aircraft structures

J.A.R. 66J.A.R. 66MODULE 11aMODULE 11a

AIRFRAME STRUCTURES

GENERAL CONCEPTS

INTRODUCTION• An aircrafts structure is designed to withstand flight forces whilst

remaining aerodynamically efficient. The forces are:

Lift

Drag

Weight

Thrust

An aircrafts structure is generally made of three different types of member

Beams - resist bending moments

Struts - compression loads

Ties - tensile loads

• Any load or bending moment acting on the aircrafts structure could be measured in pounds, kilograms or Newton's.

• The most important factor is the relationship between this load and the cross-sectional area of the member.

• THIS IS CALLED STRESS and is measured in pounds per square inch, Newton's per square meter or another suitable unit.

• Or simply : STRESS =FORCE

AREA

• In a simple beam there are two stresses trying to cause failure, stress due to bending and stress due to shear. In diagram (a) on the following slide, the bending moment is the greatest in the centre where the load is applied

• In diagram (b), however, which uses a cantilever structure, the bending moment is most at the support end.

(a) Simply supported beam (b)Cantilever beam

• In practice a beam can be tapered to evenly distribute the stress along its length.

• Shear force is greatest near the points of support and least at the centre of the span in a simply supported beam.

• In a cantilever beam, shear force does not vary along the length of the beam if the weight of the beam is ignored.

• We have already established that a strut is a member in compression and a tie is a member in tension.

• Struts are often hollow, whereas ties are mostly solid, rods or wires.

• In practice few members are one thing or the other due to the complex stresses they will take.

• When a member is subjected to a load, it will move under the effect of this load, e.g. a tie will stretch.

• This movement is called STRAIN.• STRAIN = CHANGE OF LENGTH

ORIGINAL LENGTH

A beam is a member subjected to bending, it may be supported at both ends or only at one end.

A beam supported at one end only is called a CANTILEVER.

When a beam is put under load, the surface on the outside of the bend will be in tension and the surface on the inside of the bend will be in compression.

• If the beam is a solid member there will be a plane in the middle that is neither in compression nor in tension.

• This is known as the neutral plane of the neutral axis.

• It is a requirement that all aircraft under construction must conform to minimum airworthiness requirements as laid down in BCAR’s.

• The minimum SAFETY FACTOR specified in BCAR’s is 1.5 (with provisions for castings and forgings).

• The SAFETY FACTOR can be considered as the relationship between the capability of a structural member to carry the working load and its ultimate strength (breaking point of material).

LOAD FACTORS

• A major consideration in aircraft design is the variation of loads which the structure must withstand

• A load factor is used to identify the specific number of cycles the structure will withstand before failure.

NORMAL LOAD

• This can be considered as the load imposed on the structure which is directly attributed to the weight of the aircraft in straight and level flight.

OPERATING LOAD

• An aircraft in certain modes of flight will have an increased load imposed on its structure.

• The manufacturer will impose limits on the operational capabilities of its aircraft.

• This limit (or operating load) is generally less than two thirds of the aircrafts ultimate load due to permanent distortion and damage caused as the airframe approaches its ultimate load.

ULTIMATE LOAD

• This is the maximum load that the aircrafts structure will withstand before complete failure. It covers factors such as:

• Variation and degradation of materials used• Flight outside the aircrafts permissible envelope.• This load forms the basis for minimum structural

requirements in the initial design and construction of the aircraft.

STRUCTURAL CLASSIFICATION

• The structure of an aircraft falls into three main categories:

• Primary

• Secondary

• Tertiary

PRIMARY STRUCTURE

• This can be described as any part of the aircraft structure, which upon failure, will cause the catastrophic loss of control of the aircraft, or inhibit the operation of essential services/equipment.

• Frames• Main spars• Engine mount points

SECONDARY STRUCTURE

• This can be described as sections of the structure which have a considerable amount of strength above the design specifications of the aircraft.

• Stringers

• Fuselage skin

• Wing ribs

TERTIARY STRUCTURE

• This portion of the structure can be classified as any part that does not incur structural loads.

• They are normally sections of structure that are necessary for aerodynamic reasons only.

• Fairings

• Support brackets

FAILSAFE STRUCTURE

• In the event of failure of a primary section of the airframe, there is sufficient strength in associated parts of the localised area which failure has occurred to maintain the aircraft in flight for a limited period of time.

• This is dependant on the nature of the failure , and how widespread the failure is.

• A structural failure can be described as complete or partial cracking of a portion of primary or secondary structure.

