rockets systems

7/24/2019 Rockets Systems

1/93

UNIT-1

ROCKET SYSTEMS

TYPES OF IGNITER:

The types of igniters which are commonly used are,

Gaseous Igniter

Liquid igniter

Solid igniter

GASEOUS IGNITER:

It is the old and primitive type of igniter which is not used now. In this type of igniter

the reactive gaseous mixtures are held in a very thin tube with high pressure. It is hazardous innature and reliable. Directional control can be done by using burst dampers.

Example for gaseous igniters is shock tube.

LIQUID IGNITER:

Liquid igniter is of two types. Theyare,

Liquid- Liquid type , which is known as hypergolic igniter

LiquidSolid type, which is known as hybrid igniter

CHARACTERISTICS OF HYPERGOLIC LIQUIDS:

Hypergolic liquids have a very high bulk density.

Ignition delay for these types of liquids should be less than 50 milliseconds.

These liquids are chemically instable.

They must be work well together with some of selected polymers and resins.

Their viscosity should be less than 10 centistokes.

They should have a very low vapour pressure.

They should have a very good heat transfer characteristics.

SOME COMBINATIONS OF HYPERGOLIC LIQUIDS:

FUEL OXIDIZER

KEROSINE RFNA

HYDRAZINE CHLOROFLUORINE

AMMONIA OXYGEN

HYDROGEN ClO3F


2/93

FACTORS AFFECTING IGNITION DELAY:

The factors which affect the ignition delay are,

Purity of materials

Initial temperature and pressure.

t = A

where ,

t = Time

A= Minimum possible ignition delay

E = Temperature coefficient

R = Universal Gas constant

T =Temperature

SOLID ROCKET IGNITER:

Solid rocket igniters are broadly classified as follows,

dvedgeldv

SOLID IGNITER

TOTALLY CONFINED

IGNITER

UNCONFINED IGNITER

NOZZLE IGNITER

BAG

IGNITERPOWDER CAN

IGNITER

JELLY ROLL

FILM IGNITER PYROCORE

CONDUCTING

FILM IGNITER

BASKET

IGNITER ALCO JET

PYROGEN


3/93

TOTALLY CONFINED IGNITERS:

BAG IGNITER:

It is the old and primitive type of igniter.

We dont have enough control over ignition in this type of igniter

After the ignition of fully charged bag igniter, the heat and pressure generation occurs.

The rate of heat and pressure release is very high and there is a possibility of bursting.

ADVANTAGES:

It is very easy to fabricate

The cost of production is very low.

DISADVANTAGE:

This particular system is very far from meeting the requirements of modern high performance

rocket motors.

POWDER CANIGNITER:

In this type of igniter pallets are used .Pallets are made up of black powder or metal oxidantsand aluminium powder. Here directional control is done but not sufficient. It is only suitable for


4/93

small rocket motors and not suitable for large rocket motors because of its erratic transient ignition

characteristics and it is rapturous.

ADVANTEGES:

Ease of fabrication and production cost is low.

DISADVANTAGES:

As the igniter is made of steel casing the weight is much heavier.

Only suitable for short range missions.

JELLY ROLL:

It consists of a film coated pyrotechnic and a binder. Then the film is rolled over a rod with a

squib support at the front and back. Addition to that a rubber support is given externally. Ignition is

generally started at the squib. In jelly roll the ignition transfers layer by layer. Productive cover is

used to tight the main charge.

ADVANTAGES:

These igniters are nozzle insertables.

They make efficient use of motor fuel volume.

The hardware weight is low.

DISADVANTAGES:

They are very fragile and not suitable for large rocket motor

They are difficult to manufacture and the principle of operation is complex.

They produce high shocks.

UNCONFINED IGNITERS:

Actually they are confined. They are unconfined only relative to others.

FILM IGNITER:


5/93

The film igniteris produced by painting an ignitable fuel oxidizer binder mixture directly ontothe

propellant surface. The film normally contains,

Fine metal powder aluminium powder

Per chlorate oxidizer - ammoniumper chlorate

Polymeric binder.

The film can be activated by the conventional pyrotechnic igniter. It permits the use of low

conventional ignition system and has often be used an aid to ignite the systems which handle

materials difficult to ignite.

CONDUCTING FILM IGNITER:

It contains the strips of pyrotechnic material applied directly to the propellant, which can

overlay of circuit leads. It consists of the application of thin strips within the perpendicular overlay of

actuation circuitry. A typical pyrotechnic mixture consists of metal powder,per chlorateoxidizer ,

silver conductor and the polymeric binder. Aluminium foils are used as protective layer of

conducting film igniter.

ADVANTAGES:

These igniters produce low pressure peaks

They make efficient use of space

They are intensive to electromagnetic radiation

DISADVANTAGES:

They are very difficult to apply

Quality control is difficult

They cannot be removed from the motor easily

They are very sensitive to friction and resistance

NOZZLE IGNITERS (or) BASICALLY CONTROLLED IGNITERS:

BASKET IGNITER:

This type of igniter contains pallet charges. Basket igniter are fabricated from heavy wire

mesh , perforated sheet metal or perforated glass fibre reinforced resins. The perforated container

retain the high surface area palette charge when it burns. The exhaust products ejected in a pattern

determined by the geometry of the design contains reactive products as well as reactive materials.


6/93

This system makes available in the wide choice of configuration allowing for some flame pattern

control.

ADVANTAGES:

This igniter is made efficiently strong to withstand environmental conditions.

Proper control of length and port area can furnish a controlled flame pattern and givemedium to fast ignition with low ignition charge.

DISADVANTAGES:

The hardware weight is high.

Forward attachment is often difficult

The burning area of the pallets cant be readily determined

Internal igniter pressure and mass delivery rate are difficult to determine.

PYROGEN:

A pyrogenigniter consists of small nozzle pressure chamber containing high energy fast

burning rocket propellant usually having a complex geometry.

Essentially it is a rocket motor within a rocket motor. The design is especially used in very

large motor.


7/93

ADVANTAGES:

The igniters have little or no shocks.

They eliminate the handling of large amounts of relatively hazardous metal oxide

charges.

They are adoptable to either head end or launcher mount applications.

DISADVANTAGES:

The pyrogen must itself have an igniter and itstherefore depends upon the principle used

to ignite.

ALCOJET:

There are two tubes in this igniter .In the annular space between the two tubes, we have main

charge. Booster charge present inside the tube. The booster charge is first ignited. The ignitionpasses through the perforations in the inner wall to the main charge. There are perforations in the

outer tube through which flame comes out. Since there is a control, it is a ballistically controlled

igniter.

LIVE IGNITERCOMPONENTS:

The important components of a live igniterare ,

Firing console

Squib Transfer charge

Booster charge

Main charge

Motor grain

SQUIB :

The squib is the primary element for ignition that affects the conversion of electrical impulse

from the control console to chemical reaction in the rocket motor.

The squib consists of the following parts,


8/93

1. INERT COMPONENTS :

Circuit element

Base or body

Insulation

Metal case

2.

ACTIVE COMPONENTS :

Prime charge

Booster charge

Main charge

CHARACTERISTICS OF A SQUIB:

1.

A functioning time curve

2. Pressure output characteristics

3. Thermal output characteristics

4.

Auto ignition characteristics

5. Static sensitivity characteristics

6. Shock and mechanical sensitivity characteristics

IGNITER DESIGNCONSIDERATION :

The data to be considered while designing an igniterare,

The pyrotechnic material data

Propellant ignitability data

Rocket motor data

Back up data (previous test firing data).

IGNITABILITY BOMB:

The ignitability bomb is a device used to determine the relative ignitability of the propellants at

various pressures under the direct fire of ignition materials.

INJECTORS :

An injector or ejector is a system of admitting the fuel into the combustion engine. Its function

is similar to a carburettor.

PRIMARY DIFFERENCE BETWEEN A CARBURATOR AND AN INJECTOR:

In an injector the fuel injection atomizes the fuel by forcibly pumping it through a small nozzle under

high pressure while a carburettor relies on suction created by intake air rushing through a venturi

to draw the fuel into the airstream.

FUNCTION OF AN INJECTOR:

The injectors are mainly used to meter the flow of the liquid propellant to the combustion

chamber which causes the liquids to be broken into small droplets. This process is known as


9/93

atomization. It also helps to distribute and mix the propellant in a correctly proportionate mixture of

fuel and oxidizer, which results in uniform propellant mass flow.

INJECTION HOLE PATTERNS:

The injectionhole pattern on the face of the injector is closely related to the internal manifolds orfeed passages. These hole patterns provides the distribution of propellant from the injector inlet to

all the injection holes.

A large complex manifold volume allows low passage velocities and good distribution of flow

over the chamber.A small manifold volume allows for a light weight injector and reduces the amount

of dribble afterthe main walls are shut.

