent, it is often said that induced drag is a part of lift. It arises from the downwash induced by the wingtip and trailing edge vortices which, for a given amount of lift being produced, tilts the total reaction force further backwards through the induced downwash angle. This extra rearward tilt, in effect, increases the length of the drag vector and it is this increase in drag which is known as induced drag. Obviously, the smaller the angle of induced downwash, the lower will be the induced drag.

There are a number of factors affecting induced drag:

Aspect Ratio

High aspect ratio wings produce smaller vortices and, in comparison with a wing of lower aspect ratio, proportionally less of the airflow swept by the longer span is affected by the vortices. Consequently, the induced downwash angle when averaged over the whole of the high aspect ratio wing, is smaller and the induced drag is low. To minimise lift-induced drag, gliders have very high aspect ratios.

Wing Planform

For a wing of given span, an elliptical planform produces the smallest vortices and therefore the lowest induced drag.However, for wings with straight leading and trailing edges, the judicious use of taper and washout of the wing sections toward the tips can produce a similar reduction in induced drag.

Coefficient of Lift

From the pilot's point of view, where the aspect ratio and planform of the aircraft are fixed, the important factors in determining induced drag are angle of attack, airspeed and aircraft weight.

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These are incorporated in the induced drag formula which can be seen to have a powerful effect on the amount of induced drag generated.

Angle of attack Induced drag increases as the angle of attack is increased. The strength of the vortices is determined by the pressure difference above and below the wing. When the wing is at the zero-lift angle of attack (AoA -2° / CL = 0) there are no vortices and therefore no induced drag. As the angle of attack is increased, vortices form and increase in strength up to the angle of attack for CL Max, typically 16° (varies between aircraft). Induced drag therefore increases with angle of attack to be at a maximum at the stalling angle. Airspeed Induced drag is inversely proportional to the square of the indicated airspeed (IAS). This is the opposite to the effect of airspeed on parasite drag, which is directly proportional to IAS². When the factors of angle of attack and airspeed are combined, induced drag is greatest at low airspeeds and at high angles of attack.Weight Increased weight means that higher angles of attack must be used to produce a given amount of lift for a given speed. Induced drag increases in proportion to weight squared (W²).

An alternative way of looking at induced drag is as follows. The production of vortices is an inevitable consequence of the production of lift with a wing of finite span. These vortices result in an induced downwash which is over and above the downwash necessary to produce lift. To produce a rotary motion of any fluid requres energy — an example is the energy requred to stir a large volume of water in a drum with some sort of paddle. In flight, the energy required to create the vortices must be sourced from somewhere. Ultimately, that demand is placed on the engine by requiring higher power to be used to offset the induced drag when it is desired to maintain a given speed.

Some airliners and many gliders have small fins at the wing tips winglets to minimise the vortices. Wingtip tanks have a similar effect.

Induced drag must be added to the parasitic drag to find the total drag. As discussed above, induced drag becomes less of a factor the faster the aircraft flies because at higher speeds a smaller angle of attack is required for the same amount of lift. The opposite occurs with parasitic drag (the drag caused simply by pushing the aircraft through the air), which increases with speed. The combined overall drag curve therefore shows a minimum at some airspeed — an

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aircraft flying at this speed will be at or close to its optimal efficiency.

In aerodynamics, the aspect ratio is an airplane's wing's span divided by its standard mean chord (SMC). It can be calculated more

easily, however as span squared divided by wing area:

Informally, a "high" aspect ratio indicates long, narrow wings, whereas a "low" aspect ratio indicates short, stubby wings.

Aspect ratio is a powerful indicator of the general performance of a wing. Wingtip vortices greatly deteriorate the performance of a wing, and by reducing the amount of wing tip area, making it skinny or pointed for instance, you reduce the amount of energy lost to this process, and increase the lift generated by the wing.

High aspect-ratio wings reduce the amount of induced drag relative to the amount of lift produced.

Why don't all aircraft have high aspect-ratio wings? There are several reasons:

Structural: the deflection along a high aspect-ratio wing tends to be much higher than for one of low aspect ratio, thus the stresses and consequent risk of fatigue failures are higher - particularly with swept-wing designs.

Maneuverability: a high aspect-ratio wing will have a lower roll rate than one of low aspect ratio, due to higher drag and greater moment of inertia, thus rendering them unsuitable for fighter aircraft.

Stability - low aspect ratio wings tend to be more naturally stable than high-aspect ratios. This confers handling advantages, especially at slow speeds.

Practicality - low aspect ratios have a greater useful internal volume, which can be used to house the fuel tanks, retractable landing gear and other systems.

Lift consists of the sum (technically the negative product) of all the fluid dynamic forces on a body normal (i.e. perpendicular) to the direction of the external flow around that body.

Reaction due to accelerated air

In air (or comparably in any fluid), lift is created as an airstream passes by an airfoil and is deflected downward. The force created by

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this deflection of the air creates an equal and opposite force upward on an airfoil (see Newton's third law.) The deflection of airflow downward during the creation of lift is known as downwash. (Note: Confusingly, the term "downwash" has two somewhat different meanings with regard to aircraft. See downwash for a more complete explanation.)

It is important to note that the acceleration of the air does not simply involve the air molecules "bouncing off" the bottom of the airfoil. Rather, air molecules closely follow both the top and bottom surfaces of the airfoil, and so the airflow is deflected downward. In fact, the acceleration of the air during the creation of lift can also be described as a "turning" of the airflow.

Bernoulli's principle

The force on the wing can also be examined in terms of the pressure differences above and below the wing. (This method of explanation is mathematically equivalent to the Newton's 3rd law explanation as developed above.) The relationship between the velocities and pressures above and below the wing are nearly predicted by Bernoulli's equation

Circulation

A third way of calculating lift is a mathematical construction called circulation. Again, it is mathematically equivalent to the two explanations above. It is often used by practicing aerodynamicists as a convenient quantity, but is not often useful for a layperson's understanding. The circulation is the line integral of the velocity of the air, in a closed loop around the boundary of an airfoil. It can be understood as the total amount of "spinning" (or vorticity) of air around the airfoil. When the circulation is known, the section lift can be calculated using:

where ρ is the air density, V is the free-stream airspeed, and Γ is the circulation.

The Helmholtz theorem states that circulation is conserved. When an aircraft is at rest, there is no circulation. As the flow speed increases (that is, the aircraft accelerates in the air-body-fixed frame), a vortex, called the starting vortex, forms at the trailing edge of the airfoil, due to viscous effects in the boundary layer. Eventually the vortex detaches from the airfoil and gets swept away from it rearward. The circulation in the starting vortex is equal in magnitude and opposite in direction to the circulation around the airfoil. Theoretically, the starting vortex remains connected to the

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vortex bound in the airfoil, through the wing-tip vortices, forming a closed circuit. In reality the starting vortex gets dissipated by a number of effects, as do the wing-tip vortices far behind the aircraft.

Coefficient of lift

When the coefficient of lift is known, for instance from tables of airfoil data, lift can be calculated using the Lift Equation:

where:

CL is the coefficient of lift, ρ is the density of air (1.225 kg/m3 at sea level)* V is the freestream velocity, that is the airspeed far from the lifting surface A is the surface area of the lifting surface L is the lift force produced.

This equation can be used in any consistent system. For instance, if the density is measured in kilograms per cubic metre, the velocity is measured in metres per second, and the area is measured in square metres, the lift will be calculated in newtons. Or, if the density is in slugs per cubic foot, the velocity is in feet per second, and the area is in square feet, the resulting lift will be in pounds force.

Downwash

The term downwash has two nearly unrelated meanings within the field of aerodynamics.One meaning, used most often by non-engineers, refers to the forcing of air downward during the creation of lift. This usage is most common with regard to helicopters where the effect is most dramatic.The other meaning, used most often by engineers, refers to the flow of air over the tip of a wing and is a critical component in the creation of wing tip vortices.

In fluid dynamics, Bernoulli's equation, derived by Daniel Bernoulli, describes the behavior of a fluid moving along a streamline.

v = fluid velocity along the streamline g = acceleration due to gravity on Earth h = height from an arbitrary point in the direction of gravity p = pressure along the streamline ρ = fluid density

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These assumptions must be met for the equation to apply:Inviscid flow − viscosity (internal friction) = 0 Steady flow Incompressible flow − ρ = constant. (There exists a second form of Bernoulli's equation that is applicable for compressible flow, which makes use of the thermodynamic enthalpy.)

Generally, the equation applies along a streamline. For irrotational flow, it applies throughout the entire flow field.

The decrease in pressure simultaneous with an increase in velocity, as predicted by the equation, is often called Bernoulli's principle.

Wingtip vortices are vortices that develop at the edge of a wing as it flies through the air (or potentially another fluid). Wingtip vortices dramatically reduce the lift generated by the wing, and are therefore critically important in aerospace engineering.

Cause and Effects

As a wing flies through the air, it generates a low pressure zone on top of the wing through the Bernoulli effect. Fluids naturally flow from high to low pressure and the relatively high pressure air below the wing has a natural tendency to flow to the top of the wing. The air naturally cannot flow around the leading or trailing edge of the wing due to airspeed, but it can flow around the end. Consequently, air flows from below the wing, out around the edge to the top of the wing in a circular fashion. This raises the pressure on top of the wing and lowers the overall lift that the wing can produce.

Luckily, wingtip vortices only affect the portion of the wing closest to the end. Thus, the longer a wing is, the smaller the affected fraction of it will be. As well, the shorter the chord of the wing, the less opportunity air will have to form vortices. This means that for an airplane to be most efficient, it should have a very high aspect ratio. However, increasing the wingspan reduces the manoueverability of the aircraft, which is why combat and aerobatic planes usually feature short, stubby wings despite the efficiency losses.

Another method of reducing wingtip vortices is winglets, as seen on a number of modern airliners such as the Airbus A340. Winglets work by interfering with the formation of the vortex, thereby effectively increasing the aspect ratio of the wing. Winglets can yield very worthwhile economy improvements on long distance flights

A vortex is a spinning turbulent flow (or any spiral whirling motion) with closed streamlines. The shape of media or mass rotating rapidly around a center forms a vortex. It is a flow involving rotation about

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an axis (not always oriented vertically though; sometimes possessing a horizontal axis).

Washout

The term washout can have various meanings. In aviation, washout refers to the practice of building a wing with a slight twist, reducing the angle of incidence from root to the wingtip. The effect of this twist is that when the aircraft begins to stall, the wing's root does so before the tips. When this occurs the ailerons are still in smooth airflow, allowing the pilot to maintain roll control.

The primary benefit of washout is to increase the range of speeds at which the aircraft is stable and controllable. Without washout, a stalling aircraft is in danger of rolling out of control into an unrecoverable spin.

An angle of incidence is the angle between a beam incident on a surface and the normal (line perpendicular to the surface at the point of incidence).

Another common usage is in aviation, where it refers to the angle between the wing's chord (aircraft) and the longitudinal axis of an aircraft (a fixed value). Fig.2 shows a side view of part of an aeroplane. The angle of attack, which is the angle the wing chord presents to the airflow in flight.

Stall (flight)

In aerodynamics, a stall is a condition in which an excessive angle of attack causes loss of lift due to disruption of airflow.

An aircraft in flight is usually not pointed directly into the oncoming airflow. The angle (when viewed from the side of the aircraft) between the airflow and the wing is called the angle of attack (not to be confused with the pitch angle). If a pilot allows the angle of attack to become too large, the airflow will be unable to remain attached to the wing and it will begin to separate from the wing, creating a dramatic loss of lift. This condition is known as a stall.

Stall recovery usually involves reducing the angle of attack to "break" the stall, and adding power to begin a climb

Rigorous definition

A stall is a condition in aerodynamics and aviation where the angle between the wing's chord line and the relative wind, defined as the angle of attack, exceeds the critical angle of attack. The critical angle of attack is the maximum angle of attack on the lift coefficient

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versus angle-of-attack curve, and it defines the boundary between the wing's linear and nonlinear regimes. Flow separation begins to occur at this point, decreasing lift, increasing drag, and changing the wing's pitching moment. A fixed-wing aircraft during a stall will experience buffeting, a change in pitching moment (nose up or nose down depending on tailplane configuration), and changes in most stability derivatives.

Aileron control of roll becomes less effective, whereas its (inverse) control of yaw increases, making adverse yaw even more pronounced. Roll-yaw coupling becomes more pronounced as roll due to sideslip angle predominates. Pitch and roll damping decrease due to lower dynamic pressure, and strong nonlinearities in the airflow.

Increasing the angle of attack between an airfoil and the airflow causes the lift and drag produced to increase. This can continue until a point is reached where maximum lift is generated and this is known as the stall or stall angle. Any further increase in angle does not produce a corresponding increase in lift but will in fact lead to a sudden reduction in lift, a change in pitching moment or a wing drop. This is due to flow separation occurring on the upper surface of the airfoil, and therefore the critical angle of attack is dependent not only on the geometry of the configuration, but on the Reynolds number and surface roughnes

Typical behavior of most airfoils.

Graph

The graph shows that the greatest amount of lift is produced just before the critical angle of attack is reached (which in early 20th century aviation is called the "burble point") . This angle is 17.5 degrees in this case but changes from aircraft to aircraft. The graph shows that as the critical angle of attack is exceeded, the lift

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produced by the wing decreases significantly. The aerofoil is now stalled.

Note that this graph shows the stall angle, yet in practice most pilots discuss stalling in terms of airspeed. This is because in general terms one can relate the angle of attack to airspeed - a lower speed requires a greater angle of attack to produce the necessary lift and vice versa. Thus as speed falls, AoA increases, until the critical angle is reached. The airspeed at which this occurs is the stalling speed of the aircraft in that particular configuration. Deploying flaps/slats decreases the stall speed to allow the aircraft to land at a slower speed.

Aerodynamic description of a stall

Stalling an aeroplane

An aeroplane can be made to stall in any pitch attitude or bank angle or at any airspeed but is commonly practised by pilots reducing the speed to the stall speed, at a safe altitude. Stall speed varies on different airplanes and is represented by color codes on the air speed indicator. As the plane flies at this speed the angle of attack must be increased to prevent any loss of altitude or gain in airspeed (which corresponds to the stall angle described above). The pilot will notice the flight controls have become less responsive and may also notice some buffeting, an aerodynamic vibration caused by the airflow starting to detach from the wing surface.

In most light aircraft, as the stall is reached the aircraft will start to descend (because the wing is no longer producing enough lift to support the aeroplane's weight) and the nose will pitch down. Recovery from this stalled state usually involves the pilot decreasing the angle of attack and increasing the air speed, until smooth air flow over the wing is resumed. Normal flight can be resumed once recovery from the stall is complete.

The most common stall-spin scenarios occur on takeoff (departure stall) and during landing (base to final turn). Stalls also occur during a go-around maneuver if the pilot does not properly respond to the out-of-trim situation resulting from the transition from low power setting to high power setting at low speed. Stall speed is increased when the upper wing surfaces are contaminated with ice or frost.

A special form of asymmetric stall in which the aircraft also rotates about its yaw axis is called a spin. A spin will occur if an aircraft is stalled and there is an asymmetric yawing moment applied to it. This yawing moment can be aerodynamic (sideslip angle, rudder, adverse yaw from the ailerons), thrust related (p-factor, one engine inoperative on a multi-engine non-centerline thrust aircraft), or from any number of possible sources of yaw.

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Symptoms of an approaching Stall

One symptom of an approaching stall is slow and sloppy controls. As the speed of the airplane decreases approaching the stall, there are less air particles moving over the wing and therefore less will be deflected by the control surfaces (ailerons, rudder and elevator) at this slower speed. There may be slight buffeting of the controls as the stall is reached. The stall warning will sound, in most aircraft 5 to 10 knots above the stall speed.

Stalling characteristics

Different aircraft types have different stalling characteristics. A benign stall is one where the nose drops gently and the wings remain level throughout. Slightly more demanding is a stall where one wing stalls slightly before the other, causing that wing to drop sharply, with the possibility of entering a spin. A dangerous stall is one where the nose rises, pushing the wing deeper into the stalled state and potentially leading to an unrecoverable deep stall.

Stall warning and safety devices

Airplanes can be equipped with a variety of devices to prevent or postpone a stall or to make it less (or in some cases more) severe, or to make recovery easier.

A slight twist can be introduced to the wing with the leading edge near the wing tip twisted downward. This is called washout and causes the wing root to stall before the wing tip. This makes the stall gentle and progressive. Since the stall is delayed at the wing tips, where the ailerons are, roll control is maintained when the stall begins.

The wing can be built with aerodynamic twist; the airfoil changes shape toward the wing tip in such a way that the wing tip has a lower stall speed than the wing root. This serves the same purpose as washout.

A stall strip is a small sharp-edged device which, when attached to the leading edge of a wing, encourages the stall to start there in preference to any other location on the wing. If attached close to the wing root it makes the stall gentle and progressive; if attached near the wing tip it encourages the aircraft to drop a wing when stalling.

Vortex generators, tiny strips of metal or plastic placed on top of the wing near the leading edge, lower the stall speed by preventing flow separation over the top of the wing.

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An anti-stall strake is a wing extension at the root leading edge which generates a vortex on the wing upper surface to postpone the stall.

A stick-pusher is a mechanical device which prevents the pilot from stalling an aeroplane by pushing the controls forwards as the stall is approached.

A stick-shaker is a similar device which shakes the pilot's controls to warn of the onset of stall.

A stall warning is an electronic or mechanical device which sounds an audible warning as the stall speed is approached. The majority of aircraft contain some form of this device that warns the pilot of an impending stall. The simplest such device is a 'stall warning horn', which consists of either a pressure sensor or a movable metal tab that actuates a switch, and produces an audible warning in response.

An angle of attack limiter or an "alpha" limiter is a flight computer that automatically prevents pilot input from causing the plane to rise over the stall angle. Some alpha limiters can be disabled by the pilot.

If a forward canard is used for pitch control rather than an aft tail, the canard is designed to stall at a slightly higher speed than the wing (i.e. the canard stalls first). When the canard stalls, the nose drops, lowering the angle of attack thus preventing the wing from stalling. Thus the wing virtually never stalls.

If an aft tail is used, the wing is designed to stall before the tail. In this case, the wing can be flown at higher lift coefficient (closer to stall) to produce more overall lift

Flight dynamics is the study of orientation of air and space vehicles and how to control the critical flight parameters, typically named pitch, roll and yaw.

Pitch is rotation around the lateral or transverse axis. This axis is parallel to the wings, thus the nose and tail both pitch up or down.

Roll is rotation around the longitudinal axis—an axis drawn through the body of the vehicle from tail to nose. This is also known as bank.

Yaw is rotation about the normal axis—an axis perpendicular to the pitch and roll axes. If an airplane model placed on a flat surface is spun or pivoted around the center of mass (coordinate origin) it would be described as yawing.

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Ailerons are hinged flaps attached to the trailing edge of an airplane wing, usually near the wingtips. They are used to control the aircraft in roll. The two ailerons are interconnected so that one goes down when the other goes up: the downgoing aileron increases the lift on its wing while the upgoing aileron reduces the lift on the other wing, producing a rolling moment about the aircraft's longitudinal axis. The word aileron is French for "little wing."An unwanted side-effect of aileron operation is adverse yaw - a yawing moment in the opposite direction to the turn generated by the ailerons. In other words, using the ailerons to roll an aircraft to the right would produce a yawing motion to the left. The yaw occurs because the down-going aileron will increase the angle of attack of the upgoing wing, increasing both lift and drag (Form + Induced). Conversely, the wing with the upgoing aileron will see a small increase in drag (Form), as well as the intended reduction in lift.

Adverse yaw can be countered with the aircraft's rudder (a co-ordinated turn), but can also be reduced with clever design. If the upgoing aileron moves further upwards than the downgoing aileron moves down, it will create extra profile drag on that wing and try to yaw the aircraft into the turn. This set-up is known as "differential aileron". Another solution is to use a "Frise aileron", where the up going aileron also projects a section downwards below the wing, again increasing drag on the inside of the turn.

Modern airliners tend to have a second set of inboard ailerons much closer to the fuselage, which are used at high speeds. Some aircraft use spoilers to achieve the same effect as ailerons.

Flaps are hinged surfaces on the trailing edge of an airplane wing which, when deployed, increase the lift (and drag) of a wing. They are usually used while landing to allow the aircraft to fly more slowly and to steepen the approach to the landing site.

Types include:

Plain flap - rotates on a simple hinge.

Split flap - upper and lower surfaces are separate, the lower surface operates like a plain flap, but the upper surface stays immobile or moves only slightly.

Fowler flap - slides backwards before hinging downwards, thereby increasing both camber and chord, creating a larger wing surface better tuned for lower speeds.

Slotted flap - systems made up of several individual Fowler flaps, which combine to form a single, much more powerful, flap.

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Blown flaps - systems that blow engine air over the upper surface of the flap at certain angles to improve lift characteristics.

In aviation, a planform is the shape and layout of an airplane's wing. Of all the myriad planforms used, they can typically be grouped into those used for low-speed flight, found on general aviation aircraft, and those used for high-speed flight, found on many military aircraft and airliners.

Low-speed planforms

The primary concern in low speed flight is the aspect ratio, the comparison of the length of the wing measured out from the fuselage, span, compared to the length from front to back, chord. Wings with higher aspect ratios, that is, wings that are longer and skinnier, have lower drag for any given amount of lift than a wing of the same area that is shorter and fatter. This is due to an effect known as induced drag, caused by airflow over the tip of the wing. As the size of the tip decreases compared to the wing's overall size, the magnitude of the induced drag is reduced.

There are other ways to reduce induced drag as well, mostly by changing the shape of the wing to reduce the size of the tip.A practical and simple compromise is to taper the wing towards the tip, a feature that can be found on almost all modern aircraft (including gliders).

High-speed planforms

At higher speeds nearing the speed of sound, a new form of drag appears: wave drag. Wave drag is considerably more powerful than induced drag, and must be avoided at all costs in order to improve performance. Doing so demands a wing that is as thin as possible, with a slowly changing profile over a wide chord. Of course this is basically the opposite goal to low speed wings, which presents a problem.

Just as on the lower speed designs, making the "perfect" high speed planform is difficult for practical reasons. In this case a very thin wing makes it difficult to use the internal room for storage of fuel and landing gear, as well as making the wing considerably less stiff torsionally as well as leading to increased induced drag when flying slower.