• The nature of the failure can usually be attributed to:

• Fatigue loads

• Corrosion

• Stress corrosion

• Accidental damage

• In addition to the structural strength of this type of structure, the crack propagation characteristics of the materials used are sufficiently slow.

• The design of the structure allows for ease of inspection and monitoring.

• This allows for a limited operational period where an acceptable but reduced airworthiness status is maintained and monitored for structural strength and stiffness.

FAIL-SAFE ESSENTIAL FEATURES

• The materials used should have low crack propagation and residual strength properties

• The material thickness is chosen so that stress levels are low for the critical fatigue load conditions

• The structure must have alternative load paths and crack stoppers

• Sufficient structural access for visual and NDT testing.

SAFE-LIFE STRUCTURE

• A safe-life structure possess a relatively quick crack propagation characteristic

• It is therefore essential that the safe-life component is removed before cracking commences

• In comparison to fail-safe structure, safe-life will retain its full static strength and stiffness throughout its designated operational life.

• The safe-life or fatigue life of a component must be established for every component.

• This life is expressed in terms of flying hours, number of flights or number of application loads.

• In order to establish a safe-life, it is necessary to estimate the pattern of repeated loads in service and to carry out loading tests on the components.

• Theses tests are used to establish the safe-life of individual components.

• The fatigue behaviour of a structure can be assumed as:

1. The period before crack propagation commences

2. The period of steady crack propagation

3. The period of rapid or unstable crack propagation

• While fail-safe structure can be allowed to extend over stages 1 and 2.

• Safe-life structure is restricted to stage 1 only

An example of a Safe-Life spar

• The spar illustrated on the previous slide is manufactured in two sections.

• The top section consists of a cap, riveted to the upper web plate.

• The lower section is a single extrusion consisting of a lower cap and web plate.

• If either section of this type of spar should fail, the other will take the load.

• This type of construction follows the Fail-Safe concept.

DAMAGE TOLERANCE

• As previously stated the main drawback of a Fail-Safe structure is the additional weight of the structure through the use of additional structural members.

• Damage tolerant structures eliminate the need for additional structure by distributing the load of a particular structure over a larger area.

• This concept ensures that if any part of the structure suffers cracking or partial failure, the remaining structure will be able to withstand reasonable loads without failure until the damage is detected.

• This means that during normal inspection cycles, any damage will be found before a failure occurs.

• The following slide illustrates this type of structure.

• The previous slide shows a wing to body attachment (both old and new design).

• The original design relies on the pin to accommodate the entire load. This arrangement incurs serious fatigue due to the fluctuating loads imparted on the pin.

• The original design must also possess a reinforced section of structure to accommodate the pin, therefore increasing the overall weight of the component.

• The new design incorporates a fixed joint that runs the entire depth of the spar.

• This arrangement gives us a better load transference and also minimises fatigue at the joint.

• By having numerous fasteners (attachment points). The concept of Fail-Safe is maintained.

• Fatigue will be easily identifiable by the degradation of the fasteners before a failure occurs.

MULTIPLE LOAD PATH

• This method incorporates a number of small members, each carrying a small percentage of the total load.

• In the event of failure of one load member, the other members will “take on” the load.

• This aileron attachment is a typical example of a damage tolerant structure.

• This structure also incorporates primary and secondary structure.

• A further method of achieving damage tolerance is by the use of crack limiting joints.

• This is where parts of the structure are designed with crack limiting joints.

• These also ensure that these cracks can be easily detected during normal inspection cycles.

• The following slide illustrates additional skin cleats which restrict cracks to one bay width.

Skin

cleat

cleat

stringer

skin

• The fail-safe concept is not restricted to structural components.

• If we look at a passenger cabin window, you will note that it has to withstand the pressurisation of the fuselage.

• In the next slide you should note that both the inner and outer pane are more than capable of taking these loads.

• It is therefore logical to assume that if one fails, the other is more than capable of withstanding the load, hence this may be assumed as Fail-Safe.

ZONAL AND STATION ZONAL AND STATION IDENTIFICATION SYSTEMSIDENTIFICATION SYSTEMS

LOCATION NUMBERING SYSTEM (STATION NUMBERS) FUSELAGE

STATIONS

• In order to identify a part of an airframe, a station number is used.

• The station number is often a measurement in inches or millimetres from a fixed point in the airframe. (station zero).

• Station Zero may be the nose of the aircraft, where station 500 may be 500 inches or millimetres aft from the nose.