TYPES OF INJECTORS:

IMPIN

IMPINGING STREAM PATTERN :

The types of impinging stream pattern are ,

Doublet impinging stream pattern

Triplet impinging stream pattern

Self impinging stream pattern

These impinging stream type multiholes injectors are commonly used with oxygen hydrocarbon and

storable propellants.

In this type of injectors, the propellants are injected through a number of separate holes in

impingement forms thin liquid fans that aids the atomization of liquids into droplets.

Impinging hole injectors are also used like a cell impinging patterns.

INJECTORS

IMPINGING

STREAM TYPE

DOUBLET

IMPINGING

STREAMPATTERN

TRIPLET

IMPINGIN

GSTREAM

PATTERN

SELF IMPINGING

STREAM PATTERN

NON IMPINGING

(or) SHOWER

HEAD

SHEET (or)

SPRAY TYPE

COAXIAL HOLLOW

POST INJECTOR


10/93

The two liquid stream forms like a fan which breaks into droplets. For uneven volume flow the

triplet pattern seems to be more effective.

NON- IMPINGING (or) SHOWER HEAD TYPE:


11/93

Nonimpinging (or) shower head injector employs non-impinging stream of propellants usually

emerge in normal to the face of the injector.

It releases the fuel and oxidizer on turbulence and diffusion to achieve good mixing.

This type of injectors is not used now, because it requires a large chamber volume for goodcombustion.

SHEET (or) SPRAY TYPE INJECTORS:

Sheet (or) spray type injectors give cylindrical, conical or other types of spray sheets , these

sprays generally intersect and thereby promote mixing and atomization .

By varying the width of the sheet (through an axially movable sleeve) it is possible to throttle the

flow over a wide range without excessive reduction in the pressure drop.

This type of variable area concentric tube injector was used on the descent engine of the lunar

excursion module.

THE COAXIAL HOLLOW POST INJECTOR:

The coaxial hollow post injector has been used for liquid oxygen and gaseous hydrogen injectors.


12/93

It works well when the liquid hydrogen has absorbed heat from cooling jackets and has been

gasified.This gasified hydrogen flows at a high speed of 330m/s.

The liquid oxygen flows far slowly at a speed of 33m/s ,and the differential velocity cause a shear

action which helps to break up the oxygen stream into small droplets .

The injector has a multiplicity of these coaxial posts on its face .

The coaxial hollow post injector is not used with liquid storable bipropellants in part because the

pressure drop to achieve high velocity would become too high.

DESIGN CONSIDERATION OF A LIQUID ROCKET COMBUSTION CHAMBER:

Combustion chamber which is also known as thrust chamber, where the combustion or burning of

propellants take place. The combustion temperature is much higher than the melting points of most

chamber wall materials. Therefore it is necessary to cool these walls or to stop rocket operation

before the critical wall areas become too hot. If the heat transfer is too high and thus the wall

temperatures become locally too high, then the thrust chamber will fail.

VOLUME AND SHAPE CONSIDERATIONS:

Spherical volume gives the least internal surface area and mass per unit chamber volume. It is very

expensive to build the spherical chambers.

Today most of all prefer cylindrical or slightly tapered cone frustum with a flat injector and a

converging diverging nozzle. Neglecting the effect of the corner radii, the chamber volume is given

by,

Here L is the length of the cylinder AL/At is the chamber contraction ratio, and Lcis the length of the

conical frustum.

CHAMBER VOLUME - DEFINITION:

The chamber volume is defined as the volume up to the nozzle throat section and it includes

the cylindrical chamber and converging cone frustum of the nozzle.

The volume and shape of a combustion chamber are selected after evaluating various

parameters. Some of them are as follows,

1.

The volume has to be large enough for adequate mixing, evaporation and complete

combustion of propellants.

2. Chamber volume varies for different propellants with the time delay necessary to vaporize

and activate the propellants and with the speed of the propellant combination.

3. When the chamber volume is too small, combustion is incomplete and the performance is

poor.

4.

With higher chamber pressure or with highly reactive propellants and with injectors that

give improved mixing, a smaller chamber volume is usually permissible.


13/93

5.

The chamber volume and diameter can influence the cooling requirements. If the chamber

volume and diameter are large, the heat transfer rates to the wall will be reduced, the area

exposed to heat will be large, and the walls are somewhat thicker.

6.

All inert components should have a minimum mass. The thrust chamber mass is a function

of the chamber dimensions, chamber pressure, and nozzle area ratio, and the method of

cooling.

7. Manufacturing consideration favour simple chamber geometry, such as a cylinder with a

double cone bow tie shaped nozzle, low cost materials and simple fabrication process.

8. In some applications the length of the chamber and the nozzle relate directly to the overall

length of the vehicle.A large diameter but short chamber can allow a somewhat shorter

vehicle with a lower structural inert vehicle mass.

9. The gas pressure drop for accelerating the combustion products within the chamber should

be a minimum; any pressure reduction at the nozzle inlet reduces the exhaust velocity and

the performance of the vehicle. These losses become appreciable when the chamber volume

less than three times the throat area.10.For the same thrust the combustion volume and the nozzle throat area become smaller as

the operating chamber pressure is increased. This means that the chamber length and the

nozzle length also decrease with increasing chamber pressure, the performance will go up

with chamber pressure.

PROPELLANT HAMMER:

Propellant hammer is nothing but a pressure surging present in the liquid propellant feed line.

Basically the feed lines are very thin. On sudden closure of valve, a pressure pulse is generated at the

neighbourhood of the valve. It travels back to the tank at some velocity and keeps the liquid static

pressureincreasing.

a =

Where,

a = velocity of propagation of pressure pulse

E = Modulus of elasticity of pipeline material

K = Bulk modulus of elasticity of propellant

D =Diameter of propellant feed line

t = wall thickness of feedline

Fig: Propellant hammer in the pipe line due to sudden closure of valve


14/93

changes in the momentum of the fluid in the feed line is caused by the opening or closing of valves

in the line result in pressure peaks analogous to the propellant hammer ,such situation occurs

during the rocket engine start , during the initial bleed of the rocket engine or rocket engine set

down . This situation fall under two categories.

1. Valve opening

2. Valve closure

In case of valve closure ,i.e ,

tc=valve closure time

a = velocity of propagation of pressure pulse

2L/a tc ; for fast valve closure

2L/a


15/93

Cavitation is an undesirable phenomenon because there will be increased losses in the outlet.

Cavitation occurs in the converging duct of the outlet where the fluid velocity increases and there is

a corresponding decrease in static pressure.

Fig: cavitation phenomenon due to sudden static pressure drop

SOLUTION FOR CAVITATION:

Cavitation problem can be avoided by contouring the outlet, so that the static pressure is constant

throughout the outlet. Cavitation can also suppress by avoiding high flow velocities or by using high

fluid pressures or by combination of both. The high fluid pressures in the turbo pumps are achieved

by high tank pressures, possibly in combination with booster pumps.

2. LIQUID DROP OUT:

Liquid drop out is an undesirable phenomenon in case of liquid rocket engines. Liquid dropout is

basically a depression in the liquid surface at centre of the flow lines, which occurs in higher vertical

velocity along the centre line of the outlet than along the wall exit.

Fig:Dropout inside a liquid fuel tank


16/93

Liquid dropout will not occur when the liquid surface remains stationary. This problem can be

avoided by contouring the outlet so that the axial component of velocity along a stream lineadjacent to the wall of outlet is equal to the average velocity which is obtained by dividing the flow

rate by the cross sectional area.

3. VORTEXING:

Fig: Formation of vortex inside the fuel tank

Vortexing is a phenomenon which is similar to the coriolisforce effects in bath tubs being emptied

and can be augmented if the vehicle spins or rotates during flight.

Typically a series of internal baffles is often used to reduce the magnitude of vortexing in

propellant tanks with modest side acceleration. vortexing can greatly increase the unavailable or

residual propellant , and thus cause a reduction in vehicle performance .

OUTAGE:

The amount of liquid oxidizer or propellant present in the tank at the time of completing the

operation of vehicle is called as an outage.


17/93

GEYSERING EFFECT :

The term geysering is applied to the phenomenon which occurs in a liquid propellant

system, a column of liquid in long vertical lines is expelled by the release of bubbles.

If the bubbles will swarm causing the creation of slow moving mass or a single large bubblestravels at faster velocity causing more and more bubble formation and decrease the column static

pressure rapidly.

Fig: Bubble formation inside the fuel tank due to Geysering effect

The pressure surging produced due to geysering can be large and damage the fluid lines, wall

supports and the line supports.

Geysering can be also results from the action of the release of super heat and reduced pressure

boiling in a saturated or superheated liquid column.