Solutions to this problem come in many forms, notably the use of the swept-wing and delta-wing, both of which "fool" the air into thinking it is flowing over a thinner wing with more chord.

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Speed of sound

The speed of sound c (from Latin celeritas, "velocity") varies depending on the medium through which the sound waves pass. It is usually quoted in describing properties of substances (e.g. see the article on sodium).

More commonly the term refers to the speed of sound in air. The speed varies depending on atmospheric conditions; the most important factor is the temperature. The humidity has very little effect on the speed of sound, while the static sound pressure (air pressure) has none. Sound travels slower with an increased altitude (elevation if you are on solid earth), primarily as a result of temperature and humidity changes.

Stagnation pressure

Stagnation pressure is the pressure a fluid exerts when it is motionless. Consequently, although a fluid moving at higher speed will have a lower static pressure, it may have a higher stagnation pressure. Static pressure and stagnation pressure are related by the Mach number of the fluid. In addition, there can be differences in pressure due to differences in the elevation (height) of the fluid. See Bernoulli's equation.

The pressure of a moving fluid can be measured using a Pitot probe, or one of its variations such as a Kiel probe or Cobra probe, connected to a manometer. Depending on where the inlet holes are located on the probe, it can measure static pressure or stagnation pressure.

Shock wave

In fluid dynamics, a shock wave is a nonlinear or discontinuous pressure wave. It can also be when the actual molecular or particle speed is moving faster than the wave propagation speed (space shuttle through air). They can and do transport and transmit tremendous amounts of energy (hundreds of Megawatts per square meter for shocks generated by nuclear explosions).

In compressible fluids such as air, disturbances such as the pressure changes caused by a solid object moving through the medium will propagate through the fluid as pressure waves traveling at the speed of sound. When the cause of the disturbance is moving slowly relative to the speed of sound, the pressure wave takes the form of conventional sound waves. The pressure waves enable the fluid to redistribute itself to accommodate the disturbance, and the fluid behaves similarly to an incompressible fluid.

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However, when a disturbance moves faster than the pressure waves it causes, fluid near the disturbance cannot react to it or "get out of the way" before it arrives. The properties of the fluid (density, pressure, temperature, velocity, etc.) thus change almost instantaneously as they adjust to the disturbance, creating thin disturbance waves called shock waves and shock heating.

Shock waves are not sound waves; a shock wave takes the form of a very thin membrane (sheet of energy) on the order of micro-meters in thickness. The pressure excursion within the shock wave is so extreme that it causes the speed of sound within the wave to change. Shock waves in air are heard as a loud "crack" or "snap" noise. Over time a shock wave can change from a nonlinear wave into a linear wave, degenerating into a conventional sound wave as it heats the air and loses energy. The sound wave is heard as the familiar "thud" or "thump" of a sonic boom, commonly created by the supersonic flight of aircraft.

There are two types of shock waves: normal shocks and oblique shocks. A normal shock extends perpendicular to the flow of fluid, and the flow goes from supersonic upstream of the shock wave to subsonic downstream. An oblique shock is formed at an angle to the flow, and although the component of flow perpendicular to the oblique shock goes from supersonic to subsonic in crossing the wave, the tangent component of flow is not affected, so the net flow may remain supersonic downstream of an oblique shock wave.

A cage around the engine reflects any shock waves. A spike behind the engine converts them into trust.

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To generate lift a supersonic airplane has to produce at least two shock waves: One over-pressure downwards wave, and one under-pressure upwards wave. Withcomb’s area rule states, we can reuse air displacement without generating additional shock waves. In this case the fuselage reuses some displacement of the wings.

Supersonic

Any speed over the speed of sound, which is approximately 343 m/s, 1,087 ft/s, 761 mph or 1,225 km/h in air at sea level, is said to be supersonic. Speeds greater than 5 times the speed of sound are sometimes referred to as hypersonic. The aircraft's design was revolutionary introducing many innovations which are still used on today's supersonic aircraft. The single most important development was the all-moving tailplane which allowed control to be maintained at supersonic speeds;

Tailplane

A tailplane is a small lifting surface located behind the main lifting surfaces of a fixed-wing aircraft.

An aeroplane must be in balance longitudinally in order to fly. This means that the net effect of all the forces acting on the aeroplane produces no overall pitching moment about the centre of gravity. Without a tailplane there would be only one combination of speed and centre of gravity position for which this requirement was met. The tailplane provides a balancing force to maintain equilibrium for different speeds and centre of gravity positions. Because the tailplane is located some distance from the centre of gravity, even the small amount of lift it produces can generate a large pitching moment at the centre of gravity.

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Stability

An aeroplane with a wing only is normally unstable in pitch (longitudinal stability). This means that any disturbance (such as a gust) which raises the nose produces a nose-up pitching moment which tends to raise the nose further. With the same disturbance, the presence of a tailplane produces a restoring nose-down pitching moment which counteracts the natural instability of the wing and make the aircraft longitudinally stable. A stable aeroplane can be flown "hands-off" and will maintain the same altitude and pitch attitude.

Control

A tailplane has a hinged flap called an elevator, which allows the pilot to control the amount of lift produced by the tailplane. This in turn causes a nose-up or nose-down pitching moment on the aircraft, which is used to control the aircraft in pitch.

T-tail

In aircraft a T-tail is an arrangement of the tail control surfaces with the horizontal surfaces (tailplane and elevators) mounted to the top of the fin, rather than the more common location on the fuselage at the base of the fin. The resulting arrangement looks like a T when viewed from the front, hence the name.

Pros

The tailplane surfaces are kept well out of the airflow behind the wing, giving smoother flow, more predictable design characteristics, and better pitch control. This is especially important for planes operating at low speed, where clean airflow is required for control.

The effective distance between wing and tailplane can be increased without a significant increase in the weight of the aircraft. The distance between the two planes gives the "leverage" by which the tailplane can control the aircraft's pitch attitude - with a greater distance, smaller, lighter tailplanes and elevators can be used.

The tail surfaces are mounted well out of the way of the rear fuselage, permitting this site to be used for the aircraft's engines.

Cons

The aircraft will tend to be much more prone to a dangerous deep stall condition, where blanking of the airflow over the tailplane and elevators by a stalled wing can lead to total loss of pitch control. For

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similar reasons, T-tailed aircraft can be much more difficult to recover from a fully-developed spin.

The fin must be made considerably stronger and stiffer to support the forces generated by the tailplane. This inevitably makes it heavier as well.

The control runs to the elevators are more complex. The elevator surfaces are much more difficult to casually inspect from the ground. In aircraft fitted with an ejector seat and a

Elevator (aircraft)

Elevators are control surfaces, usually at the rear of an aircraft, which control the aircraft's orientation by changing the pitch of the aircraft, and so also the angle of attack of the wing. An increased angle of attack will cause a greater lift to be produced by the profile of the wing, and (if no power is added or available), a slowing of the aircraft. A decreased angle of attack will produce an increase in speed (a dive). There may be separate elevators on each side, operating in unison. The elevator or elevators may be the only pitch control surface present, or may be hinged to a fixed or adjustable surface called a stabilizer.In some aircraft the elevator is in the front, ahead of the wing; this type of configuration is called a canard The canard type is more efficient, since the forward surface produces upward lift. The main wing is also less likely to stall, as the forward control surface is configured to stall before the wing, causing a pitch down and reducing the angle of attack of the wing.

Stabilizer (aircraft)

For aircraft, the horizontal stabilizer is a fixed or adjustable surface from which an elevator may be hinged, while a vertical stabilizer (also called a fin) is fixed to the aircraft and supports the rudder. For aircraft with a v-tail each stablizer/fin will support a "ruddervator", combining the functions of the rudder and the elevator.

Rudder

A rudder is a device used to steer a ship or other watercraft. In its simplest form, a rudder is a flat sheet of material attached with hinges to the ship's stern. A tiller - basically, a stick or pole - is attached to the top of the rudder to allow it to be turned in different directions.

Deep stall

Deep stall is a dangerous condition that affects certain aircraft designs, notably those with a T-tail configuration. In these designs,

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the turbulent wake of a stalled main wing "blanks" the horizontal stabilizer, rendering the elevators ineffective and preventing the aircraft from recovering from the stall

Spin (flight)

In flying, a spin is a special case of a stall, with the aircraft descending rapidly and rotating about a vertical axis. It is characterized by low airspeed, a high rate of descent, and high yaw and roll rates. In most aircraft, a spin is a stable condition that will continue until the aircraft descends into the ground unless the pilot takes action to recover from it.

A spining aircraft has a large positive angle of attack (resulting in the stall) and usually a large nose-down pitch angle. A spin can be entered in any attitude, however; a negative angle of attack in a spin is called an "inverted spin". A spin in which the nose is more-or-less level with the horizon is a "flat spin." Flat spins can be very difficult to recover from because there is little or no smooth airflow over the control surfaces. In many aircraft a flat spin is unrecoverable.

A spin will occur if an aircraft is stalled and there is an asymmetric yawing moment applied to it. This yawing moment can be aerodynamic (sideslip angle, rudder, adverse yaw from the ailerons), thrust related (p-factor, one engine inoperative on a multi-engine

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non-centerline thrust aircraft), or from any number of possible sources of yaw.

This may happen during an uncoordinated turn or other maneuver. An aircraft may be deliberately spun for purposes of training, test flying, or aerobatics. A spin is usually entered by flying the aircraft into a stall condition. As the stall point is reached, the rudder (and sometimes opposite aileron) is used to yaw the aircraft. It is a common misconception that the outboard wing is still flying while the inner wing is stalled; in reality both wings are stalled.

Fuselage

In an aircraft, the fuselage is the main body section that holds crew and passengers or cargo. In single engine aircraft it will usually contain an engine, athough in some amphibious aircraft the single engine is mounted on a pylon attached to the fuselage. The fuselage also serves to position control and stabilization surfaces in specific relationships to lifting surfaces, required for aircraft stability and maneuverability.

Fuselages are constructed using three types of structures:

A box truss structure. The structural elements resemble those of a bridge, with emphasis on using linked trianglular elements. The aerodyamic shape is completed by additional elements called formers and stringers and is then covered with fabric and painted. Most early aircraft used this technique with wood and wire trusses and this type of structure is still in use in many lightweight aircraft using welded steel tube trusses. This method is especially suitable for amateur built aircraft kits, where a complete welded truss structure is delivered with the fitting of other components, covering, and finishing completed by the user, as it ensures that a robust, uniform load bearing structure is within the completed aircraft.

A monocoque shell. In this, the exterior surface of the fuselage is also the primary structure. A typical early form of this was built using moulded plywood, where the layers of plywood are formed over a "plug" or within a mold, A later form of this structure uses fiberglass cloth impregnated with polyester or epoxy resin. A simple form of this used in some amateur built aircraft uses rigid expanded foam plastic with a fiberglass covering, eliminating the necessity of fabricating molds, but requiring more effort in finishing.

Semi-monocoqe. This is the preferred method of constructing an all aluminum fuselage. First, a series of formers in the shape of the fuselage cross sections are held in position on a rigid fixture. These formers are then joined with lightweight longitudinal elements called stringers. These are in turn covered with a skin of sheet

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aluminum, attached by riveting or by bonding with special adhesives. The fixture is then disassembled and removed from the fuseleage, which is then fitted out with wiring, controls, and interior equipment such as seats and luggage bins. Most modern large aircraft are built using this technique, but use several large sections constructed in this fashion which are then joined with fasteners to form the complete fuselage. As the accuracy of the final product is determined largely by the costly fixture, this form is suitable for series production, where a large number of identical aircraft are to be produced.

Both monocoque and semi-monocoque are referred to as "stressed skin" structures as all or a portion of the load is taken by the surface covering.

Aerodynamic heating

Aerodynamic heating is the heating of a solid body produced by passage of air or other gases over the body. It is caused by friction and by compression processes and significant chiefly at high speeds.

Mach number

Mach number (Ma) (pronounced as "mack" in International English or "mock" in the American English) is defined as a ratio of speed to the speed of sound in the medium in case. The Mach number is commonly used both with objects travelling at high speed in a fluid, and with high-speed fluid flows inside channels such as nozzles, diffusers or wind tunnels. As it is defined as a ratio of two speeds, it is a dimensionless number. At standard sea level conditions, Mach 1 is 1,225 km/h (765.6 MPH) in the atmosphere.

Since the speed of sound increases as the temperature increases, the actual speed of an object travelling at Mach 1 will depend on the fluid temperature around it.

It can be shown that the Mach number is also the ratio of inertial forces (also referred to aerodynamic forces) to elastic force

Critical mach

Critical mach is a aeronautics term that refers to the speed at which some of the airflow on a wing becomes supersonic. When this occurs the distribution of forces on the wing changes suddenly and dramatically, typically leading to a strong nose-down force on the aircraft. This effect led to a number of accidents in the 1930s and 1940s, when aircraft in a dive would hit critical mach and continue to push over into a steeper and steeper dive. This problem is often lumped in with the catch-all phrase compressibility.

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Wings generate much of their lift due to the Bernoulli effect; by speeding up the airflow over the top of the wing, the air has less density on top than on the bottom, leading to a net upward force. The relative difference in speed is due largely to the wing's shape, so the difference in speed remains a fairly constant ratio over a wide range of speeds.

But if the air speed on the top of the wing is faster than on the bottom, there will be some speed where the air on top reaches the speed of sound. This is the critical mach. When this happens shock waves form on the upper wing at the point where the flow becomes supersonic, typically behind the midline of the chord. Shock waves generate lift of their own, so the lift of the wing suddenly moves rearward, twisting it down. This effect is known as mach tuck.

The actual speed of critical mach varies from wing to wing. In general a thicker wing will have a lower critical mach, because a thicker wing accelerates the airflow more than a thinner one.

Today a compromise design is used, the swept-wing. This design "fools" the air into thinking it's flowing over a thin wing, which is in fact fairly thick. Swept-wings are used on almost all aircraft that fly in the transonic, and is a common feature of almost all airliners and modern fighter aircraft.

It is possible to see the mach line on an airliner visually, as these aircraft fly beyond the critical mach in crusing flight. The shock wave extends vertically from the wing, and the change in density is enough to make it operate as a lens. By looking at straight lines running parallel to the wing you can often spot a discontinuity where the line "jumps". Roads are an excellent marker for this.

Chord (aircraft)

In reference to aircraft, chord refers to the distance between the front and back of a wing, measured in the direction of the normal airflow. These front and back points are referred to as the leading edge and trailing edge.

Standard mean chord (SMC) is defined as wing area divided by wing span.

,

where S is the wing area and b is the span of the wing.

Mean aerodynamic chord (MAC) is defined as

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The ratio of the chord of a wing to its width (or span) is known as the aspect ratio an important indicator of the lift-induced drag the wing will create. In general planes with higher aspect ratios - wide skinny wings - will have less drag. This is why gliders have long wings.

Swept wing

A swept-wing is a wing planform used on high-speed aircraft that spend a considerable portion of their flight time in the transonic. Simply put, a swept-wing is a wing that is bent back at some angle, instead of sticking straight out from the fuselage. They were initially used only on fighter aircraft, but have since become almost universal on all jets, including airliners and business jets. As an aircraft approaches the speed of sound, an effect known as wave drag starts to appear. This happens because the air which would normally follow a streamline around the aircraft no longer has time to "know" about the approaching object and simply hits it directly. This results in greatly increased drag. Research into the nature of this effect led to the conclusion that it was reduced by having the profile of the aircraft change as slowly as possible, what we today refer to as fineness ratio. To account for this in their designs, aerospace engineers use the Whitcomb area rule, leading to long highly-tapered profiles.

When a swept-wing travels at high speed, the airflow has little time to react and simply flows over the wing. However at lower speeds there is more time for motion and a strong streamline, and with the front of the wing angled, some of the air is pushed to the side towards the wing tip. At the wing root, by the fuselage, this has little noticeable effect, but as you move towards the tip the airflow is pushed sidewise not only by the wing, but the sidewise moving air beside it. By the time you reach the tip the airflow is moving along the wing instead of over it, a problem known as spanwise flow.

The problem with spanwise flow is that the lift of the wing is generated by the airflow over it from front to rear. As an increasing amount travels spanwise, the amount flowing front to rear is

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reduced, leading to a loss of lift. Normally this is not much of a problem, but as the plane slows for landing the tips can actually drop below the stall point even at speeds where stalls should not occur. When this happens the tip stalls, and since the tip is swept to the rear, the net lift moves forward. This causes the plane to pitch up, leading to more of the wing stalling, leading to more pitch up, and so on.

The solution to this problem took on many forms. One was the addition of a strip of metal known as a wing fence on the upper surface of the wing to redirect the flow to the rear (see the MiG-15 as an example), another closely related design was to add a dogtooth notch to the leading edge (Avro Arrow). Other designs took a more radical approach, including the XF-91 Thunderceptor's wing that grew thicker towards the tip to provide more lift there, and the British-favoured compound sweep or scimitar wing that reduced the sweep along the span, used on their V Bombers.

Modern solutions to the problem no longer require "custom" designs such as these, but are taken as a whole with the need for shorter takeoff and landing than the early large jets. The addition of leading edge slats and large compound flaps to the wings have largely resolved the issue. On fighter designs, the addition of leading edge extensions, included for high manoeuvrability, also serve to add lift during landing and reduce the problem.

The swept-wing also has several more mundane problems. One is that for any given length of wing, the actual span from tip-to-tip is shorter than the same wing that isn't swept. Low speed drag is strongly correlated with the aspect ratio, the span compared to chord, so a swept wing always has more drag at lower speeds. Another concern in the torque generated at the fuselage, as much of the wing's lift lies behind where the root connects to the plane. Finally, while it is fairly easy to run the main spars of the wing right through the fuselage in a straight wing design to use a single continuous piece of metal, this is not possible on the swept wing because the spars will meet at an angle.

Streamline

In fluid dynamics, a streamline is a line which is everywhere tangent to the velocity of the flow. This can be contrasted with a pathline, which is the trajectory that an imaginary infinitesimally small point would make if it followed the flow of the fluid in which it was embedded, and a streakline, which is the current location of all fluid particles that have passed through a particular spatial point in the past. In steady (time-independent) flow, the streamlines, pathlines, and streaklines coincide. A scalar function whose contours define the streamlines is known as the streamfunction.

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Streamlines are frame-dependent. That is, the streamlines observed in one inertial reference frame are different from those observed in another inertial reference frame. For instance, the streamlines in the air around a aircraft wing are defined differently for the passengers in the aircraft than for an observer on the ground. When possible, fluid dynamicists try to find a reference frame in which the flow is steady, so that they can use experimental methods of creating streaklines to identify the streamlines. In the aircraft example, the observer on the ground will observe unsteady flow, and the observers in the aircraft will observe steady flow, with constant streamlines.

By definition, streamlines defined at a single instant in a flow do not intersect. They cannot begin or end inside the fluid.

A region bounded by streamlines is called a stream tube. Because the streamlines are tangent to the flow velocity, fluid that is inside a stream tube must remain forever within that same stream tube.

Knowledge of the streamlines can be useful in fluid dynamics. For example, Bernoulli's principle, which expresses conservation of mechanical energy, is only valid along a streamline. Also, the curvature of a streamline is an indication of the pressure change perpendicular to the streamline. The instantaneous center of curvature of a streamline is in the direction of increasing pressure, and the magnitude of the pressure gradient can be calculated from the curvature of the streamline.

Engineers often use dyes in water or smoke in air in order to see streaklines, and then use the patterns to guide their design modifications, aiming to reduce the drag. This task is known as streamlining, and the resulting design is referred to as being streamlined. Streamlined designs, like steam locomotives, streamliners and human bodies are often esthetically pleasing to the eye. The Streamline Moderne style, an 1930s and 1940s offshoot of Art Deco, brought flowing lines to architecture and design of the era.

The same terms have since become common vernacular to describe any process that smooths an operation. For instance, it is common to hear references to streamlining a business practice, or operation.

Control reversal

Control reversal is an adverse affect on the controllability of aircraft. To the pilot it appears that the controls have reversed themselves; in order to roll to the left, for instance, they have to push the control stick to the right, opposite of the normal direction.

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There are several causes for this problem: pilot error, effects of high speed flight, incorrectly connected controls, and various coupling forces on the aircraft.

Pilot error is the most common cause of control reversal. In unusual attitudes it is not uncommon for the pilot to become disoriented and start feeding in incorrect control movements in order to regain level flight. This is particularly common when using helmet mounted display systems, which introduce graphics that remain steady in the pilot's view, notably when using a particular form of attitude display known as an inside-out display.

Incorrectly connected controls is another common cause of this problem. It is a recurring problem after maintenance on aircraft, notably homebuilt designs that are being flown for the first time after some minor work. However it is not entirely uncommon on commercial aircraft, and has been the cause of several near-accidents.

Another version of the problem occurs when the amount of airflow over the wing becomes great enough that the force generated by the ailerons is enough to twist the wing itself. For instance when the aileron is deflected upwards in order to make that wing move down, the wing twists in the opposite direction. The net result is that the airflow is directed down instead of up and the wing moves upward, opposite of what was expected. This form of control reversal is often lumped in with a number of "high speed" effects as compressibility.

Wing fence

A Polish Sukhoi Su-20, with large wing fences on inner wings.

Wing fences, also known as boundary layer fences and potential fences are fixed aerodynamic devices attached to aircraft wings. Wing fences are flat metal plates fixed to the upper surfaces (and often wrapping around the leading edge) parallel to the airflow. They work by obstructing the cross flow along the the wing and preventing the entire wing from stalling at once. They are commonly seen on swept-wing aircraft, as they remedy the stall characteristics of swept

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wings.

Critical Mach number

The Critical Mach number (Mcr) is the maximum Mach number (airspeed in relation to the speed of sound - Mach 1.0) which a subsonic aircraft can attain whilst still remaining controllable by the pilot.

At the Critical Mach number, local airflow over the airframe reaches the speed of sound (due to the airflow speeding-up to go around various curvatures in the aircraft structure) and creates shock waves sufficient to affect the airflow over the control surfaces, resulting in a loss of control, although the aircraft itself may still be flying subsonically.