• Alternatively, if the centre of the airframe is used as zero, then 500f would be 500 inches forward of zero, or, 500a would be 500 inches aft of zero.

WATER LINES

• Vertical measurements are made along water lines and are measured from water line zero.

• In the following slide we can see that the cabin floor is WL-16 which means that it is 16 inches below the water line.

An example of a water line

BUTTOCK LINES

• Distances to the right or the left of the centre line of the fuselage are measured by buttock lines

• The tip of the horizontal stabiliser is located at BL108.88 which means that it is 108.88 inches from the fuselage centre line.

Wing station may also be measured from the fuselage centre line or possibly from the wing root

IN SUMMATION

• Fuselage stations are used for locations fore and aft along the fuselage.

• Water lines locate positions vertically on the fuselage.

• Buttock lines locate points to the right or left of the centre line of the fuselage.

• Wing stations are measurements along the span of the wing, with wing station zero being the centre line of the fuselage.

ZONING

• The following diagrams show how an aircraft is divided into Major Zones and Major Sub-Zones.

• All zones are identified by three digit numbers.

• The first digit defines the major zone.• The remaining two digits denote the major

sub-zone.

Major zones

Major sub-zones

ACCESS DOORS AND PANELS

• Each access door or Panel is identified by by a numerical/alpha number.

• The numeric component of the number is the three digit number which identifies the smallest zone in which the door is located.

• The two element alpha suffix designates the location of the door within the zone.

• The first alpha letter identifies the door within the zone in a logical sequence, i.e.: inboard to outboard.

• The second alpha letter identifies the door in its relation to the aircraft as follows:

• T = top

• B = bottom

• L = left hand (port)

• R = right hand (starboard)

• Z = internal

• For example : access door 230 ER would be the fifth access door aft of frame 74 on the fuselage right side.

AIRCRAFT STRUCTUREAIRCRAFT STRUCTURE

DESIGN AND PHILOSOPHYDESIGN AND PHILOSOPHY

• If we take the Airbus as an example of a typical modern liner design:

• The structure is designed to be both fail-safe and damage tolerant.

• The basic material used is a high strength aluminium alloy, with certain components and fittings made from steel or titanium.

• Glass fibre laminates and non-perforated honeycomb structures are also used in small areas of the structure.

• A safe structural life is established by a fatigue test programme that exceeds the expected working life of the airframe.

• Through the use of this programme, the manufacturer is able to identify any possible weak points in the primary structure and issue modifications.

FATIGUE TESTING

• These tests are carried out with the use of rigs on separate sections of the structure.

• By carrying out these tests on separate sections it ensures that the tests start early in the production life, and any tests or inspection problems do not restrict work on other sections of the airframe

• The fatigue test programmes cover a total of 120,000 flights

• The tests are divided between random operational loadings and investigations of crack propagation.

• The first stage of 96,000 simulated flights give a margin of 100% over the design target for economic repair.

• The second stage involves a series of 24,000 simulated flights checking crack propagation.

• These are carried out on selected areas of the structure that have been weakened by saw cuts.

• The object of this exercise is to compare crack propagation with calculations to establish the tolerance to damage of the structure.

TEST LOADING OF THE SECTTIONS

• These simulated flight and ground loads are applied to the structure through dual action hydraulic jacks that transmit both compression and tension loads.

• The simulated pressurisation loads are applied by compressed air .

• For safety and time saving, the internal fuselage is filled to 75% of its volume with polystyrene blocks.

• The whole test programme is computer controlled, comparing actual with calculated loads. Any significant difference automatically halting the test.

FATIGUE

• As we have already established, a continuous increasing load on a piece of material will eventually reach its ultimate static load and fracture it.

• If, however the load is significantly less than the ultimate static load, but is applied in continuous cycles, the material will stretch, but return to its original shape when the cycle is complete.

• When visually inspected, the material will show no sign of degradation of its structure.

• However, the material is actually work hardening and will eventually after a certain number of cycles, weaken and crack.

• This phenomenon of fracturing after a series of cyclic loads is known as fatigue.

FATIGUE TESTING

• The fatigue testing of an airframe is carried out on a structure taken off the production line.

• It is then subjected to a series of tests designed to simulate the airframes working life.

• By this method, an equivalent of many thousands of actual flying hours can be reproduced on the airframe in a short period of time.

• The actual testing and recording of flight forces, weights, height, airspeed, etc are controlled by computer.

• After each period of testing is complete. The structure is scrutinised for signs of failure or material degradation.