PROPELLANT SLOSH:

SLOSH-DEFINITION:

Slosh refers to the movement of liquid inside an object, which is typically undergoing motion.

Fig: Sloshing of a liquid inside a glass


18/93

EXPLANATION:

Sometimes the liquid contains in the propellant tank may oscillate back and forth and this liquid

motion is generally referred as propellant slosh.Propellants slosh generally occurs in space crafttanks, rockets (especially in upper stages), then cargo slosh in ships and trucks transporting liquids

(for example oil and gasoline)

The resulting oscillatory forces and moments on the tank walls are not negligible and must be

considered in the dynamic analysis of the missiles.

When the tank is partly empty, sloshing can uncover the tank outlet and allow gas bubbles to enter

into the propellant discharge line. These bubbles can cause combustion problem in the thrust

chamber, the aspirating of bubbles or the uncovering of the tank outlets by liquids therefore needs

to be avoided. Sloshing can also shifts in vehicles centre of gravity and makes the flight control

difficult.

Fig: Sloshing of liquid inside a rectangular fuel tank

In the missiles the dynamic excitation during the powered flight is strongly offered by the sloshing

motion of the liquids in the tanks.


19/93

The associated frequencies during sloshing can be accurately determined for the design of autopilot

because they may be within the autopilot effective control frequency.

The effect of propellant slosh in the structural dynamics of the missile is generally idealized

mathematically based knowledge. The fundamental mode of propellant motion plays a very

significant role inthe study of structural dynamics.

METHOD TO AVIOD PROPELLANT SLOSH:

The propellant is replaced for analytical purposes by a mass mounted within the tank, a frictional

guide which is perpendicular to the tank axis. The motion of the equivalent mass along the guide is

restrained by a mass less spring.

There are several types of slosh suppression devices has been employed successfully to increase the

damping of liquid sloshing induced by vehicle motions. The devices include rigid ring baffles (Of

various geometries and orientation), cruciform baffles, deflectors, flexible flat ring baffle, floating

can, positive expulsion bags and diaphragms. Gel, packed fibres, and foams have been employed in

non space applications, but are not now being used for space vehicles.

Various ring baffle configuration used for suppression of sloshing in cylindrical & spherical tanks


20/93

PROPELLANT FEED SYSTEM:

Liquid propellants are required to be injected at a pressure slightly above the combustor pressure.

There are two types of feed systems can be employed for this function. They are,

1. Gas pressure feed system

2. Turbo pump feed system

The pressure feed system is much simpler and widely used for low thrust and short range

operations. The latter is used in large engines.

GAS PRESSURE FEED SYSTEM:

The gas pressure feed system is quite simple. An inert gas is separately carried at a pressure muchhigher than the injection pressure; this is used to exert the required pressure in the propellant tanks.

The pressurizing gas is chosen on the basis of its chemical properties, density, pressure and the total

weight of the gas and the tank. A gas which is ideal for one propellant unsuitable for another.

Nitrogen, Helium and air have been used for pressurization. The propellants under high pressure are

forced to flow into the thrust chamber through valves, feed lines and injectors. Several regulating

and check valves are used for filling draining, starting and checking the flow of propellants.

In this type of systems there are no moving parts such as turbines and pumps are used. Therefore

this system is considerably simpler. However, the pressurization of the propellant tanks requires

them to be comparatively much heavier and introduces a weight penalty besides other problems.

Therefore this system is unsuitable for large rocket and long range missions.


21/93

Pressure for injection can also be generated within the propellant tank by introducing a small

quantity of a gas, which reacts exothermally with the propellant, this produces high pressure gas

required to force the propellant into the combustor.

TURBO PUMP FEED SYSTEM:

In the turbo pump feed system, the propellants are pumped into the combustor by gas turbinedriven by centrifugal pumps.


22/93

The turbines derive the power from the expansion of hot gases .The gases are generated separately

by the gas generator. Figure above depicts a general arrangement of a turbo pump system. In order

to achieve flexibility in choosing the design and operating parameters the fuel and oxidizer pumps

can be separately by their turbines.

The turbine operates on a separate gas stream generated from the propellants in an independent

gas generator. A pressurizing gas can be used to increase the pressure of the propellants at the

pump suctions to avoid cavitation and the resulting instability in pump operation.

Generally turbine speeds are high , therefore propellant pumps can be driven at optimum speeds

through reduction gear with an additional weight penalty. The working gas for the turbine can also

be generated at optimum temperature and pressure. The generator also has its own injection and

ignition systems. The flow of propellants to the gas generator occurs due to the action of

pressurizing gases. If the gas pressurization is not employed to the propellants can be bled from the

delivery lines of the pumps. The propellant flow required for driving the turbines is of the order of

1.5 to 5% of the main flow. The turbine exhaust is also expanded through an exhaust nozzle to

provide an additional thrust.

An auxiliary power unit is also needed in a rocket engine. A single turbine can develop sufficient

power to drive the propellant pumps as well as the electric generator. Besides working on high


23/93

energy gases bled from the main thrust chamber or combustor it can also employ its own combustor

with a gas pressure feed system. An alternative method which is comparatively simpler is to

generate the working gases by burning solid propellants in a manner similar to the solid propellant

rocket.

The turbines and pumps for rocket applications are designed to meet some special requirements.

There are enormous temperature differences with a turbine inlet at a high temperature of the

propellants are highly reactive. Therefore the sealing arrangement in propellant pumps should be

perfect and resistant to corrosion.

Both positive displacement and turbo pumps can be used for delivering propellants from the tank to

the combustion chamber. However centrifugal pumps are widely used.

VALVES AND PIPE LINES:

VALVES:

Valves control the flows of liquids and gases and pipes conduct these fluids to the intended

components. There are no rocket engines without them. There are many different types of valves.

All have to be reliable, light weight, leak proof, and must withstand intensive vibrations and very

loud noises.

With many of these valves, any leakage or valve failure can cause a failure of the rocket unit

itself. Allvalves are tested for two qualities prior to installation; they are tested for leaks - through

the seat and alsothrough the glands--and for functional soundness or performance.

The propellant valves in high thrust units handle relatively large flows at high service pressures.Therefore, the forces necessary to actuate the valves are large. Hydraulic or pneumatic pressure,

controlled bypilot valves, operates the larger valves. These

Classification of Valves Used in Liquid Propellant Rocket Engines

1. Fluid valve:

For carrying fuel, oxidizer,cold pressurized gas, and hot turbine gas this type of valve is used.

2. Application or Use:

The valves which are mainly used for propellant control are

Thrust chamber valve (dual or single),bleed valve, drain valve, filling valves, by-pass valve,preliminary stage flow valve, pilot valve, safety valve; overboard dump valve, regulator

valve, gas generator control valve, sequence control valve.

3. Mode of Actuation:

The valves are operated by different means of actuation. The different modes are,

Automatically operated (by solenoid, pilot valve, trip mechanism, pyrotechnic, etc.)

Manually operated

Pressure-operated by air, gas, propellant, or hydraulic fluid (e.g., check valve, tank

vent valve, pressure regulator, relief valve)

4. The flow magnitude determines the size of the valve.


24/93

5. Valve Types:

Normally open, normally closed, normally partly open, two-way, three-way,

with/without valve position feedback, ball valve, gate valve, butterfly type,spring loaded valve.

PIPES (or) LINES:

The various fluids in a rocket engine are conveyed by pipes or lines, usually made of metal

and are joined byfittings or welds. Their design must provide thermal expansion and provide support

to minimize vibrationeffects. For gimballed thrust chambers it is necessary to provide flexibility in

the piping to allow the thrust axis tobe rotated through a small angle, typically +3 to 10 . This

flexibility is provided by flexible pipe joints and or byallowing pipes to deflect when using two or

more right-angle turns in the lines. Sudden closing of valves can cause propellant hammer in the

pipelines, leading to unexpected pressure rises which can be destructive to propellant system

components. The friction of the pipe and the branching ofpipelines reduce this maximum pressure.

Propellant hammer can also occur when admitting the initial flow of high-pressure

propellant intoevacuated pipes. The pipes are under vacuum to remove air and prevent the forming

of gas bubbles in the propellant flow, which can cause combustion problems.

COOLING OF THRUST CHAMBER:

NEED FOR COOLING:

The primary objective of cooling is to prevent the chamber and nozzle walls from becoming too hot,

so they will no longer able to withstand the imposed loads and stresses, thus causing the chamber or

nozzle to fail. Most materials lose strength and become weaker as temperature is increased. Coolingthus reduces the wall temperatures to an acceptable limit.

METHODS OF COOLING THETHRUST CHAMBER:

The cooling methods of a thrust chamber are briefly classified as below,


25/93

Now a days there are two most cooling methods are commonly used. They are, Active cooling

system and Passive cooling system.

