liquid rocket combustion chamber

8/10/2019 Liquid Rocket Combustion Chamber

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LIQUID PROPELLANTROCKET COMBUSTION

CHAMBER


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Liquid propellant rocket combustionchamber:

The combustion chamber is thatpart of a thrust chamber where thecombustion or burning of the propellant

takes place.The combustion temperature is

much higher than the melting points of most

chamber wall materials.Therefore it is necessary either to

cool these walls or to stop rocket operation

before the critical wall areas become toohi h the thrust chamber will fail.


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Combustion of liquid propellants:The combustion of liquid propellants is

very efficient in well designed thrustchambers, precombustion chamber, or gasgenerators.

Efficiencies of 95% to 99.5% are typicalcompared to turbojets or furnaces, which canrange from 50 to 97%.

This is due to very high reaction rate atthe high combustion temperatures and thethrust mixing of fuel and oxidizer reaction

species by means of good injectiondistribution and as turbulence.


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COMBUSTION CHAMBER ZONES1.Injection/Atomization zone2.Rapid combustion zone3.Stream-tube combustion zone4.Transonic-flow zone5.Supersonic expansion zone


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Injection/atomization zone In this zone the liquids are atomized

into a large number of small droplets.heat is transformed to the droplets

by radiation from the very hot rapidcombustion zone and by convection frommoderately hot gases in the first zone.

it is heterogenous, it contains liquid

and vapourized propellant as well as someburning hot gases.


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chemical reactions occur in this zone,but the rate of heat generation is relatively

low, in part because the liquids and gasesare still relatively low, in part because theliquid and gases are still relatively cold.

In part because vapourization nearthe droplets fuel rich and fuel lean regionswhich do not burn as quickly.

some hot gases in combustion zone

are re-circulated back from the rapidcombustion zone and they can create localgas velocities that flow across the injectorphase.


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Stream tube combustion zone:In this oxidation reactions continue, but at

a lower rate, and some additional heat isreleased.

However chemical reactions continue

because the mixture tends to be driventowards an equilibrium composition.Streamlines are formed and there is

relatively little turbulent mixing acrossstreamline boundaries.The residence time in this zone is very

short compared to the residence time in theother two zones.


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The residence time of the propellantmaterial in the combustion chamber is very

short, usually less than 100 milliseconds.Combustion in a liquid rocket engine is

very dynamic with the volumetric heat

release being approximately 370 MJ/m3-sec,which is much higher than in turbojets.


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DESIGN CONSIDERATION OFCOMBUSTION CHAMBER

1.Combustion chamber shape2.Combustion chamber volume3.Chamber wall load and stresses


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Combustion chamber shape: A long chamber with small cross section

entails high nonisentropic pressure losses.Long chamber also dictate a longer thrust

chamber envelope and impose space

limitations on the injector design toaccommodate the desired number of injectionelements.

With short chamber a large cross section,the propellant atomization and vaporizationzone occupies a relatively large portion of thechamber volume, while mixing andcombustion zone becomes too short for


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Other factors such as heat transfer,combustion stability, weight also be consideredin determining the final combustion chamberconfiguration.

Three geometrical shapes that have beenused in combustion chamber design.

There are spherical, near spherical,cylindrical chamber.While spherical and near spherical chambers

were used in early European designs, the

cylindrical chamber has been employed mostfrequently in the united states, compared to acylindrical chamber of same volume, a sphericalor near spherical chamber offers theadvantages of less cooling surfaces,and weight.


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A spherical has the best surfaces tovolume ratio of all the geometric choices andfor the same material strength and chamberpressure.

The minimum wall thickness required forpressure load is about half that of a cylinder.

Spherical chamber gives the least internalsurfaces area and mass per unit chambervolume.

However , the spherical chamber is moredifficult manufacture and has provided poorerperformance of other respects.

Today we prefer a cylindrical chamber(orslightly tapered cone frustrum)with a flat


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Typical combustion chamber characteristicslength for various propellant combinations.

