Propulsion 2

7/24/2019 Propulsion 2

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saac Newton stated in his third law of motion that "for every action there is an equal and

opposite reaction." It is upon this principle that a rocket operates. Propellants are combined in a

combustion chamber where they chemically react to form hot gases which are then acceleratedand ejected at high velocity through a nozzle thereby imparting momentum to the engine. !he

thrust force of a rocket motor is the reaction eperienced by the motor structure due to ejection

of the high velocity matter. !his is the same phenomenon which pushes a garden hose backwardas water flows from the nozzle or makes a gun recoil when fired.

!hrust

Thrustis the force that propels a rocket or spacecraft and is measured in pounds kilograms or

Newtons. Physically speaking it is the result of pressure which is eerted on the wall of thecombustion chamber.

#igure $.$ shows a combustion chamber with an

opening the nozzle through which gas can escape. !he

pressure distribution within the chamber is asymmetric%that is inside the chamber the pressure varies little but

near the nozzle it decreases somewhat. !he force due to

gas pressure on the bottom of the chamber is not

compensated for from the outside. !he resultantforceFdue to the internal and eternal pressure

difference the thrust is opposite to the direction of the

gas jet. It pushes the chamber upwards.

!o create high speed ehaust gases the necessary high

temperatures and pressures of combustion are obtained

by using a very energetic fuel and by having themolecular weight of the ehaust gases as low aspossible. It is also necessary to reduce the pressure of

the gas as much as possible inside the nozzle by creating

a large section ratio. !he section ratio or epansionratio is defined as the area of the eit Aedivided by the

area of the throatAt.

!he thrustFis the resultant of the forces due to the pressures eerted on the inner and outer

walls by the combustion gases and the surrounding atmosphere taking the boundary between theinner and outer surfaces as the cross section of the eit of the nozzle. &s we shall see in the net

section applying the principle of the conservation of momentum gives

where qis the rate of the ejected mass flowPathe pressure of the ambient atmospherePethe

pressure of the ehaust gases and Vetheir ejection speed. !hrust is specified either at sea level or

in a vacuum.


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'onservation of (omentum

!he linear momentum)p* or simply momentum of a particle is the product of its mass and its

velocity. !hat is

Newton epressed his second law of motion in terms of momentum which can be stated as "the

resultant of the forces acting on a particle is equal to the rate of change of the linear momentumof the particle". In symbolic form this becomes

which is equivalent to the epressionF=ma.

If we have a system of particles the total momentumPof the system is the sum of the momenta

of the individual particles. +hen the resultant eternal force acting on a system is zero the total

linear momentum of the system remains constant. !his is called the principle ofconservation oflinear momentum. ,et-s now see how this principle is applied to rocket mechanics.

'onsider a rocket drifting in gravity free space. !he rocket-s engine is fired for time tand

during this period ejects gases at a constant rate and at a constant speed relative to the rocket

)ehaust velocity*. &ssume there are no eternal forces such as gravity or air resistance.

#igure $.)a* shows the situation at time t. !he rocket and fuel have a total massMand the

combination is moving with velocity vas seen from a particular frame of reference. &t a time

tlater the configuration has changed to that shown in #igure $.)b*. & mass Mhas been ejectedfrom the rocket and is moving with velocity uas seen by the observer. !he rocket is reduced to

massM- Mand the velocity vof the rocket is changed to v+ v.

/ecause there are no eternal forces dP/dt=0. +e can write for the time interval t


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whereP2is the final system momentum #igure $.)b* andPis the initial system momentum

#igure $.)a*. +e write

If we let tapproach zero v/ tapproaches dv/dt the acceleration of the body. !he quantityMis the mass ejected in t% this leads to a decrease in the massMof the original body.

0ince dM/dt the change in mass of the body with time is negative in this case in the limit the

quantity M/ tis replaced by -dM/dt. !he quantity u-!v+ v"is Vrel the relative velocity of the

ejected mass with respect to the rocket. +ith these changes equation )$.1* can be written as

!he right2hand term depends on the characteristics of the rocket and like the left2hand term hasthe dimensions of a force. !his force is called the thrust and is the reaction force eerted on the

rocket by the mass that leaves it. !he rocket designer can make the thrust as large as possible by

designing the rocket to eject mass as rapidly as possible )dM/dtlarge* and with the highestpossible relative speed )Vrellarge*.

In rocketry the basic thrust equation is written as

where qis the rate of the ejected mass flow Veis the ehaust gas ejection speedPeis thepressure of the ehaust gases at the nozzle eitPais the pressure of the ambient atmosphere

andAeis the area of the nozzle eit. !he product qVe which we derived above )Vrel# dM/dt* is

called the momentum or velocity thrust. !he product !Pe-Pa"Ae called the pressure thrust is theresult of unbalanced pressure forces at the nozzle eit. &s we shall see latter maimum thrust

occurs whenPe=Pa.

'lick herefor eample problem 3$.$

)use your browser-s "back" function to return*

4quation )$.5* may be simplified by the definition of an effective e$haust %as velocit&' ('defined

as

4quation )$.5* then reduces to
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Impulse 6 (omentum

In the preceding section we saw that Newton-s second law may be epressed in the form

(ultiplying both sides by dtand integrating from a time tto a time t2 we write

!he integral is a vector known as the linear impulse or simply the impulse of the forceFduring

the time interval considered. !he equation epresses that when a particle is acted upon by a

forceFduring a given time interval the final momentump2of the particle may be obtained byadding its initial momentumpand the impulse of the forceFduring the interval of time.

+hen several forces act on a particle the impulse of each of the forces must be considered.

+hen a problem involves a system of particles we may add vectorially the momenta of all the

particles and the impulses of all the forces involved. +hen can then write

#or a time interval t we may write equation )$.$7* in the form

,et us now see how we can apply the principle of impulse and momentum to rocket mechanics.

'onsider a rocket of initial massMwhich it launched vertically at time t87. !he fuel isconsumed at a constant rate qand is epelled at a constant speed Verelative to the rocket. &t

time t the mass of the rocket shell and remaining fuel isM-qt and the velocity is v. 9uring the

time interval t a mass of fuel q tis epelled. 9enoting by uthe absolute velocity of theepelled fuel we apply the principle of impulse and momentum between time tand time t+ t.

Please note this derivation neglects the effect of air resistance.


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+e write

+e divide through by tand replace u-!v+ v"with Ve the velocity of the epelled mass relativeto the rocket. &s tapproaches zero we obtain

0eparating variables and integrating from t87 v87 to t=t' v=v we obtain

which equals

!he term -%tin equation )$.$:* is the result of 4arth-s gravity pulling on the rocket. #or a rocket

drifting in space -%tis not applicable and can be omitted. #urthermore it is more appropriate toepress the resulting velocity as a change in velocity or ;. 4quation )$.$:* thus becomes

'lick herefor eample problem 3$.

Note thatMrepresents the initial mass of the rocket andM-qtthe final mass. !herefore equation

)$.$5* is often written as


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where mo/mfis called the mass ratio. 4quation )$.$onstantin 4. !siolkovsky )$?:


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"fuel2rich" miture ratio.Mi$ture ratiois defined as the mass flow of oidizer divided by the

mass flow of fuel.

'onsider the following reaction of kerosene)$*with oygen

Biven the molecular weight of '$C5is $ in 0I

units or 1@


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It should be pointed out that in the combustion process there will be a dissociation of molecules

among the products. !hat is the high heat of combustion causes the separation of molecules into

simpler constituents that are then capable of recombining. 'onsider the reaction of kerosene withoygen. !he true products of combustion will be an equilibrium miture of atoms and molecules

consisting of ' 'D 'D C C CD CD D and D. 9issociation has a significant effect on

flame temperature.

If you wish to learn more about the thermodynamics of rockets engines please consider readingthe appendi =ocket !hermodynamics.

Dr you can skip all the science and just look up the numbers you need. 0ee Propellant

'ombustion 'hartsto find optimum miture ratio adiabatic flame temperature gas molecularweight and specific heat ratio for some common rocket propellants.

)$* In dealing with combustion of liquid hydrocarbon fuels it is convenient to epress the

composition in terms of a single hydrocarbon even though it is a miture of many

hydrocarbons. !hus gasoline is usually considered to be octane '?C$? and kerosene isconsidered to be dodecane '$C5.

)* 0pecific heat or heat capacity represents the amount of heat necessary to raise the

temperature of one gram of a substance one degree '. 0pecific heat is measured at

constant2pressure 'P or at constant2volume ';. !he ratio 'PE';is called the specific

heat ratio represented by )or .

0pecific Impulse

!he specific impulse of a rocket,sp is the ratio of the thrust to the flow rate of the weightejected that is

whereFis thrust qis the rate of mass flow and%ois standard gravity )@.?755: mEs*.

0pecific impulse is epressed in seconds. +hen the thrust and the flow rate remain constant

throughout the burning of the propellant the specific impulse is the time for which the rocket

engine provides a thrust equal to the weight of the propellant consumed.

#or a given engine the specific impulse has different values on the ground and in the vacuum of

space because the ambient pressure is involved in the epression for the thrust. It is therefore

important to state whether specific impulse is the value at sea level or in a vacuum.

!here are a number of losses within a rocket engine the main ones being related to theinefficiency of the chemical reaction )combustion* process losses due to the nozzle and losses

due to the pumps. Dverall the losses affect the efficiency of the specific impulse. !his is the ratio
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of the real specific impulse )at sea level or in a vacuum* and the theoretical specific impulse

obtained with an ideal nozzle from gases coming from a complete chemical reaction. 'alculated

values of specific impulse are several percent higher than those attained in practice.

'lick herefor eample problem 3$.5

#rom 4quation )$.?* we can substitute q(forFin 4quation )$.A* thus obtaining

4quation )$.1* is very useful when solving 4quations )$.$?* through )$.$*. It is rare we are

given the value of (directly however rocket engine specific impulse is a commonly givenparameter from which we can easily calculate (.