ACTIVE COOLING SYSTEM:

In liquid rocket motor, the nozzle and chamber walls are exposed to hot combustionproducts. Usually these walls are provided with ducts

The four most important active cooling methods are,

1. Regenerative cooling

2. Film cooling

3. Transpiration cooling

4. Dump cooling

REGENERATIVE COOLING:

It is one of the most efficient and sophisticated means of cooling. This method is used in

many of the large rocket engines. The thrust chamber and nozzle wall contains passages through

which one of the propellants, usually the fuel flows. The passages may either formed by a simple,

double wall construction, by composing the thrust chamber and nozzle of a bundle of coolant tubes,

or by milling out the coolant ducts in the wall of the chamber and nozzle. The coolant passing at high

pressures through the ducts then it is injected into the combustion chamber. In some cases, if the

coolant is at a super critical pressure, it is possible to use the absorbed energy to drive a turbo pump

unit before the coolant is injected into the combustion chamber.

The size of the coolant ducts and coolant flow rate are determined by the following considerations:

the total amount of heat absorbed should not raise the bulk temperature to the boiling point, or to

such a level that propellant decomposition takes place, the local heat transfer rate should not

exceed the maximum nucleate boiling heat transfer rate, the pressure in the cooling jacket should

not become too low.

Coolant boiling is accomplished with the formation of large vapor bubbles and a strong decrease in

density and cooling capacity. Moreover, a blockage of the flow may occur. Propellant decomposition

may form deposits on the hot walls of the cooling jacket, thus effectively reducing the conductivity

of the wall, and hence the heat transfer rate.


26/93

Local nucleate boiling strongly increases the heat transfer rate, however if film boiling takes place,

an insulating vapor film at the wall reduces the possible heat flexures strongly. If the fluids are at

super critical pressures, neither boiling nor nucleate or film boiling will occur and high heat transfer

rates are possible.

Regenerative cooling is very effective as nearly all heat energy that has transferred to the wall is fed

back into the thrust chamber and hence is available for propulsion. This requires a complicated

construction and there is a large pressure drop along the coolant jacket, hence needed very high

pump pressure. Moreover, some propellants only allow low wall temperatures otherwise

decomposition may take place.

FILM COOLING:

Film cooling method is suited when it is used with the combination of other methods

such as regenerative cooling or insulation cooling. Pure film cooling permits a relatively simple

chamber and nozzle design. The coolant is injected along the gas side wall surface by means of

tangential slots. The coolant forms a cool boundary layer between the gas side wall surface and hot

gases. As this boundary layer gradually mixes with the main flow, its temperature rises and

downstream of the slot new coolant has to be injected.

DUMP COOLING:

Dump cooling resembles regenerative cooling, but after having performed its coolingfunction, the coolant is dumped overboard at the nozzle exit. Many o the restrictions for


27/93

egenerativecooling also hold for dump cooling. The heated, gasified coolant can be accelerated to

supersonic speeds thus providing a small extra thrust. The method is especially suited for low

pressure engines, using low molecular weight propellants, but yields a performance loss as

compared to regenerative cooling. On the other hand, the construction is simpler as compared to

regeneratively cooled engines.

PASSIVE COOLING SYSTEMS:

Among these systems, the most important ones are: insulation cooling, heat sink cooling,

ablative cooling and radiation cooling.

INSULATION COOLING:

This method is not a real method of cooling by itself; it is mostly used in combination with

other cooling techniques such as, heat sink, radiation and regenerative cooling. A very special

material is pyrolytic graphite. This material has high and low conductivity directions. While the

conductivity parallel to the layer planes is in the order of 2x103w/m.k, the conductivity

perpendicular to the layer plane is only 5.75w/m.k. this make it is possible to conduct the heat in

preferred directions, and so to avoid the heating of critical parts.

HEAT SINK COOLING:

Heat sink cooling is mostly used in solid rockets. The method consists of applying a piece of

solid material with good conductivity and a high specific heat capacity to certain hot spots. The heat

sink absorbs the heat from the hot gases, thereby raising its own temperature but keeping the wall

relatively cool. This method is only suitable for short duration applications, but is sometimes used in

combination with insulation cooling for small liquid rocket engines.

ABLATIVE COOLING:

Ablative cooling consists of covering hot gas side of the engine wall with a material that decomposes

endo thermally at high temperatures, while forming a insulating char layer. It is often used incombination with radiation and insulation cooling and chosen for upper stage motors and reaction


28/93

control engines for the sake of simplicity. It is also an effective means to keep the temperature of

variable thrust motors within an acceptable range. Regenerative cooling often poses a problem for

variable thrust motors, because of the variable chamber pressure and flow rate. Therefore, ablative

cooling offers a simple and efficient way to keep the engine wall relatively cool.

RADIATION COOLING:

Radiation cooling is often used in upper stage engines and reaction control engines in

combination with insulation and ablative cooling. The hot walls radiate the heat to the surroundings.

As the radiative heat flux is proportional to T4, the material temperature must be high to obtain a

large radiative heat flux.

Refractory metals, such as molybdenum, niobium can withstand high temperature without

losing their strength. Some refractory metals easily react with the combustion products. As the

melting point of their oxides or compounds often is much lower than that of the metals, coatings

have to be applied on many cases. The refractory alloys based on titanium, niobium andmolybdenum have found successful applications as nozzle construction materials. Wolfram

(tungsten) alloys have found applications for nozzle inserts.

COMBUSTION SYSTEM OF SOLID ROCKETS:

PHYSICAL AND CHEMICAL PROCESS:

The combustion in the solid propellant motor involves exceedingly complex reaction

taking place in the solid, liquid & gas phase of a heterogeneous mixture.

Visual observations and measurements of flames in simple experiments such as

strand burner test give an insight into the combustion processes. For double base propellants, the

combustion flame structure appears to be homogeneous and one-dimensional along the burning

direction. When the heat from the combustion melts, decomposes and vaporizes the propellant at

the burning surface, the resulting gases seems to be already premixed.

Burn rate catalysts seem to affect the primary combustion zone rather than the

processes in the condensed phase. They catalyze the reaction at or near the surface, increase or

decrease the heat input to the surface, the change the amount of propellant that is burned.

Solid Fuel Geometry


29/93

Solid Fuel Geometry Dependent Thrust-Time Curve

(Typical solid propellant grain configurations and the corresponding thrust-time curves)

BASIC CONCEPTS:

A simple solid rocket motor consists of a casing, nozzle, grain (propellant charge), and

igniter.

The grain behaves like a solid mass, burning in a predictable fashion and producing

exhaust gases. The nozzle dimensions are calculated to maintain a design chamber pressure, while

producing thrust from the exhaust gases.

IGNITION PROCESS:

Solid propellant ignition consist of a series of complex rapid events, which

starts on receipt of a signal and include heat generation, transfer of the heat from the igniter to the

motor grain surface, spreading the flame over the entire burning surface area, filling the chamber

free volume with gas and elevating the chamber pressure without series abnormalities such as over

pressure, combustion oscillation, damaging shock waves, hang fire extinguishment and chuffing.


30/93

Satisfactory attainment of equilibrium chamber pressure with full gas flow depends on:

Characteristics of the igniter and the gas temperature.

Motor propellant composition and surface ignitability.

Heat transfer by radiation and convection between gas and grain surface.

Grain flame spreading rate.

The dynamics of filling the motor free volume with the hot gas.

Ignitibility of a propellant is affected by,

The propellant formulation.

The initial temperature of the propellant.

The surrounding pressure.

The mode of heat transfer.

Grain surface toughness.

Age of the propellant.

The velocity of the hot igniter gas.

The cavity volume and configuration

The variables determining grain-relative performance are core surface area and specific

impulse.

Surface area is the amount of propellant exposed to interior combustion flames, existing

in a direct relationship with thrust.

An increase in surface area will increase thrust but will reduce burn-time since the

propellant is being consumed at an accelerated rate.

The optimal thrust is typically a constant one, which can be achieved by maintaining a

constant surface area throughout the burn.

Examples of constant surface area grain designs include: end burning, internal-core and

outer-core burning, and internal star core burning.

Solid propellants are either "composites" composed mostly of large, distinct

macroscopic particles or which are a homogeneous mixture of one or more primary

ingredients


31/93

Often, the ingredients of a double base propellant have multiple roles such as RDX which

is both a fuel and oxidizer or nitrocellulose which is a fuel, oxidizer and plasticizer.

Grain:

Solid fuel grains are usually molded from a thermoset elastomer (which doubles asfuel), additional fuel, oxidizer, and catalyst. HTPB is commonly used for this purpose.

Ammonium perchlorate is the most common oxidizer used today.

The fuel is cast in different forms for different purposes. Slow, long burning rockets

have a cylinder shaped grain, burning from one end to the other. Most grains, however, are cast

with a hollow cross section, burning from the inside out (and outside in, if not case bonded), as well

as from the ends.