S.no Propellant combination Combustion chamber(inch)

1. Chlorine triflouride/hydrazine base fuel 20-35

2. Liquid flourine/hydrazine 24-38

3. Hydrogen peroxide/RP-I 60-70

4. Liquid flouring/liquid hydrogen 25-30

5. Nitric acid/hydrazine base fuel 30-35

6. Liquid oxygen/liquid hydrogen 30-40

7. Liquid oxygen/RP-1 40-50


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2.Combustion chamber volume:The chamber volume is defined as the

volume upto the nozzle throat section and itincludes the cylindrical chamber and theconverging cone frustrum of the nozzle.

Neglecting the effect of corner radii, thechamber volume v c isvc = A1L + A 1LC (1+

L cylindrical length A1/At = chamber contraction ratioLC Length of conical frustrum


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The theorically required combustionchamber volume is a function of the mass

flow rate of the propellants, the averagedensity of the combustion products and thestay time needed for efficient combustion.

VC = mVt sm propellant mass flow rateVC = Average specific volume

ts propellant stay time


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Stay time (t s):It is the average valve of the time spent

by each molecule or atom within thechamber.

stay times have the values of .001 to

.040 sec for different types of thrustchambers and propellant.

Characteristics chamber length (L C or L *):

The characteristics chamber length isdefined as the length that a chamber of thesame volume would have if it were a

straight tube and had no converging nozzle


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L* = V C/AtVC chamber volume

A t

nozzle throat areaTypical values for L * are between 0.8 and 3meters (i.e. 2.6 to 10ft) for severalbipropellants and higher for somemonopropellants.

The volume and shape are selected afterevaluating the following parameters:

1. The volume has to be large enough foradequate mixing, evaporation and completecombustion of propellants, chamber volumesvary for different propellants with the timedelay necessary to vaporize and activate thepropellants and with the speed of reaction ofthe propellant combination.


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When the chamber volume is too small,combustion is incomplete and the

performance is poor.With higher chamber pressures or withhighly reactive propellants, and withinjectors that give improved mixing, asmaller chamber volume is usuallypermissible.

2. The chamber diameter and volume can

influence the cooling requirements. If thechamber volume and the chamber diameterare large , the heat transfer rates to thewalls willbe reduced, the area exposed toheat will be large, and the walls are


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3.All inert components should have minimummass. The thrust chamber mass is a functionof the chamber dimensions, chamber pressureand nozzle area ratio and the method of cooling.

4.Manufacturing consideration favor simplechamber geometry, such as a cylindrical with adouble cone bow-tie shaped nozzle, low costmaterials and simpler fabrication processes.

5.In some applications the length of the chamberand the nozzle relate directly to the overall

length of the vehicle. A large diameter but short chamber can allowa somewhat shorter vehicle with lower structuralinert vehicle mass.


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6.The gas pressure drop for accelerating thecombustion products within the chamber shouldbe minimum; any pressure reduction at thenozzle inlet reduces the exhaust velocity and theperformance of the vehicle.These losses become appreciable when thechamber area loss than three times the throatarea.

7.For the same thrust combustion volume and thenozzle throat area become smaller as the

operating chamber pressure is increased.This means that the chamber length and thenozzle length also decrease with increasingchamber pressure, the performance will go upwith chamber pressure.


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3.Chamber Walls Loads And Stresses:The analysis of loads and stresses is

performed on all propulsion componentsduring their engineering design.

Its purpose is to assure the propulsion

designed and the flight vehicle user that,The components are strong enough to

carry all the loads, so that they can fulfill their

intended functions1. Potential failures have been identifiedtogether with the possible remedies orredesigns.


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2. Their masses have been reduced to a practicalminimum.In this section we concentrate on the loads andstresses in the walls of thrust chambers, wherehigh heat fluxes and large thermal stressescomplicate the stress analysis.