&nother important figure of merit for evaluating rocket performance is the characteristic e$haustvelocit&' ()pronounced "' star"* which is a measure of the energy available from the

combustion process and is given by

wherePcis the combustion chamber pressure andAtis the area of the nozzle throat. 9eliveredvalues of (range from about $AAA mEs for monopropellant hydrazine up to about A57 mEs for

cryogenic oygenEhydrogen.

=ocket 4ngines

& typical rocket engine consists of the nozzle the combustion chamber and the injector asshown in #igure $.1. !he combustion chamber is where the burning of propellants takes place at

high pressure. !he chamber must be strong enough to contain the high pressure generated by and

the high temperature resulting from the combustion process. /ecause of the high temperature

and heat transfer the chamber and nozzle are usually cooled. !he chamber must also be ofsufficient length to ensure complete combustion before the gases enter the nozzle.


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Nozzle

!he function of the nozzle is to

convert the chemical2thermal energygenerated in the combustion chamber

into kinetic energy. !he nozzleconverts the slow moving high

pressure high temperature gas in thecombustion chamber into high

velocity gas of lower pressure and

temperature. 0ince thrust is theproduct of mass and velocity a very

high gas velocity is desirable.

Nozzles consist of a convergent anddivergent section. !he minimum flow area between the convergent and divergent section is

called the nozzle throat. !he flow area at the end of the divergent section is called the nozzle eit

area. !he nozzle is usually made long enough )or the eit area is great enough* such that thepressure in the combustion chamber is reduced at the nozzle eit to the pressure eisting outside

the nozzle. It is under this conditionPe=PawherePeis the pressure at the nozzle eit andPais

the outside ambient pressure that thrust is maimum and the nozzle is said to be adapted also

called optimum or correct epansion. +henPeis greater thanPa the nozzle is under2etended.+hen the opposite is true it is over2etended.

+e see therefore a nozzle is designed for the altitude at which it has to operate. &t the 4arth-s

surface at the atmospheric pressure of sea level )7.$ (Pa or $1.< psi* the discharge of the

ehaust gases is limited by the separation of the jet from the nozzle wall. In the cosmic vacuumthis physical limitation does not eist. !herefore there have to be two different types of engines

and nozzles those which propel the first stage of the launch vehicle through the atmosphere andthose which propel subsequent stages or control the orientation of the spacecraft in the vacuumof space.

!he nozzle throat areaAt can be found if the total propellant flow rate is known and the

propellants and operating conditions have been selected. &ssuming perfect gas law theory we

have

where qis the propellant mass flow ratePtis the gas pressure at the nozzle throat Ttis the gastemperature at the nozzle throat*is the universal gas constant and )is the specific heat

ratio.Ptand Ttare given by


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wherePcis the combustion chamber pressure and Tcis the combustion chamber flametemperature.

'lick herefor eample problem 3$.* and the highheat transfer rates )up to $5 kFEcm2s* encountered in a combustion chamber present a formidable


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challenge to the designer. !o meet this challenge several chamber cooling techniques have been

utilized successfully. 0election of the optimum cooling method for a thrust chamber depends on

many considerations such as type of propellant chamber pressure available coolant pressurecombustion chamber configuration and combustion chamber material.

Regenerative coolingis the most widely used method of cooling a thrust chamber and isaccomplished by flowing high2velocity coolant over the back side of the chamber hot gas wall to

convectively cool the hot gas liner. !he coolant with the heat input from cooling the liner is thendischarged into the injector and utilized as a

propellant.

4arlier thrust chamber designs such as the ;2 and=edstone had low chamber pressure low heat flu

and low coolant pressure requirements which could

be satisfied by a simplified "double wall chamber"

design with regenerative and film cooling. #or

subsequent rocket engine applications howeverchamber pressures were increased and the cooling

requirements became more difficult to satisfy. Itbecame necessary to design new coolant

configurations that were more efficient structurally

and had improved heat transfer characteristics.

!his led to the design of "tubular wall" thrustchambers by far the most widely used design

approach for the vast majority of large rocket engine

applications. !hese chamber designs have been

successfully used for the !hor Fupiter &tlas C2$ F2 #2$ =02< and several other &ir #orceand N&0& rocket engine applications. !he primary advantage of the design is its light weight

and the large eperience base that has accrued. /ut as chamber pressures and hot gas wall heatflues have continued to increase )K$77 atm* still more effective methods have been needed.

Dne solution has been "channel wall" thrust chambers so named because the hot gas wall

cooling is accomplished by flowing coolant through rectangular channels which are machined or

formed into a hot gas liner fabricated from a high2conductivity material such as copper or acopper alloy. & prime eample of a channel wall combustion chamber is the 00(4 which

operates at 71 atmospheres nominal chamber pressure at A577 > for a duration of :7 seconds.

Ceat transfer and structural characteristics are ecellent.

In addition to the regeneratively cooled designs mentioned above other thrust chamber designshave been fabricated for rocket engines using dump cooling film cooling transpiration cooling

ablative liners and radiation cooling. &lthough regeneratively cooled combustion chambers have

proven to be the best approach for cooling large liquid rocket engines other methods of coolinghave also been successfully used for cooling thrust chamber assemblies. 4amples includeH


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ump cooling which is similar to regenerative cooling because the coolant flows through small

passages over the back side of the thrust chamber wall. !he difference however is that after

cooling the thrust chamber the coolant is discharged overboard through openings at the aft endof the divergent nozzle. !his method has limited application because of the performance loss

resulting from dumping the coolant overboard. !o date dump cooling has not been used in an

actual application.

!ilm coolingprovides protection from ecessive heat by introducing a thin film of coolant orpropellant through orifices around the injector periphery or through manifolded orifices in the

chamber wall near the injector or chamber throat region. !his method is typically used in high

heat flu regions and in combination with regenerative cooling.

"ranspirationcooling provides coolant )either gaseous or liquid propellant* through a porous

chamber wall at a rate sufficient to maintain the chamber hot gas wall to the desired temperature.

!he technique is really a special case of film cooling.

+ith ablative cooling combustion gas2side wall material is sacrificed by melting vaporizationand chemical changes to dissipate heat. &s a result relatively cool gases flow over the wall

surface thus lowering the boundary2layer temperature and assisting the cooling process.

+ith radiation cooling heat is radiated from the outer surface of the combustion chamber or

nozzle etension wall. =adiation cooling is typically used for small thrust chambers with a high2temperature wall material )refractory* and in low2heat flu regions such as a nozzle etension.

0olid =ocket (otors

0olid rockets motors store propellants in solid form. !he fuel is typically powdered aluminum

and the oidizer is ammonium perchlorate. & synthetic rubber binder such as polybutadieneholds the fuel and oidizer powders together. !hough lower performing than liquid propellantrockets the operational simplicity of a solid rocket motor often makes it the propulsion system of

choice.

Solid Fuel eometr!

& solid fuel-s geometry determines the area and contours of its eposed surfaces and thus itsburn pattern. !here are two main types of solid fuel blocks used in the space industry. !hese are

cylindrical blocks with combustion at a front or surface and cylindrical blocks with internal

combustion. In the first case the front of the flame travels in layers from the nozzle end of the

block towards the top of the casing. !his so2called end burner produces constant thrustthroughout the burn. In the second more usual case the combustion surface develops along the

length of a central channel. 0ometimes the channel has a star shaped or other geometry to

moderate the growth of this surface.


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!he shape of the fuel block for a rocket is chosen for the particular type of mission it will

perform. 0ince the combustion of the block progresses from its free surface as this surface

grows geometrical considerations determine whether the thrust increases decreases or staysconstant.

#uel blocks with a cylindrical channel )$* develop their thrust progressively. !hose with a

channel and also a central cylinder of fuel )* produce a relatively constant thrust which reduces

to zero very quickly when the fuel is used up. !he five pointed star profile )A* develops arelatively constant thrust which decreases slowly to zero as the last of the fuel is consumed. !he

-cruciform- profile )1* produces progressively less thrust. #uel in a block with a -double anchor-

profile ):* produces a decreasing thrust which drops off quickly near the end of the burn. !he-cog- profile )5* produces a strong inital thrust followed by an almost constant lower thrust.

"urn #ate

!he burning surface of a rocket propellant grain recedes in a direction perpendicular to thisburning surface. !he rate of regression typically measured in millimeters per second )or inches

per second* is termed 5urn rate. !his rate can differ significantly for different propellants or forone particular propellant depending on various operating conditions as well as formulation.

>nowing quantitatively the burning rate of a propellant and how it changes under various

conditions is of fundamental importance in the successful design of a solid rocket motor.


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Propellant burning rate is influenced by certain factors the most significant beingH combustion

chamber pressure initial temperature of the propellant grain velocity of the combustion gases

flowing parallel to the burning surface local static pressure and motor acceleration and spin.!hese factors are discussed below.

/urn rate is profoundly affected by chamber pressure. !he usual representation of thepressure dependence on burn rate is the 0aint2=obert-s ,aw

where ris the burn rate ais the burn rate coefficient nis the pressure eponent andPcisthe combustion chamber pressure. !he values of aand nare determined empirically for a

particular propellant formulation and cannot be theoretically predicted. It is important to

realize that a single set of a' nvalues are typically valid over a distinct pressure range.(ore than one set may be necessary to accurately represent the full pressure regime of

interest.

4ample a' nvalues are :.57:@J )pressure in (Pa burn rate in mmEs* and 7.A:

respectively for the 0pace 0huttle 0=/s which gives a burn rate of @.A1 mmEs at theaverage chamber pressure of 1.A (Pa.

J N&0& publications gives a burn rate coefficient of 7.7A?55: )pressure in P0I burn

rate in inchEs*.

!emperature affects the rate of chemical reactions and thus the initial temperature of the

propellant grain influences burning rate. If a particular propellant shows significant

sensitivity to initial grain temperature operation at temperature etremes will affect the

time2thrust profile of the motor. !his is a factor to consider for winter launches foreample when the grain temperature may be lower than "normal" launch conditions.