The thrust profile over time can be controlled by grain geometry. For example, a star

shaped hole down the center of the grain will have greater initial thrust because of the additionalsurface area. As the star points are burned up, the surface area and thrust are reduced.

Casing:

The casing may be constructed from a range of materials. Cardboard is used for model

engines. Steel is used for the space shuttle boosters. Filament wound graphite epoxy casings are

used for high performancemotors.

Nozzle:

A Convergent Divergent design accelerates the exhaust gas out of the nozzle to producethrust. Sophisticated solid rocket motors use steerable nozzles for rocket control.

COMBUSTION MECHANISM OF SOLID PROPELLANTS:

Some solid rocket propellants are mixed at the molecular level. A double base

propellant made from nitrocellulose and nitro-glycerin. The dominant difference is the break in

temperature slope at the solid gas interface. The solid usually requires some heat input to gasify and

this heat is the heat of pyrolysis.

Consequently, the gas phase heat transfer at the interface goes towards providing

both the latent heat and continued heat transfer into the solid.

Advantages:

Solid propellant rockets are much easier to store and handle than liquid propellant

rockets.

High propellant density makes for compact size as well.


32/93

These features plus simplicity and low cost make solid propellant rockets ideal for

military applications.

These features plus simplicity and low cost make solid propellant rockets ideal for

military applications whenever large amounts of thrust are needed and cost is an issue.

Disadvantages:

Relative to liquid fuel rockets, solid fuel rockets have lower specific impulse.

The propellant mass ratios of solid propellant upper stages is usually in the .91 to

.93 range which is as good or better than that of most liquid propellant upper stages

but overall performance is less than for liquid stages because of the solids' lower

exhaust velocities.

Solid rockets cannot be throttled in real time.

Solid fuel rockets are intolerant to cracks and voids.

COMBUSTION INSTABILITY

Combustion instability occurs when normal velocity (Vn) is not equal to the combustion velocity or

flame velocity(Vf).

There are 2 types of combustion instability:

1) Set of acoustic resonance, which can occur with any rocket motor.

2) Vortex shedding phenomenon, which only with particular type of propellant grains.

These two types of problems, mainly occurs only when the rocket combustion is not

controlled. It causes excessive pressure vibration forces or excessive heat transfer.

The combustion in liquid rocket is never perfectly smooth, there are some

fluctuations of pressure, temperature, and velocities are present.

ROUGH COMBUSTION:

Rough combustion is defined as the Combustion that gives greater pressure fluctuation at a

chamber wall location which occurs at completely random intervals is called rough combustion.

POGO OSCILLATION:

Periodic variations of thrust, caused by combustion instability or longitudinal vibrations of

structures between the tanks and the engines which modulate the propellant flow, are known as

"pogo oscillations" or "pogo", named after the Pogo stick.

Three different types of combustion instabilities occur. Some of them are,

CHUGGING:

Chugging, the first type of combustion instability occurs mostly from the elastic nature of the feed

systems and due to low frequency in the feed system which ranges from 100-400 HZ. This can cause


33/93

cyclic variation in thrust, and the effects can vary from merely annoying to actually damaging the

payload or vehicle. Chugging can be minimized by using gas-filled damping tubes on feed lines of

high density propellants.

BUZZING:

This is the intermediate type of instability and its frequency ranges from 400-1000HZ. This can be

caused due to insufficient pressure drop across the injectors. It generally is mostly annoying, rather

than being damaging. However, in extreme cases combustion can end up being forced backwards

through the injectors. This can cause explosions with monopropellants.

SCREECHING (OR) SCREAMING (OR) SQUEALING:

This is the third type of instability which has higher frequency of range 1000HZ and above. It is

mostly perplexing which occurs both liquid and solid propellant rockets. This type is most destructing

and has capability of destroying the engine much less than 1 sec.

POPPING:

Popping is an undesirable random high amplitude pressure disturbance that occurs during steady

state operation of a rocket engine with hypergolic propellant. Its one of the pressure source

triggering high frequency, instability in a rocket engine.

ELIMINATION OF POPPING:

The elimination of popping is usually achieved by re-design of the injector rather than the

application of baffles and absorbers.


34/93

ANSWER THE FOLLOWING IN SHORT

1. What is an igniter?

2. Name the types of igniters

3. What are the types of liquid igniter?

4. State any 5 characteristics of hypergolic liquids

5. State any 4 combinations of hypergolic ignition6. Name the factors which affect the ignition delay and also give its equation with its terms

7. Classify solid rockets

8. Give the advantage and disadvantages of a bag igniter

9. Give the disadvantage of a basket igniters

10. What are the important components of a pyrotechnic igniter?

11. Give any 4 characteristics of squib

12. What are the factors to be considered while designing an igniter?

13. Distinguish between pressure feed system and pump feed system

14. What is geysering effect?

15. What is meant by outage?

16. Define combustion instability17. Define cavitation and how cavitation will be avoided?

18. What are the problems to be avoided while designing a fuel tank outlet?

19. What is meant by liquid drop out?

20. What is an injector?

21. What are the types of injectors?

22. Define atomization of fuel

23. What is meant chugging?

24. Define the term buzzing

25. Define the term screeching

26. Define the term popping

27. What is an ignitability bomb?28. Name the components of live igniters

29. What are the methods used for cooling of thrust chambers?

30. Define the term rough combustion

31. Draw a neat sketch of solid rocket combustion chamber

32. Name the types of valves which are used in rockets.

33. What are the modes of actuations used for operating a valve in a rocket?

34. Give the applications of valves in a rocket.

DETAILED ANSWERS:

1.

Classify the different types of igniters with neat sketches.2.

What is an injector? What is the main difference between an injector and a carburettor?

Classify its various types with neat sketches.

3. Distinguish between Helium pressure feed system and centrifugal pump feed system.

Pictorially represent Helium pressure feed system and explain in detail.

4.

Elaborately explain the propellant pump feed system with an appropriate sketch.

5.

What are the important design considerations in the section of liquid rocket combustion

chamber volume and shape? List and explain them briefly.

6.

What is the need for cooling a thrust chamber? What are the different methods of cooling of

thrust chambers? Explain them briefly and draw appropriate sketches wherever necessary.

7. What is propellant slosh? Discuss its effects on flight vehicles and explain how it is

controlled?


35/93

8.

What are problems generally faced by a designer while designing the liquid propellant tank

outlet design? Explain them briefly

9. Explain the phenomenon, propellant hammer in a liquid propellant rocket engine with an

appropriate sketch.

10.What is geysering effect? When and where does it occur? Explain your answer with a neat

sketch.11.

Define combustion instability and explain briefly about the various types of combustion

instability.

12.Elucidate the combustion mechanism of a solid propellant rocket.

C


36/93

UNIT-2

AERODYNAMICS OF ROCKETS AND MISSILES

AIRFRAME COMPONENTS OF A MISSILE:

The components or parts which are experienced bythe course of air are known as airframe components

.The body of the missile can be divided into three major sect ions .They are

Nose or Fore body

Midsection or Main body

The aft or Boat tail section

Fins

MAJOR COMPONENTS OF A MISSILE

NOSE (or) FORE BODY:

It is the first and foremost component of a missile which experiences air while travelling through the

atmosphere. Several types of nose sections were used in various types of missiles. Some of the types are,

Conical fore body

Ogival fore body

Hemispherical fore body

CONICAL FORE BODY:

These types of fore body are used in missiles, which are intended to fly at supersonic speeds. The missile,

while travelling in the atmosphere oblique shock is formed at the tip of the wedge and apex of the cone. There

are various aerodynamic and thermodynamic changes are noticeable in the flow characteristics of air, in the

case of conical nose.


37/93

CONICAL NOSE OF A SUPERSONIC MISSIE

OGIVAL FORE BODY:

Ogival nose configuration is used more frequently than the conical nose. An ogive is similar to a cone

except that the plan form shaped is formed by an arc of a circle instead of a straight line as shown in figure. The

ogival shape has several advantages over the conical section.

ADVANTAGES:

1. Slightly greater volume for a given base and length(L/D ratio)

2. A blunter nose provides structural superiority.

3.

Slightly lower drag.

HEMISPHERICAL FORE BODY:

This type of nose is used on some of the missiles, particularly those which use IR (infrared) seekers.

This type of nose imposes an extremely high drag penalty on the missile. The use of this type of nose on missiles

indicates the extent to which an aerodynamicist must compromise to achieve an optimum and feasible missile

system.

MID SECTION:

In most missile configurations, the mid section is in cylindrical shape. The shape is advantageous from the

stand points of drag, ease of manufacturing, and the load carrying capability. The zero-lift drag of a cylindrical

body is caused by skin friction force only. At low angle of attack, a very small amount of normal force is

developed on the body, this results from the carryover fromthe nose section.