Vortex generator

After-market Micro Dynamics vortex generators mounted on the wing of a Cessna 182K

A vortex generator is an aerodynamic surface, basically a small vane, that creates a vortex. They can be found in many devices, but the term is most often used in aircraft design.

Vortex generators are added to the front of a swept-wing in order to maintain steady airflow over the control surfaces at the rear of the wing. They are typically rectangular or triangular, about a centimetre or two in size, and run in lines chordwise at about the thickest part of the wing. They can be seen on the wings and vertical tails of many airliners.

The purpose of the generators are to stick out of the stagnant air near the surface of the wing, and into the freely moving air outside the boundary layer. This layer is typically quite thin, but dramatically reduces speed of the airflow towards the rear of the wing. The generators mix the free stream with the stagnant air to

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get it moving again, providing considerably more airflow at the rear of the wing and thereby providing the control surfaces with more power. This process is typically referred to as re-energizing the boundary layer.

Air jet vortex generators work on a different principle. They direct a jet of air into the boundary layer, thereby re-energising it.

Wingspan

The wingspan (or just span) of an airplane is the distance from the left wingtip to the right wingtip. For example, the Boeing 777 has a wingspan of about 60 m (200 feet). Planes with a longer wingspan are generally more efficient because they suffer less induced drag and their wingtip vortices do not affect the wing as much. However, the long wings mean that the plane has a greater moment of inertia about its longitudinal axis and therefore cannot roll as quickly and is less manouverable. Thus, combat aircraft and aerobatic planes usually opt for shorter wingspans to increase manouverability. Since the amount of lift that a wing generates is proportional to the area of the wing, planes with short wings must correspondingly have a longer chord. An aircraft's ratio of its wingspan to chord is therefore very important in determining its characteristics, and aerospace engineers call this value the aspect ratio of a wing.

Dihedral

In geometry, the dihedral is the angle between two planes. See dihedral angle.

Dihedral is the upward angle of an aircraft's (or bird's) wings from root to tip, as viewed from directly in front of or behind the aircraft. Downward angled wings are said to have anhedral.

Dihedral on the wings and tailplane of a Boeing 737

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The purpose of dihedral is to confer stability in the roll axis. A popular but erroneous explanation for how it works is that if the aircraft is perturbed such that one wing is lowered relative to the other, dihedral causes the lower wing to increase its surface area relative to the airflow, thus increasing its lift. This acts to oppose the original roll motion. An alternative way to visualize this is to imagine that the aircraft is sitting in the bottom of a shallow V-shaped "slot" in the air, thanks to the angle of the wings. This position is naturally stable. This explanation is often put forward in many books "explaining" aeronautical principles, but it is wholly false.

The true explanation for the action of dihedral is this: If a disturbance causes an aircraft to roll away from its normal position, the aircraft will sideslip in the direction of the down-going wing. This creates an airflow component along the length of the wing from tip to root. The dihedral angle can be seen as presenting a positive angle of attack to this lateral flow, hence generating some additional lift. It is this lift which restores the aircraft to its normal attitude. The apparent increase in surface area is in fact an illusion and contributes no additional lift.

Most aircraft in the civilian or transport sector use dihedral for stability. Military combat aircraft, in contrast, often have flat wings or anhedral. This reduces inherent stability but increases manoeuvrability. Many military aircraft are in fact inherently unstable, and only fly due to the constant vigilance of on-board computers.

Anhedral on a Harrier GR7

A side effect of dihedral can be roll-coupling, a tendency for an aircraft to "corkscrew" through the air under certain conditions. This rolling motion, called a dutch roll, is unpleasant to experience for those flying, and can lead to loss of control or can overstress an aircraft. A certain amount of anhedral can combat this effect. Pronounced anhedral is also often seen on aircraft with a high mounted wing, such as the BAe 146, Lockheed Galaxy and others. In

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such designs, the high mounted wing itself confers roll stability (due to the pendulum effect of the fuselage, engines, etc), so additional dihedral is not required. In fact such designs can be excessively stable, so the anhedral is added to cancel out some of the roll stability to ensure that the aircraft can be easily manoeuvred.

An alternative to dihedral for the wing as a whole is to cant the wingtips or outer section of the wing upwards instead. This has the same effect. It is commonly seen on gliders, and some other aircraft. The McDonnell Douglas F-4 Phantom II is one such example, unique among fighters for having dihedral wingtips.

Dutch roll

Dutch roll is one of an aircraft's flight dynamic modes (others include phugoid, short period, and spiral divergence). It involves a coupling of roll and yaw which is normally well damped in most light aircraft. Some aircraft with well-damped dutch roll modes can experience a degradation in damping as airspeed and altitude increase. Dutch roll stability can be artificially increased by the installation of a yaw-damper (commonly referred to incorrectly as a yaw-dampener).

The Dutch Roll mode can be excited by any use of aileron or rudder, but for flight test purposes it is usually excited with a rudder doublet or singlet. Some larger aircraft are better excited with aileron inputs. Periods can range from a few seconds for light aircraft to a minute or more for airliners.

The name comes from the movement that (Dutch) skaters make when skating on ice.

Dutch roll is also the name (considered by professionals to be a misnomer) given to a coordination maneuver generally taught to student pilots to help them improve their crosswind-landing technique. The airplane is alternately rolled as much as 60-degrees left and right while opposite rudder is applied to keep the nose of the airplane pointed at a fixed point. (This technique is more commonly referred to as a slip. If the airspeed is allowed to decay the aircraft can stall, and the crossed controls can cause it to spin.)

Phugoid

A phugoid is one of the flight dynamics modes of an aircraft (others include short period, dutch roll, and spiral divergence). It consists of a (theoretically) constant angle of attack exchange of airspeed and altitude. It can be excited by an elevator singlet resulting in a pitch increase with no change in trim from the cruise condition. As speed decays, the nose will drop below the horizon. Speed will increase, and the nose will climb above the horizon. Periods can vary from

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under 30 seconds for light aircraft to more than a few minutes for larger aircraft. Microlight aircraft typically show a phugoid period of 15-25 seconds, and it has been suggested that birds and model aeroplanes show convergence between the phugoid and short period modes. A classical model for the phugoid period can be simplified to about (0.85 x speed in knots) seconds, but this only really works for larger aircraft.

Commonly known as porpoising, phugoids are often demonstrated to student pilots as an example of the speed stability of the aircraft and the importance of proper trimming. When it occurs, it's a pure nuisance mode, and in lighter aeroplanes (typically showing a shorter period) it can be a cause of Pilot Induced Oscillation, or PIO.

An interesting characteristic of the phugoid is that it occurrs at effectively constant Angle of Attack (AoA), although in practice AoA actually varies by a few tenths of a degree. This means that the stalling AoA is never exceeded, and it's possible (in the <1g section of the cycle) to fly at speeds below the known stalling speed.

The name apparently is an example of poor Latin translation by Lanchester, a British aerodynamicist who first predicted it. The Latin verb for "to flee" (fugio) was used when what was desired was the Latin verb for "to fly".

For a dramatic example of phugoids, read about the United Airlines Flight 232 incident, where an engine failure caused total hydraulic system failure. The crew flew the aircraft with throttle only. Suppressing the phugoid tendency was particularly difficult.

Airfoil

An airfoil (in American English, or aerofoil in British English) is the shape of a wing or blade (of a propeller or ship's screw or sail) as seen in cross-section. It is used to provide lift or downforce, depending on its application. Subsonic-flight airfoils have a characteristic shape with a rounded leading edge, followed by a sharp trailing edge, and often with camber.

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The historical evolution of airfoil sections, 1908 - 1944, NASA

Lift and Drag curves for a typical airfoil

An inverted airfoil will create a downward pressure on an automobile or other motor vehicle, improving its traction and keeping it on the ground. The term "lift" can mean a force generated in any direction in any medium. Any thin object with a positive angle of attack, such as a flat plate or the deck of a bridge, will generate lift. Airfoils though are more efficient, generating lift with the least drag. A lift and drag curve obtained in wind tunnel testing is show on the right.

Airfoil design is a major facet of aerodynamics. Various airfoils serve different flight regimes. A supercritical airfoil, with its low camber, reduces transonic drag divergence, while a symmetric airfoil may better suit frequent inverted flight. Supersonic airfoils are much more angular in shape and can have a very sharp leading edge. While sharper leading edged airfoils produce stiffer and lighter wings, large rounder edges increase wing volume for fuel. Moveable high-lift devices, flaps and slats are fitted to airfoils on most aircraft. New airfoil design techniques continue to develop.

Various systems have been devised to describe and characterise airfoils — the most common and prevalent is the NACA system. Before this, various ad-hoc systems were used. An example of a general purpose airfoil that finds wide application, and predates the NACA system is the Clark-Y.

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Slats are small aerodynamic surfaces on the leading edge of an airplane wing which, when deployed, allow the wing to operate at a higher angle of attack. Lift is a product of angle of attack and speed, so by deploying slats an aircraft can fly slower or take off and land in a shorter distance. They are usually used while landing or performing manoeuvres which take the aircraft close to the stall, but are usually retracted in normal flight to minimise drag.

The position of the leading edge slats on an airliner (Airbus A310). In this picture, the slats are extended.

Types include:

Automatic - the slat lies flush with the wing leading edge until reduced aerodynamic forces allow it to extend by way of springs when needed. This type is typically used on light aircraft.

Fixed - the slat is permanently extended. This is rarely used, except on specialist low-speed aircraft (see: slot).

Powered - the slat extension can be controlled by the pilot. This is commonly used on airliners.

The chord of the slat is typically only a few percent of the wing chord. They may extend over the outer third of the wing or may cover the entire leading edge. Slats work by increasing the camber of the wing, and also by opening a small gap (the slot) between the slat and the wing leading edge, allowing a small amount of high-

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pressure air from the lower surface to reach the upper surface, where it helps postpone the stall.

The slat has a counterpart found in the wings of some birds, the alula – a feather or group of feathers which the bird can extend under control of its "thumb".

Subsonic aerodynamics

In a subsonic aerodynamic problem, all of the flow speeds are less than the speed of sound. This class of problems encompasses nearly all internal aerodynamic problems, as well as external aerodynamics for most aircraft, model aircraft, and automobiles.

In solving a subsonic problem, one decision to be made by the aerodynamicist is whether or not to incorporate the effects of compressibility. Compressibility is a description of the amount of change of density in the problem. When the effects of compressibility on the solution are small, the aerodynamicist may choose to assume that density is constant. The problem is then an incompressible problem. When the density is allowed to vary, the problem is called a compressible problem. In air, compressibility effects can be ignored when the Mach number in the flow does not exceed 0.3. Above 0.3, the problem should be solved using compressible aerodynamics.

Transonic aerodynamics

Transonic aerodynamic problems are defined as problems in which both supersonic and subsonic flow exist. Normally the term is reserved for problems in which the characteristic Mach number is very close to one.

Transonic flows are characterized by shock waves and expansion waves. A shock wave or expansion wave is a region of very large changes in the flow properties. In fact, the properties change so quickly they are nearly discontinuous across the waves.

Transonic problems are arguably the most difficult to solve. Flows behave very differently at subsonic and supersonic speeds, therefore a problem involving both types is more complex than one in which the flow is either purely subsonic or purely supersonic.

Supersonic aerodynamics

Supersonic aerodynamic problems are those involving flow speeds greater than the speed of sound. Calculating the lift on the Concorde during cruise can be an example of a supersonic aerodynamic problem.

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Supersonic flow behaves very differently from subsonic flow. The speed of sound can be considered the fastest speed that "information" can travel in the flow. Gas travelling at subsonic speed diverts around a body before striking it, it can be said to "know" that the body is there. Air cannot divert around a body when it is travelling at supersonic speeds. It subsonic flow and a diffuser in supersonic flow). Subsonic flow additional shock waves. In this case the fuselage reuses some displacement of the wings. ]]

C.P.C.I.C.F. from ATR MANUAL2.2. CORROSION CLASSIFICATIONThe corrosion is the deterioration of a metal due to chemical or electrochemical action that converts it into metallic compound such oxide, hydroxide or sulfate.The corrosion is more likely to occur or to spread as the airplane ages and, if not controlled, can reduce the capability of the structure to carry the required loads.Four conditions must exist before corrosion can occur:a. Presence of a metal tending to corrode (anode).b. Presence of dissimilar conductive material (cathode) with less tendencyto corrode.c. Presence of conductive liquid.d. Electrical contact between the anode and the cathode.The elimination of any of these conditions will stop corrosion.

2.2. CORROSION CLASSIFICATIONCorrosion damage classifications are defined as follows for a given area of the aircraft and a given operating environment:NO CORROSION· No corrosion findings between successive inspections.MILD CORROSION· Corrosion findings characterized by discoloration or pitting with depth of approximately 0.025mm (0.001 inch).· Mild corrosion findings whose cumulative material during successive inspection exceeds the allowable damage limits.· Corrosion findings due to accidental causes as corrosive liquid spillage or others.NOTE: ALLOWABLE DAMAGE LIMITS ARE GIVEN IN THE APPROPRIATE SRM CHAPTERS.LARGER ALLOWABLE LIMITS MAY, HOWEVER, BE APPROVED FOR PARTICULAR CASES.MODERATE CORROSION· Corrosion findings similar to mild corrosion but with some blister or flaking. Depth of the corrosion may be as deep as 0.25 mm (0.01 inch) maximum.NOTE: FINDINGS OF THE "MODERATE CORROSION" TYPE OUT OF ALLOWABLE DAMAGE LIMITS SHOULD PROMOTE ACTION ON THE OTHER AIRCRAFT OF THEOPERATOR FLEET AND BE REPORTED TO THE

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MANUFACTURER.

SEVERE CORROSION· Corrosion findings with general appearance as moderate corrosion but with severe blistering exfoliation or flaking. Depth of the corrosion is greater than 0.25 mm (0.01 inch).NOTE: FINDINGS OF THE "SEVERE CORROSION" TYPE SHOULD PROMOTE IMMEDIATE ACTION ON THE OTHER AIRCRAFT OF THE OPERATOR FLEET AND BE REPORTED TO THE MANUFACTURER AND REGULATORY AUTHORITY2.3. CORROSION PRONE AREASome areas of the aircraft are prone to corrosion due to the particular structural detail, to dissimilar metal, build-up of moisture, engine exhaust gas deposit, accumulation of water, debris, loose fasteners, hydraulic fluids, ineffective drain holes plugged by dirt, grease, abrasion, etc.The main areas prone to corrosion are· Door areas.· Lavatories, galley and luggage compartment understructure.· Internal surface of fuselage lower panels.· Landing gear wheel wells.· Joint with steel, C.RE.S, nickel or titanium fasteners.· Batteries compartment.· Rear pressure bulkhead.· Electrical connectors.· Lap joints and butt joints.· Trailing edge. open areas.These areas should be checked for corrosion whenever possible and the causes that favour corrosion eliminated (water accumulation, spillage of any kind, dirt, plugged drain holes, etc.).2.5. CORROSION INSPECTION AND MEASUREMENTThe extent and the depth of any corrosion must be clearly identified. The visual inspection is the most common means to detect corrosion If the visual inspection is deemed not effective (hidden corrosion suspected) adequate inspection technique or disassembling shall be performed.The most common means of inspection, other than visual inspection, to detect corrosion are:Eddy current inspection.Eddy current (primarily low frequency) can be used to detect thinning due to corrosion and cracks in multilayered structure. Low frequency eddy current can be used to detect corrosion in underlying structure because the Eddy current will penetrate in the second layer with sufficient sensitivity for approximate results.X-RAY inspection.The X-RAY technique is effective for severe or moderate-to severe corrosion but its use is limited for mild-to-moderate corrosion. In any case X-RAY requires qualified and certified personnel to obtain reliable results.Ultrasonic inspection

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Ultrasonic inspection provides sensitive detection capability for corrosion damage detection when access is available to a surface with a continuous bulk of material exposed to corrosion. Ultrasonic inspection is commonly used to detect exfoliation, stress corrosion cracks and general thinning of material. Trained personnel must conduct the examination if any useful information has to be derived from indicating devices. Use of calibration block may be required.

If the corrosion findings are MODERATE CORROSION, the Corrosion Control and Prevention Program is considered not effective. Adjustment (decrease of the threshold/interval of inspections), based on the specific experience, shall be made to theProgram to maintain the corrosion findings during successive inspections at MILD CORROSION or NO CORROSION.· If the corrosion findings are SEVERE CORROSION, immediate actions to define the causes of such corrosion shall be implemented, inspection extended to the rest of the fleet shall be performed and adjustment to the Corrosion Control and Prevention Program maybe made to bring the level of corrosion between successive inspections at MILD CORROSION or NO CORROSION.· Depending of the corrosion occurrence related to the environment and operational severity conditions, the frequency of the inspections given in the Baseline Maintenance Program shall be adjusted on the base of the operator specific experience.CREVI CESYMPTOMS:Severe local corrosion along faying surface.CAUSE:Penetration of oxygen and corrosive agent into a joint.PREVENTION:Efficient sealing of faying surfaces from corrosiveve substances.FRETTI NGSYMPTOMS:Destruction of natural protective film over large surfaces and loss of metal from surface followed by dark coloured oxidation.CAUSE:Abrasion of metal under load in humid environmental conditions.PREVENTION:Detail design and protective treatment, material selection.GALVANI CSYMPTOMS:Powder-like white or grey deposits.CAUSE:Two dissimilar metals in contact.PREVENTION:Detail design, protective treatment, special assembly techniques. (Sealing, electrical insulation of metals).PI TTI NGSYMPTOMS:Holes in metal surface.

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CAUSE:Halogen ions present in attacking electrolyte (corrosive agent), destroying surface treatment.PREVENTION:Protective treatment.I NTERGRANULARSYMPTOMS:Normally only perceived by cracking.CAUSE:Chemical action along grain boundaries within the material Difference in electrical potential between grain and grain boundaries.PREVENTION:Material selection and protective treatment.

EXFOLI ATI ONSYMPTOMS:Flaking and loss of metal thickness.CAUSE:Swelling and flaking at grain ends exposed by machining.PREVENTION:Pre-heat treatment and material selection.FILI FORMSYMPTOMS:Paint bulging and longitudinal propagation of blisters on surface.CAUSE:Paint damage.PREVENTION:Corrosion resistant primer, restoration of paint system.

MI CRO- BI OLOGI CALSYMPTOMS:Local surface attack or formation of deposits such as fungi.CAUSE:Growth of micro-organisms in moisture traps.PREVENTION:Detail design, protective treatment and assembly techniques, use of inhibitors in primers, etc.STRESSSYMPTOMS:Normally only perceived by cracking with fast crack propagation leaving bare metal subject to corrosion.CAUSE:Residual stress from manufacturing process or stress concentrations due to design features.PREVENTION:Material selection and handing care, detail design and assembly techniques, background surface protection.5. REMOVAL OF CORROSIONC. Operating procedure

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(1) Surface preparation· Masking of non-corroded adjacent areas by installing plastic screens secured with adhesive tape.· Cleaning and degreasing with cotton cloth moistened in solvent.· Paint stripping:- Chemical stripping (external areas, isolated parts, areas with fixed boundaries).Use paint stripper COMORCAP B7 or equivalent stripping shall be followed by rinsing with solvent.- Mechanical stripping: (in box-type structures or difficult-toreach areas) use emery cloth or scotchbrite pads.Rinse off with clear water.(2) Removal of corrosionThis operation comprises two successive phases:(a) Mechanical action (See SRM Chapter 51-21-58)Use of steel wire brushes is prohibited.· Rub down using scotchbrite pads or emery cloth. Grade selection will depend on corrosion extent. Finish off using the finest grade.· For extensive corrosion use nylon brushes (recommended), cutters, grinding wheels.NOTE 1: Take care not to heat the surface by an excessive rotation speed.NOTE 2: Refer to specific SRM chapters for information concerning permissible damage depth.NOTE 3: For large surfaces and in the event of filiform corrosion, VACUBLAST treatment with glass beads is recommended.b) Chemical action (Ref. NOTE below)· Remove dust and degrease with solvent.· Using a brush, apply chromic acid anhydride solution (chromic acid (10%) + demineralized water) to the damaged area.· Allow solution to act for 10 mn.· Rinse off with demineralized water and rub down with a nylon brush to eliminate dark yellow coloration.· Wipe off with cloth.Dry off with dry oil-free air.NOTE: If rinsing cannot be performed correctly and in cases where the chromic acid anhydride solution may reach inaccessible structural stackings, do not apply chemical action.(3) Checking of corrosion removalUsing a magnifying glass, check for evidence of corrosion and for presence of cracks.In doubt, conduct dye-penetrant or Eddy-current inspection.(4) Neutralization· Apply potassium dichromate solution (3% + demineralized water) with brush.· Allow solution to act until surface is dry.· Rinse off with demineralized water and rub down with a nylon brush to eliminate yellow traces.(5) Final step

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Overall rinsing with demineralized water and drying off with dry oilfree air.(6) Paint touch-upThe time elapsed between the drying phase and paint application shall be as short as possible.If this operation is delayed, the area treated against corrosion shall be protected to prevent any external contamination.· Perform paint touch-up as follows:- Wash primer P99 or A 166 ASTRAL/SIKKENS.NOTE:Wash primer can be replaced by application of Alodine 1200.In this case, after Alodine application rinse off with clear water (do not rub) and dry off with dry oil-free air.· Apply protective finish scheme relevant to the zone:- primer,- top coat,- livery.5.3. TREATMENT OF CORROSION ON STEELSAccomplishment of this process is applicable to steel types having the following characteristics:· Low-alloy steels with main added element < 5% and total alloy elements < 10%.· Bare metal condition (no cadmium plating, no chop process).· Tensile strength < 130 hb or 1300 MPa.A. Equipment and materials· Non-metallic pad (scotchbrite very fine grade).· Abrasive paper (fine 400 grade).· Deoxidizing agents (equivalent products):- Rust removing phosphating agent DERCAM SARL DERCAM- Chlorinated or ketonic solvent (MEK, baltane).B. Operating procedure(1) Surface preparation· Masking of non-corroded adjacent areas by installation of plastic screens secured with adhesive tape.· Mechanical removal of corrosion by rubbing with scotchbrite and abrasive pads. (Ref. SRM 51-21-58).(2) Deoxidation· Remove dust and clean with solvent.· Brush apply either DERCAM or ARDROX Type 140 (these two products are ready for use) and rub well into surface using scotchbrite pad. If part is removable, immerse it in the deoxidizing solution.· Allow to act for 30 min. and repeat application.IMPORTANT NOTE:Within one hour, the action of the product causes an attenuation of the corrosion but the attack is not sufficient to reach the uncorroded metal substrate. Several hours are then necessary to obtain complete neutralization of corrosion products and superficial phosphating of metal (white or greyish film).(3) Checking of corrosion removal and phosphating action

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· Using suitable light source, make certain that all areas have been treated.(4) Final steps· Overall rinsing off with demineralized water.· Drying off with dry oil-free air.(5) Paint applicationAs the phosphate film is a treatment for painting, apply suitable protection scheme to the component.NOTE: Installation of part (sealed or with PR sealant) or application of protective finishing shall be performed immediately after corrosion treatment.