#or most propellants certain levels of local combustion gas velocity )or mass flu*

flowing parallel to the burning surface leads to an increased burning rate. !his

"augmentation" of burn rate is referred to as erosive 5urnin% with the etent varying with

propellant type and chamber pressure. #or many propellants a threshold flow velocityeists. /elow this flow level either no augmentation occurs or a decrease in burn rate is

eperienced )ne%ative erosive 5urnin%*.

!he effects of erosive burning can be minimized by designing the motor with a

sufficiently large port2to2throat area ratio )&portE&t*. !he port area is the cross2section areaof the flow channel in a motor. #or a hollow2cylindrical grain this is the cross2section

area of the core. &s a rule of thumb the ratio should be a minimum of for a grain ,E9

ratio of 5. & greater &portE&tratio should be used for grains with larger ,E9 ratios.

In an operating rocket motor there is a pressure drop along the ais of the combustion

chamber a drop that is physically necessary to accelerate the increasing mass flow of

combustion products toward the nozzle. !he static pressure is greatest where gas flow is


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zero that is at the front of the motor. 0ince burn rate is dependant upon the local

pressure the rate should be greatest at this location. Cowever this effect is relatively

minor and is usually offset by the counter2effect of erosive burning.

/urning rate is enhanced by acceleration of the motor. +hether the acceleration is a result

of longitudinal force )e.g. thrust* or spin burning surfaces that form an angle of about 572@7owith the acceleration vector are prone to increased burn rate.

It is sometimes desirable to modify the burning rate such that it is more suitable to a certain grainconfiguration. #or eample if one wished to design an end burner grain which has a relatively

small burning area it is necessary to have a fast burning propellant. In other circumstances a

reduced burning rate may be sought after. #or eample a motor may have a large ,E9 ratio togenerate sufficiently high thrust or it may be necessary for a particular design to restrict the

diameter of the motor. !he web would be consequently thin resulting in short burn duration.

=educing the burning rate would be beneficial.

!here are a number of ways of modifying the burning rateH decrease the oidizer particle sizeincrease or reduce the percentage of oidizer adding a burn rate catalyst or suppressant and

operate the motor at a lower or higher chamber pressure. !hese factors are discussed below.

!he effect of the oidizer particle size on burn rate seems to be influenced by the type of

oidizer. Propellants that use ammonium perchlorate )&P* as the oidizer have a burn

rate that is significantly affected by &P particle size. !his most likely results from thedecomposition of &P being the rate2determining step in the combustion process.

!he burn rate of most propellants is strongly influenced by the oidizerEfuel ratio.

Gnfortunately modifying the burn rate by this means is quite restrictive as the

performance of the propellant as well as mechanical properties are also greatly affectedby the DE# ratio.

'ertainly the best and most effective means of increasing the burn rate is the addition of

a catal&stto the propellant miture. & catalyst is a chemical compound that is added in

small quantities for the sole purpose of tailoring the burning rate. & burn

ratesuppressantis an additive that has the opposite effect to that of a catalyst L it is usedto decrease the burn rate.

#or a propellant that follows the 0aint2=obert-s burn rate law designing a rocket motor to

operate at a lower chamber pressure will provide for a lower burning rate. 9ue to the

nonlinearity of the pressure2burn rate relationship it may be necessary to significantlyreduce the operating pressure to get the desired burning rate. !he obvious drawback is

reduced motor performance as specific impulse similarly decays with reducing chamber

pressure.

Product eneration #ate


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!he rate at which combustion products are generated is epressed in terms of the regression

speed of the grain. !he product generation rate integrated over the port surface area is

where qis the combustion product generation rate at the propellant surface pis the solidpropellant densityA5is the area of the burning surface and ris the propellant burn rate.

'lick herefor eample problem 3$.$7

If the propellant density is unknown it can be derived from the mass fraction and density of the

individual constituents as followsH

where 6is the mass fraction and the subscript idenotes the individual constituents. !his is

the idealdensity% the actualdensity is typically @12@


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particles depends upon frictional drag in the gas flow which necessitates a differential

velocity. !he net result is that the condensed2phase particles eit the nozzle at a lower

velocity than the gases. !his is referred to asparticle velocit& la%.

Chamber Pressure

!he pressure curve of a rocket motor ehibits transient and steady state behavior. !he transient

phases are when the pressure varies substantially with time L during the ignition and start2up

phase and following complete )or nearly complete* grain consumption when the pressure fallsdown to ambient level during the tail2off phase. !he variation of chamber pressure during the

steady state burning phase is due mainly to variation of grain geometry with associated burn rate

variation. Dther factors may play a role however such as nozzle throat erosion and erosive burnrate augmentation.

(onopropellant 4ngines

/y far the most widely used type of propulsion for spacecraft attitude and velocity control ismonopropellant hydrazine. Its ecellent handling characteristics relative stability under normal

storage conditions and clean decomposition products have made it the standard. !he generalsequence of operations in a hydrazine thruster isH

+hen the attitude control system signals for thruster operation an electric solenoid valve

opens allowing hydrazine to flow. !he action may be pulsed )as short as : ms* or long

duration )steady state*.

!he pressure in the propellant tank forces liquid hydrazine into the injector. It enters as a

spray into the thrust chamber and contacts the catalyst beds.

!he catalyst bed consists of alumina pellets impregnated with iridium. Incoming

hydrazine heats to its vaporizing point by contact with the catalyst bed and with the hot

gases leaving the catalyst particles. !he temperature of the hydrazine rises to a pointwhere the rate of its decomposition becomes so high that the chemical reactions are self2

sustaining.

/y controlling the flow variables and the geometry of the catalyst chamber a designer

can tailor the proportion of chemical products the ehaust temperature the molecular

weight and thus the enthalpy for a given application. #or a thruster application wherespecific impulse is paramount the designer attempts to provide A7217 ammonia

dissociation which is about the lowest percentage that can be maintained reliably. #orgas2generator application where lower temperature gases are usually desired thedesigner provides for higher levels of ammonia dissociation.

#inally in a space thruster the hydrazine decomposition products leave the catalyst bed

and eit from the chamber through a high epansion ratio ehaust nozzle to produce

thrust.


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(onopropellant hydrazine thrusters typically produce a specific impulse of about A7 to 17

seconds.

Dther suitable propellants for catalytic decomposition engines are hydrogen peroide and nitrousoide however the performance is considerably lower than that obtained with hydrazine 2

specific impulse of about $:7 s with CDand about $


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'lick herefor eample problem 3$.$

+e define thepa&load fractionas the ratio of payload mass to initial mass or mpl/mo.

#or a multistage vehicle with dissimilar stages the overall vehicle payload fraction depends on

how the ; requirement is partitioned among stages. Payload fractions will be reduced if the; is partitioned suboptimally. !he optimal distribution may be determined by trial and error. &

; distribution is postulated and the resulting payload fraction calculated. !he ; distribution is

varied until the payload fraction is maimized. Dnce the ; distribution is selected vehiclesizing is accomplished by starting with the uppermost or final stage )whose payload is the actual

deliverable payload* and calculating the initial mass of this assembly. !his assembly then forms

the payload for the previous stage and the process repeats until all stages are sized. =esults revealthat to maimize payload fraction for a given ; requirementH

$. 0tages with higher Ispshould be above stages with lower Isp.

. (ore ; should be provided by the stages with the higher Isp.

A. 4ach succeeding stage should be smaller than its predecessor.1. 0imilar stages should provide the same ;.

!hrust is the forcewhich moves any aircraft through the air. !hrust is generated by

thepropulsion systemof the aircraft. 9ifferent propulsion systems develop thrust in different

ways but all thrust is generated through some application of Newton-s third lawof motion. #or

every action there is an equal and opposite reaction. In any propulsion system a &or'in( )luidis

accelerated by the system and the reaction to this acceleration produces a force on the system. &general derivation of thethrust equationshows that the amount of thrust generated depends on

the mass flowthrough the engine and theeit velocityof the gas.

9uring and following +orld +ar II there were a number of rocket2 powered aircraft built to

eplore high speed flight. !he M2$& used to break the "sound barrier" and the M2$: were

rocket2powered airplanes. In a roc'et en(ine fuel and a source of oygen called an oidizer

are mied and eploded in acombustion chamber. !hecombustionproduces hot ehaust which

is passed through a nozzleto accelerate the flow andproduce thrust.#or a rocket the accelerated

gas or &or'in( )luid, is the hot ehaust produced during combustion. !his is a different

working fluid than you find in a turbine engine or apropellerpowered aircraft. !urbine engines

and propellers use air from the atmosphere as the working fluid but rockets use the combustion

ehaust gases. In outer space there is no atmosphere so turbines and propellers can not work

there. !his eplains why a rocket works in space but a turbine engine or a propeller does not

work.
http://www.braeunig.us/space/problem.htm#1.12https://www.grc.nasa.gov/www/k-12/airplane/thrust1.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/bgp.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/bgp.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/newton3.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/thrsteq.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/thrsteq.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/thrsteq.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/mflow.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/nozzleh.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/nozzleh.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/nozzleh.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/lowhyper.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/sound.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/burner.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/burner.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/combst1.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/combst1.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/nozzle.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/rockth.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/rockth.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/turbine.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/propeller.htmlhttp://www.braeunig.us/space/problem.htm#1.12https://www.grc.nasa.gov/www/k-12/airplane/thrust1.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/bgp.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/newton3.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/thrsteq.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/mflow.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/nozzleh.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/lowhyper.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/sound.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/burner.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/combst1.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/nozzle.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/rockth.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/turbine.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/propeller.html


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!here are two main categories of rocket engines% liuid roc'etsand solid roc'ets. In a liquid

rocketthe propellants the fuel and the oidizer are stored separately as liquids and are pumped

into the combustion chamber of the nozzle where burning occurs. In a solid rocketthe

propellants are mied together and packed into a solid cylinder. Gnder normal temperature

conditions the propellants do not burn% but they will burn when eposed to a source of heat

provided by an igniter. Dnce the burning starts it proceeds until all the propellant is ehausted.