Some of the information on safety factors andstress analysis apply also to all propulsionsystem, including solid propellant motors andelectric propulsion.

The safety factors are very small in rocketpropulsion when compared to commercialmachining, where these factors can be 2-6 timeslarger.several load conditions are considered for each


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a. Maximum expected working load is thelargest likely operating load under all likely

operating conditions or transients.b. The design limit load is typically 1.20 times

the maximum expected load to provide a

margin.c. The damaging load can be based on theyield load or the ultimate load of theendurance limit load, whichever gives thelowest value.

The yield load causes a permanent set ordeformation, and is typically set as 1.10times the design limit load.


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d. The proof test load is applied to engines theircomponents during development and

manufacturing inspection.It is often equal to the design limit load, providethis load condition can be stimulated in alaboratory.

The walls of all the thrust chambers aresubjected to rapid and axial loads from thechamber pressures, flight accelerations(axial

and transverse), vibrations and thermalstresses.

They also have withstand a momentary

ignition pressure surge or shock, often due to


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Thus surge can exceed the nominal chamberpressure.

In addition the chamber walls have to transmit

thrust loads as well as forces and in someapplications also moments, imposed by thrustvector control devices.

walls also have to survive a thermal shocknamely the initial thermal stresses at rapidstarting.

When walls are cold or at ambient temperaturethey experience higher gas heating rates than afterthe walls have been heated.

These loads are different for almost everydesign and unit has to be considered individually indetermining the wall strengths.


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COMBUSTION INSTABILITYIt is defined in terms of the nature of

pressure fluctuations in the combustionchamber.

combustion chamber pressurefluctuations present during normal, stableoperation of a rocket engine system.

there are usually quite random

showing frequently spectra that areessentially continious in nature,with few, if anrecongnizable peaks.


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Combustion instabilities cause1.Pressure peaks which increases the

burning rate and results in enormous pressurebuild up.

2. increased heat transfer rates andhigher wall temperatures.

3. vibration of the structure and thesensitive electronic instruments and the pay

load.4. uncomfortable to the passengers

inside.


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If the process of rocket combustion isnot controlled,

then combustion instabilities can occurwhich can vary quickly cause excessive heattransfer.

the aim is to prevent occurrence of thisinstability and to maintain reliable operation.


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Types of combustion instabilities

s.no type Worddescription Frequency range Cause relationship

1 Low frequency Chugging(feed systeminstability)

10-400 Linked with pressureinteractions betweenpropellant feed system.If not the entire vehicleand combustionchamber

2 Intermediatefrequency

Acoustic orbuzzing

400-1000 Linked with mechanicalvibrations of propusionstructure, injector

manifold flow eddies,fuel/oxidizer ratiofluctuations andpropellant feed systemresonances

3 Highfrequency

Screaming/screeching/

Above1000

Linked with combustionprocess forces


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CHUGGINGChugging a low frequency instability

arise from an interaction between thepropellant feed system and the combustionchamber of the rocket motor.

Propellant flow ratedisturbance(50Hz) gives rise to low frequencylongitudinal combustion instability, producing alongitudina motion of vibration in the vehicle.

This vehicle instability is called pogoinstability.eg. Space launch vehicle/ballisticmissiles


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BUZZINGThe inter-mediate type of instability which

represents pressure perturbations greaterthan 5%Of the mean combustion chamber.

It has been associated in resonances ofvarious parts of the feed system and mountingstructure and vortex formation in the gas flowingaround corners.

Most have low amplitude.It is often more noisy annoying than

damaging, although the occurrence of buzzingmay initiate high frequency instabilities.(mediumsize en ine


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SCREECHINGScreeching or screaming has frequency

and is most persisting and most common inthe development of new engines.

it involves high frequency oscillations,

this is due to the interaction between acousticsof the chamber and the combustion process.

high frequency instability occurs in two

modes, longitudinal and transverse.trigerring the high frequency instability

is a rocket combustion phenomenon called

i

liquid rocket combustion chamber

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