Aluminum alloys (2024 and 7075) are generally used on ATR, with different heat treatment.The 7000 series aluminum alloys used on the ATR are heat treated to produce either T6, T73 or T76 tempers. T73 has excellent stress corrosion and exfoliation corrosion resistance. T76temper has strength between T73 and T6, with high exfoliation corrosion resistance and intermediate stress corrosion resistance.8. PROTECTION8.1. PAINT SYSTEMSSURFACES TREATMENTS:Aluminum alloys = The metal is generally electro chemically anodized with Chromic Acid Anodizing (CAA) or with Phosphoric Acid Anodizing (PAA) Clad parts in thefuselage are treated with alodine 1200 or wash primer (Chemical Conversion Coating - CCC).Titanium alloy = This totally corrosion resistant metal is not protected unless in contact with a different material and in this case, it is sand blasted and painted. Small parts aretreated by sulfuric acid anodizing or ionic vapor deposit processes.Steel = Generally, cadmium plating.FOR CONTACT SURFACES:· No relative motion:- Corrosion resistant steel/corrosion resistant steel: passivation,- Alloy steel/alloy steel: cadmium plating,- Steel/aluminum: cadmium plating of the steel part,- Aluminum alloy/carbon fiber:. glass fiber/kevlar/tedlar layer and primer on the carbon part, . anodizing + primer on the aluminum alloy part, . wet assembly of fasteners + interfay sealant.· Relative motion: chromium plating of the steel part.Parts not accessible to cadmium plating: phosphating, plus grease or oil protection.Stainless steels: no protection, except for contacts with other materials.IN ADDITION:· All bonded parts are treated by phosphoric acid anodizing (instead of being only pickled as was common practice on last generation aircraft).

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· The wing tanks are treated by chromic acid anodizing and painted (primer loaded with chromate).

Turbojet

In a turbojet all of the air passing through the engine goes through the combustion chambers. Generally turbojets are arranged around a central shaft, running the length of the engine, with the compressor and turbine connected to the shaft at opposite ends. In the middle of the engine is a combustion area, typically in the form of a number of individual "flame tubes" or "cans". The combustion area is either annular or can-annular (a series of burner cans arranged in a ring), with annular predominating in larger more modern engines

Compressor

The compressor adds energy to the air flow, at the same time squeezing it into a smaller space (increasing its pressure), slowing it down, and increasing its temperature

Fuel burning

The burning process in the cans is significantly different from that in a piston engine. In the piston engine the burning gases are confined to a small volume and, as the fuel burns, the pressure increases dramatically. In a turbojet the air and fuel mixture passes, unconfined, through a can. As the mixture burns its volume increases dramatically and the pressure actually decreases (in the convergent duct) as the gases accelerate towards the rear of the engine.

In detail, the fuel-air mixture must be brought almost to a stop so that a stable flame can be maintained, this occurs just after the beginning of the combustion chamber. The aft part of this flame front is allowed to progress rearward in the engine. This ensures that the rest of the fuel is burned as the flame becomes hotter when it leans out, and because of the shape of the combustion chamber the flow is accelerated rearwards. Some pressure drop is unavoidable, as it is the reason why the expanding gases travel out the rear of the engine rather than out the front. Less than 25% of the air is involved in combustion, in some engines as little as 12%, the rest acting as a reservoir to soak up the heating effect of the fuel burning.

Temperature

Another difference between piston engines and jet engines is that the peak flame temperature in a piston engine is experienced only

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momentarily, and for a small portion of the entire cycle. The can in a jet engine is exposed to the peak flame temperature continuously and operates at a pressure high enough that a stoichiometric fuel-air ratio would melt the can and everything downstream. Instead, jet engines run a very lean mixture, so lean that it would not normally support combustion. A central core of the flow is mixed with enough fuel to burn readily. The cans are carefully shaped to maintain a layer of fresh unburned air between the metal surfaces and the central core. This unburned air mixes into the burned gases to bring the temperature down to something the turbine can tolerate.

Turbine

After the cans, the gases are allowed to expand through the turbine. In the first stage the turbine is largely a reaction turbine (similar to a pelton wheel) and rotates because of the impact of the hot gas stream. Later stages are convergent ducts that accelerate the gas rearward and gain energy from that process. Pressure drops, and energy is transferred into the shaft. The turbine's rotational energy is used to drive the compressor to compress the intake air and some shaft power is extracted to drive accessories like fuel, oil, and hydraulic pumps. The pressure drop through the turbine is much lower than the pressure rise through the compressor because the flow volume in the turbine is so much higher (since fuel has been added), which in turn is due to the higher temperature. In a turbojet almost two thirds of all the power generated by burning fuel is used by the compressor to compress the air for the engine.

Temperature

The efficiency of a jet engine is strongly dependent on the pressure drop through the turbine and nozzle. To achieve the largest possible drop, the engine operates at the highest possible compression ratio. Higher compression ratios imply higher compressor outlet temperatures and thus higher flame temperatures. The tolerable temperature limit is set by the turbine blades—usually the first stage. Modern turbine blades are single crystal metals with hollow interiors. Cooler air from the compressor is blown through the hollow interior of the blades. In a modern engine the turbine inlet temperature will typically be around 1,700 °C, higher than the melting temperature of the blade material (around 1,600 °C). Even higher temperature operation will require not only better materials but also some means of eliminating the oxides of nitrogen that form at such high combustion temperatures.

After the turbine, the gases are allowed to expand and accelerate further through the exhaust nozzle. In some turbojets the gases may actually transition to supersonic flow in the nozzle, in which case the nozzle will be a converging-diverging nozzle. A subsonic nozzle converges all the way to the end. Some supersonic military jets have

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variable nozzles that can change from subsonic to supersonic flow in different flight regimes.

Theory of operation of Turbine

A working fluid contains potential energy (pressure head) and kinetic energy (velocity head). The fluid may be compressible or non-compressible. Several physical principles are employed by turbines to collect this energy;

Silicon nitride turbine wheel for use in small turbogenerators

Impulse turbines change the direction of flow of a high velocity fluid jet. The resulting impulse spins the turbine and leaves the fluid flow with diminished kinetic energy. There is no pressure change of the fluid in the turbine blades. Pressure head is changed to velocity head by accelerating the fluid with a nozzle, prior to hitting the turbine blades. Pelton wheels and de Laval turbines use this concept. Impulse turbines do not require a pressure casement around the runner, since the fluid jet is prepared by a nozzle prior to hitting the turbine. Newton's second law describes the transfer of energy for impulse turbines.

Reaction turbines develop torque by reacting to the fluid's pressure or weight. The pressure of the fluid changes as it passes through the turbine. A pressure casement is needed to contain the working fluid as it acts on the turbine runner, or the turbine must be fully immersed in the fluid flow (wind turbines). The casing contains and directs the working fluid, and for water turbines, maintains suction imparted by the draft tube. Francis turbines and most steam turbines use this concept. For compressible working fluids, multiple turbine stages may by used to efficiently harness the expanding gas. Newton's third law describes the transfer of energy for reaction turbines.

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Turbine designs will use both these concepts to varying degrees whenever possible. Wind turbines use a foil to generate lift from the moving fluid and impart it to the rotor (this is a form of reaction), they also gain some energy from the impulse of the wind, by deflecting it at an angle. Crossflow turbines are designed as an impulse machine, with a nozzle, but in low head applications maintain some efficiency through reaction, like a traditional water wheel.

The primary numerical classification of a turbine is its specific speed. This number describes the speed of the turbine at its maximum efficiency with respect to the power and flow rate. The specific speed is derived to be independent of turbine size. Given the fluid flow conditions and the desired shaft output speed, the specific speed can be calculated and an appropriate turbine design selected.

The specific speed, along with some fundamental formulas can be used to reliably scale an existing design of known performance to a new size with corresponding performance.

Combustion chamber

A combustion chamber is part of an engine in which fuel is burned

Turbofan

CFM56-3 turbofan, lower half, side view.

Boeing 747 jet engine up close

The turbofan is a type of airplane engine which is evolved from the axial-flow turbojet engine essentially by increasing the size of the first-stage compressor to the point where it acts as a ducted

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multiple thin propeller (or fan) blowing air past the "core" of the engine.

If the propeller is better at low speeds, and the turbojet is better at high speeds, it might be imagined that at some speed range in the middle a mixture of the two is best. Such an engine is the turbofan (originally termed bypass turbojet by the inventors at Rolls Royce). Turbofans essentially increase the size of the first-stage compressor to the point where they act as a ducted fan (or propeller) blowing air past the "core" of the engine. The difference between a ducted fan and a propeller is that the duct slows the air before it arrives at the fan. As both propeller and fan blades must operate subsonically to be efficient, ducted fans allow efficient operation at higher vehicle speeds.

The bypass ratio (the ratio of bypassed air mass to combustor air mass) is an important parameter for turbofans.

Turbofans (especially high bypass engines) are relatively quiet compared to turbojets. The noise of a jet engine is strongly related to the velocity of the air coming out the exhaust. A turbofan has a larger mass flow of air for a given thrust than a turbojet, so the exhaust velocity will be slower and hence the turbofan engine will be quieter than an equivalent turbojet. Jet aircraft are often considered loud, but a conventional piston engine or a turboprop engine delivering the same power would be much louder. (NASA has a web page with details on jet noise.)

Advantages-----1) increased thrust at fwd. speed similer to turboprop result in short take off, weight falls between turbojet and turboprop, ground clearanes are less then turboprop, TSFC and sp. Wt. fall bet. Turbojet and turboprop resulting in good operating economy.superior to turbojet in hot da performance.

Low-bypass turbofans

Early turbojet engines were very fuel-inefficient, as their compression ratio was limited. Improved materials, and the introduction of twin compressors such as in the Pratt & Whitney JT3C engine, increased the compression ratio and thus the thermodynamic efficiency of engines, but led to a poor propulsive efficiency, as pure turbojets have a low-mass, high velocity exhaust.

.Imagine a retrofit situation where a new low bypass ratio, mixed exhaust, turbofan is replacing an old turbojet, in a particular military application. Say the new engine is to have the same airflow and net thrust (i.e. same specific thrust) as the one it is replacing. A bypass flow can only be introduced if the turbine inlet temperature is allowed to increase, to compensate for a correspondingly smaller

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core flow. Improvements in turbine cooling/material technology would facilitate the use of a higher turbine inlet temperature, despite increases in cooling air temperature, resulting from a probable increase in overall pressure ratio.

Efficiently done, the resulting turbofan would probably operate at a higher nozzle pressure ratio than the turbojet, but with a lower exhaust temperature to retain datum net thrust. Since the temperature rise across the whole engine (intake to nozzle) would be lower, the (dry power) fuel flow would also be reduced, resulting in a better specific fuel consumption (SFC).

High-bypass turbofan engines

The introduction of variable compressor stators enabled high pressure ratio compressors to work surge-free at all throttle settings. This innovation made its debut in the General Electric J79, a single-shaft turbojet for supersonic military aircraft. When variable stators were combined with multiple compressors, dramatic increases in overall pressure ratio became possible. Higher turbine inlet temperatures (through improvements in turbine cooling/material technology) enabled relatively small mass flow gas generators to be employed, thus making high-bypass turbofan engines feasible, with bypass ratios of 5 or more.

The first high-bypass turbofan engine was the General Electric TF39, built to power the Lockheed C-5 Galaxy military transport aircraft. The civil General Electric CF6 engine used a related design. Other high-bypass turbofans are the Pratt & Whitney JT9D, the three-shaft Rolls-Royce RB211 and the CFM International CFM56.

The tremendously higher thrust provided by high-bypass turbofan engines also made civil wide-body aircraft practical and economical. In addition to the vastly increased thrust, these engines are also generally quieter. This is not so much due to the higher bypass ratio, but as to the use of low pressure ratio, single stage, fans, which significantly reduce specific thrust and, thereby, jet velocity. The combination of a higher overall pressure ratio and turbine inlet temperature improves thermal efficiency. This together with a lower specific thrust (better propulsive efficiency) leads to a lower specific fuel consumption.

For reasons of fuel economy, and also of reduced noise, almost all of today's jet airliners are powered by high-bypass turbofans.

The Soviet Union's engine technology was less advanced than the West's and its first wide-body aircraft, the Ilyushin Il-86, was powered by low-bypass engines. The Yakovlev Yak-42, a medium-

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range, rear-engined aircraft seating up to 120 passengers was the first Soviet aircraft to use high-bypass engines.

Technical Discussion

Specific Thrust (net thrust/intake airflow) is an important parameter for turbofans and jet engines in general. Imagine a fan (driven by an appropriately sized electric motor) operating within a pipe, which is connected to a propelling nozzle. Fairly obviously, the higher the Fan Pressure Ratio (discharge pressure/inlet pressure), the higher the jet velocity and the corresponding specific thrust. Now imagine we replace this set-up with an equivalent turbofan - same airflow and same fan pressure ratio. Obviously, the core of the turbofan must produce sufficient horsepower to drive the fan via the Low Pressure (LP) Turbine. If we choose a low (HP) Turbine Inlet Temperature for the gas generator, the core airflow needs to be relatively high to compensate. The corresponding bypass ratio is therefore relatively low. If we raise the Turbine Inlet Temperature, the core airflow can be smaller, thus increasing bypass ratio. Raising turbine inlet temperature tends to increase thermal efficiency and, therefore, improve fuel efficiency.

Naturally, as altitude increases there is a decrease in air density and, therefore, the net thrust of an engine. There is also a flight speed effect, termed Thrust Lapse Rate. The net thrust (Fn) of an jet engine (a mixed exhaust turbofan, say) is basically:

Fn = m * (Vjfe - Va)

where:

m intake mass flow

Vjfe fully expanded jet velocity (in the exhaust plume)

Va aircraft flight velocity

With a high specific thrust (e.g. fighter) engine, the jet velocity is relatively high, so intuitively one can see that increases in flight velocity have less of an impact upon net thrust than a medium specific thrust (e.g. trainer) engine, where the jet velocity is lower.

Thrust growth on civil turbofans is usually obtained by increasing fan airflow, thus preventing the jet noise becoming too high. However, the larger fan airflow requires more horsepower from the core. This can be achieved by raising the Overall Pressure Ratio (combustor inlet pressure/intake delivery pressure) to induce more airflow into the core and by increasing turbine inlet temperature.

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Together, these parameters tend to increase core thermal efficiency and improve fuel efficiency.

Recent developments in blade technology

The turbine blades in a turbofan engine are subject to high heat and stress, and require special fabrication. New material construction methods and material science have allowed blades, which were originally polycrystalline (regular metal), to be made from lined up metallic crystals and more recently mono-crystalline blades, which can operate at higher temperatures with less distortion.

Turbofan engine manufacturers

Bypass ratio

In aeronautical engineering, and jet engine design in particular, bypass ratio is a common measurement that compares the amount of air deliberately "blown past" the engine to that moving through the core. For instance, an engine that blows two kilograms of air around the engine for every kilogram that passes through it is said to have a bypass ratio of 2. Higher bypass ratios generally infer better specific fuel consumption as an increasing amount of thrust is being generated without burning more fuel.

Jet engines are generally able to create considerably more energy than they can use in moving air through the engine core. This is because the limiting factor is the temperature at the turbine face, and that is a function of the total amount of fuel burned. Increasing airflow, and thus thrust, would imply burning more fuel and generating higher temperatures. It is possible to increase the airflow by burning "too much" fuel or adding water in front of the turbine to cool it, but both methods lead to incomplete combustion and very poor fuel efficiency. This was nevertheless common for some time to produce added thrust on takeoff, which is why older aircraft appear so smoky in films.

Today almost all jet engines include some amount of bypass. For "low speed" operations like airliners modern engines use bypass ratios up to 17, while for "high speed" operations like fighter aircraft the ratios are much lower, around

Specific fuel consumption

Specific fuel consumption, often shortened to SFC, is an engineering term that is used to describe the fuel efficiency of an engine design.

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It measures the amount of fuel needed to provide a given power for a given period.

SFC is dependent on the engine design, with differences in the SFC between different engines tending to be quite small. For instance, typical gasoline engines will have a SFC of about 0.5 lb/(hp·h) (0.3 kg/(kW·h) = 83 g/MJ), regardless of the design of a particular engine. One exception to the rule is that the SFC within a particular class of engine will vary based on the compression ratio, an engine with a higher compression ratio will deliver a better SFC because it extracts more power from the fuel. Diesel engines have better SFCs than gasoline largely because they have much higher compression ratios, the way they burn their fuel is actually less efficient.

Types

There are a large number of types of jet engines, which get propulsion from a high speed exhaust jet. Some examples are as follows:

Type Description Advantages Disadvantages

water jet Squirts water out the back of a boat

Can run in shallow water, powerful, less harmful to wildlife

Can be less efficient than a propeller

Thermojet Most primitive airbreathing jet engine

Very inefficient and underpowered

Turbojet Generic term for simple turbine engine

Simplicity of design Basic design, misses many improvements in efficiency and power

Turbofan Power tapped off exhaust used to drive bypass fan

Quieter due to greater mass flow and lower total exhaust speed, more efficient for a useful range of subsonic airspeeds for same reason

Greater complexity (additional ducting, usually multiple shafts), large diameter engine, need to contain heavy blades. More subject to FOD and ice damage. Different degrees of bypass are possible - this is the design most commonly used on

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commercial airliners

Rocket Carries own propellant onboard, emits jet for propulsion

Very few moving parts, Mach 0 to Mach 25+, efficient at very high speed (> Mach 10.0 or so), thrust/weight ratio over 100, relatively simple, no air inlet, doesn't require atmosphere, high compression ratio, very high speed exhaust

very low specific impulse- typically 100-450 seconds. Typically requires carrying oxidiser onboard which increases risks.

Ramjet Intake air is compressed entirely by speed of oncoming air and duct shape (divergent)

Very few moving parts, Mach 0.8 to Mach 5+, efficient at high speed (> Mach 2.0 or so), lightest of all airbreathing jets (thrust/weight ratio up to 30 at optimum speed)

Must have a high initial speed to function, inherently inefficient at slow speeds due to poor compression ratio, difficult to arrange shaft power for accessories, difficult to engineer to be efficient over a wide range of airspeeds.

Turboprop (Turboshaft similar)

Strictly not a jet at all- a gas turbine engine is used as powerplant to drive (propeller) shaft

High efficiency at lower subsonic airspeeds(300 knots plus), high shaft power to weight

Limited top speed (aeroplanes), somewhat noisy, complexity of propeller drive, very large yaw (aeroplane) if engine fails

Propfan Turboprop engine drives one or more propellers. much like a turbofan but without ductwork

Higher fuel efficiency, some designs are less noisy than turbofans, could lead to higher-speed commercial aircraft, popular in the 1980s during

Development of propfan engines has been very limited, typically more noisy than turbofans, complexity

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fuel shortages,

Pulsejet Air enters a divergent-duct inlet, the front of the combustion area is shut, fuel injected into the air ignites, exhaust vents from other end of engine

Very simple design, commonly used on model aircraft

Noisy, inefficient (low compression ratio), works best at small scale, valves need to be replaced very often

Pulse detonation engine

Similar to a pulsejet, but combustion occurs as a detonation instead of a deflagration, may or may not need valves

Maximum theoretical engine efficiency

Extremely noisy, parts subject to extreme mechanical fatigue, hard to start detonation, not practical for current use

Integral rocket ramjet

Essentially a ramjet where intake air is compressed and burnt with the exhaust from a rocket

Mach 0 to Mach 4.5+ (can also run exoatmospheric), good efficiency at Mach 2 to 4

Similar efficiency to rockets at low speed or exoatmospheric, inlet difficulties, a relatively undeveloped and unexplored type, cooling difficulties

Scramjet Intake air is compressed but not slowed to below supersonic, intake, combustion and exhaust occur in a single constricted tube

can operate at very high Mach numbers (Mach 8 to 15)[1]

still in development stages, must have a very high initial speed to function (Mach >6), cooling difficulties, inlet difficulties, very poor thrust/weight ratio (~2), airframe difficulties, testing difficulties

Turborocket An additional oxidizer such

Very close to existing designs,

Airspeed limited to same range as

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as oxygen is added to the airstream to increase max altitude

operates in very high altitude, wide range of altitude and airspeed

turbojet engine, carrying oxidizer like LOX can be dangerous

Precooled jets / LACE

Intake air is chilled to very low temperatures at inlet

Very high thrust/weight ratios are possible (~14) together with good fuel efficiency over a wide range of airspeeds, mach 0-5+

Exists only at the lab protoyping stage. Examples include RB545, SABRE, ATREX

Air intakes

See also: Inlet cone

Supersonic inlets: Normal shock is not isentroph

For aircraft travelling at supersonic speeds, a design complexity arises, since the air ingested by the engine must be below supersonic speed, otherwise the engine will "choke" and cease working. This subsonic air speed is achieved by passing the approaching air through a deliberately generated shock wave (since one characteristic of a shock wave is that the air flowing through it is slowed). Therefore, some means is needed to create a shockwave ahead of the intake.

The earliest types of supersonic aircraft featured a central shock cone, called an inlet cone, which was used to form the shock wave.. The same approach can be used for air intakes mounted at the side of the fuselage, where a half cone serves the same purpose with a semicircular air intake, as seen on the F-104 Starfighter and BAC TSR-2. A more sophisticated approach is to angle the intake so that one of its edges forms a leading blade. A shockwave will form at this blade, and the air ingested by the engine will be behind the

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shockwave and hence subsonic. The Century series of US jets featured a number of variations on this approach, usually with the leading blade at the outer vertical edge of the intake which was then angled back inwards towards the fuselage. Typical examples include the Republic F-105 Thunderchief and F-4 Phantom.