+ith a liquid rocket you can stop the thrust by turning off the flow of propellants% but with a

solid rocket you have to destroy the casing to stop the engine. ,iquid rockets tend to be heavier

and more comple because of the pumps and storage tanks. !he propellants are loaded into the

rocket just before launch. & solid rocket is much easier to handle and can sit for years before

firing.

Dn this slide we show a picture of an M2$: rocket2powered airplane at the upper left and a

picture of a rocket engine test at the lower right. #or the picture at the right we only see the

outside of the rocket nozzle with the hot gas eiting out the bottom. !he M2$: was powered by aliquid rocket engine and carried a single pilot to a height of more than 57 miles above the earth.

!he M2$: flew more than si times the speed of sound nearly 17 years ago. !he speed record for

a piloted aircraft is only eceeded today by the 0pace 0huttle. !he altitude record is only topped

by the 0pace 0huttle and the recent 0pace 0hip $ which also used rocket propulsion.

#+C-T P#+P.LSI+N

=ockets )and jet engines* work much like a balloon filled with air.

If you fill a balloon with air and hold the neck closed the pressure inside the balloon is slightlyhigher than the surrounding atmosphere. Cowever there is no net force on the balloon in any

direction because the internal pressure on the balloon is equal in all directions.

If you release the neck of the balloon it acts like a hole with no surface area for the internal

pressure to act on. !here is now an imbalanced force on the balloon and the internal pressure on

the front of the balloon is greater than the internal pressure on the back of the balloon.

!his results in a net force acting forward on the balloonthrust. !he balloon flies forward

under the influence of the thrust and the air coming out of the back of the balloon is the equaland opposite reaction to the thrust.
https://www.grc.nasa.gov/www/k-12/airplane/lrockth.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/lrockth.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/srockth.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/srockth.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/lrockth.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/lrockth.htmlhttps://www.grc.nasa.gov/www/k-12/airplane/srockth.html


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& roc'et en(ineis a type ofjet engineO$that uses only stored rocket propellantmass for forming

its high speed propulsivejet. =ocket engines are reaction enginesobtaining thrust in accordance

withNewton-s third law.(ost rocket engines areinternal combustion engines although non2

combusting forms also eist. ;ehicles propelled by rocket engines are commonly called rockets.

0ince they need no eternal material to form their jet rocket engines can perform in

avacuumand thus can be used to propell spacecraftandballistic missiles.

=ocket engines as a group have the highest thrust are by far the lightest but are the least

propellant efficient )have the lowest specific impulse* of all types of jet engines. !he ideal

ehaust is hydrogen the lightest of all gasses but chemical rockets produce a mi of heavier

species reducing the ehaust velocity. =ocket engines become more efficient at high velocities

)Dberth effect*. 0ince they do not benefit from air they are best suited for uses in space and the

high atmosphere.

Contents

Ohide

$ !erminology

Principle of operation

o .$ Introducing propellant into a combustion chamber
https://en.wikipedia.org/wiki/Jet_enginehttps://en.wikipedia.org/wiki/Jet_enginehttps://en.wikipedia.org/wiki/Jet_enginehttps://en.wikipedia.org/wiki/Rocket_engine#cite_note-1https://en.wikipedia.org/wiki/Rocket_propellanthttps://en.wikipedia.org/wiki/Jet_(fluid)https://en.wikipedia.org/wiki/Reaction_enginehttps://en.wikipedia.org/wiki/Reaction_enginehttps://en.wikipedia.org/wiki/Newton's_third_lawhttps://en.wikipedia.org/wiki/Newton's_third_lawhttps://en.wikipedia.org/wiki/Newton's_third_lawhttps://en.wikipedia.org/wiki/Internal_combustion_enginehttps://en.wikipedia.org/wiki/Rockethttps://en.wikipedia.org/wiki/Vacuumhttps://en.wikipedia.org/wiki/Vacuumhttps://en.wikipedia.org/wiki/Spacecrafthttps://en.wikipedia.org/wiki/Ballistic_missilehttps://en.wikipedia.org/wiki/Specific_impulsehttps://en.wikipedia.org/wiki/Hydrogenhttps://en.wikipedia.org/wiki/Oberth_effecthttps://en.wikipedia.org/wiki/Rocket_enginehttps://en.wikipedia.org/wiki/Rocket_engine#Terminologyhttps://en.wikipedia.org/wiki/Rocket_engine#Principle_of_operationhttps://en.wikipedia.org/wiki/Rocket_engine#Introducing_propellant_into_a_combustion_chamberhttps://en.wikipedia.org/wiki/Jet_enginehttps://en.wikipedia.org/wiki/Rocket_engine#cite_note-1https://en.wikipedia.org/wiki/Rocket_propellanthttps://en.wikipedia.org/wiki/Jet_(fluid)https://en.wikipedia.org/wiki/Reaction_enginehttps://en.wikipedia.org/wiki/Newton's_third_lawhttps://en.wikipedia.org/wiki/Internal_combustion_enginehttps://en.wikipedia.org/wiki/Rockethttps://en.wikipedia.org/wiki/Vacuumhttps://en.wikipedia.org/wiki/Spacecrafthttps://en.wikipedia.org/wiki/Ballistic_missilehttps://en.wikipedia.org/wiki/Specific_impulsehttps://en.wikipedia.org/wiki/Hydrogenhttps://en.wikipedia.org/wiki/Oberth_effecthttps://en.wikipedia.org/wiki/Rocket_enginehttps://en.wikipedia.org/wiki/Rocket_engine#Terminologyhttps://en.wikipedia.org/wiki/Rocket_engine#Principle_of_operationhttps://en.wikipedia.org/wiki/Rocket_engine#Introducing_propellant_into_a_combustion_chamber


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o . 'ombustion chamber

o .A =ocket nozzles

o .1 Propellant efficiency

o .: /ack pressure and optimal epansion

o .5 !hrust vectoring

A Dverall performance

o A.$ 0pecific impulse

o A. Net thrust

o A.A ;acuum Isp

o A.1 !hrottling

o A.: 4nergy efficiency

o A.5 !hrust2to2weight ratio

1 'ooling

: (echanical issues

5 &coustic issues

o 5.$ 'ombustion instabilities

o 5. 4haust noise

< !esting

? 0afety

@ 'hemistry

$7 Ignition

$$ Plume physics
https://en.wikipedia.org/wiki/Rocket_engine#Combustion_chamberhttps://en.wikipedia.org/wiki/Rocket_engine#Rocket_nozzleshttps://en.wikipedia.org/wiki/Rocket_engine#Propellant_efficiencyhttps://en.wikipedia.org/wiki/Rocket_engine#Back_pressure_and_optimal_expansionhttps://en.wikipedia.org/wiki/Rocket_engine#Thrust_vectoringhttps://en.wikipedia.org/wiki/Rocket_engine#Overall_performancehttps://en.wikipedia.org/wiki/Rocket_engine#Specific_impulsehttps://en.wikipedia.org/wiki/Rocket_engine#Net_thrusthttps://en.wikipedia.org/wiki/Rocket_engine#Vacuum_Isphttps://en.wikipedia.org/wiki/Rocket_engine#Vacuum_Isphttps://en.wikipedia.org/wiki/Rocket_engine#Throttlinghttps://en.wikipedia.org/wiki/Rocket_engine#Energy_efficiencyhttps://en.wikipedia.org/wiki/Rocket_engine#Thrust-to-weight_ratiohttps://en.wikipedia.org/wiki/Rocket_engine#Coolinghttps://en.wikipedia.org/wiki/Rocket_engine#Mechanical_issueshttps://en.wikipedia.org/wiki/Rocket_engine#Acoustic_issueshttps://en.wikipedia.org/wiki/Rocket_engine#Combustion_instabilitieshttps://en.wikipedia.org/wiki/Rocket_engine#Exhaust_noisehttps://en.wikipedia.org/wiki/Rocket_engine#Testinghttps://en.wikipedia.org/wiki/Rocket_engine#Safetyhttps://en.wikipedia.org/wiki/Rocket_engine#Chemistryhttps://en.wikipedia.org/wiki/Rocket_engine#Ignitionhttps://en.wikipedia.org/wiki/Rocket_engine#Plume_physicshttps://en.wikipedia.org/wiki/Rocket_engine#Combustion_chamberhttps://en.wikipedia.org/wiki/Rocket_engine#Rocket_nozzleshttps://en.wikipedia.org/wiki/Rocket_engine#Propellant_efficiencyhttps://en.wikipedia.org/wiki/Rocket_engine#Back_pressure_and_optimal_expansionhttps://en.wikipedia.org/wiki/Rocket_engine#Thrust_vectoringhttps://en.wikipedia.org/wiki/Rocket_engine#Overall_performancehttps://en.wikipedia.org/wiki/Rocket_engine#Specific_impulsehttps://en.wikipedia.org/wiki/Rocket_engine#Net_thrusthttps://en.wikipedia.org/wiki/Rocket_engine#Vacuum_Isphttps://en.wikipedia.org/wiki/Rocket_engine#Throttlinghttps://en.wikipedia.org/wiki/Rocket_engine#Energy_efficiencyhttps://en.wikipedia.org/wiki/Rocket_engine#Thrust-to-weight_ratiohttps://en.wikipedia.org/wiki/Rocket_engine#Coolinghttps://en.wikipedia.org/wiki/Rocket_engine#Mechanical_issueshttps://en.wikipedia.org/wiki/Rocket_engine#Acoustic_issueshttps://en.wikipedia.org/wiki/Rocket_engine#Combustion_instabilitieshttps://en.wikipedia.org/wiki/Rocket_engine#Exhaust_noisehttps://en.wikipedia.org/wiki/Rocket_engine#Testinghttps://en.wikipedia.org/wiki/Rocket_engine#Safetyhttps://en.wikipedia.org/wiki/Rocket_engine#Chemistryhttps://en.wikipedia.org/wiki/Rocket_engine#Ignitionhttps://en.wikipedia.org/wiki/Rocket_engine#Plume_physics


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$ !ypes of rocket engines

o $.$ Physically powered

o $. 'hemically powered

o $.A 4lectrically powered

o $.1 !hermal

$.1.$ Preheated

$.1. 0olar thermal

$.1.A /eamed thermal

$.1.1 Nuclear thermal

o $.: Nuclear

$A Cistory of rocket engines

$1 0ee also

$: =eferences

$5 4ternal links

Terminology[edit]

Cere "rocket" is used as an abbreviation for "rocket engine".