• The results of the tests, together with the data available for the structural materials, enable an accurate prediction to be made on how long, in terms of aircraft life and total fatigue, the aircraft can safely remain in service.

• The aircraft , or parts of it, is then scheduled for replacement before its safe life is exhausted.

• Some aircraft, in particular military aircraft, use a fatigue indexing system.

• The safe life or fatigue life of a particular aircraft is given in terms of a fatigue index.

• A higher fatigue index may be given to aircraft if the airframe has been subjected to further tests or if major structural parts have been replaced.

• To determine the amount of fatigue index used for a particular flight, requires a fatigue meter to be fitted.

• The fatigue meter will measure the varying loads imposed on the airframe by the action of manoeuvres, landings or atmospheric conditions and converts these into “G” loads, which are recorded by the fatigue meter.

• On landing, the “G” load figures are recorded and a calculation of the amount of fatigue life consumed during the flight is made.

STRUCTURAL SURVEYS

• Structural surveys and NDT programmes are carried out to ensure the continuing integrity of an airframe.

• As the aircraft is maintained to an approved maintenance schedule, these surveys provide a continuous assessment of the airframes structural condition.

• With long life aircraft, a more in-depth inspection may be required by the manufacturers.

• This may take the form of a structural integrity programme or structural surveys, where more in-depth inspections may need to be carried out such as NDT testing.

STRUCTURAL SURVEYS

• Information for structural surveys can be found in CAA Airworthiness notice 89.

• This document outlines inspection requirements relating to structural re-assessment, and also the development of a continuing structural inspection programme for older transport aircraft.

STRUCTURAL RE-ASSESSMENT

• This involves:-– The identification of structural parts which

contribute significantly to carrying flight, ground and pressure loads whose failure would affect the structural integrity of the aircraft. These items are know as structurally significant items. Typical examples are spars, bulkheads, landing gear, beams, etc.

– The establishment of a procedure for developing programmes that provide a high probability of detecting fatigue damage before the residual structural strength falls below the fail-safe requirements.

• A programme such as this is normally directed at older types of aircraft, as these are normally the aircraft with the highest number of flight cycles and are therefore most likely to experience initial fatigue damage.

• The inspection programme is normally directed to aircraft that have high cycle figures and is usually requested by the aircraft manufacturer.

• The areas of the structure that the inspection is concentrated on is decided on by the manufacturer.

• Additionally, structural sampling may be required by the manufacturer, where a section of structure is removed and laboratory tests are carried out to determine the exact degradation of the construction materials.

NDT INSPECTION PROGRAMMES

• NDT inspections may be called up at various times during an aircrafts life. For example:– Scheduled maintenance– Implementation of service bulletins– Special structural survey checks

• The use of NDT equipment and methods are carried out to detect degradation within aircraft structures and components before it causes major problems.

• To achieve this, NDT inspections are carried out at set times. These times or frequencies can be dependant upon cycles, hours, landings, etc, and are normally incorporated into maintenance schedules.

• NDT checks may also be called up during special structural survey checks. The most commonly used forms of NDT checks are:-

• Visual

• Penetrant (dye)

• Magnetic particle

• X-ray

• Ultrasonic

• Eddy Current

• One or more of the afore mentioned methods may be called upon by the manufacturer in order to establish the best course of repair to the particular component or section of airframe.

• There is no best inspection method. Most NDT procedures are used when they are best suited to the task in hand and are dependant on time-scale, access, requirements, etc.

AIRCRAFT DRAINAGE

• Drainage and drain paths are provided in aircrafts structures to prevent the collection of water and other fluids within the structure.

• The collection of such fluids could cause/accelerate corrosion, or fire.

• Airframe drainage may be separated into two main categories:-

– External drains

– Internal drains

EXTERNAL DRAINS

• Exterior drains are located on exterior surfaces of the fuselage, wing and empennage to dump fluid overboard.

• In non-pressurised areas the drain ports are always open, but in pressurised areas, air leakage has to be considered.

• There are three main types of drain valve used for this purpose

• Two of the types mentioned rely upon pressurised air within the cabin to keep the valves closed.

• Both are open when the aircraft is stationary on the ground, allowing fluids to drain overboard.

• During flight the valves close, thus preventing any air losses from the cabin area.

• One of the above valves uses a rubber seal, while the other uses a spring seal.

• You should also note the use of a levelling compound, normally a rubberised sealant which fills the cavity and brings the level up to the drain hole to ensure all fluids drain out.

• This type of valve uses cabin pressure air to close off the drain path by moving the plunger down to seal the drain.