Later this evolved so that the leading edge was at the top horizontal edge rather than the outer vertical edge, with a pronounced angle downwards and rearwards. This approach simplified the construction of the intakes and permitted the use of variable ramps to control the airflow into the engine. Most designs since the early 1960s now feature this style of intake, for example the F-14 Tomcat, Panavia Tornado and the Concorde.

Axial-flow compressor

The axial flow compressor is an improvement on the centrifugal compressor previously used in turbine engines,though small and micro turbines use centrifugal compressors with relative advantages (in terms of pressure ratios achieveable per stage of compression). The key improvement is that axial flow compressors work without radically changing the direction of gas flow. It can handle large volume of airflow but it is more susceptible to FOD, IT is expensive and heavy in weight as compered to centrifugal compressor.

Diagram of an axial flow compressor

An axial flow compressor typically has a set of fixed inlet guide vanes to condition the incoming gas. There are then multiple compressor stages, each consisting of a set of rotating blades (much like a propeller) that force the gas to the rear, and then a set of fixed stator blades that condition the air ready for the next compressor stage.

The gas conditioning done by the stator blades is needed to ensure reasonable efficiency. Without the stator blades the gas would rotate with the rotor blades giving a big drop in efficiency.

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Axial flow compressors are typically used in the compression stage of turbine engines. Their disadvantages (low increase in pressure at each stage) are outweighed by their advantages (multi-stages are very compact, they do continuous compression, and they are easy to drive).

A spool is defined as group of compressor stage, a shaft and one or more turbine stage linked mechanically and rotating at same speed.

Centrifugal compressor

Also called a radial blower, squirrel cage, or squirrel wheel compressor, a centrifugal compressor consists of an axle to which is mounted a cylindrical assembly of compressor blades. The compressor operates by using the centrifugal force applied to an air mass to achieve compression. Centrifugal compressors are used throughout industry because they have few moving parts, are very energy efficient, and give higher airflow than a similarly sized reciprocating compressor. Their primary drawback is that they cannot achieve the high compression ratio of reciprocating compressors without multiple stages. Centrifugal compressors are more suited to continuous-duty applications such as ventilation fans, air movers, cooling units, and other uses that require high volume but fairly low pressures. While technically centrifugal blowers can operate in reverse, due to blade design and other factors their efficiency is greatly reduced. Centrifugal blowers are used in some small jet turbine engines; when similar blowers are used in pipelines they are sometimes called jets.

Big centrifugal compressors are used for gas transportation in gas pipelines all around the world. They have the following operating limits: Minimum Operating Speed: is the minimum speed for sustentation, below this value the compressor stops or goes to the called "Idle Speed". Maximum Allowable Speed: is the maximum design speed for the compressor, beyond this value the vibrations increase rapidly, becoming dangerous for the equipment. Stonewall or Choke: this occurs when the velocity of the gas approaches its sonic speed somewhere in the compressor (it may occur at the impeller inlet or at the vaned diffuser inlet). It is generally not detrimental to the compressor. Surge: normally occurs at about 50% of design inlet capacity at design speed, is the point at which the impeller cannot add enough power to overcome the discharge pressure. This causes flow reversal (surge), high vibration, temperature increases, and rapid changes in axial thrust that can damage the labyrinth seals or even the driver.

Advantages ------

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1) Low weight 2) ruggedness and therefore resistance to FOD 3) SIMPLICITY AND LOW COST 4) High compressor ratio per stage( with a limited no. of stages)

Free Power turbine--- has no mechanical connectionto the primary and gas generator turbine, which, in this turn only used to turn compressor in order to to supply high enegy gases to drive free power turbine. Design lend itself to variable speed opertion better then single shaft and it produces high torque at low free power turbine speed. It also has an advantage of requiring no clutch when starting or when load is applied. On the other hand single or fixed shaft engine when used as turboprop allow rapid response rate. The free shaft engine even aat running at idle power is running at same rpm as it is at 100%. All that is required to obtain maximum power to increase fuel flow and propeeler blade angle. Also fixed shaft will burn less fuel as compered to free turbine as there is no fludic coupling to create inefficienices.

Turboprop

A diagram showing how a turboprop works.

A Turboprop (Turbo-propeller) or turboshaft engine is a type of gas turbine. It differs from a turbofan in that the design is optimized to produce rotating shaft power in order to drive a propeller, instead of thrust from the exhaust gas.

A jet engine consists of a set of compressor fans that compress the intake air, a flameholder where the combustion happens, and another set of fans (a set of turbine stages) at the rear to catch some of the hot exhaust and use it to drive the initial compressor fans.

By adding another turbine stage to the engine, all of the jet exhaust can be used for rotary force rather than jet thrust. Coupling this second (or third) turbine stage to a propeller makes for a very efficient engine due to the inherent efficiency of a propeller at low speeds. This is called a turboprop, and can be found on many

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smaller commuter planes, cargo planes, and helicopters (where it is often known as a turboshaft).

While most modern turbojet and turbofan engines use axial-flow compressors, turboprop engines usually contain at least one state of centrifugal compression.

A Rolls-Royce RB.50 Trent on a test rig at Hucknall, in March 1945

Propellers lose efficiency as aircraft speed increases, which is why turboprops are not used on higher-speed aircraft. However, turboprops are far more efficient than piston-driven propeller engines.

Advantges – 1) high propulsive efficieney at low air speeds, wich result in shorter take off but fals of rapidly as airspeed increases.. Engine is able to develop high thrust at low airspeeds because propeller can accelarate large quantity of air at zero fwd. velocity of A/C. 2) heavy in weght 3) lowest TSFC.4) large frontal area of propeller and engine neccisates use of longer landing gear for low wing A/C but does not increase parasite drag5) possibility of efficient reverse thrust.

To overcome the loss in efficieny by formation of shock wave propeller blade of small diameter multi bladed wide chord propeller are more efficient.

Brayton cycle

The Brayton cycle is a cyclic process generally associated with the gas turbine. Like other internal combustion power cycles it is an open system, though for thermodynamic analysis it is a convenient fiction to assume that the exhaust gases are reused in the intake, enabling analysis as a closed system.

Model

A Brayton-type engine consists of three components:

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A gas compressor

A mixing chamber

An expander

In the original 19th century Brayton engine ambient air is drawn into a piston compressor, where it is pressurized; a theoretically isentropic process. The compressed air then runs through a mixing chamber where fuel is added, a constant-pressure process. The heated, pressurized air and fuel mixture is then ignited in an expansion cylinder and gives up its energy, expanding through a piston/cylinder; another theoretically isentropic process. Some of the work extracted by the piston/cylinder is used to drive the compressor through a crankshaft arrangement.

The term Brayton cycle has more recently been given to the gas turbine engine. This also has three components:

A gas compressor

A burner (or combustion chamber)

An expansion turbine

Ambient air is drawn into the compressor, where it is pressurized—a theoretically isentropic process. The compressed air then runs through a combustion chamber, where fuel is burned, heating that air—a constant-pressure process, since the chamber is open to flow in and out. The heated, pressurized air then gives up its energy, expanding through a turbine (or series of turbines)—another theoretically isentropic process. Some of the work extracted by the turbine is used to drive the compressor.

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Since neither the compression nor the expansion can be truly isentropic, losses through the compressor and the expander represent sources of inescapable working inefficiencies.

In general, increasing the compression ratio is the most direct way to increase the overall power output of a Brayton system.

ENGINE CONSTRUCTION  ( Page 1 of 2) AIR INLET DUCTAn engine's air inlet duct is normally considered an airframe part and made by aircraft manufacturer . During flight operation , it is very important to the engine performance . Engine thrust can be high only if the inlet duct supplies the engine with the required airflow at the highest posible pressure . The inlet duct has two engine functions and one aircraft function .       First : it must be able recover as much of the total pressure of the free air stream as posible and deliver this pressure to the front of the engine compressor .       Second : the duct must deliver air to the compressor under all flight conditions with a little turbulance .       Third : the aircraft is concerned , the duct must hold to a minimum of the drag.The duct also usually has a diffusion section just ahead of the compressor to change the ram air velocity into higher static pressure at the face of the engine . This is called ram recovery . The inlet duct is built generally in the divergent shape (subsonic diffuser).

Supersonic DuctThe supersonic duct proplems start when the aircraft begins to fly at or near the speed of sound. At this speeds sonic shock waves are developed which , if not controlled , will give high duct loss in pressure and airflow , and will set up vibrating conditions in the inlet duct called inlet " buzz " . Buzz is an airflow instability caused by the shock wave rapidly being alternately swallowed and expelled at the inlet of the duct. Air enters the compressor section of engine must be slow to subsonic velocity. At supersonic speeds the inlet does the

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job by slowing the air with minimize energy loss and the temperature rise.At transonic speeds the inlet duct is designed to keep shock waves out of the duct. This is done by locating the inlet duct behind a spike or probe which create the shock wave infront of inlet duct. This normal shock wave will produce a pressure rise and velocity decrease to subsonic speeds .

At higher mach numbers, the single normal shock wave is very strong and causes a great reduction in the total pressure recoverd by the duct and excessive air temperature rise inside the duct. The oblique shock wave will be used to slow the supersonic velocity down but still supersonic , the normal shock wave will drop the velocity to subsonic before the air enter to the compressor. Each reduce in velocity will increase a pressure. At very high mach number , the inlet duct must set up one or moreoblique shocks and a normal shock. COMPRESSORThe combustion of fuel and air at normal atmospheric pressure will not produce sufficient energy enough to produce useful work . The energy released by combustion is proportional to the mass of air consumed and its pressure. Therefore , higher pressure are needed to increase the efficiency of the combustion cycle . On the jet engines must rely upon some other means of compression .

Although centrifugal compressors are used in many jet engine , the efficiency level of a single stage is relatively low . The multistage of centrifugal compressor is better , but still do not compare with those axial flow compressors . Some small modern turboshaft and turboprop engines achieve good results by using a combination of axial flow and centrifugal compressor. Centrifugal compressorCentrifugal compressors operate by taking in outside air near their hub and rotating it by means of an impeller . The impeller , which is

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usually an aluminum alloy , guides the air toward the outer circumference of the compressor , building up the velocity of the air by means of high rotational speed of the impeller . The compressor consists of three main parts:1) Impeller 2) A Diffuser 3) A Comprssor ManifoldAir leaves the Impeller at high speed , and flows through the diffuser which converts high velocity , kinetic energy to low velocity , high pressure energy . The diffuser also serves to direct airflow to the compressor manifold which acts as collector ring. They also delivery air to the manifold at a velocity and pressure which will be satisfactory for use in the burner section of the engine.

Axial compressorThe air in an axial compressor flows in an axial direction through a series of rotating rotor blades and stationary stator vanes. The flow path of an axial compressor decreases in cross-section area in the direction of flow , reducing the volume of the air as compression progresses from stage to stage of compressor blades .

  

The air being delivered to the face of compressor by the air inlet duct, the incoming air passes through the inlet guide vanes . Air upon entering the first set of ratating blades and flowing in axial direction, is deflected in the direction of rotation . The air is arrested and turn as it is passed on to a set of stator vanes , following which it is again picked up by another set of rotating blades , and so on , through the compressor . The pressure of the air increases each time that it passes through a set of rotors and stators . The aerodynamic principles are applied to the compressor blade design in order to increase efficiency . The blades are treated as lifting surfaces like aircraft wings or propeller blades . The cascade

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effect is a primary consideration in determining the airfoil section , angle of attack , and the spacing between blades to be used for compressor blade disign . The blade must be designed to withstand the high centrifugal forces as well as the aerodynamic loads to which they are subjected . The clearance between the rotating blades and their outer case is also very important . The rotor assembly turns at extreamely high speed , and must be rigid , well aligned and well balance . Compressor Surge and Compressor StallThis characteristic has been called both " Surge " and " Stall " in the past , but is more properly called SURGE when it is response of the entire engine. The word stall applies to the action occuring at each individual compressor blade. Compressor surge , also called Compressor stall , is a phenomenon which is difficult to understand because it is usually caused by complex combination of factors . The basic cause of compressor surge is fairly simple , each blade in an axial flow compressor is a miniature airplane wing which , when subjected to a higher angle of attack , will stall just as an airplane stalls. Surge may define as results from an unstable air condition within the compressor. Pilot or engine operator has no instrument to tell him that one or more blades are stalling. He must wait until the engine surges to know that. The unstable condition of air is often caused from air piling up in the rear stages of the compressor. Surge may become sufficiently pronounce to cause lound bangs and engine vibration. In most case , this condition is of short duration , and will either correct itself or can be corrected by retarding the throttle or power lever to Idle and advanncing it again , slowly. Among other things , to minimize the tendency of a compressor to surge , the compressor can be "unload" during certain operating conditions by reducing the pressure ratio across the compressor for any giving airflow. One method of doing this is by bleeding air from the middle or toward the rear of the compressor. In dual axial compressor engines , air is often bled from between the low and the high pressure compressor. Air bleed ports are located in the compressor section. These ports are fitted with automatic , overboard bleed valves which usually operate in a specified range of engine RPM. Some large engine have been provided with variable-angle stators ( variable stators) in a few of the forward compressor stages. The angle of these vanes change automatically to prevent the choking of the downstream compressor stages as engine operating conditions vary.

DIFFUSER SECTION

The diffuser has an expanding internal diameter to decrease the velocity and increase the static pressure of air . The air leaving compressor , then through a diffuser section . The diffuser prepares the air for entry the combustion section at low velocity to permit proper mixing with fuel . Ports are built in the diffuser case through

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which compressor discharge air is bled off from the aircraft engine .

   

On dual compressor engines , bleed air for service functions is also taken from additional ports located between the low and high compressors , or at intermediate stages in the high pressure compressor case . Air is bled from most engine vented over board out of the primary air flow path during certain engine operating conditions to prevent compressor surge .This is called over board and must not be confused with the air remove from the engine to perform service function .

FUEL MANIFOLDS and NOZZLES

Fuel is introduced into the air stream at the front of the burners in spray form , suitable for rapid mixing with air for combustion. The fuel is carried from outside the engine , by manifold system , to nozzles mounted in the burner cans .

Primary and secondary fuel manifolds are often used on large engines . The primary manifold provides sufficient fuel for low thrust operation. At high thrust , the secondary , or main manifold cuts in , and fuel commences to flow through both primary and secondary elements of double-orifice nozzle. Usually , primary fuel is sprayed through a single orifice at the center of nozzle. Secondary fuel is sprayed through a number of orifices in a ring around the center orifice.

COMBUSTION CHAMBERS OR BURNER SECTION

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There are three basic types of burner systems in use today. They are can type , annular type and can-annular type. Fuel is introduced at the front end of the burner. Air flows in around the fuel nozzle and through the first row of combustion air holes in the liner. The air entering the forward section of the liner tends to recirculate and move up stream against the fuel spray. During combustion , this action permits rapid mixing and prevents flame blowout which acts as a continuous pilot for the rest of the burner.

   

There are usually has only two igniter plugs in an engine. The igniter plug is usually locate in the up stream region of the burner. About 25 percent of the air actually takes part in the combustion process. The gases that result from the combustion have temperatures of 3500 degree F. Before entering the turbine , the gases must be cooled to approximately half this value , up to the designed of turbine materials involved. Cooling is done by diluting the hot gases with secondary air that enters through a set of relative large holes located toward the rear of the liner.

TURBINE SECTION

The turbine in all modern jet engines , regardless of the type of

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compressor used , are of axial flow design.

The turbine extract kinetic energy from the expanding gases as the gases come from the burner , converting this energy into shaft horsepower to drive the compressor and the engine accessory. Nearly three fourths of all energy available from the product of combustion is needed to drive the compressors.

The turbine wheel is one of the most highly stressed parts in the engine. Not only must it operateat temperature 1700 degree F, but it must do so under severe centrifugal loads imposed by high rotational speeds of over 40000 rpm for small engines to 8000 rpm for a larger engines.The engine speed and turbine inlet temperature must be accurately controlled to keep the turbine within safe operating limits.The turbine assembly is made of two main parts , the disk and the blades. The disk or wheel is statically and dynamically balanced and unit specially alloyed steel usually containing large percentages of chromium , nickle , and cobalt. The blades are attached to the disk by means of a " fir tree " design to allow for different rates of expansion between the disk and the blade while still holding the blade firmly against centrifugal loads. The blade is kept from moving axially either by rivets , special locking tabs or devices , or another turbine stage.

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The blade is shrouded at the tip. The shrouded blades form a band around the perimeter of the turbine which serves to reduce blade vibrations. The shrouds improve the airflow characteristics and increase the efficiency of the turbine. The shrouds also serve to cut down gas leakage around the tips of the turbine blades.

EXHAUST DUCT OR EXHAUST PIPE

A larger total thrust can be obtained from the engine if the gases are discharged from the aircraft at a higher velocity than is permissible at the turbine outlet. An exhaust duct is therefore added , both to collect and straighten the gas flow as it comes from the turbine and to increase the velocity of the gases before they are discharged from the exhaust nozzle at the rear of the duct.

Increasing the velocity of the gases increases their momentum and increase the thrust produced.The duct is essentially a simple , stainless steel , conical or cylinder pipe .

The tail cone helps smooth the flow. A conventional convergent type of exhaust duct is capable of keeping the flow through the duct

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constant at velocity not to exceed Mach 1.0 at the exhaust nozzle.

AFTER BURNING

The afterburner , whose operation is much like a ram-jet , increases thrust by adding fuel to the exhaust gases after they have passed through the turbine section. At this point there is still much uncombined oxygen in the exhaust. Only approximately 25 percent of the air passing through the engine is consumed by the combustion. The remainder or 75 percent , of the air is capable of supporting additional combustion if more fuel is added. The resultant increase in the temperature and velocity of gases therefore boosts engine thrust. Most afterburners will produce an approximately 50 percent more thrust. Afterburning or " hot " operation or " reheating " is used only for a time limited operation of takeoff , climb , and maximum burst speed.

JET ENGINE EQUATION

Since Fuel flow adds some mass to the air flowing through the engine , this must be added to the basic of thrust equation . Some formular do not consider the fuel flow effect when computing thrust because the weight of air leakage is approximately equal to the weight of fuel added . The following formular is applied when a nozzle of engine is " choked " , the pressure is such that the gases are treveling through it at the speed of sound and can not be further accelerated . Any increase in internal engine pressure will pass out through the nozzle still in the form of pressure . Even this pressure energy cannot turn into velocity energy but it is not lost .

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FACTORS AFFECTING THRUST

The Jet engine is much more sensitive to operating variables . Those are:1.) Engine rpm.2.) Size of nozzle area.3.) Weight of fuel flow.4.) Amount of air bled from the compressor.5.) Turbine inlet temperature.6.) Speed of aircraft (ram pressure rise).7.) Temperature of the air.8.) Pressure of air 9.) Amount of humidity.Note ; item 8,9 are the density of air .

ENGINE STATION DESIGNATIONS

Station designations are assigned to the varius sections of gas turbine engines to enable specific locations within the engine to be easily and accurately identified. The station numbers coincide with position from front to rear of the engine and are used as subscripts when designating different temperatures and pressures at the front , rear , or inside of the engine. For engine configurations other than the picture below should be made to manuals published by the engine manufacturer.

N = Speed ( rpm or percent )N1 = Low Compressor Speed N2 = High Compressor Speed N3 = Free Turbine Speed P = Pressure T = Temperature t = Total

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EGT = Exhaust Gas Temperature EPR = Engine Pressure Ratio ( Engine Thrust in term of EPR ). Pt7 / Pt2Ex.: Pt 2 = Total Pressure at Station 2 ( low pressure compressor inlet )   Pt 7 = Total Pressure at Station 7 ( turbine discharge total pressure )

Requirements \ Commonly Used Materials

This graphic demonstrates the how the temperature fluxuates

during take-off, flight, and landing. Hot times are

experienced during take-off and landing.

Displays how man hours are expected out of an engine today.

Fan

Requirements:

Fan1. High strength.2. Lightweight (Safety precaution in case it blows up).3. Be able to handle a direct blow without breaking (bird strike).4. Temperature range: ~ -50 - 100° F

Commonly used material:

Blades - Polymer Composite or Titanium alloys.Containment - Nickel-based alloys, Polymer Composite, or Titanium alloys.

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Fan containment1. Absorbent.2. Compact.3. Preferably a layered structure.4. Temperature range: 400 - 500° F

Compressor

Requirements:

1. 200 to 300 hot hours.2. Temperature range: 800 - 1200° F

Disk1. High strength.2. Resist centrifugal stress.3. Resist fatigue.

Commonly used material:

Blades - Titanium alloys (cold side). Nickel-based alloy or Titanium alloy (hotter end).Disk - Titanium alloys (cold) and Nickel-based alloy (hot).

Combustor

Requirements:

Combustor1. 18,000 to 20,000 hours.2. 9,000 hot hours.3. Average temperature around 2,800° F.

Combustor liner1. Stresses due to thermal gradient heat.2. Transient stresses due to takeoff and cool down situations.3. Resist oxidation.

Commonly used material:

Currently - Nickel-based alloy.Future - Ceramic composite.

Turbine

Requirements:

1. Rotational strength.2. Pressure loading.3. High temperatures.4. Resist Creep.

Commonly used material:

Disk - Nickel-based alloyBlades - Single crystal Nickel-based alloy with thermal barrier coating.

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5. Resist Oxidation.6. Temperature range: 1000 - 2000° F

Mixer

Requirements:

1. High Temperatures.2. Temperature range: 1000 - 1200° F

Commonly used material:

Nickel-based alloy

Nozzle

Requirements:

1. High Temperature.2. Temperature range: 1200 - 2400° F

Commonly used material:

Nickel-based alloyTitanium alloyCeramic matrix composite

1.10. TURBOJET

The turbojet is the engine in most common use today in high-speed, high-altitude aircraft, not in Army aircraft. With this engine, air is drawn in by a compressor which raises internal pressures many times over atmospheric pressure. The compressed air then passes into a combustion chamber where it is mixed with fuel to be ignited and burned. Burning the fuel-air mixture expands the gas, which is accelerated out the rear as a high-velocity jet-stream. In the turbine section of the engine, the hot expanded gas rotates a turbine wheel which furnishes power to keep the compressor going. The gas turbine engine operates on the principle of intake, compression, power, and exhaust, but unlike the reciprocating engine, these events are continuous. Approximately two-thirds of the total energy developed within the combustion chamber is absorbed by the turbine wheel to sustain operation of the compressor. The remaining energy is discharged from the rear of the engine as a high velocity jet, the reaction to which is thrust or forward movement of the engine. The turbojet is shown schematically in figure 1.3.