Chemical roc'etsare powered by eothermicchemical reactions of the propellant.

Thermal roc'etsuse an inert propellant heated by a power source such as solaror nuclear

powerorbeamed energy.

Solid$)uel roc'ets)or solid$propellant roc'etsor motors* are chemical rockets which use

propellant in a solid state.

Liuid$propellant roc'etsuse one or more liquid propellants fed from tanks.
https://en.wikipedia.org/wiki/Rocket_engine#Types_of_rocket_engineshttps://en.wikipedia.org/wiki/Rocket_engine#Physically_poweredhttps://en.wikipedia.org/wiki/Rocket_engine#Chemically_poweredhttps://en.wikipedia.org/wiki/Rocket_engine#Electrically_poweredhttps://en.wikipedia.org/wiki/Rocket_engine#Thermalhttps://en.wikipedia.org/wiki/Rocket_engine#Preheatedhttps://en.wikipedia.org/wiki/Rocket_engine#Solar_thermalhttps://en.wikipedia.org/wiki/Rocket_engine#Beamed_thermalhttps://en.wikipedia.org/wiki/Rocket_engine#Nuclear_thermalhttps://en.wikipedia.org/wiki/Rocket_engine#Nuclearhttps://en.wikipedia.org/wiki/Rocket_engine#History_of_rocket_engineshttps://en.wikipedia.org/wiki/Rocket_engine#See_alsohttps://en.wikipedia.org/wiki/Rocket_engine#Referenceshttps://en.wikipedia.org/wiki/Rocket_engine#External_linkshttps://en.wikipedia.org/w/index.php?title=Rocket_engine&action=edit&section=1https://en.wikipedia.org/wiki/Exothermichttps://en.wikipedia.org/wiki/Thermal_rockethttps://en.wikipedia.org/wiki/Solar_thermal_rockethttps://en.wikipedia.org/wiki/Nuclear_thermal_rockethttps://en.wikipedia.org/wiki/Nuclear_thermal_rockethttps://en.wikipedia.org/wiki/Beamed_propulsionhttps://en.wikipedia.org/wiki/Solid-fuel_rockethttps://en.wikipedia.org/wiki/Liquid-propellant_rockethttps://en.wikipedia.org/wiki/Rocket_engine#Types_of_rocket_engineshttps://en.wikipedia.org/wiki/Rocket_engine#Physically_poweredhttps://en.wikipedia.org/wiki/Rocket_engine#Chemically_poweredhttps://en.wikipedia.org/wiki/Rocket_engine#Electrically_poweredhttps://en.wikipedia.org/wiki/Rocket_engine#Thermalhttps://en.wikipedia.org/wiki/Rocket_engine#Preheatedhttps://en.wikipedia.org/wiki/Rocket_engine#Solar_thermalhttps://en.wikipedia.org/wiki/Rocket_engine#Beamed_thermalhttps://en.wikipedia.org/wiki/Rocket_engine#Nuclear_thermalhttps://en.wikipedia.org/wiki/Rocket_engine#Nuclearhttps://en.wikipedia.org/wiki/Rocket_engine#History_of_rocket_engineshttps://en.wikipedia.org/wiki/Rocket_engine#See_alsohttps://en.wikipedia.org/wiki/Rocket_engine#Referenceshttps://en.wikipedia.org/wiki/Rocket_engine#External_linkshttps://en.wikipedia.org/w/index.php?title=Rocket_engine&action=edit&section=1https://en.wikipedia.org/wiki/Exothermichttps://en.wikipedia.org/wiki/Thermal_rockethttps://en.wikipedia.org/wiki/Solar_thermal_rockethttps://en.wikipedia.org/wiki/Nuclear_thermal_rockethttps://en.wikipedia.org/wiki/Nuclear_thermal_rockethttps://en.wikipedia.org/wiki/Beamed_propulsionhttps://en.wikipedia.org/wiki/Solid-fuel_rockethttps://en.wikipedia.org/wiki/Liquid-propellant_rocket


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/!brid roc'etsuse a solid propellant in the combustion chamber to which a second liquid or

gas oidiseror propellant is added to permit combustion.

%onopropellant roc'etsuse a single propellant decomposed by a catalyst. !he most common

monopropellants arehydrazineand hydrogen peroide.

Principle of operation[edit]

=ocket engines produce part of their thrust due to unopposed pressure on the combustion

chamber

=ocket engines produce thrust by the epulsion of ehaust which has been accelerated to a high2

speed.

!he ehaust must be a fluid usually a gas created by high pressure )$7277bar* combustion of

solid or liquidpropellants consisting of fueland oidisercomponents within a combustion

chamber.)&n eception iswater rockets which use water pressurised by compressed air carbon

dioide nitrogen or manual pumping.*

!he ehaust is then passed through a supersonicpropelling nozzlewhich uses heat energy of the

gas to accelerate the ehaust to very high speed and the reaction to this pushes the engine in the

opposite direction.

In rocket engines high temperatures and pressures are highly desirable for good performance asthis permits a longer nozzle to be fitted to the engine which gives higher ehaust speeds as well

as giving better thermodynamic efficiency.
https://en.wikipedia.org/wiki/Hybrid_rockethttps://en.wikipedia.org/wiki/Oxidizing_agenthttps://en.wikipedia.org/wiki/Monopropellant_rockethttps://en.wikipedia.org/wiki/Monopropellant_rockethttps://en.wikipedia.org/wiki/Catalysthttps://en.wikipedia.org/wiki/Hydrazinehttps://en.wikipedia.org/wiki/Hydrogen_peroxidehttps://en.wikipedia.org/wiki/Hydrogen_peroxidehttps://en.wikipedia.org/w/index.php?title=Rocket_engine&action=edit&section=2https://en.wikipedia.org/wiki/Fluidhttps://en.wikipedia.org/wiki/Bar_(unit)https://en.wikipedia.org/wiki/Rocket_propellanthttps://en.wikipedia.org/wiki/Rocket_propellanthttps://en.wikipedia.org/wiki/Fuelhttps://en.wikipedia.org/wiki/Oxidizing_agenthttps://en.wikipedia.org/wiki/Combustion_chamberhttps://en.wikipedia.org/wiki/Combustion_chamberhttps://en.wikipedia.org/wiki/Combustion_chamberhttps://en.wikipedia.org/wiki/Water_rockethttps://en.wikipedia.org/wiki/Water_rockethttps://en.wikipedia.org/wiki/Carbon_dioxidehttps://en.wikipedia.org/wiki/Carbon_dioxidehttps://en.wikipedia.org/wiki/Nitrogenhttps://en.wikipedia.org/wiki/Propelling_nozzlehttps://en.wikipedia.org/wiki/Hybrid_rockethttps://en.wikipedia.org/wiki/Oxidizing_agenthttps://en.wikipedia.org/wiki/Monopropellant_rockethttps://en.wikipedia.org/wiki/Catalysthttps://en.wikipedia.org/wiki/Hydrazinehttps://en.wikipedia.org/wiki/Hydrogen_peroxidehttps://en.wikipedia.org/w/index.php?title=Rocket_engine&action=edit&section=2https://en.wikipedia.org/wiki/Fluidhttps://en.wikipedia.org/wiki/Bar_(unit)https://en.wikipedia.org/wiki/Rocket_propellanthttps://en.wikipedia.org/wiki/Fuelhttps://en.wikipedia.org/wiki/Oxidizing_agenthttps://en.wikipedia.org/wiki/Combustion_chamberhttps://en.wikipedia.org/wiki/Combustion_chamberhttps://en.wikipedia.org/wiki/Water_rockethttps://en.wikipedia.org/wiki/Carbon_dioxidehttps://en.wikipedia.org/wiki/Carbon_dioxidehttps://en.wikipedia.org/wiki/Nitrogenhttps://en.wikipedia.org/wiki/Propelling_nozzle


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Introducin( propellant into a combustion chamberOedit

=ocket propellant is mass that is stored usually in some form of propellant tank prior to being

ejected from a rocket engine in the form of a fluid jet to produce thrust.

'hemical rocket propellants are most commonly used which undergo eothermic chemical

reactions which produce hot gas which is used by a rocket for propulsive purposes. &lternatively

a chemically inert reaction masscan be heated using a high2energy power source via a heat

echanger and then no combustion chamber is used.

& solid rocket motor.

0olid rocketpropellants are prepared as a miture of fuel and oidising components called -grain-

and the propellant storage casing effectively becomes the combustion chamber. ,iquid2fuelled

rocketstypically pump separate fuel and oidiser components into the combustion chamber

where they mi and burn. Cybrid rocketengines use a combination of solid and liquid or gaseous

propellants. /oth liquid and hybrid rockets use in7ectorsto introduce the propellant into the

chamber. !hese are often an array of simplejets2 holes through which the propellant escapes

under pressure% but sometimes may be more comple spray nozzles. +hen two or more

propellants are injected the jets usually deliberately cause the propellants to collide as this

breaks up the flow into smaller droplets that burn more easily.