BOAT TAIL:

The tapered portion of the aft section of a body is called the boat tail. The purpose of boat tail is to decrease

the drag of a body which has a squared off base. The mid section has relatively large base pressure and

consequently high drag values because of large base area. By boat tailing the rear portion of the body, the base

area is reduced and thus the base drag is reduced. However, the decrease in base drag may be partially nullified

by the boat tail drag.


38/93

FINS:

The purpose of putting fins on the rocket is to provide stability, provide lift and control the flight path of the

missile. The plan form of fins of a rocket is of different types. They are of clipped tip delta, rectangular,

triangular, trapezoidal etc.

AERODYNAMIC SURFACES OF MISSILES:

1. SUPERSONIC WING CROSS SECTIONAL SHAPES :

The various supersonic wings cross sectional shapes are,

1. Double wedge

2. Modified double wedge and

3. Biconvex

1. Double wedge :

The double wedge offers a least drag but lacks strength.

2. Modified double wedge:

The modified double wedge has relatively low drag and comparatively stronger than the latter one.

3.

Biconvex:

The biconvex causes considerable drag but it is the strongest of the three designs. The biconvex shape

has a slight advantage in minimum drag for unit cross sectional strength in addition to the absence of

sharp corner. The sharp corners affect the flow conditions over the surface. The biconvex section also

provides larger wedge angles at the leading and trailing edges.

SUPERSONIC WING PLAN FORMS:

(a) CLIPPED TIP DELTA (b) DELTA (or) TRIANGULAR (c) RECTANGULAR


39/93

(d) RECTAGULAR WITH RAKE

The main difference between the subsonic and supersonic types of wing plan forms is the symmetry about the

chord and sharpness of the leading edge. For the supersonic case, the need for sharp leading edge is to encounter

the type of flow and pressure distribution while travelling faster than speed of sound.

AERODYNAMIC CONTROLS OF A MISSILE:

Aerodynamic control is the connecting link between the guidance system and the flight path of the missile.

Effective control of flight path requires smooth and exact operation of the control surfaces of the missile. They

must have the best possible design configuration for the intended speed of the missile. The control surface must

move with enough force to produce the necessary change of direction. The adjustments they make must

maintain the balance and centre of gravity of the missile. The control surface must also be positioned to meet

variations in lift and drag at different flight speeds. All these conditions contribute to the flight stability of the

missile.

ARRAGEMENTS OF CONTROL SURFACES IN A MISSILE

(a) CONVENTIONAL (b) H TYPE (or) DOUBLE RUDDER (c) V-TAIL

The types of aerodynamic controls of a missile are,

1. Canard control

2. Wing control

3. Tail control

4. Unconventional control

1. CANARD CONTROL:

Canard control is also quite commonly used, especially on short-range air-to-air missiles. The primary

advantage of canard control is better maneuverability at low angles of attack, but canards tend to become

ineffective at high angles of attack because of flow separation that causes the surfaces to stall. Since canards are

ahead of the centre of gravity, they cause a destabilizing effect and require large fixed tails to keep the missilestable. These two sets of fins usually provide sufficient lift to make wings unnecessary.


40/93

1. a.SPLIT CANARD:

A further subset of canard control missiles is the split canard. Split canards are a relatively new

development that has found application on the latest generation of short-range air-to-air missiles. The

term split canard refers to the fact that the missile has two sets of canards in close proximity, usually one

immediately behind the other. The first canard is fixed while the second set is movable.

The advantage of this arrangement is that the first set of canards generates strong, energetic

vortices that increase the speed of the airflow over the second set of canards making them more effective.

In addition, the vortices delay flow separation and allow the canards to reach higher angles of attack before

stalling. This high angle of attack performance gives the missile much greater maneuverability compared to

a missile with single canard control.

AERODYNAMIC CHARACTERISTICS OF CANARD CONTROL:

1. The canard control missile has the advantage of small control surfaces for longitudinal control and it places

the portion of the control equipment well forward in the body out of the way of the main propulsion and

guidance unit.

2. This type tends to give low drag as much as the main lifting surfaces fixed and it can be made of large

sweep back type where in the lift to drag ratio can be optimized.

3. The canard control surfaces are deflected in the positive manner that is the leading edge upward to

provide a positive angle of attack of the missile and this is in turn places the control surfaces at quite large

angle of attack relative to the free stream especially when the missiles pitched to large angles.

4. This change tends to increase loads and hinge moments on the control surfaces. High control surface rates

and hence high power will be required, to increase the angle of attack to acquire the required maneuver.

WING CONTROL:

Wing control was one of the earliest forms of missile control developed, but it is becoming less

commonly used on today's designs. Most missiles using wing control are longer-range missiles. The primary

advantage of wing control is that the deflections of the wings produce a very fast response with little motion of

the body. This feature results in small seeker tracking error and allows the missile to remain locked on target

even during large maneuvers.

The major disadvantage is that the wings must usually be quite large in order to generate both

sufficient lift and control effectiveness, which makes the missiles rather large overall. In addition, the wings

generate strong vortices that may adversely interact with the tails causing the missile to roll. This behaviour is

known as induced roll, and if the effect is strong enough, the control system may not be able to compensate.


41/93

TAIL CONTROL:

Tail control is probably the most commonly used form of missile control, particularly for longer range air-to-air

missiles and surface-to-air missiles. The primary reason for this application is because tail control provides

excellent maneuverability at the high angles of attack often needed to intercept a highly maneuverable aircraft.

Missiles using tail control are also often fitted with a non-movable wing to provide additional lift and improve

range. Some good examples of such missiles are air-to-ground weapons like Maverick and AS.30 as well as

surface-to-surface missiles like Harpoon and Exocet. Tail control missiles rarely have canards.

UNCONVENTIONAL CONTROL:

The surface of a missile that create a jet exhaust perpendicular to the vehicle surface and produce

an effect similar to thrust Unconventional control systems is a broad category that includes a number of

advanced technologies. Most techniques involve some kind ofthrust vectoring.Thrust vectoring is defined as a

method of deflecting the missile exhaust to generate a component of thrust in a vertical and/or horizontal

direction. This additional force points the nose in a new direction causing the missile to turn. Another technique

that is just starting to be introduced is called reaction jets. Reaction jets are usually small ports in vectoring.

These techniques are most often applied to high off-boresight air-to-air missiles to provide exceptional

maneuverability. The greatest advantage of such controls is that they can function at very low speeds or in avacuum where there is little or no airflow to act on conventional fins. The primary drawback, however, is that

they will not function once the fuel supply is exhausted.

Note that most missiles equipped with unconventional controls do not rely on these controls alone

for maneuverability, but only as a supplement to aerodynamic surfaces like canards and tail fins.

Classification of Missile

Missiles are generally classified on the basis of their Type, Launch Mode, Range, Propulsion,

Warhead and Guidance Systems.

Type:

1. Cruise Missile

2. Ballistic Missile

Launch Mode:

1. Surface-to-Surface Missile

2. Surface-to-Air Missile

3. Surface (Coast)-to-Sea Missile

4. Air-to-Air Missile

5. Air-to-Surface Missile

6. Sea-to-Sea Missile

7.

Sea-to-Surface (Coast) Missile

8. Anti-Tank Missile

Range:

1. Short Range Missile

2. Medium Range Missile

3. Intermediate Range Ballistic Missile

4. Intercontinental Ballistic Missile
http://www.aerospaceweb.org/question/propulsion/q0095.shtmlhttp://www.aerospaceweb.org/question/propulsion/q0095.shtml


42/93

Propulsion:

1. Solid Propulsion

2. Liquid Propulsion

3. Hybrid Propulsion

4. Ramjet

5. Scramjet

6.

Cryogenic

Warhead:

1. Conventional

2. Strategic

Guidance Systems:

1. Wire Guidance

2. Command Guidance

3. Terrain Comparison Guidance

4.

Terrestrial Guidance5. Inertial Guidance

6. Beam Rider Guidance

7. Laser Guidance

8. RF and GPS Reference

On the basis of Type:

(i) Cruise Missile:A cruise missile is an unmanned self-propelled (till the time of impact) guided vehicle that

sustains flight through aerodynamic lift for most of its flight path and whose primary mission is to place an

ordnance or special payload on a target. They fly within the earths atmosphere and use jet engine

technology. These vehicles vary greatly in their speed and ability to penetrate defences.Cruise missiles can

be categorised by size, speed (subsonic or supersonic), range and whether launched from land, air, surface

ship or submarine.