1.12. ADVANTAGES OF TURBINE ENGINES

Keeping in mind the basic theory of turbine engines, compare the advantages and disadvantages of the turbine engine with the piston or reciprocating engine. The advantages are covered in the

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subparagraphs below, and disadvantages are discussed in the next section.

Turbine engines have a higher power-to-weight Power-to-weight ratio. ratio than reciprocating engines. An example of this is the T55-L-l11. It weighs approximately 650 pounds and delivers 3, 750 shaft horsepower. The power-to-weight ratio for this engine is 5.60 shp per pound, where the average reciprocating engine has a power-to-weight ratio of approximately .67 shp per pound.

Less maintenance. Maintenance per hour of operation is especially important in military operations. Turbine engines require less maintenance per flying hour than reciprocating engines generally do. As an aircraft maintenance officer, this advantage will appeal to you because of a greater aircraft availability and lower maintenance hour to flying hour ratio. The turbine engine also has fewer moving parts than a reciprocating engine; this is also an advantage over the reciprocating engine.

Less drag. Because of the design, the turbine engine has a smaller frontal area than the reciprocating engine. A reciprocating engine requires a large frontal area which causes a great deal of drag on the aircraft. Turbine engines are more streamlined in design, causing less drag. Figure 1.6 shows one of the two nacelles that contain reciprocating engines in the old CH-37 cargo helicopter. Figure 1.7 shows the smaller frontal area of the turbine engines that power the CH-47 Chinook helicopter. Because of this, the engine nacelles are more streamlined in design, causing less drag.

Cold weather starting. The turbine engine does not require any oil dilution or preheating of the engine before starting. Also, once started, the reciprocating engine takes a long time to warm up to operating temperatures, whereas the turbine engine starts readily and is up to operating temperature immediately.

Low oil consumption. The turbine engine, in general, has a lower rate of oil consumption than the reciprocating engine. The turbine engine does not require the oil reservoir capacity to be as large as the reciprocating engine's; because of this, a weight and economy factor is an additional advantage.

1.13. DISADVANTAGES OF TURBINE ENGINES

Just like everything else, along with the advantages or the good, we have to take the disadvantages or the bad. This also holds true with the turbine engine. The disadvantages of the turbine engine are discussed in the following subparagraphs.

Foreign object damage. One of the major problems faced by the turbine engine is foreign object damage (FOD). A turbine

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engine requires tremendous quantities of air. This air is sucked into the engine at extremely high velocities, and it will draw up anything that comes near the inlet area. The turbine engines used in Army aircraft are fitted with filters around the engine inlet to prevent foreign objects from entering the engine and damaging the compressor vanes. However, even with this precaution, FOD is still a menace to turbine engine operation, as shown in figure 1.8.

High temperatures. In the combustion chamber, the temperature is raised to about 3, 500° F. in the hottest part of the flame. Because this temperature is above the melting point of most metals, proper cooling and flame dilution must be employed at all times to insure that the engine is not damaged.

Slow acceleration. The acceleration rate of a turbine engine is very slow in comparison with that of a reciprocating engine. The pilot must be aware of the time lag in the turbine engine acceleration between the instant when power is requested and when power is available.

High fuel consumption. Turbine engines are very uneconomical when it comes to the amount of fuel they consume. The Lycoming T53 turbine engine, for instance, uses approximately 1.5 gallons per minute of fuel. Compare it to a reciprocating engine of approximately the same horsepower which has a fuel consumption rate of 1 gallon per minute.

Cost. The initial cost of a turbine engine is very high when compared to the cost of a reciprocating engine. For example the T53-L-13B engine costs about $63,000, and the cost of a reciprocating engine of approximately the same horsepower is $20,000.

1.19. COMPRESSOR SECTION

The compressor is the section of the engine that produces an increase in air pressure. It is made up of rotating and stationary vane assemblies. The first stage compressor rotor blades accelerate the air rearward into the first stage vane assemblies. The first stage vane assemblies slow the air down and direct it into the second stage compressor rotor blades. The second stage compressor rotor blades accelerate the air rearward into the second stage vane assemblies, and so on through the compressor rotor blades and vanes until air enters the diffuser section. The highest total air velocity is at the inlet of the diffuser. As the air passes rearward through the diffuser, the velocity of the air decreases and the static pressure increases. The highest static pressure is at the diffuser outlet.

The compressor rotor may be thought of as an air pump. The volume of air pumped by the compressor rotor is basically proportional to the rotor rpm. However, air density, the weight of a given volume of

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air, also varies this proportional relationship. The weight per unit volume of air is affected by temperature, compressor air inlet pressure, humidity, and ram air pressure*. If compressor air inlet temperature is increased, air density is reduced. If compressor air inlet pressure is increased, air density is increased. If humidity increases, air density is decreased. Humidity, by comparison with temperature, and pressure changes, has a very small effect on density. With increased forward speed, ram air pressure increases and air temperature and pressure increase.

*ram air pressure - free stream air pressure provided by the forward motion of the engine.

Compressor efficiency determines the power necessary to create the pressure rise of a given airflow, and it affects the temperature change which takes place in the combustion chamber. Therefore, the compressor is one of the most important components of the gas turbine engine because its efficient operation is the key to overall engine performance. The following subparagraphs discuss the three basic compressors used in gas turbine engines: the centrifugal-flow, the axial-flow, and axial-centrifugal-flow compressors. The axial-centrifugal-flow compressor is a combination of the other two and operates with characteristics of both.

Centrifugal-flow compressor. Figure 1.12 shows the basic components of a centrifugal-flow compressor: rotor, stator, and compressor manifold.

Figure 1.12. Typical Single-stage Centrifugal Compressor

As the impeller (rotor) revolves at high speed, air is drawn into the blades near the center. Centrifugal force accelerates this air and causes it to move outward from the axis of rotation toward the rim of the rotor where it is forced through the diffuser section at high velocity and high kinetic energy. The pressure rise is produced by reducing the velocity of the air in the diffuser, thereby converting

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velocity energy to pressure energy. The centrifugal compressor is capable of a relatively high compression ratio per stage. This compressor is not used on larger engines because of size and weight.

Because of the high tip speed problem in this design, the centrifugal compressor finds its greatest use on the smaller engines where simplicity, flexibility of operation, and ruggedness are the principal requirements rather than small frontal area and ability to handle high airflows and pressures with low loss of efficiency.

Axial-flow compressor. The air is compressed, as the name implies, in a direction parallel to the axis of the engine. The compressor is made of a series of rotating airfoils called rotor blades, and a stationary set of airfoils called stator vanes. A stage consists of two rows of blades, one rotating and one stationary. The entire compressor is made up of a series of alternating rotor and stator vane stages as shown in figure 1.13.

Figure 1.13. Axial-flow Compressor.

Axial flow compressors have the advantage of being capable of very high compression ratios with relatively high efficiencies; see figure 1.14. Because of the small frontal area created by this type of compressor, it is ideal for installation on high-speed aircraft. Unfortunately, the delicate blading and close tolerances, especially toward the rear of the compressor where the blades are smaller and more numerous per stage, make this compressor highly susceptible to foreign-object damage. Because of the close fits required for

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efficient air-pumping and higher compression ratios, this type of compressor is very complex and very expensive to manufacture. For these reasons the axial-flow design finds its greatest application where required efficiency and output override the considerations of cost, simplicity, and flexibility of operation. However, due to modern technology, the cost of the small axial-flow compressor, used in Army aircraft, is coming down.

Figure 1.14. Compressor Efficiencies and Pressure Ratios.

Axial-centrifugal-flow compressor. The axial-centrifugal-flow compressor, also called the dual compressor, is a combination of the two types, using the same operating characteristics. Figure 1.15 shows the compressor used in the T53 turbine engine. Most of the gas turbine engines used in Army aircraft are of the dual compressor design. Usually it consists of a five- or seven-stage axial-flow compressor and one centrifugal-flow compressor. The dual compressors are mounted on the same shaft and turn in the same direction and at the same speed. The centrifugal compressor is mounted aft of the axial compressor. The axial compressor contains numerous air-foil-shaped blades and vanes that accomplish the task of moving the air mass into the combustor at an elevated pressure.

Figure 1.15. Axial-Centrifugal-Flow Compressor.

As the air is drawn into the engine, its direction of flow is changed by the inlet guide vanes. The angle of entry is established to ensure that the air flow onto the rotating compressor blades is within the stall-free (angle of attack) range. Air pressure or velocity is not changed as a result of this action. As the air passes from the trailing edge of the inlet guide vanes, its direction of flow is changed due to the rotational effect of the compressor. This change of airflow direction is similar to the action that takes place when a car is driven during a rain or snow storm. The rain or snow falling in a vertical direction strikes the windshield at an angle due to the horizontal velocity of the car.

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In conjunction with the change of airflow direction, the velocity of the air is increased. Passing through the rotating compressor blades, the velocity is decreased, and a gain in pressure is obtained. When leaving the trailing edge of the compressor blades, the velocity of the air mass is again increased by the rotational effect of the compressor. The angle of entry on to the stationary stator vanes results from this rotational effect as it did on the airflow onto the compressor.

Passing through the stationary stator vanes the air velocity is again decreased resulting in an increase in pressure. The combined action of the rotor blades and stator vanes results in an increase in air -pressure; combined they constitute one stage of compression. This action continues through all stages of the axial compressor. To retain this pressure

buildup, the airflow is delivered, stage by stage, into a continually narrowing airflow path. After passing from the last set of stator vanes the air mass passes through exit guide vanes. These vanes direct the air onto the centrifugal impeller.

The centrifugal impeller increases the velocity of the air mass as it moves it in a radial direction.

Compressor stall. Gas turbine engines are designed to avoid the pressure conditions that allow engine surge to develop, but the possibility of surge still exists in engines that are improperly adjusted or have been abused. Engine surge occurs any time the combustion chamber pressure exceeds that in the diffuser, and it is identified by a popping sound which is issued from the inlet. Because there is more than one cause for surge, the resultant sound can range from a single carburetor backfire pop to a machinegun sound.

Engine surge is caused by a stall on the airfoil surfaces of the rotating blades or stationary vanes of the compressor. The stall can occur on individual blades or vanes or, simultaneously, on groups of them. To understand how this can induce engine surge, the causes and effects of stall on any airfoil must be examined.

All airfoils are designed to provide lift by producing a lower pressure on the convex (suction) side of the airfoil than on the concave (pressure) side. A characteristic of any airfoil is that lift increases with an increasing angle of attack, but only up to a critical angle. Beyond this critical angle of attack, lift falls off rapidly. This is due largely to the separation of the airflow from the suction surface of the airfoil, as shown in the sketch. This phenomenon is known as stall. All pilots are familiar with this condition and its consequences

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as it applies to the wing of an aircraft. The stall that takes place on the fixed or rotating blades of a compressor is the same as the stalling phenomenon of an aircraft wing.

1.20. COMPRESSOR CONSTRUCTION

Centrifugal - flow compressors are usually made of titanium. The diffuser is generally manufactured of a stainless steel alloy. A close fit is important between the compressor and its case to obtain maximum compressor efficiency. Correct rotor assembly balancing is essential for safe operation because of the high rpm. Balancing the rotor can be accomplished by removing metal from specified areas of the compressor or by using balancing weights installed in holes in the hub of the compressor. On some engines where the compressor and turbine wheel are balanced as a unit, special bolts and nuts having slight variations in weight are used.

Axial-flow compressors are constructed of many different materials, depending upon the load and temperature under which the unit must operate. The rotor blades are generally cast of stainless-steel alloy. Some manufacturers use mdybdenum coated titanium blades to dampen vibrations on some stages of rotor blades. The clearance between the rotor blades and the outer case is most important. Some companies coat the inner surface of the compressor case with a soft material that can be worn away by the blades as they expand because of the heat generated from compressing the air. This type of compressor uses the "wear-fit" method to form its own clearance between the compressor case and the rotor blade tip.

Methods of attaching the blade to the disk or hub vary between manufacturers, with the majority using some variation of the dove-tail method to hold the rotor blades to the disk. Various other methods are used to anchor the blades in place. Some blades do not have a tight fit in the disk, but rather are seated by centrifugal force during engine operation. By allowing the blades to move, vibrational stress is reduced during start and shutdown. Stator vanes, shown in figure 1.16, can be either solid or hollow construction, and are connected together at their tips by a shroud. This shrouding serves two purposes. First, it provides support, and second, it provides the necessary air seal between rotating and stationary parts. Most manufacturers use the split compressor cases, while some others favor a weldment, forming a continuous case. The advantages of the split case lie in the fact that the compressor and stator blades are readily available to inspection. The one-piece case offers simplicity and strength because it is one piece; in most instances, it is a principal structural part of the engine and is usually made of cast aluminum, magnesium, or steel. Figures 1.16 and 1.17 show shrouded compressor stators in both the split case and the one-piece case.

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1.21. COMBUSTION SECTION

Today, three basic combustion chambers are in use. They are the annular combustion chamber, the can type, and the combination of the two called the can-annular. Variations of these basic systems are used in a number of engines. The three systems are discussed individually in the following subparagraphs. The most commonly used gas turbine engine in Army aircraft is the annular reverse-Row type. The combustion section contains the combustion chambers, igniter plugs, and fuel nozzles or vaporizing tubes. It is designed to burn a fuel-air mixture and deliver the combusted gases to the turbine at a temperature which will not exceed the allowable limit at the turbine inlet.

Fuel is introduced at the front end of the burner in a highly atomized spray from the fuel nozzles. Combustion air flows in around the fuel nozzle and mixes with the fuel to form a correct fuel-air mixture. This is called primary air and represents approximately 25 percent of total air taken into the engine. The fuel-air mixture which is to be burned is a ratio of 15 parts of air to 1 part of fuel by weight. The remaining 75 percent of the air is used to form an air blanket around the burning gases and to lower the temperature. This temperature may reach as high as 3500° F. By using 75 percent of the air for cooling, the temperature operating range can be brought down to about half, so the turbine section will not be destroyed by excessive heat. The air used for burning is called primary air- and that for cooling is secondary air. The secondary air is controlled and directed by holes and louvers in the combustion chamber liner.

Igniter plugs function only during starting, being cut out of the circuit as soon as combustion is self-supporting. On engine shutdown, or, if the engine fails to start, the combustion chamber drain valve, a pressure-actuated valve, automatically drains any remaining unburned fuel from the combustion chamber. All combustion chambers contain the same basic elements: a casing or outer shell, a perforated inner liner or flame tube, fuel nozzles, and some means of initial ignition. The combustion chamber must be of light construction and is designed to burn fuel completely in a high velocity airstream. The combustion chamber liner is an extremely critical engine part because of the high temperatures of the flame. The liner is usually constructed of welded high-nickel steel. The most severe operating periods in combustion chambers are encountered in the engine idling and maximum rpm ranges. Sustained operation under these conditions must be avoided to prevent combustion chamber liner failure.

The annular-type combustion chamber shown in figure 1.18 is used in engines of the axial-centrifugal-flow compressor

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design. The annular combustion chamber permits building an engine of a small and compact design. Instead of individual combustion chambers, the primary compressed air is introduced into an annular space formed by a chamber liner around the turbine assembly. A space is left between the outer liner wall and the combustion chamber housing to permit the flow of secondary cooling air from the compressor. Primary air is mixed with the fuel for combustion. Secondary (cooling) air reduces the temperature of the hot gases entering the turbine to the proper level by forming a blanket of cool air around these hot gases.

1. ANNULAR TYPE COMBUSTION CHAMBER LINER2. COMBUSTION CHAMBER HOUSING ASSEMBLYFigure 1.18. Annular-type Combustion Chamber.

The annular combustion chamber offers the advantages of a larger combustion volume per unit of exposed area and material weight, a smaller exposed area resulting in lower pressure losses through the unit, and less weight and complete pressure equalization.

The can-type combustion chamber is one made up of individual combustion chambers. This type of combustion chamber is so arranged that air from the compressor enters each individual chamber through the adapter. Each individual chamber is composed of two cylindrical tubes, the combustion chamber liner and the outer combustion chamber, shown in figure 1.19. Combustion takes place within the liner. Airflow into the combustion area is controlled by small louvers located in the inner dome, and by round holes and elongated louvers along the length of the liner. Airflow into the combustion area is

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controlled by small louvers located in the inner dome, and by round holes elongated louvers along the length of the liner.

Figure 1.19. Can-type Combustion Chamber (Cutaway).

Through these openings flows the air that is used in combustion and cooling. This air also prevents carbon deposits from forming on the inside of the liner. This is important, because carbon deposits can block critical air passages and disrupt airflow along the liner walls causing high metal temperatures and short burner life.

Ignition is accomplished during the starting cycle. The igniter plug is located in the combustion liner adjacent to the start fuel nozzle. The Army can-type engine employs a single can-type combustor.

Can-annular combustion chamber. This combustion chamber uses characteristics of both annular and can-type combustion chambers. The can-annular combustion chamber consists of an outer shell, with a number of individual cylindrical liners mounted about the engine axis as shown in figure 1.20. The combustion chambers are completely surrounded by the airflow that enters the liners through various holes and louvers. This air is mixed with fuel which has been sprayed under pressure from the fuel nozzles. The fuel-air mixture is ignited by igniter plugs, and the flame is then carried through the crossover tubes to the remaining liners. The inner casing assembly is both a support and a heat shield; also, oil lines run through it.

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Figure 1.20. Can-Annular Combustion Chamber.

 

1.22. TURBINE SECTION

A portion of the kinetic energy of the expanding gases is extracted by the turbine section, and this energy is transformed into shaft horsepower which is used to drive the compressor and accessories. In turboprop and turboshaft engines, additional turbine rotors are designed to extract all of the energy possible from the remaining gases to drive a powershaft.

Types of turbines. Gas turbine manufacturers have concentrated on the axial-flow turbine shown in figure 1.21. This turbine is used in all gas-turbine-powered aircraft in the Army today. However, some manufacturers are building engines with a radial inflow turbine, illustrated in figure 1.22. The radial inflow turbine has the advantage of ruggedness and simplicity, and it is relatively inexpensive and easy to manufacture when compared to the axial-flow turbine. The radial flow turbine is similar in design and construction to the centrifugal-flow compressor described in paragraph 1.19a. Radial turbine wheels used for small engines are well suited for a higher range of specific speeds and work at relatively high efficiency.

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Figure 1.21. Axial-flow Turbine Rotor.

Figure 1.22. Radial Inflow Turbine.

The axial-flow turbine consists of two main elements, a set of stationary vanes followed by a turbine rotor. Axial-flow turbines may be of the single-rotor or multiple-rotor type. A stage consists of two main components: a turbine nozzle and a turbine rotor or wheel, as shown in figure 1.21. Turbine blades are of two basic types, the impulse and the reaction. Modern aircraft gas turbines use blades that have both impulse and reaction sections, as shown in figure 1.23.

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Figure 1.23. Impulse-Reaction Turbine Blade.

The stationary part of the turbine assembly consists of a row of contoured vanes set at a predetermined angle to form a series of small nozzles which direct the gases onto the blades of the turbine rotor. For this reason, the stationary vane assembly is usually called the turbine nozzle, and the vanes are called nozzle guide vanes.

Single-rotor turbine. Some gas turbine engines use a single-rotor turbine, with the power developed by one rotor. This arrangement is used on engines where low weight and compactness are necessary. A single-rotor, single-stage turbine engine is shown in figure 1.24, and a multiple-rotor, multiple-stage turbine engine is shown in figure 1.25.

Figure 1.24. Single-rotor,Single-stage Turbine.

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Figure 1.25. Multiple-rotor,Multiple-stage Turbine.

Multiple-rotor turbine. In the multiple-rotor turbine the power is developed by two or more rotors. As a general rule, multiple-rotor turbines increase the total power generated in a unit of small diameter. Generally the turbines used in Army aircraft engines have multiple rotors. Figure 1.26 illustrates a multistage, multiple-rotor turbine assembly.

Figure 1.26. Multirotor - Multistage Turbine.

1.23. TURBINE CONSTRUCTION

The turbine rotor is one of the most highly stressed parts in the engine. It operates at a temperature of approximately 1,700° F. Because of the high rotational speeds, over 40,000 rpm for the smaller engines, the turbine rotor is under severe centrifugal loads. Consequently, the turbine disk is made of specially alloyed steel, usually containing large percentages of chromium, nickel, and

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cobalt. The turbine rotor assembly is made of two main parts, the disk and blades.

Nozzle vanes may be either cast or forged. Some vanes are made hollow to allow cooling air to flow through them. All nozzle assemblies are made of very high-strength steel that withstands the direct impact of the hot gases flowing from the combustion chamber.

The turbine blades are attached to the disk by using the "fir tree" design, shown in figure 1.27, to allow for expansion between the disk and the blade while holding the blade firmly to the disk against centrifugal loads. The blade is kept from moving axially either by rivets or special locking devices. Turbine rotors are of the open-tip type as shown in figure 1.27, or the shroud type as shown in figure 1.28.

Figure 1.27. Turbine Wheel Open Tip.

Figure 1.28. Turbine Blade "Fir Tree Root" Shroud.

The shroud acts to prevent gas losses over the blade tip and excessive blade vibrations. Distortion under severe loads tends to twist the blade toward low pitch, and the shroud helps to reduce this tendency. The shrouded blade has an aerodynamic advantage in that thinner blades can be used with the support of the shroud. Shrouding, however, requires that the turbine run cooler or at reduced rpm because of the extra mass at the tip.

Blades are forged or cast from alloy steel and machined and carefully inspected before being certified for use. Manufacturers stamp a "moment weight" number on the blade to retain rotor balance when replacement is necessary. Turbine blade maintenance and replacement are covered in a separate lesson.