• When cabin pressure falls, the spring assists the valve to open and drain any fluid to atmosphere.

• To enable external drains to function correctly, some means must be provided within the airframe structure to ensure that all fluids are directed towards the external drain points, this is achieved by using internal paths and drain holes.

• The internal structure is provided with tubes channels, dams and drain holes to direct the flow of fluid towards external drain points.

• A typical example of this would be holes drilled in the stringers to allow the flow of fluid to the drain.

• Other methods of draining are shown in the next slide where drain tubes remove the fluid from the pressure cabin and direct them through tubes in the wing to the body fairing and dump the fluids overboard.

LIGHTENING STRIKES

• Lightening is the discharge of electricity between highly charged cloud formations or the cloud formation and the ground.

• If an aircraft were to be struck by either type of lightening, it would result in very high voltages being passed through the structure.

• All separate parts of the structure are electrically bonded together to direct the lightening strike away from sensitive parts of the airframe such as fuel tanks or flying controls.

• Lightening strikes may have two effects on an aircraft:-– Strike damage– Static discharge damage

STRIKE DAMAGE

• This is the point where the charge enters the aircraft.

• Strike damage is generally found at the wing tips and other extremities of the structure.

• It is normally in the form of small circular holes in the exterior skin, either in clusters or spread out over a wide area and often shows signs of discolouration, blisters on radomes and cracks in glass fibre.

STATIC DISCHARGE

• This is the point where the charge leaves the aircraft.

• Static discharge damage is in the form of localised pitting and burning at trailing edges.

• Because of the costs involved in replacing entire control surfaces through discharge, manufacturers place “static wicks” at known discharge points.

• The replacement of a static wick over an entire control surface is less costly and less time consuming.

INSPECTION

• Since lightening and turbulence both occur in thunderstorms, an inspection for lightening damage will often be carried out following a report of flight through severe turbulence.

• The areas stated in the manufacturers guidelines should be examined for signs of strike or static discharge damage.

• Bonding strips and static discharge wicks should be examined for burning and disintegration.

• All control surfaces including flaps, spoilers and tabs should be inspected for damage at their hinge bearings.

• Unsatisfactory bonding may have allowed static discharge and tracking across the bearings, causing burning, break-up or seizure.

• A check for roughness and control surface resistance to movement at each bearing will usually indicate damage at such points.

• In addition, the following inspections should be carried out:-

• Examination of engine cowlings for signs of burning or pitting.

• Examination of fuselage skin and rivets generally for signs of burning and pitting.

• If the landing gear was extended at the time of strike, examine the lower parts for static discharge damage and also check for residual magnetism.

• The inspections outlined should be followed by functional checks of the radio and radar equipment, instruments, compasses, electrical circuit and flying controls, in accordance with the relevant chapters of the approved maintenance manual.

• On some aircraft a bonding resistance check on radomes may also be specified.

FLIGHT THROUGH SEVERE TURBULANCE

• If an aircraft has flown through conditions of severe turbulence, the severity may be difficult to establish.

• An indication may be found in the accelerometer or fatigue meter fitted to some aircraft.

• However, these instruments are designed to record steady loads and force peaks and recorded data through turbulence may be exaggerated due to instrument inertia and should not be taken as actual loads

• Generally if a reading that exceeds –0.5g and +2.5g is recorded on a transport aircraft, then some damage may be found.

• With other types of aircraft (e.g.. Aerobatic, semi-aerobatic), accelerometers and fatigue meters are seldom fitted and reported.

• In this instance flight through severe turbulence should always be investigated.

• Severe turbulence may cause excessive vertical or lateral forces on the aircraft structure and the effects may be increased by the inertia of heavy components such as engines, fuel tanks, water tanks and cargo.

• Damage may be expected at main assembly points such as the wing-to-fuselage joints, tail-to-fuselage joints and engine mountings.

• Skin wrinkles, pulled rivets or visible distortion may also be evident.

AIRCRAFT GROUNDING

• The most common type of ground used at all principal airports is the grid system, in which any number of individual grounds are linked to a common ground.

• Should a grid system not be available, then an individual ground will be used.

• If a grid system is employed, it must be of an approved type.

WARNING!!

• Do not use head-sets or handle electrical connections to the aircraft during severe electrical disturbances since YOU MAY BECOME A GROUND CONNECTION YOURSELF!!

• Static grounding of aircraft is achieved by attaching approved grounding cables to specified locations on the landing gear assemblies.

• The exact locations will be specified in the aircraft maintenance manual, Chapter 20 (ATA system).