Depending upon the speed such missiles are classified as:

1) Subsonic cruise missile

2) Supersonic cruise missile

3) Hypersonic cruise missile

Subsonic cruise missile

Subsonic cruise missile flies at a speed lesser than that of sound. It travels at a speed of around

0.8 Mach. The well-known subsonic missile is the American Tomahawk cruise missile. Some other examples

are Harpoon of USA and Exocet of France.

Supersonic cruise missile

This missile travels at a speed of around 2-3 Mach i.e.; it travels a kilometre approximately in a

second. The modular design of the missile and its capability of being launched at different orientations

enable it to be integrated with a wide spectrum of platforms like warships, submarines, different types ofaircraft, mobile autonomous launchers and silos. The combination of supersonic speed and warhead mass

provides high kinetic energy ensuring tremendous lethal effect. BRAHMOSis the only known versatile

supersonic cruise missile system which is in service.

Hypersonic cruise missile

This missile travels at a speed of more than 5 Mach. Many countries are working to develop

hypersonic cruise missiles. BrahMos Aerospace is also in the process of developing a hypersonic cruise

missile, BRAHMOS-II, which would fly at a speed greater than 5 Mach.


43/93

(ii) Ballistic Missile:

A ballistic missile is a missile that has a ballistic trajectory over most of its flight path, regardless of

whether or not it is a weapon-delivery vehicle. Ballistic missiles are categorised according to their range,

maximum distance measured along the surface of earth's ellipsoid from the point of launch to the point of

impact of the last element of their payload. These missiles carry a huge payload. The carriage of a deadly

warhead is justified by the distance the missile travels. Ballistic missiles can be launched from ships andland based facilities. For example, Prithvi I, Prithvi II, Agni I, Agni II and Dhanush ballistic missiles are

currently operational in the Indian defence forces.

On the basis of Launch Mode:

(i) Surface-to-Surface Missile: A surface-to-surface missile is a guided projectile launched from a hand-held,

vehicle mounted, trailer mounted or fixed installation. It is often powered by a rocket motor or sometimes

fired by an explosive charge since the launch platform is stationary.

(ii) Surface-to-Air Missile:A surface-to-air missile is designed for launch from the ground to destroy aerialtargets like aircrafts, helicopters and even ballistic missiles. These missiles are generally called air defence

systems as they defend any aerial attacks by the enemy.

(iii) Surface (Coast)-to-Sea Missile:A surface (coast)-to-sea missile is designed to be launched from land to

ship in the sea as targets.

(iv) Air-to-Air Missile:An air-to-air missile is launched from an aircraft to destroy the enemy aircraft. The

missile flies at a speed of 4 Mach.

(v) Air-to-Surface Missile:An air-to-surface missile is designed for launch from military aircraft and strikes

ground targets on land, at sea or both. The missiles are basically guided via laser guidance, infrared

guidance and optical guidance or via GPS signals. The type of guidance depends on the type of target.

(vi) Sea-to-Sea Missile:A sea-to-sea missile is designed for launch from one ship to another ship.

(vii) Sea-to-Surface (Coast) Missile:A sea-to-surface missile is designed for launch from ship to land based

targets.

(viii) Anti-Tank Missile:An anti-tank missile is a guided missile primarily designed to hit and destroy

heavily-armoured tanks and other armoured fighting vehicles. Anti-tank missiles could be launched from

aircraft, helicopters, tanks and also from shoulder mounted launcher.

On the basis of Range:

This type of classification is based on maximum range achieved by the missiles. The basic classification is as

follows:

(i) Short Range Missile

(ii) Medium Range Missile

(iii) Intermediate Range Ballistic Missile

(iv) Intercontinental Ballistic Missile

On the basis of Propulsion:

(i) Solid Propulsion:Solid fuel is used in solid propulsion. Generally, the fuel is aluminium powder. Solid

propulsion has the advantage of being easily stored and can be handled in fuelled condition. It can reach

very high speeds quickly. Its simplicity also makes it a good choice whenever large amount of thrust is

needed.

(ii) Liquid Propulsion:The liquid propulsion technology uses liquid as fuel. The fuels are hydrocarbons. The

storage of missile with liquid fuel is difficult and complex. In addition, preparation of missile takes


44/93

considerable time. In liquid propulsion, propulsion can be controlled easily by restricting the fuel flow by

using valves and it can also be controlled even under emergency conditions. Basically, liquid fuel gives high

specific impulse as compared to solid fuel.

(ii) Hybrid Propulsion:There are two stages in hybrid propulsion - solid propulsion and liquid propulsion.

This kind of propulsion compensates the disadvantages of both propulsion systems and has the combined

advantages of the two propulsion systems.

(iii) Ramjet:A ramjet engine does not have any turbines unlike turbojet engines. It achieves compression of

intake air just by the forward speed of the air vehicle. The fuel is injected and ignited. The expansion of hot

gases after fuel injection and combustion accelerates the exhaust air to a velocity higher than that at the

inlet and creates positive push. However, the air entering the engine should be at supersonic speeds. So,

the aerial vehicle must be moving in supersonic speeds. Ramjet engines cannot propel an aerial vehicle

from zero to supersonic speeds.

(iv) Scramjet:Scramjet is an acronym for Supersonic Combustion Ramjet. The difference between scramjet

and ramjet is that the combustion takes place at supersonic air velocities through the engine. It is

mechanically simple, but vastly more complex aerodynamically than a jet engine. Hydrogen is normally the

fuel used.

(v) Cryogenic:Cryogenic propellants are liquefied gases stored at very low temperatures, most frequently

liquid hydrogen as the fuel and liquid oxygen as the oxidizer. Cryogenic propellants require special insulated

containers and vents which allow gas to escape from the evaporating liquids. The liquid fuel and oxidizer

are pumped from the storage tanks to an expansion chamber and injected into the combustion chamber

where they are mixed and ignited by a flame or spark. The fuel expands as it burns and the hot exhaust

gases are directed out of the nozzle to provide thrust.

On the basis of Warhead:

(i) Conventional Warhead:A conventional warhead contains high energy explosives. It is filled with a chemi

al explosive and relies on the detonation of the explosive and the resulting metal casing fragmentation as

kill mechanisms.

(ii) Strategic Warhead:In a strategic warhead, radio active materials are present and when triggered they

exhibit huge radio activity that can wipe out even cities. They are generally designed for mass annihilation.

On the basis of Guidance Systems:

(i) Wire Guidance: This system is broadly similar to radio command, but is less susceptible to electronic

counter measures. The command signals are passed along a wire (or wires) dispensed from the missile after

launch.

(ii) Command Guidance:Command guidance involves tracking the projectile from the launch site or

platform and transmitting commands by radio, radar, or laser impulses or along thin wires or optical fibres.

Tracking might be accomplished by radar or optical instruments from the launch site or by radar or

television imagery relayed from the missile.

(iii) Terrain Comparison Guidance:Terrain Comparison (TERCOM) is used invariably by cruise missiles. The

system uses sensitive altimeters to measure the profile of the ground directly below and checks the result

against stored information.

(iv) Terrestrial Guidance:This system constantly measures star angles and compares them with the pre-

programmed angles expected on the missiles intended trajectory. The guidance system directs the control

system whenever an alteration to trajectory is required.

(v) Inertial Guidance: This system is totally contained within the missile and is programmed prior to launch.

Three accelerometers, mounted on a platform space-stabilised by gyros, measure accelerations along three

mutually perpendicular axes; these accelerations are then integrated twice, the first integration giving

velocity and the second giving position. The system then directs the control system to preserve the pre-

programmed trajectory. These systems are used in the surface-to-surface missiles and in cruise missiles.


45/93

(vi) Beam Rider Guidance:The beam rider concept relies on an external ground or ship-based radar station

that transmits a beam of radar energy towards the target. The surface radar tracks the target and also

transmits a guidance beam that adjusts its angle as the target moves across the sky.

(vii) Laser Guidance:In laser guidance, a laser beam is focused on the target and the laser beam reflects off

the target and gets scattered. The missile has a laser seeker that can detect even miniscule amount of

radiation. The seeker provides the direction of the laser scatters to the guidance system. The missile islaunched towards the target, the seeker looks out for the laser reflections and the guidance system steers

the missile towards the source of laser reflections that is ultimately the target.

(viii) RF and GPS Reference:RF (Radio Frequency) and GPS (Global Positioning System) are examples of

technologies that are used in missile guidance systems. A missile uses GPS signal to determine the location

of the target. Over the course of its flight, the weapon uses this information to send commands to control

surfaces and adjusts its trajectory. In a RF reference, the missile uses RF waves to locate the target.

FORCES ACTING ON A MISSIE WHILE PASSING THROUGH ATMOSPHERE:

DERIVATION:


46/93


47/93


48/93

For finned rockets, the coefficient cdis usually smaller than dcl/dthat is only 2-4% of dcl/dand therefore can

be neglected.