2.7. FUEL NOZZLES

On most gas turbine engines, fuel is introduced into the combustion chamber through a fuel nozzle that creates a highly atomized and accurately shaped spray of fuel suitable for rapid mixing and combustion. Most engines use either the simplex or the duplex

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nozzle. The exception to this is the Lycoming T53-L-11 engine which uses vaporizer tubes in place of fuel nozzles. Each type of nozzle is discussed in the following subparagraphs.

Simplex nozzle. Figure 2.4 illustrates a typical simplex nozzle; as its name implies, it is simpler in design than the duplex nozzle. Its big disadvantage lies in the fact that a single orifice cannot provide a satisfactory spray pattern with the changes in fuel pressure.

Duplex nozzle. Because the fuel-flow divider and the duplex nozzle work hand in hand, the description of these units is combined. The chief advantage of the duplex nozzle is its ability to provide good fuel atomization and proper spray pattern at all fuel pressures. For the duplex nozzle to work, there must be a fuel-flow divider to separate the fuel into low (primary) and high (secondary) pressure supplies. Single-entry duplex nozzles have an internal flow divider and require only a single fuel manifold, while, as shown in figure 2.5, dual-entry fuel nozzles require a double fuel manifold. The flow divider, whether self-contained in each nozzle, or installed separately with the manifold, is usually a spring-loaded valve set to open at a specific fuel pressure. When the pressure is below this value, the flow divider directs fuel to the primary manifold. Pressures above this value cause the valve to open and fuel is allowed to flow in both manifolds. A fuel flow divider is shown in figure 2.6. In addition, an air shroud surrounding the nozzle, as shown in figure 2.7, cools the nozzle tip and improves combustion by retarding the accumulation of carbon deposits on the face. The shroud also helps to contain the flame in the center of the liner.

5.3 Propeller types

There are four common families of propeller, which I will introduce to you. They are fixed pitch, ground adjustable, in flight adjustable and constant speed. The last two are both examples of variable pitch propellers.

What is pitch?

Propeller theory includes a variety of concepts that may at times be called pitch. Pitch can refer to the blade angle with respect to a flat plane, the distance that a propeller will advance through the air for each rotation or the amount of "bite" that the blade has on the air. Essentially these concepts all describe the same thing. To use our automobile analogy, pitch is like the gear ratio of the gearbox. The important thing to note with pitch, is that it is available in a wide variety of degrees, or 'amounts', much like different gear ratios. To demonstrate, consider the following examples:

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A fine pitch propeller has a low blade angle, will try to move forward a small distance through the air with each rotation, and will take a 'small' bite of the air. It requires relatively low power to rotate, allowing high propeller speed to be developed, but achieving only limited airspeed. This is like having a low gear in your automobile.

A coarse pitch propeller has a high blade angle, will try to advance a long distance through the air with each rotation, and will take a big 'bite' of the air. It requires greater power to rotate, limiting the propeller speed that can be developed, but achieving high airspeeds. This is like having a high gear in your automobile.

Fixed Pitch Propeller

With a fixed pitch propeller, the pitch of the propeller is fixed from manufacture. The performance of your aircraft is determined on the day your propeller is fitted, and is going to be limited within the constraints of the propeller. An analogy with an automobile is as though you had only one gear. Often when choosing a fixed pitch propeller for your aircraft, manufacturers give you a choice of either a climb or a cruise prop. A climb propeller has a relatively fine pitch and a cruise propeller has a relatively coarse pitch. This is like a car manufacturer giving you a choice of a low or a high gear. Either you will be really slow off the mark, or your engine is going to have to be red-lined to get anywhere at a reasonable speed.

Ground Adjustable Propeller

Many propellers manufactured and sold for ultralight and experimental aircraft are ground adjustable. These propellers have the advantage of being able to have their pitch set before each flight if required, taking into account the type of flying you intend to do. More usually however they are used as a low cost way to try out various pitches and settle on the propeller pitch that best suits your aircraft and your style of flying. This can be compared to having a gearbox in your car that you can only change before you set out on your journey.

Variable Pitch Propeller

With a variable pitch propeller, you really have choices. To use the automobile analogy again, your car now has a real gearbox that you can change gear with on the go. (I hope that your car can do this at least!) In addition, rather than being limited to 4 or 5 gears, you can utilise any pitch along the continuum from

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maximum to minimum. The pitch of the propeller may be controlled in flight to provide improved performance in each phase of flight. Typically you would take-off in a fine pitch (low gear) allowing your engine to develop reasonable revs, before increasing the pitch (change up gears) as you accelerated to your cruising speed. You'll end up with the propeller at a relatively coarse pitch, (high gear) allowing the miles to pass beneath you at a rapid rate, while your engine is gently ticking over at a comfortable speed. This feature of a variable pitch propeller will provide you with performance advantages, including:

Reduced take-off roll and Improved climb performance. Fine pitch allows the engine to reach maximum speed and hence maximum power at low air speeds. Vital for take-off, climb, and for a go-around on landing.

Improved fuel efficiency and greater range. Coarse pitch allows the desired aircraft speed to be maintained with a lower throttle setting and slower propeller speed, so maintaining efficiency and improving range.

Higher top speed. Coarse pitch will ensure your engine does not overspeed while the propeller absorbs high power, producing a higher top speed.

Steeper descent and shorter landing roll. With a fine pitch and low throttle setting, a slow turning propeller is able to add to the aircraft's drag, so slowing the aircraft quicker on landing.

Variable pitch propellers actually come in a variety of versions. These different versions refer to the different ways that they are controlled, and include:

Two-position propeller. In flight adjustable propeller. Automatic propeller. Constant speed propeller.

A couple of these are now of historic interest only, so lets concentrate on the two most common options these days; the in flight adjustable operation and the constant speed propeller.

The in flight adjustable propeller allows the pilot to directly vary the pitch of the propeller to the desired setting. Combined with the throttle control, this control allows a wide variety of power settings to be achieved. A range of airspeeds can be maintained while keeping the engine speed within limits. While rare in larger aircraft, the in flight adjustable propeller is the most common type of

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variable pitch propeller that is encountered in sport aviation.

When operated in manual mode, the Airmaster propeller is an example of an in flight adjustable propeller.

Constant Speed Propeller

The constant speed propeller is a special case of variable pitch, which is considered in a family of its own, and offers particular operating benefits.

With constant speed control, the pitch of the variable pitch propeller is changed automatically by a governor. After the pilot sets the desired engine/propeller speed with the propeller speed control, the governor acts to keep the propeller speed at the same value. If the governor detects the propeller speed increasing, it increases the pitch a little to bring the speed back within limits. If the governor detects the propeller speed decreasing, it decreases the pitch a little to bring the speed again back within limits. This operation may be compared to an automatic gearbox in an automobile, where the gears are changed automatically to keep the engine operating at a reasonable speed.

(The governor or constant speed unit [CSU] may be an electronic device which detects the rotational speed of a slip-ring incorporated in the propeller hub and controls operation of a servomotor/leadscrew pitch change actuator in the hub assembly: or it may be an hydraulic fly-ball governor attached to the engine, using engine oil to operate a hydraulic pitch change piston in the hub assembly. In the first case the cockpit control device is likely to be knobs and switches, in the hydraulic system the governor is likely to be cable operated from a cockpit lever.      . . . JB)

A constant speed propeller will automatically deliver you the advantages outlined above for variable pitch propellers, with almost no control required from the pilot. Once a propeller/engine speed is selected, the pilot is able to control the power purely with the throttle (actually controlling manifold pressure, which then determines power output) and the controller will act to keep the propeller/engine speed at the selected setting.

While allowing the pilot to ignore the propeller for most of the time, the pilot must still choose the most appropriate engine/propeller speed for the different phases of flight.

Take-off, go-around and landing. A high speed setting is used when maximum power is needed for a short time such as on take-off. The high speed setting may also be used to keep the propeller pitch low during approach and landing, to provide the desired drag and be ready for a go-around should it be required.

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Climb and high speed cruise. A medium speed setting is used when high power is needed on a continuous basis, such as during an extended climb, or high speed cruise.

Economic cruise. A low speed setting is used for a comfortable cruise with a low engine speed. This operation produces low fuel consumption and longer range, while the advantages of low noise and low engine wear are also enjoyed.

Special Pitch Modes

As well as the ability to vary the pitch of the propeller to optimise the aircraft performance, some variable pitch propellers have some other special modes of operation that can be very useful in certain circumstances:

Feather. A feathering propeller can alter the pitch of the blades up to almost 90 degrees. That is, the blade pitch is changed so that they have their leading edge pointing right into the direction of flight, offering minimum resistance to the airflow. This mode allows the propeller rotation to be stopped, without adding excessive drag to the aircraft. Feather may be used to improve the performance of the aircraft after the failure of an engine, but more usually in light aircraft it is used in motor glider applications. Here the engine is used to gain altitude, before the engine is switched off, the propeller feathered, and then gliding flight commenced.

Reverse. A reversing pitch propeller can alter the pitch of the blades to a negative angle. That is, the blade pitch is changed so that they have their leading edge pointing slightly opposite to the direction of flight. This mode allows reverse thrust to be developed by the propeller. In larger commuter and transport aircraft this feature is often used to slow the aircraft rapidly after landing, but in sport aircraft it is more usually used to enhance manoeuvring on the ground. A popular application is in seaplanes, where the ability to manoeuvre backwards, and sometimes to reduce the thrust to nothing, is especially useful.

5.4 Propeller theory

The forces. Propeller blades are constructed using aerofoil sections to produce an aerodynamic force, in a similar manner to a wing. Consequently the blades are subject to the same aerodynamics – induced drag, parasite drag, wingtip vortices, lift/drag ratios at varying aoa, pressure distribution changing with aoa etc. There is a difference in application because, in flight, the propeller has rotational velocity added to the translational [forward] velocity thus the flight path of any blade section is a spiral – a helical flight path.

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The diagram at left represents a blade section in flight and rotating around the shaft axis. Because of the different application it doesn't serve much purpose to express the resultant aerodynamic force as we would for a wing, with the components acting perpendicular (lift) and parallel (drag) to that flight path, as in the upper figure. So we represent the aerodynamic force component acting forward and aligned with the aircraft's longitudinal axis as the thrust force, and that component acting parallel to the direction of rotation as the propeller torque force.

As you see in the lower figure the component of the lift acting in the rotational plane has now been added to the drag to produce the 'propeller torque force' vector. The remaining forward acting portion of lift is then the thrust. That is why propeller efficiency is usually no greater than 80 – 85%, not all the lift can be used as thrust and the propeller torque force consumes quite a bit of the shaft horse power. The propeller torque and the engine torque will be in balance when the engine is operating at constant rpm in flight.

There are other forces acting on the blades during flight, turning moments that tend to twist the blades and centrifugal force for example. The air inflow at the face of the propeller disc also affects propeller dynamics.

Blade angle and pitchAlthough all parts of the propeller, from the hub to the blade tips, have the same forward velocity, the rotational velocity – and thus the helical path of any blade station – will depend on its distance from the hub centre. Consequently, unless adjusted, the angle of attack, will vary along the length of the blade. Propellers operate most efficiently when the aoa at each blade station is consistent (and, for propeller efficiency, that giving the best lift drag ratio) over most of the blade, so a twist is built into the blades to achieve a more or less uniform aoa.

The blade angle is the angle the chord line of the aerofoil makes with the propeller's rotational plane and is expressed in degrees. Because of the twist the blade angle will vary throughout its length so normally the standard blade angle is measured at the blade station 75% of the distance from the hub centre to the blade tip. The angle between the aerofoil chord line and the helical flight path (the relative airflow) at the blade station is, of course, the angle of attack and the angle between the helical flight path and the rotational plane is the angle of advance or helix angle. The aoa and helix angle vary with rotational and forward velocity.

The basic dimensions of propellers for light aircraft are usually stated in the form of number of blades, diameter and pitch with the latter values given in inches. e.g. 3 blade 64" × 38". The pitch

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referred to is the geometric pitch which is calculated, for any blade station but usually the 75% radius position, thus:

Geometric pitch = the circumference (2 π r) of the propeller disc at the blade station multiplied by the tangent of the blade angle. Thus it is the distance the propeller – and aircraft – would advance during one revolution of the propeller if the blade section followed a path extrapolated along the blade angle. e.g. For a blade station 24 inches from the hub centre [0.75r] and a 14° blade angle, the circumference = 2 × 3.14 × 24 = 150 inches and tangent 14° = 0.25. Thus the geometric pitch is 150 × 0.25 = 38 inches. Propellers are usually designed so that all blade stations have much the same geometric pitch.

Designers may establish the ideal pitch of a propeller which is the theoretical advance per revolution which would cause the blade aerofoil to be at the zero lift aoa; thus it would generate no thrust and, ignoring drag, is the theoretical maximum achievable aircraft speed.

The velocity that the propeller imparts to the air flowing through its disc is the slipstream and slip used to be described as the difference between the velocity of the air behind the propeller ( i.e. accelerated by the propeller) and that of the aircraft. Nowadays slip has several interpretations, most being aerodynamically unsatisfactory, but you might consider it to be the difference, expressed as a percentage, between the ideal pitch and the advance per revolution when the the propeller is working at maximum efficiency in conversion of engine power to thrust power. Slip in itself is not a measure of propeller efficiency; as stated previously propeller efficiency is the ratio of the thrust power (thrust × aircraft velocity) output to the engine power input.

The runaway propeller

As a propeller system increases in complexity then the possibilities for malfunction increase. A problem associated with constant speed propellers is governor failure during flight which, in most installations, will cause the propeller blades to default to a fine pitch limit. This greatly reduces the load on the power plant, and the engine will immediately overspeed, particularly if in a shallow dive. The rpm of an overspeeding engine – sometimes referred to as a 'runaway prop' – will quickly go way past red-line rpm and, unless immediate corrective action is taken, the engine is likely to self destruct and/or the propeller blades depart the hub due to the increased centrifugal force.

The corrective action is to immediately close the throttle and reduce

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to minimum flight speed by pulling the nose up. (But see 'Recovery from flight at excessive speed'). Once everything is settled down fly slowly, consistent with the fine pitch setting, to a suitable airfield using minimum throttle movements. (The constant speed propeller fitted to a competition aerobatic aircraft usually defaults to the coarse pitch limit to prevent overspeeding but an immediate landing is required.)

  

  Blade Face is the surface of the propeller blade that corresponds to the lower surface of an airfoil.

 Thrust Face is the curved surface of the airfoil.

      Blade Shank (Root) is the section of the blade nearest the hub.

     Blade Tip is the outer end of the blade fartest from the hub.     Plane of Rotation is an imaginary plane perpendicular to the shaft. It is the plane that contains the circle in which the blades

rotate.

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      Blade Angle is formed between the face of an element and the plane of rotation. The blade angle throughout the length of the blade is not the same. The reason for placing the blade element

sections at different angles is because the various sections of the blade travel at different speed. Each element must be designed as

part of the blade to operate at its own best angle of attack to create thrust when revolving at its best design speed

      Blade Element are the airfoil sections joined side by side to form the blade airfoil. These elements are placed at different angles in

rotation of the plane of rotation.      The reason for placing the blade element sections at different

angles is because the various sections of the blade travel at different speeds. The inner part of the blade section travels slower

than the outer part near the tip of the blade. If all the elements along a blade is at the same blade angle, the relative wind will not strike the elements at the same angle of attack. This is because of the different in velocity of the blade element due to distance from

the center of rotation.      The blade has a small twist (due to different angle in each

section) in it for a very important reason. When the propeller is spinning round, each section of the blade travel at different speed, The twist in the peopeller blade means that each section advance forward at the same rate so stopping the propeller from bending.

      Thrust is produced by the propeller attached to the engine driveshaft. While the propeller is rotating in flight, each section of the blade has a motion that combines the forward motion of the aircraft with circular movement of the propeller. The slower the speed, the steeper the angle of attack must be to generate lift.

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Therefore, the shape of the propeller's airfoil (cross section) must chang from the center to the tips. The changing shape of the airfoil (cross section) across the blade results in the twisting shape of the

propeller.

      Relative Wind is the air that strikes and pass over the airfoil as

the airfoil is driven through the air.     Angle of Attack is the angle between the chord of the element

and the relative wind. The best efficiency of the propeller is obtained at an angle of attack around 2 to 4 degrees.

     Blade Path is the path of the direction of the blade element moves.

      Pitch refers to the distance a spiral threaded object moves

forward in one revolution. As a wood screw moves forward when turned in wood, same with the propeller move forward when turn in

the air.     Geometric Pitch is the theoritical distance a propeller would

advance in one revolution.

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      Effective Pitch is the actual distance a propeller advances in one

revolution in the air. The effective pitch is always shorter than geometric pitch due to the air is a fluid and always slip.

Forces and stresses acting on a propeller in flight The forces acting on a propeller in flight are :

     1. Thrust is the air force on the propeller which is parallel to the directionof advance and induce bending stress in the propeller.     2. Centrifugal force is caused by rotation of the propeller and

tends to throw the blade out from the center.     3. Torsion or Twisting forces in the blade itself, caused by the

resultant of air forces which tend to twist the blades toward a lower blade angle.

 The stress acting on a propeller in flight are :

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     1. Bending stresses are induced by the trust forces. These stresses tend to bend the blade forward as the airplane is moved

through the air by the propeller.     2. Tensile stresses are caused by centrifugal force.

     3. Torsion stresses are produced in rotating propeller blades by two twisting moments. one of these stresses is caused by the air

reaction on the blades and is called the aerodynamic twisting moment. The another stress is caused by centrifugal force and is

called the centrifugal twisting moment.Control and Operation   (page 1)

Propeller Control          basic requirement: For flight operation, an engine is demanded

to deliver power within a relatively narrow band of operating rotation speeds. During flight, the speed-sensitive governor of the

propeller automatically controls the blade angle as required to maintain a constant r.p.m. of the engine.

           Three factors tend to vary the r.p.m. of the engine during operation. These factors are power, airspeed, and air density. If the

r.p.m. is to maintain constant, the blade angle must vary directly with power, directly with airspeed, and inversely with air density.

The speed-sensitive governor provides the means by which the propeller can adjust itself automatically to varying power and flight

conditions while converting the power to thrust.      Fundamental Forces : Three fundamental forces are used to

control blade angle . These forces are:           1. Centrifugal twisting moment, centrifugal force acting on a rotating blade which tends at all times to move the blade into low

pitch.           2. Oil at engine pressure on the outboard piston side, which

supplements the centrifugal twisting moment toward low pitch.           3. Propeller Governor oil on the inboard piston side, which

balances the first two forces and move the blades toward high pitch            Counterweight assembly (this is only for counterweight

propeller) which attached to the blades , the centrifugal forces of the counterweight will move the blades to high pitch setting

      Constant Speed, Counterweight Propellers The Counterweight type propeller may be used to operate either as

a controllable or constant speed propeller. The hydraulic counterweight propeller consists of a hub assembly, blade assembly,

cylinder assembly, and counterweight assembly.            The counterweight assembly on the propeller is attached to the blades and moves with them. The centrifugal forces obtained

from rotating counterweights move the blades to high angle setting. The centrifugal force of the counterweight assembly is depended on

the rotational speed of the propellers r.p.m. The propeller blades have a definite range of angular motion by an adjusting for high and

low angle on the counterweight brackets.      Controllable : the operator will select either low blade angle or high blade angle by two-way valve which permits engine oil to flow

into or drain from the propeller.

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      Constant Speed : If an engine driven governor is used, the propeller will operate as a constant speed. The propeller and engine speed will be maintained constant at any r.p.m. setting within the

operating range of the propeller.

      Governor Operation (Constant speed with counterweight ) the Governor supplies and controls the flow of oil to and from the

propeller. The engine driven governor receives oil from the engine lubricating system and boost its pressure to that required to operate

the pitch-changing mechanism. It consists essentially of :      1. A gear pump to increase the pressure of the engine oil to the

pressure required for propeller operation.      2. A relief valve system which regulates the operating pressure

in the governor.       3. A pilot valve actuated by flyweights which control the flow of

oil through the governor      4. The speeder spring provides a mean by which the initial load

on the pilot valve can be changed through the rack and pulley arrangement which controlled by pilot.

      The governor maintains the required balance between all three control forces by metering to, or drain from, the inboard side of the propeller piston to maintain the propeller blade angle for constant

speed operation.      The governor operates by means of flyweights which control the position of a pilot valve. When the propeller r.p.m. is below that for which the governor is set through the speeder spring by pilot , the

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governor flyweight move inward due to less centrifugal force act on flyweight than compression of speeder spring. If the propeller r.p.m.

is higher than setting , the flyweight will move outward due to flyweight has more centrifugal force than compression of speeder spring . During the flyweight moving inward or outward , the pilot

valve will move and directs engine oil pressure to the propeller cylinder through the engine propeller shaft.

 Principles of Operation (Constant Speed with Counterweight

Propellers)       The changes in the blades angle of a typical constant speed with

counterweight propellers are accomplished by the action of two forces, one is hydraulic and the other is mechanical.

      1. The cylinder is moved by oil flowing into it and opposed by centrifugal force of counterweight. This action moves the

counterweight and the blades to rotate toward the low angle positon.

      2. When the oil allowed to drain from the cylinder , the centrifugal force of counterweights take effect and the blades are

turned toward the high angle position.      3. The constant speed control of the propeller is an engine

driven governor of the flyweight typeControl and Operation   (page 2)Governor Operation Condition

On-Speed Condition      The on-speed condition exists when the propeller operation

speed are constant . In this condition, the force of the flyweight (5) at the governor just balances the speeder spring (3) force on the

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pilot valve (10) and shutoff completely the line (13) connecting to the propeller , thus preventing the flow of oil to or from the

propeller.

  The pressure oil from the pump is relieved through the relief valve

(6). Because the propeller counterweight (15) force toward high pitch is balanced by the oil force from cylinder (14) is prevented

from moving, and the propeller does not chang pitch Under-Speed Condition

      The under-speed condition is the result of change in engine r.p.m. or propeller r.p.m.which the r.p.m. is tend to lower than

setting or governor control movement toward a high r.p.m. Since the force of the flyweight (5) is less than the speeder spring (3) force , the pilot valve (10) is forced down. Oil from the booster pump flows through the line (13) to the propeller. This forces the cylinder (14)

move outward , and the blades (16) turn to lower pitch, less power is required to turn the propeller which inturn increase the engine

r.p.m. As the speed is increased, the flyweight force is increased also and becomes equal to the speeder spring force. The pilot valve is move up, and the governor resumes its on-speed condition which

keep the engine r.p.m. constant.