Combustion chamberOedit

Main article8 (om5ustion cham5er

#or chemical rockets the combustion chamber is typically just a cylinder andflame holdersare

rarely used. !he dimensions of the cylinder are such that the propellant is able to combust

thoroughly% different rocket propellantsrequire different combustion chamber sizes for this to

occur. !his leads to a number called H
https://en.wikipedia.org/w/index.php?title=Rocket_engine&action=edit&section=3https://en.wikipedia.org/wiki/Reaction_masshttps://en.wikipedia.org/wiki/Solid_rockethttps://en.wikipedia.org/wiki/Liquid_rockethttps://en.wikipedia.org/wiki/Liquid_rockethttps://en.wikipedia.org/wiki/Hybrid_rockethttps://en.wikipedia.org/wiki/Liquid-fuel_rocket#Injectorshttps://en.wikipedia.org/wiki/Jet_(nozzle)https://en.wikipedia.org/wiki/Jet_(nozzle)https://en.wikipedia.org/w/index.php?title=Rocket_engine&action=edit&section=4https://en.wikipedia.org/wiki/Combustion_chamberhttps://en.wikipedia.org/wiki/Flame_holderhttps://en.wikipedia.org/wiki/Flame_holderhttps://en.wikipedia.org/wiki/Rocket_propellanthttps://en.wikipedia.org/w/index.php?title=Rocket_engine&action=edit&section=3https://en.wikipedia.org/wiki/Reaction_masshttps://en.wikipedia.org/wiki/Solid_rockethttps://en.wikipedia.org/wiki/Liquid_rockethttps://en.wikipedia.org/wiki/Liquid_rockethttps://en.wikipedia.org/wiki/Hybrid_rockethttps://en.wikipedia.org/wiki/Liquid-fuel_rocket#Injectorshttps://en.wikipedia.org/wiki/Jet_(nozzle)https://en.wikipedia.org/w/index.php?title=Rocket_engine&action=edit&section=4https://en.wikipedia.org/wiki/Combustion_chamberhttps://en.wikipedia.org/wiki/Flame_holderhttps://en.wikipedia.org/wiki/Rocket_propellant


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whereH

is the volume of the chamber

is the area of the throat

,J is typically in the range of :L57 inches )7.51L$.: m*.

!he combination of temperatures and pressures typically reached in a combustion chamber is

usually etreme by any standards. Gnlike in airbreathing jet engines no atmospheric

nitrogen is present to dilute and cool the combustion and the temperature can reach

true stoichiometricratios. !his in combination with the high pressures means that the rate ofheat conduction through the walls is very high.

#oc'et nozzlesOedit

Main article8*oc)et en%ine no99le

!ypical temperatures )!* and pressures )p* and speeds )v* in a 9e ,aval Nozzle

!he large bell or cone shaped epansion nozzle gives a rocket engine its characteristic shape.
https://en.wikipedia.org/wiki/Airbreathing_jet_enginehttps://en.wikipedia.org/wiki/Stoichiometrichttps://en.wikipedia.org/w/index.php?title=Rocket_engine&action=edit&section=5https://en.wikipedia.org/wiki/Rocket_engine_nozzlehttps://en.wikipedia.org/wiki/Airbreathing_jet_enginehttps://en.wikipedia.org/wiki/Stoichiometrichttps://en.wikipedia.org/w/index.php?title=Rocket_engine&action=edit&section=5https://en.wikipedia.org/wiki/Rocket_engine_nozzle


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In rockets the hot gas produced in the combustion chamber is permitted to escape from the

combustion chamber through an opening )the "throat"* within a high epansion2ratio -de

,aval- nozzle.

+hen sufficient pressure is provided to the nozzle )about .:2A above ambient pressure* thenozzlecho)esand a supersonic jet is formed dramatically accelerating the gas converting

most of the thermal energy into kinetic energy.

!he ehaust speeds vary depending on the epansion ratio the nozzle is designed to give but

ehaust speeds as high as ten times thespeed of sound at sea level airare not uncommon.

&bout half of the rocket engine-s thrust comes from the unbalanced pressures inside the

combustion chamber and the rest comes from the pressures acting against the inside of the

nozzle )see diagram*. &s the gas epands )adiabatically* the pressure against the nozzle-s

walls forces the rocket engine in one direction while accelerating the gas in the other.

Propellant e))icienc!Oedit

:ee also8 :pecific impulse

#or a rocket engine to be propellant efficient it is important that the maimum pressures

possible be created on the walls of the chamber and nozzle by a specific amount of

propellant% as this is the source of the thrust. !his can be achieved by all ofH

heating the propellant to as high a temperature as possible )using a high energy fuel

containing hydrogen and carbon and sometimes metals such as aluminiumor even using

nuclear energy*

using a low specific density gas )as hydrogen rich as possible*

using propellants which are or decompose to simple molecules with few degrees of

freedom to maimise translational velocity

=ocket thrust is caused by pressures acting in the combustion chamber and nozzle. #rom

Newton-s third law equal and opposite pressures act on the ehaust and this accelerates it to

high speeds.
https://en.wikipedia.org/wiki/Rocket_engine_nozzleshttps://en.wikipedia.org/wiki/Rocket_engine_nozzleshttps://en.wikipedia.org/wiki/Choked_flowhttps://en.wikipedia.org/wiki/Choked_flowhttps://en.wikipedia.org/wiki/Speed_of_soundhttps://en.wikipedia.org/wiki/Speed_of_soundhttps://en.wikipedia.org/wiki/Adiabatic_processhttps://en.wikipedia.org/w/index.php?title=Rocket_engine&action=edit&section=6https://en.wikipedia.org/wiki/Specific_impulsehttps://en.wikipedia.org/wiki/Aluminiumhttps://en.wikipedia.org/wiki/Aluminiumhttps://en.wikipedia.org/wiki/Rocket_engine_nozzleshttps://en.wikipedia.org/wiki/Rocket_engine_nozzleshttps://en.wikipedia.org/wiki/Choked_flowhttps://en.wikipedia.org/wiki/Speed_of_soundhttps://en.wikipedia.org/wiki/Adiabatic_processhttps://en.wikipedia.org/w/index.php?title=Rocket_engine&action=edit&section=6https://en.wikipedia.org/wiki/Specific_impulsehttps://en.wikipedia.org/wiki/Aluminium


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0ince all of these things minimise the mass of the propellant used and since pressure is

proportional to the mass of propellant present to be accelerated as it pushes on the engine

and since fromNewton-s third lawthe pressure that acts on the engine also reciprocally acts

on the propellant it turns out that for any given engine the speed that the propellant leaves

the chamber is unaffected by the chamber pressure )although the thrust is proportional*.Cowever speed is significantly affected by all three of the above factors and the ehaust

speed is an ecellent measure of the engine propellant efficiency. !his is termed e$haust

velocit& and after allowance is made for factors that can reduce it the e))ecti0e ehaust

0elocit!is one of the most important parameters of a rocket engine )although weight cost

ease of manufacture etc. are usually also very important*.

#or aerodynamic reasons the flow goes sonic )"chokes"* at the narrowest part of the nozzle

the -throat-. 0ince thespeed of soundin gases increases with the square root of temperature

the use of hot ehaust gas greatly improves performance. /y comparison at room

temperature the speed of sound in air is about A17 mEs while the speed of sound in the hot

gas of a rocket engine can be over $


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!o maintain this ideal of equality between the ehaust-s eit pressure and the ambient

pressure the diameter of the nozzle would need to increase with altitude giving the pressure

a longer nozzle to act on )and reducing the eit pressure and temperature*. !his increase is

difficult to arrange in a lightweight fashion although is routinely done with other forms of

jet engines. In rocketry a lightweight compromise nozzle is generally used and somereduction in atmospheric performance occurs when used at other than the -design altitude- or

when throttled. !o improve on this various eotic nozzle designs such as theplug

nozzlestepped nozzles the epanding nozzleand the aerospikehave been proposed each

providing some way to adapt to changing ambient air pressure and each allowing the gas to

epand further against the nozzle giving etra thrust at higher altitudes.

+hen ehausting into a sufficiently low ambient pressure )vacuum* several issues arise. Dne

is the sheer weight of the nozzlebeyond a certain point for a particular vehicle the etra

weight of the nozzle outweighs any performance gained. 0econdly as the ehaust gases

adiabatically epand within the nozzle they cool and eventually some of the chemicals can

freeze producing -snow- within the jet. !his causes instabilities in the jet and must be

avoided.

Dn a 9e ,aval nozzle ehaust gas flow detachment will occur in a grossly over2epanded

nozzle. &s the detachment point will not be uniform around the ais of the engine a side

force may be imparted to the engine. !his side force may change over time and result in

control problems with the launch vehicle.

Thrust 0ectorin(OeditMain article8 Thrust vectorin%

;ehicles typically require the overall thrust to change direction over the length of the burn. &

number of different ways to achieve this have been flownH

!he entire engine is mounted on a hingeor gimbaland any propellant feeds reach the

engine via low pressure fleible pipes or rotary couplings.

Fust the combustion chamber and nozzle is gimballed the pumps are fied and high

pressure feeds attach to the engine.

(ultiple engines )often canted at slight angles* are deployed but throttled to give the

overall vector that is required giving only a very small penalty.

Cigh2temperature vanes protrude into the ehaust and can be tilted to deflect the jet.
https://en.wikipedia.org/wiki/Plug_nozzlehttps://en.wikipedia.org/wiki/Plug_nozzlehttps://en.wikipedia.org/wiki/Stepped_nozzleshttps://en.wikipedia.org/wiki/Stepped_nozzleshttps://en.wikipedia.org/wiki/Expanding_nozzlehttps://en.wikipedia.org/wiki/Aerospike_enginehttps://en.wikipedia.org/wiki/De_Laval_nozzlehttps://en.wikipedia.org/w/index.php?title=Rocket_engine&action=edit&section=8https://en.wikipedia.org/wiki/Thrust_vectoringhttps://en.wikipedia.org/wiki/Hingehttps://en.wikipedia.org/wiki/Gimbalhttps://en.wikipedia.org/wiki/Gimbalhttps://en.wikipedia.org/wiki/Plug_nozzlehttps://en.wikipedia.org/wiki/Plug_nozzlehttps://en.wikipedia.org/wiki/Stepped_nozzleshttps://en.wikipedia.org/wiki/Expanding_nozzlehttps://en.wikipedia.org/wiki/Aerospike_enginehttps://en.wikipedia.org/wiki/De_Laval_nozzlehttps://en.wikipedia.org/w/index.php?title=Rocket_engine&action=edit&section=8https://en.wikipedia.org/wiki/Thrust_vectoringhttps://en.wikipedia.org/wiki/Hingehttps://en.wikipedia.org/wiki/Gimbal


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Overall performance[edit]

=ocket technology can combine very high thrust )meganewtons* very high ehaust speeds

)around $7 times the speed of sound in air at sea level* and very high thrustEweight ratios

)K$77*simultaneousl&as well as being able to operate outside the atmosphere and whilepermitting the use of low pressure and hence lightweight tanks and structure.