LATERAL AERODYNAMIC DAMPING MOMENT OF THE ROCKET:

The lateral angular velocity gives rise to an additional aerodynamic moment which is proportional to the

angular velocity Mand so directed that it tends to reduce the angular velocity and this moment is known as

lateral aerodynamic damping moment.


49/93

LONGITUDINAL AERODYNAMIC MOMENT:

In case of symmetric finned rockets the component Mxof the aerodynamic moment along a

longitudinal axis is zero. However it is not for the rockets with slanted fins. In such a rocket each fin is mounted to

make certain angle with longitudinal axis in such a way that on rotation of missile through 360/n degrees (n-

number of fins). Each fin assumes the position occupied by the adjacent fin prior rotation when fins are slantmounted, a moment arises during the flight, which tends to rotate the rocket about its axis of symmetry.

For simplicity study the motion of such rocket at an angle of attack, =0.

Since each fin encounters the airflow at an angle, the fins are acted upon by a lift force L, perpendicular to the

rocket axis. The centre of pressure of this force i.e the point of intersection of its line of action with the plane of

the fin is at a distance of rc.

If there are n number of fins the total longitudinal moment is n.L 1.rc. since the lift force is proportional to the air

density and the square of the velocity and further to the fin angle .

LONGITUDINAL AERODYNAMIC DAMPING MOMENT:


50/93

LONGITUDINAL AERODYNAMIC DAMPING MOMENT

DRAG ESTIMATION:

The drag of the rocket vehicle can be split into following components,

WAVE DRAG:

Wave is mainly due to the presence of shock waves and dependent on the Mach number. The wave

drag is connected with the shock wave, and hence occurring only at supersonic speeds.

1. The amount of wave drag for the conical body is estimated as,

Where, is the half cone angle in radians.

2. The wave drag of an isolated, rectangular wing span, b, with a double wedge profile is estimated as,

Where, is the half wedge angle. Both wave drag coefficients are strongly dependent on anddecrease with increasing Mach numbers.

VISCOUS DRAG:

The viscous drag is formed due to friction. It is the main drag component at subsonic speeds. It can be

estimated by considering the friction drag coefficient CDf, for a flat plate of equal length and equal

wetted area as a rocket vehicle. For a laminar boundary layer we may estimate,

For a turbulent boundary layer we may estimate as,

These coefficients are based on the wetted area as a reference area. For most large rockets, one may

assume the boundary layer to be turbulent. Transition from laminar to turbulent takes place around

Re=106based on body length, so that for small vehicles, still a major portion of the boundary layer may


51/93

be laminar. Surface roughness may cause a transition from laminar to turbulent at lower Reynolds

number.

INDUCED DRAG:

Induced drag is a result of the development of lift.

In a subsonic case, the induced drag, based on the projected wing area, S w, is

At supersonic speed, it can be well approximated by,

BASE DRAG:

Base drag is strongly affected by the shape of the vehicle, and the presence of a jet.

The total drag is found by the addition of all components. It turns out that the total drag coefficient canbe well approximated for preliminary calculations by,

INTERFERENCE DRAG:

Interference drag is due to the interaction of various flow fields.

ROUGHNESS DRAG:It is mainly due to the surface roughness such as rivets and welds.

ROCKET DISPERSION:

DISPERSION-DEFINITION

Dispersion is defined as the measure of deviation of the rockets trajectory from the standard

nominal trajectory.

For a rocket, dispersion arises from three different sources. They are,

1. Events that occur at launching,

2. Events during burning after launching, and

3. Events after burning.

For rockets, most of the dispersion arises during the burning period after launching.

FACTORS CAUSING DISPERSION:

The factors that induce dispersion of rockets trajectory are,

The propellant mass and composition Inaccuracy

The rocket total mass, axial and lateral Moments of inertia and resultant centre of gravity Inaccuracies

Launcher deflection The thrust force of the rocket engine: because of the tolerance in rocket engine design, propellant

properties, and manufacturing

Thrust and fin misalignments: It is an important source of dispersion in case of unguided rockets.

Atmospheric disturbances such as wind profile, tail wind, cross wind, and gusts, variation in atmospheric

density.

METHODS TO ESTIMATE DISPRSION:

There are some methods to estimate the dispersion of trajectory for a rocket. They are,


52/93

Root Mean Square Method

Monte Carlo method

Method of covariance matrix

1.

The Root Mean Square Method:

The Root Mean Square Method simulates the rocket trajectory perturbing one parameter at

time and the results are compared with the nominal results. The sum of squares deviations for all parameters is

square of total deviation.

2. Monte Carlo method:

Monte Carlo method of dispersion removes smaller dispersion parameters. Each input

parameter is selected randomly in the defined ranges and used in the simulation of trajectory.

3. The Method of covariance matrix:

In probability theory and statistics, covariance is a measure of how much two variables change

together. A covariance matrix is a matrix whose element in the i,j position is the covariance between the i th and

j th elements of a random vector (that is, of a vector of random variables). Each element of the vector is a scalar

random variable, either with a finite number of observed empirical values or with a finite or infinite number of

potential values specified by a theoretical joint probability distribution of all the random variables.

TYPES OF ROCKET DISPERSION:

There are two types of rocket dispersion, such as

I.

In plane dispersion (or) Range dispersionII. Lateral dispersion (or) Out of plane dispersion

IN-PLANE DISPERSION (or) RANGE DISPERSION:

In the absence of perturbing forces giving rise to rocket dispersion, the trajectory of the rocket would lie in

the launch plane. But practically such factors are generally active and try to produce that cause. The dispersion of

the rocket, which may be of any type sometimes the rocket, can suffer both types.

LATERAL DISPERSION (or) OUT OF PLANE DISPERSION:

If the perturbing forces are active, the axis of the rocket will deviate from the target to the trajectory of themass centre by an angle known as angle of attack. Since the thrust is directed along the axis of the rocket, the

deviation gives rise to a thrust component normal to the trajectory. The trajectory thus departs from the

intended path and put the rocket away from the target.

MINIMIZATION OF ROCKET DISPERSION:

For trajectory vehicles, dispersion can be minimized by means of a special guidance system. The guided

system is thought to be known as the brain of a rocket or a missile.

The distinct tasks of a guidance system are as follows:

1. It maintains the missile in proper attitude. Using instruments like gyros, the control system correct the

problems experienced through rotation and translation motion of a rocket.

2. The control guidance system also helps in tracking the positions, computing the tracking information,

correcting the signals and then steering the rocket in a correct orbit and thus helps in minimizing the

dispersion.


53/93

Short answers:

1. What are the airframe components of a missile?

2. Name the different types of nose cones in a missile

3. What are the advantages of ogival fore body?

4. What are the aerodynamic controls present in a missile?

5. What do you mean by canard control of guided missiles?

6. What is the important of boat tail in missiles?

7. Sketch the different shapes of supersonic wing cross section

8. Sketch the various supersonic wing plan forms

9. How do you classify the missiles based on navigation and mission?

10.Mention the aerodynamic characteristics of air to surface missile

11.What is meant by lateral aerodynamic damping moment?

12.Distinguish between body up wash and body downwash in missile aerodynamics

13.What are the forces which act on a trajectory vehicle while passing through the atmosphere?

14.What are the different types of drag which acts on a missile in atmosphere?

15.W hat is meant rocket by rocket dispersion?

16.What are the types of rocket dispersion?

17.What are the factors which causes rocket dispersion?

18.How do you minimize rocket dispersion?

19.What are the methods used to estimate rocket dispersion?

BRIEF ANSWERS:

1. Explain the various airframe components and various aerodynamic controls of a missile. Draw sketches

wherever necessary.

2. Explain the forces acting on a missile while passing through the atmosphere. Elucidate your answer with

a neat sketch.

3. What are the different types of drag which acts on a missile in atmosphere? Clearly explain what wave

drag is. What is its relative importance in the total drag estimation of a supersonic missile? How is wave

drag coefficient estimated for double wedge, modified double wedge and biconvex profiles of

supersonic airfoils?

4. With the help of a neat sketch clearly explain how fins impart stability to a rocket in flight which is in

atmosphere.

5. What are the various wings cross sectional shapes that are generally used for supersonic missiles?

Sketch such shapes and mention their advantages and limitations.

6. With a neat sketch clearly explain the lateral aerodynamic moment of a rocket and briefly elucidate the

variation of lateral aerodynamic moment coefficient variation with angle of attack. How does thus

variation affects the stability of the rocket flight?

7. With a neat sketch clearly explain the longitudinal aerodynamic moment of a rocket.

8. List any four basic aerodynamic design considerations for the development of air to air missiles. What

factors limits the range of such missiles?

9. Classify the various types of missiles.

10.What is rocket dispersion? How it is classified? What are the important factors that cause dispersion?

How it can be m

rockets systems

Documents