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 Over-Speed Condition

      The over-speed condition which occurs when the aircraft altitude change or engine power is increased or engine r.p.m. is tend to

increase and the governor control is moved towards a lower r.p.m. In this condition, the force of the flyweight (5) overcomes the speeder spring (3) force and raise the pilot valve (10) open the propeller line (13) to drain the oil from the cylinder (14). The counterweight (15)

force in the propeller to turn the blades towards a higher pitch. With a higher pitch, more power is required to turn the propeller which inturn slow down the engine r.p.m. As the speed is reduced, the

flyweight force is reduced also and becomes equal to the speeder spring force. The pilot valve is lowered, and the governor resumes

its on-speed condition which keep the engine r.p.m. constant.

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 Flight Operation

This is just only guide line for understanding . The engine or aircraft manufacturers' operating manual should be consulted for each

particular aircrat.      Takeoff : Placing the governor control in the full forward position

. This position is setting the propeller blades to low pitch angle Engine r.p.m. will increase until it reaches the takeoff r.p.m. for

which the governor has been set. From this setting , the r.p.m. will be held constant by the governor, which means that full power is

available during takeoff and climb.      Cruising : Once the crusing r.p.m. has been set , it will be held constant by the governor. All changes in attitude of the aircraft,

altitude, and the engine power can be made without affecting the r.p.m. as long as the blades do not contact the pitch limit stop.

      Power Descent : As the airspeed increase during descent, the governor will move the propeller blades to a higher pitch inorder to

hold the r.p.m. at the desired value.       Approach and Landing : Set the governor to its maximum cruising r.p.m. position during approach. During landing, the

governor control should be set in the high r.p.m. position and this move the blades to full low pitch angle.

Control and Operation   (page 3)Hydromatic Propellers

Basic Operation Principles : The pitch changing mechanism of

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hydromatic propeller is a mechanical-hydraulic system in which hydraulic forces acting upon a piston are transformed into

mechanical forces acting upon the blades.

 Piston movement causes rotation of cam which incorporates a bevel gear (Hamilton Standard Propeller) . The oil forces which act upon

the piston are controled by the governor       Single Acting Propeller: The governor directs its pump output against the inboard side of piston only, A single acting propeller

uses a single acting governor. This type of propeller makes use of three forces during constant speed operation , the blades

centrifugal twisting moment and this force tends at all times to move the blades toward low pitch , oil at engine pressure applied against the outboard side of the propeller piston and this force to supplement the centrifugal twisting moment toward the low pitch during constant speed operation., and oil from governor pressure

applied against the inboard side of the piston . The oil pressure from governor was boosted from the engine oil supply by governor pump and the force is controlled by metering the high pressure oil to or

draining it from the inboard side of the propeller piston which balances centrifugal twisting moment and oil at the engine pressure.

      Double Acting Propeller: The governor directs its output either side of the piston as the operating condition required. Double acting

propeller uses double acting governor. This type of propeller , the governor pump output oil is directed by the governor to either side

of the propeller piston.

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       Principle Operation of Double Acting :

         Overspeed Condition : When the engine speed increases above the r.p.m. for which the governor is set . Oil supply is boosted in

pressure by thr engine driven propeller governor , is directed against the inboard side of the propeller piston. The piston and the

attached rollers move outboard. As the piston moves outboard , cam and rollers move the propeller blades toward a higher angle , which

inturn, decreases the engine r.p.m.          Underspeed Condition : When the engine speed drops below

the r.p.m. for which the governor is set. Force at flyweight is decrease and permit speeder spring to lower pilot valve, thereby open the oil passage allow the oil from inboard side of piston to

drain through the governor. As the oil from inboard side is drained , engine oil from engine flows through the propeller shaft into the outboard piston end. With the aid of blade centrifugal twisting

moment, The engine oil from outboard moves the piston inboard. The piston motion is transmitted through the cam and rollers . Thus,

the blades move to lower angle The Feathering System

Feathering : For some basic model consists of a feathering pump, reservoir, a feathering time-delay switch, and a propeller feathering light. The propeller is feathered by moving the control in the cockpit against the low speed stop. This causes the pilot vave lift rod in the

governor to hold the pilot valve in the decrease r.p.m. position regardless of the action of the governor flyweights. This causes the

propeller blades to rotate through high pitch to the feathering position.

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       Some model is initiated by depressing the feathering button. This action, auxiliary pump, feather solinoid, which positions the

feathering valve to tranfer oil to feathering the propeller. When the propeller has been fully feathered, oil pressure will buildup and

operate a pressure cutout switch which will cause the auxiliary pump stop. Feathering may be also be accomplished by pulling the engine

emergency shutdown handle or switch to the shutdown position. Unfeathering : Some model is accomblished by holding the

feathering buttn switch in the out position for about 2 second . This creates an artificial underspeed condition at the governor and

causes high-pressure oil from the feathering pump to be directed to the rear of the propeller piston. As soon as the piston has moved

inward a short distance, the blades will have sufficient angle to start rotation of the engine. When this occurs , the un-feathering switch

can be released and the governor will resume control of the propeller.

 

  

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To move an airplane through the air, thrust is generated by some kind of propulsion system. Beginning with the Wright brothers' first flight, many airplanes have used internal combustion engines to turn propellers to generate thrust. Today, most general aviation or private airplanes are powered by internal combustion (IC) engines, much like the engine in your family automobile. When discussing engines, we must consider both the mechanical operation of the machine and the thermodynamic processes that enable the machine to produce useful work. On this page we consider the thermodynamics of a four-stroke IC engine.

To understand how a propulsion system works, we must study the basic thermodynamics of gases. Gases have various properties that we can observe with our senses, including the gas pressure p, temperature T, mass, and volume V that contains the gas. Careful, scientific observation has determined that these variables are related to one another, and the values of these properties determine the state of the gas. A thermodynamic process, such as heating or compressing the gas, changes the values of the state variables in a manner which is described by the laws of thermodynamics. The work done by a gas and the heat transferred to a gas depend on the beginning and ending states of the gas and on the process used to change the state. It is possible to perform a series of processes, in which the state is changed during each process, but the gas eventually returns to its original state. Such a series of processes is called a cycle and forms the basis for understanding engine operation.

On this page we discuss the Otto Thermodynamic Cycle which is used in all internal combustion engines. The figure shows a p-V diagram of the Otto cycle. Using the engine stage numbering system, we begin at the lower left with Stage 1 being the

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Centrifugal force Causes the greatest stress Tends to pull the blades from the hub The greater the RPM, the greater the stress

1. Thrust bending forceBends the blade forward at the tips due to high thrust and thin blade cross section

2. Torque bending force

o Tends to bend the blade back against the direction of rotation

o Not significant unless you are considering the power pulses of a recip engine.

3. Aerodynamic twisting moment (ATM)

o Tends to twist blade to a higher blade angle o Caused because center of lift is ahead of center of

rotation. o More apparent at higher angles of attack, and is used by

some design to aid in feathering. (Center of lift moves forward as angle of attack increases.)

4. Centrifugal twisting moment (CTM)

o Tends to decrease blade angle o Caused by all parts of a rotating mass attempting to

move in the same plane of rotation o Stronger force than ATM o Used on some designs to decrease blade angle

Vibrational force During operation, aerodynamic and mechanical forces are

present setting up blade vibrations. These blade vibrations cause the blades to flex which, in turn, causes work hardening and metal fatigue.

Aerodynamic forces cause vibrations primarily at the tip where transonic speeds occur. (example: 72" prop. 2850 RPM-tip velocity is 610 mph)

Mechanical vibrations are induced by engine power pulses o This is more detrimental than aerodynamic vibrations o Engine-induced vibration is from power pulse from

engine o This vibration tends to bend the blades back and forth.

The blades can withstand this, unless something happens to increase the force acting on it.

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o Recognized causes are: 1. Rough running engine 2. Over speed 3. Over boost or excessive M.P. 4. Surface damage 5. Unacceptable straightening 6. Continued operation at an unfavorable RPM

A Prop is like tuning fork. It has natural vibration frequencies, which are determined by shape and thickness of blade.

o If engine vibrations coincide with prop frequency, a situation called resonance peak occurs.

o Resonance peak can be so excessive that if continued, prop will fail.

o Resonance peak can cause tip to vibrate over a gap of several inches.

o Designers fine tune and match prop to engine and airframe

o Red band on tachometer indicates RPM range of unfavorable resonance. Don't operate in red band. May be deceptively smooth or quiet. Tachometer inaccuracy can lead to problems here.

o Engineers attempt to design prop so that these resonance peaks occur outside of the engine operating range. This is why it is critical that only approved prop, engine and airframe combinations be used

Properties of jet fuel

Fluidity     Obviously, jet fuel must be able to flow freely from fuel tanks in the wings to the engine through an aircraft's fuel system. Fluidity is a general term that deals with the ability of a substance to flow, but it is not a defined physical property. Viscosity and freezing point are the physical properties used to quantitatively characterize the fluidity of jet fuel. Jet fuel is exposed to very low temperatures both at altitude – especially on polar routes in wintertime – and on the ground at locations subject to cold weather extremes. The fuel must retain its fluidity at these low temperatures or fuel flow to the engines will be reduced or even stop.

Viscosity    Viscosity is a measure of a liquid's resistance to flow under pressure, generated either by gravity or a mechanical source.

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"Thin" liquids, like water or gasoline, have low viscosities; "thick" liquids, like maple syrup or motor oil, have higher viscosities. The viscosity of a liquid increases as its temperature decreases.

Jet fuel at high pressure is injected into the combustion section of the turbine engine through nozzles. This system is designed to produce a fine spray of fuel droplets that evaporate quickly as they mix with air. The spray pattern and droplet size are influenced by fuel viscosity. If it is too high, an engine can be difficult to relight in flight. For this reason, jet fuel specifications place an upper limit on viscosity.

Fuel viscosity influences the pressure drop in the fuel system lines. Higher viscosities result in higher line pressure drops, requiring the fuel pump to work harder to maintain a constant fuel flow rate. Fuel viscosity also influences the performance of the fuel system control unit.

Freezing Point    Because it is a mixture of more than a thousand individual hydrocarbons, each with its own freezing point, jet fuel does not become solid at one temperature the way water does. As the fuel is cooled, the hydrocarbon components with the highest freezing points solidify first, forming wax crystals. Further cooling causes hydrocarbons with lower freezing points to solidify. Thus, the fuel changes from a homogenous liquid, to a liquid containing a few hydrocarbon (wax) crystals, to a slush of fuel and hydrocarbon crystals, and, finally, to a near-solid block of hydrocarbons. The freezing point of jet fuel is defined as the temperature at which the last wax crystal melts, when warming a fuel that has previously been cooled until wax crystals form. Thus the freezing point of fuel is well above the temperature at which it completely solidifies.  The primary criterion for fuel system performance is pumpability – the ability to move fuel from the fuel tank to the engine. Pumpability is influenced both by fuel fluidity and fuel system design. In lieu of a fuel system flow simulation test, the industry uses freezing point as an indicator of a fuel's low-temperature pumpability. Jet fuel typically remains pumpable approximately 4ºC to 15ºC (8ºF to 27ºF) below its freezing point.1

The U.S. Air Force is evaluating the use of additives that may prevent the formation of large wax crystals that are responsible for reduced fuel flow.

Volatility     Volatility is a fuel's tendency to vaporize. Two physical properties are used to characterize fuel volatility: vapor pressure and distillation profile. A more volatile fuel has a higher vapor

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pressure and lower initial distillation temperatures.

Volatility is important because a fuel must vaporize before it can burn. However, too high a volatility can result in evaporative losses or fuel system vapor lock.

Volatility is one of the major differences between kerosene-type and wide-cut jet fuel. Kerosene-type jet fuel is relatively non-volatile. It has a Reid vapor pressure2 of about 1 kiloPascal (kPa)[0.14 pound per square inch (psi)]. Wide-cut jet fuel has a Reid vapor pressure as high as 21 kPa (3 psi).

Wide-cut jet fuel is better suited for cold weather applications because it has a lower viscosity and freezing point than kerosene-type jet fuel. In such applications, evaporative losses are less of a concern.

Non-corrosivity     Jet fuel contacts a variety of materials during distribution and use. It is essential that the fuel not corrode any of these materials, especially those in aircraft fuel systems. Typically, fuel tanks are aluminum, but fuel systems also contain steel and other metals. Fuel tanks may also have sealants or coatings, and elastomers are used in other sections of the fuel system. Engine and airframe manufacturers conduct extensive fuel compatibility testing before approving a material for fuel system use.

Corrosive compounds potentially present in jet fuel include organic acids and mercaptans. The specifications limit these classes of compounds. By-products of microbial growth also can be corrosive (see Microbial Growth).

Contamination from trace amounts of sodium, potassium, and other alkali metals in the fuel can cause corrosion in the turbine section of the engine.

Four-stroke cycle

The Otto cycle is characterized by four strokes, or straight movements alternately, back and forth, of a piston inside a cylinder:

1. intake (induction) stroke2. compression stroke3. power (combustion) stroke4. exhaust stroke

The cycle begins at top dead centre (TDC), when the piston is furthest away from the crankshaft. On the first stroke (intake) of the piston, a mixture of fuel and air is drawn into the cylinder through

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the intake (inlet) port. The intake (inlet) valve (or valves) then close(s) and the following stroke (compression) compresses the fuel-air mixture.

The air-fuel mixture is then ignited, usually by a spark plug for a gasoline or Otto cycle engine or by the heat and pressure of compression for a Diesel cycle or compression ignition engine, at approximately the top of the compression stroke. The resulting expansion of burning gases then forces the piston downward for the third stroke (power) and the fourth and final stroke (exhaust) evacuates the spent exhaust gases from the cylinder past the then-open exhaust valve or valves, through the exhaust port.

Valve timing

In its original configuration, the four-stroke engine relies entirely on the piston's motion to draw in fuel and air (naturally aspirated engine) and to force out the exhaust gasses. As the piston descends on the intake (inlet) stroke, the increasing volume within the cylinder causes a partial vacuum which draws in the air/fuel mixture. This relies on atmospheric pressure. The intake valve then closes, the piston ascends, and the mixture is compressed and ignited, causing the piston to descend again. As the exhaust valve opens, the piston ascends once more and forces the exhaust gases out. This was the technique used in early four-stroke engines. It was soon discovered, however, that at rotational speeds approaching 100 revolutions per minute (RPM) or greater, the exhaust gasses could not change direction quickly enough to exit past the exhaust valve by the piston's motion alone.

At high rotational speeds, consistent flow through the intake and exhaust ports is maintained by allowing the intake and exhaust valves to be open simultaneously at top dead center (known as valve overlap). The momentum of the exhausting gas maintains the outward flow and creates a suction effect on the cylinder known as scavenging, helping to draw the intake charge into the cylinder. In order to retain efficiency, however, the exhaust valve must be closed soon enough so that too much fuel/air mixture from the intake port is not drawn into the engine's exhaust, wasting fuel. In a high-power situation such as racing, where high engine speeds and forced induction are common, this wasted fuel charge can serve to cool the exhaust valve and prevent detonation.

After ignition of the fuel/air charge, as the piston approaches bottom dead center, combustion slows. Just before the charge is finished burning, the exhaust valve is opened at approximately twenty degrees of crankshaft rotation before bottom dead centre (BDC). This allows the still expanding gas inside the cylinder to push out through the exhaust port, starting exhaust flow and giving the exhaust flow momentum. Though a small amount of force is lost

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through the exhaust port that could be driving the piston, the force that the piston must exert on the gas to exhaust them from the cylinder is reduced, resulting in increased efficiency.

Exhaust systems in many situations are a compromise between cost of production, optimum flow, low emissions, and low noise levels. Also, exhaust gas must be kept away from the air that the engine's driver or pilot or operator breathes. Restrictions in an exhaust system, including emissions equipment, mufflers, and simple exhaust tubing can restrict proper exhaust flow. In multi-cylinder applications, in which many cylinders share a common exhaust pipe, pressure waves created by cylinders exhausting gas can impede flow of exhaust from other cylinders. Since this prevents exhaust gas from exiting the cylinder, the overlap of the intake valve can result in reversion, when exhaust gas enters the intake port. The internal pressure problems due to a multi-cylinder engine sharing a common intake plenum can be overcome by using a carburetor or injector for each cylinder.

Accomplishing maximum volumetric efficiency for a given engine is not a formulaic process. Variables such as flow rates , overlap, valve lift, porting specifications and the location of valve events create a large set of variables. Different intake and exhaust equipment is tested at different speeds and loads, and the end result is usually a compromise between power, emissions, and cost, except in situations where maximum power is desired regardless of cost or emissions (such as racing.) The new volumetric efficiency and valve run are in animations

[edit] Valve train

The valves are typically operated by a camshaft, which is a rod with a series of projecting cams (lobes), each with a carefully calculated profile designed to push the valve open by the required degree at the right moment and to hold it open as required as the camshaft rotates. Between the valve stem and the cam is a tappet, a cam follower, which accommodates variations in the line of contact of the cam. The location of the camshaft varies, as does the quantities. Some engines have overhead cams, or even dual overhead cams, as in the illustration below, in which the camshaft(s) directly actuate(s) the valves through a tappet. This design is typically capable of higher engine speeds due to fewer moving parts in the valve train. In other engine designs, the cam shaft is placed in the crankcase and its motion transmitted by a push rod, rocker arms, and valve stems.

Valve clearance adjustmentThis article or section is not written in the formal tone expected of an encyclopedia article.

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Please improve it or discuss changes on the talk page. See Wikipedia's guide to writing better articles for suggestions.

The valve clearance refers to the small gap between the valve lifter and the valve stem (or the rocker arm and the valve stem) that acts as an expansion joint in the valve train. Less expensive engines have the valve clearance set by grinding the end of the valve stem during engine assembly and is not adjustable afterwards. More expensive engines have an adjustable valve clearance although the clearance must be inspected periodically and adjusted if required. Incorrect valve clearance will adversely affect running of the engine and may result in burned valves and engine damage.

If the valve clearance is too wide the engine will be noisy and can also cause undue wear to the camshaft and valve lifter contact areas. The pushrods can also bend. If the clearance becomes wide enough valve timing will be changed and the result will be poor engine performance. If the valve clearance is too narrow it can cause problems.

A narrow valve clearance will not allow for heat expansion and will result in the failure of the valve to close on its seat. This results in the failure of the combustion chamber to seal and thus poor compression and power. The valve will also become hot and it can melt.

Some valve clearances are adjusted when the engine is cold, others when the engine is hot, according to manufacturer specifications. Some engines have different clearances on the exhaust and intake valves. Since the exhaust valves become hotter they will expand more so the exhaust valves will normally have the larger of the two clearances.

Valve clearance is measured when the piston is at Top Dead Centre of the compression stroke as then all the cylinder's valves are in the closed position. The valve lifter will be resting on the heel of the cam lobe.

Overhead engines adjust the valve clearance with an adjustable rocker arm or by placing shims between the can follower and the valve stem. With the valve cover removed a feeler gauge in accordance with the specification must pass through the clearance space. If the feeler gauge will not fit in, then the clearance is too small. If the blade of the feeler gauge fits in too loose then the clearance is too big. The feeler gauge should fit in and out with a slight drag.

Most modern engines have hydraulic valve lifters that do not need any valve clearance to be set.

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Brake horsepower (bhp)

Brake horsepower (bhp) is the measure of an engine's horsepower without the loss in power caused by the gearbox, generator, differential, water pump and other auxiliaries. Thus the prefix "brake" refers to where the power is measured: at the engine's output shaft, as on an engine dynamometer. The actual horsepower delivered to the driving wheels is less. An engine would have to be retested to obtain a rating in another system. The term "brake" refers to the use of a band brake to measure torque during the test (which is multiplied by the engine speed in revs/sec and the circumference of the band to give the power).

Shaft horsepower (shp)

Shaft horsepower is the power delivered to the propellor shaft of a ship or turboprop airplane. This may be measured, or estimated from the indicated horsepower given a standard figure for the losses in the transmission (typical figures are around 10%). This metric is uncommon in the automobile industry, through drivetrain losses can be significant.

[edit] Effective horsepower (ehp)

Effective horsepower is the power converted to useful work. In the case of a vehicle this is the power actually turned into forward motion.

In automobiles, effective horsepower is often referred to as wheel horsepower. Most automotive dynamometers measure wheel horsepower and then apply a conversion factor to calculate net or brake horsepower at the engine. Wheel horsepower will often be 5-15% lower than the bhp ratings because of a loss through the drivetrain.

Labyrinth seal

From Wikipedia, the free encyclopedia

Jump to: navigation, searchFor more uses of the word labyrinth, see Labyrinth (disambiguation)

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A simple labyrinth seal

A labyrinth seal is a mechanical seal that fits around a shaft to prevent the leakage of oil or other fluids.

A labyrinth seal is composed of many straight threads that press tightly inside another shaft, or stationary hole, so that the fluid has to pass through a long and difficult path to escape. Sometimes 'threads' exist on the outer and inner portion. These interlock, to produce the long characteristic path to slow leakage. For labyrinth seals on a rotating shaft, a very small clearance must exist between the tips of the labyrinth threads and the running surface.

Labyrinth seals on rotating shafts provide non-contact sealing action by controlling the passage of fluid through a variety of chambers by centrifugal motion, as well as by the formation of controlled fluid vortices. At higher speeds, centrifugal motion forces the liquid towards the outside and therefore away from any passages. Similarly, if the labyrinth chambers are correctly designed, any liquid that has escaped the main chamber, becomes entrapped in a labyrinth chamber, where it is forced into a vortex-like motion. This acts to prevents its escape, and also acts to repel any other fluid. Because these labyrinth seals are non-contact, they do not wear out.

Turbines use labyrinth seals due to the lack of friction, which is necessary for high rotational speeds.

Labyrinth seals are also found on pistons, which use them to store oil and seal against combustion explosions, as well as on other non-rotating shafts. In these applications, it is the long and difficult path and the formation of controlled fluid vortices plus some limited contact-sealing action that creates the seal.

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