=ockets can be further optimised to even more etreme performance along one or more of

these aes at the epense of the others.

Speci)ic impulseOedit

Main article8 :pecific impulse

T!pical per)ormances o) common propellants

Propellant mi2acuum Isp

3seconds4

-))ecti0e ehaust

0elocit! 3m5s4

liuid o!(en5

liuid h!dro(en1:: 115

liuid o!(en5

'erosene3#P$14A:?Ocitation needed A:$7

nitro(en tetroide5

h!drazineA11Ocitation needed AA5@

n.b. &ll performances at a nozzle epansion ratio of 17

!he most important metric for the efficiency of a rocket engine is impulseper unit

ofpropellant this is called specific impulse)usually written *. !his is either measured as

a speed )the effective e$haust velocit& in metresEsecond or ftEs* or as a time )seconds*. &n

engine that gives a large specific impulse is normally highly desirable.
https://en.wikipedia.org/w/index.php?title=Rocket_engine&action=edit&section=9https://en.wikipedia.org/wiki/Meganewtonhttps://en.wikipedia.org/w/index.php?title=Rocket_engine&action=edit&section=10https://en.wikipedia.org/wiki/Specific_impulsehttps://en.wikipedia.org/wiki/Liquid_oxygenhttps://en.wikipedia.org/wiki/Liquid_hydrogenhttps://en.wikipedia.org/wiki/Kerosenehttps://en.wikipedia.org/wiki/RP-1https://en.wikipedia.org/wiki/Wikipedia:Citation_neededhttps://en.wikipedia.org/wiki/Wikipedia:Citation_neededhttps://en.wikipedia.org/wiki/Nitrogen_tetroxidehttps://en.wikipedia.org/wiki/Hydrazinehttps://en.wikipedia.org/wiki/Wikipedia:Citation_neededhttps://en.wikipedia.org/wiki/Wikipedia:Citation_neededhttps://en.wikipedia.org/wiki/Impulse_(physics)https://en.wikipedia.org/wiki/Impulse_(physics)https://en.wikipedia.org/wiki/Propellanthttps://en.wikipedia.org/wiki/Specific_impulsehttps://en.wikipedia.org/wiki/Specific_impulsehttps://en.wikipedia.org/w/index.php?title=Rocket_engine&action=edit&section=9https://en.wikipedia.org/wiki/Meganewtonhttps://en.wikipedia.org/w/index.php?title=Rocket_engine&action=edit&section=10https://en.wikipedia.org/wiki/Specific_impulsehttps://en.wikipedia.org/wiki/Liquid_oxygenhttps://en.wikipedia.org/wiki/Liquid_hydrogenhttps://en.wikipedia.org/wiki/Kerosenehttps://en.wikipedia.org/wiki/RP-1https://en.wikipedia.org/wiki/Wikipedia:Citation_neededhttps://en.wikipedia.org/wiki/Nitrogen_tetroxidehttps://en.wikipedia.org/wiki/Hydrazinehttps://en.wikipedia.org/wiki/Wikipedia:Citation_neededhttps://en.wikipedia.org/wiki/Impulse_(physics)https://en.wikipedia.org/wiki/Propellanthttps://en.wikipedia.org/wiki/Specific_impulse


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!he specific impulse that can be achieved is primarily a function of the propellant mi )and

ultimately would limit the specific impulse* but practical limits on chamber pressures and

the nozzle epansion ratios reduce the performance that can be achieved.

Net thrustOedit

Main article8 Thrust

/elow is an approimate equation for calculating the net thrust of a rocket engineH O

whereH

8 ehaust gas mass flow

8 effective ehaust velocity

8 actual jet velocity at nozzle eit plane

8 flow area at nozzle eit plane )or the plane where the jet leaves the nozzle if

separated flow*

8 static pressure at nozzle eit plane

8 ambient )or atmospheric* pressure

0ince unlike a jet engine a conventional rocket motor lacks an air intake there is no-ram drag- to deduct from the gross thrust. 'onsequently the net thrust of a rocket

motor is equal to the gross thrust )apart from static back pressure*.

!he term represents the momentum thrust which remains constant at a

given throttle setting whereas the term represents the pressure

thrust term. &t full throttle the net thrust of a rocket motor improves slightly with

increasing altitude because as atmospheric pressure decreases with altitude the

pressure thrust term increases. &t the surface of the 4arth the pressure thrust may be

reduced by up to A7depending on the engine design. !his reduction drops roughlyeponentially to zero with increasing altitude.

(aimum efficiency for a rocket engine is achieved by maimising the momentum

contribution of the equation without incurring penalties from over epanding the

ehaust. !his occurs when . 0ince ambient pressure changes with

altitude most rocket engines spend very little time operating at peak efficiency.
https://en.wikipedia.org/w/index.php?title=Rocket_engine&action=edit&section=11https://en.wikipedia.org/wiki/Thrusthttps://en.wikipedia.org/wiki/Rocket_engine#cite_note-2https://en.wikipedia.org/w/index.php?title=Rocket_engine&action=edit&section=11https://en.wikipedia.org/wiki/Thrusthttps://en.wikipedia.org/wiki/Rocket_engine#cite_note-2


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If the pressure of the ehaust jet varies from atmospheric pressure nozzles can be

said to be )top to bottom*H

.nderepanded

6mbient

+0erepanded

rossl! o0erepanded

If under or overepanded then loss of efficiency occurs. Brossly overepanded

nozzles lose less efficiency but can cause mechanical problems with the nozzle.

Cowever slightly overepanded nozzles will produce more thrust than critically

epanded nozzles if boundary layer separation does not occur. =ockets become

progressively more underepanded as they gain altitude. Note that almost all rocket

engines will be momentarily grossly overepanded during startup in an atmosphere.OA

2acuum IspOedit

9ue to the specific impulse varying with pressure a quantity that is easy to compare

and calculate with is useful. /ecause rockets chokeat the throat and because the

supersonic ehaust prevents eternal pressure influences travelling upstream it turns

out that the pressure at the eit is ideally eactly proportional to the propellant
https://en.wikipedia.org/wiki/Rocket_engine#cite_note-3https://en.wikipedia.org/w/index.php?title=Rocket_engine&action=edit&section=12https://en.wikipedia.org/wiki/Choked_flowhttps://en.wikipedia.org/wiki/Rocket_engine#cite_note-3https://en.wikipedia.org/w/index.php?title=Rocket_engine&action=edit&section=12https://en.wikipedia.org/wiki/Choked_flow


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flow provided the miture ratios and combustion efficiencies are maintained. It is

thus quite usual to rearrange the above equation slightlyHO1

and so define the vacuum ,spto beH

whereH

8 the speed of sound constant at the throat

8 the thrust coefficient constant of the nozzle )typically about *

&nd henceH

Throttlin(Oedit

=ockets can be throttled by controlling the propellant

combustion rate )usually measured in kgEs or lbEs*. In liquid

and hybrid rockets the propellant flow entering the chamber is

controlled using valves in solid rocketsit is controlled by

changing the area of propellant that is burning and this can be

designed into the propellant grain )and hence cannot be

controlled in real2time*.

=ockets can usually be throttled down to an eit pressure of

about one2third of ambient pressure )often limited by flow

separation in nozzles* and up to a maimum limit determined

only by the mechanical strength of the engine.

In practice the degree to which rockets can be throttled varies

greatly but most rockets can be throttled by a factor of without

great difficulty%Ocitation neededthe typical limitation is combustion

stability as for eample injectors need a minimum pressure to

avoid triggering damaging oscillations )chugging or combustion

instabilities*% but injectors can often be optimised and tested for

wider ranges. 0olid rockets can be throttled by using shaped
https://en.wikipedia.org/wiki/Rocket_engine#cite_note-4https://en.wikipedia.org/w/index.php?title=Rocket_engine&action=edit&section=13https://en.wikipedia.org/wiki/Solid_rockethttps://en.wikipedia.org/wiki/Solid_rockethttps://en.wikipedia.org/wiki/Wikipedia:Citation_neededhttps://en.wikipedia.org/wiki/Wikipedia:Citation_neededhttps://en.wikipedia.org/wiki/Rocket_engine#cite_note-4https://en.wikipedia.org/w/index.php?title=Rocket_engine&action=edit&section=13https://en.wikipedia.org/wiki/Solid_rockethttps://en.wikipedia.org/wiki/Wikipedia:Citation_needed


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grains that will vary their surface area over the course of the

burn.

-ner(! e))icienc!Oedit

Further information8*oc)et ; 3ner%& efficienc&

=ocket vehicle mechanical efficiency as a function of vehicle

instantaneous speed divided by effective ehaust speed. !hese

percentages need to be multiplied by internal engine efficiency to

get overall efficiency.

=ocket engine nozzles are surprisingly efficient heat enginesfor

generating a high speed jet as a consequence of the high

combustion temperature and highcompression ratio.=ocket

nozzles give an ecellent approimation to adiabatic

epansionwhich is a reversible process and hence they give

efficiencies which are very close to that of the'arnot cycle.

Biven the temperatures reached over 57 efficiency can be

achieved with chemical rockets.

#or a vehicleemploying a rocket engine the energetic efficiency

is very good if the vehicle speed approaches or somewhat

eceeds the ehaust velocity )relative to launch*% but at lowspeeds the energy efficiency goes to 7 at zero speed )as with

alljet propulsion.* 0ee =ocket energy efficiencyfor more details.

Thrust$to$&ei(ht ratioOedit

Main article8 thrust-to-6ei%ht ratio
https://en.wikipedia.org/w/index.php?title=Rocket_engine&action=edit&section=14https://en.wikipedia.org/wiki/Rocket#Energy_efficiencyhttps://en.wikipedia.org/wiki/Rocket#Energy_efficiencyhttps://en.wikipedia.org/wiki/Heat_engineshttps://en.wikipedia.org/wiki/Compression_ratiohttps://en.wikipedia.org/wiki/Compression_ratiohttps://en.wikipedia.org/wiki/Adiabatic_expansionhttps://en.wikipedia.org/wiki/Adiabatic_expansionhttps://en.wikipedia.org/wiki/Adiabatic_expansionhttps://en.wikipedia.org/wiki/Carnot_cyclehttps://en.wikipedia.org/wiki/Carnot_cyclehttps://en.wikipedia.org/wiki/Jet_propulsionhttps://en.wikipedia.org/wiki/Jet_propulsionhttps://en.wikipedia.org/wiki/Rocket#Energy_efficiencyhttps://en.wikipedia.org/w/index.php?title=Rocket_engine&action=edit&section=15https://en.wikipedia.org/wiki/Thrust-to-weight_ratiohttps://en.wikipedia.org/w/index.php?title=Rocket_engine&action=edit&section=14https://en.wikipedia.org/wiki/Rocket#Energy_efficiencyhttps://en.wikipedia.org/wiki/Heat_engineshttps://en.wikipedia.org/wiki/Compression_ratiohttps://en.wikipedia.org/wiki/Adiabatic_expansionhttps://en.wikipedia.org/wiki/Adiabatic_expansionhttps://en.wikipedia.org/wiki/Carnot_cyclehttps://en.wikipedia.org/wiki/Jet_propulsionhttps://en.wikipedia.org/wiki/Rocket#Energy_efficiencyhttps://en.wikipedia.org/w/index.php?title=Rocket_engine&action=edit&section=15https://en.wikipedia.org/wiki/Thrust-to-weight_ratio


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=ockets of all the jet engines indeed of essentially all engines

have the highest thrust to weight ratio. !his is especially true for

liquid rocket engines.

!his high performance is due to the small volume ofpressurevesselsthat make up the enginethe pumps pipes and

combustion chambers involved. !he lack of inlet duct and the

use of dense liquid propellant allows the pressurisation system to

be small and lightweight whereas duct engines have to deal with

air which has a density about one thousand times lower.

7etor #oc'et en(ine%ass

3'(4

%ass

3lb4

Thrust

3'N4

Th

3

=9271$7nuclear rocket

engineO:O5777 1177 A:.

F:?jet engine )0=2


45/102

7etor #oc'et en(ine%ass

3'(4

%ass

3lb4

Thrust

3'N4

Th

3

00(4rocket engine

)0pace 0huttle*O$A$


46/102

(ost other jet engines have gas turbines in the hot ehaust. 9ue

to their larger surface area they are harder to cool and hence

there is a need to run the combustion processes at much lower

temperatures losing efficiency. In addition duct

enginesO6hich


47/102

lower than the chamber the high temperatures seen there.

)0ee rocket nozzlesabove for temperatures in nozzle*.

In rockets the coolant methods includeH

$. uncooled )used for short runs mainly during testing*

. ablativewalls )walls are lined with a material that is

continuously vaporised and carried away*.

A. radiative cooling)the chamber becomes almost white hot

and radiates the heat away*

1. dump cooling )a propellant usually hydrogenis passed

around the chamber and dumped*

:. regenerative cooling)liquid rocketsuse the fuel or

occasionally the oidiser to cool the chamber via a

cooling jacket before being injected*

5. curtain cooling )propellant injection is arranged so the

temperature of the gases is cooler at the walls*


48/102

may be achieved by making the coolant velocityin the channels

as high as possible.

In practice regenerative cooling is nearly always used in

conjunction with curtain cooling andEor film cooling.

,iquid2fueled engines are often run fuel rich which lowers

combustion temperatures. !his reduces heat loads on the engine

and allows lower cost materials and a simplified cooling system.

Tet this can also increaseperformance by lowering the average

molecular weight of the ehaust and increasing the efficiency

with which combustion heat is converted to kinetic ehaust

energy.

Mechanical issues[edit]

=ocket combustion chambers are normally operated at fairly

high pressure typically $7277 bar )$ to 7 (Pa $:72A777 psi*.

+hen operated within significant atmospheric pressure higher

combustion chamber pressures give better performance by

permitting a larger and more efficient nozzle to be fitted without

it being grossly overepanded.

Cowever these high pressures cause the outermost part of thechamber to be under very large hoop stressesL rocket engines

arepressure vessels.

+orse due to the high temperatures created in rocket engines the

materials used tend to have a significantly lowered working

tensile strength.

In addition significant temperature gradients are set up in the

walls of the chamber and nozzle these cause differential

epansion of the inner liner that create internal stresses.

Acoustic issues[edit]

!he etreme vibration and acoustic environment inside a rocket

motor commonly result in peak stresses well above mean values
https://en.wikipedia.org/wiki/Velocityhttps://en.wikipedia.org/wiki/Air-fuel_ratiohttps://en.wikipedia.org/w/index.php?title=Rocket_engine&action=edit&section=17https://en.wikipedia.org/wiki/Hoop_stresshttps://en.wikipedia.org/wiki/Pressure_vesselhttps://en.wikipedia.org/wiki/Pressure_vesselhttps://en.wikipedia.org/wiki/Internal_stresseshttps://en.wikipedia.org/w/index.php?title=Rocket_engine&action=edit&section=18https://en.wikipedia.org/wiki/Velocityhttps://en.wikipedia.org/wiki/Air-fuel_ratiohttps://en.wikipedia.org/w/index.php?title=Rocket_engine&action=edit&section=17https://en.wikipedia.org/wiki/Hoop_stresshttps://en.wikipedia.org/wiki/Pressure_vesselhttps://en.wikipedia.org/wiki/Internal_stresseshttps://en.wikipedia.org/w/index.php?title=Rocket_engine&action=edit&section=18


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especially in the presence oforgan pipe2like resonances and gas

turbulence.Ocitation needed

Combustion instabilitiesOedit

!he combustion may display undesired instabilities of sudden or

periodic nature. !he pressure in the injection chamber may

increase until the propellant flow through the injector plate

decreases% a moment later the pressure drops and the flow

increases injecting more propellant in the combustion chamber

which burns a moment later and again increases the chamber

pressure repeating the cycle. !his may lead to high2amplitude

pressure oscillations often in ultrasonic range which may

damage the motor. Dscillations of U77 psi at : kCz were the

cause of failures of early versions of the !itan IImissile second

stage engines. !he other failure mode is a deflagration to

detonation transition% the supersonicpressure waveformed in the

combustion chamber may destroy the engine.O$


50/102

variation in thrust and the effects can vary from merely

annoying to actually damaging the payload or vehicle. 'hugging

can be minimised by using gas2filled damping tubes on feed

lines of high density propellants.

"uzzin(

!his can be caused due to insufficient pressure drop across the

injectors. It generally is mostly annoying rather than being

damaging. Cowever in etreme cases combustion can end up

being forced backwards through the injectors L this can cause

eplosions with monopropellants.

Screechin(

!his is the most immediately damaging and the hardest to

control. It is due to acoustics within the combustion chamber that

often couples to the chemical combustion processes that are the

primary drivers of the energy release and can lead to unstable

resonant "screeching" that commonly leads to catastrophic

failure due to thinning of the insulating thermal boundary layer.

&coustic oscillations can be ecited by thermal processes such

as the flow of hot air through a pipe or combustion in a chamber.

0pecifically standing acoustic waves inside a chamber can beintensified if combustion occurs more intensely in regions where

the pressure of the acoustic wave is maimal.O$?O$@O7O$0uch

effects are very difficult to predict analytically during the design

process and have usually been addressed by epensive time

consuming and etensive testing combined with trial and error

remedial correction measures.

0creeching is often dealt with by detailed changes to injectors or

changes in the propellant chemistry or vaporising the propellantbefore injection or use of Celmholtz damperswithin the

combustion chambers to change the resonant modes of the

chamber.

!esting for the possibility of screeching is sometimes done by

eploding small eplosive charges outside the combustion
https://en.wikipedia.org/wiki/Rocket_engine#cite_note-18https://en.wikipedia.org/wiki/Rocket_engine#cite_note-19https://en.wikipedia.org/wiki/Rocket_engine#cite_note-20https://en.wikipedia.org/wiki/Rocket_engine#cite_note-21https://en.wikipedia.org/wiki/Helmholtz_damperhttps://en.wikipedia.org/wiki/Rocket_engine#cite_note-18https://en.wikipedia.org/wiki/Rocket_engine#cite_note-19https://en.wikipedia.org/wiki/Rocket_engine#cite_note-20https://en.wikipedia.org/wiki/Rocket_engine#cite_note-21https://en.wikipedia.org/wiki/Helmholtz_damper


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chamber with a tube set tangentially to the combustion chamber

near the injectors to determine the engine-s impulse responseand

then evaluating the time response of the chamber pressure2 a fast

recovery indicates a stable system.

-haust noiseOedit

Main article8 acoustic si%nature

#or all but the very smallest sizes rocket ehaust compared to

other engines is generally very noisy. &s thehypersonicehaust

mies with the ambient air shock wavesare formed. !he 0pace

0huttlegenerates over 77 d/)&*of noise around its base.

!he 0aturn ;launch was detectable on seismometersa

considerable distance from the launch site.Ocitation needed!he sound

intensityfrom the shock waves generated depends on the size of

the rocket and on the ehaust velocity. 0uch shock waves seem

to account for the characteristic crackling and popping sounds

produced by large rocket engines when heard live. !hese noise

peaks typically overload microphones and audio electronics and

so are generally weakened or entirely absent in recorded or

broadcast audio reproductions. #or large rockets at close range

the acoustic ef

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