ROCKET THEORY - United States Air Force

AU Space Reference Guide Second Edition, 8/99

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Appendix E

ROCKET THEORY

Rocketry encompasses a wide range of topics, each of which takes many years of studyto master. This chapter provides an initial foundation toward the study of rocket theoryby addressing the physical laws governing motion/propulsion, rocket performanceparameters, rocket propulsion techniques, reaction masses (propellants), chemical rocketsand advanced propulsion techniques.

PROPULSION BACKGROUND

Rockets are like other forms ofpropulsion in that they expend energy toproduce a thrust force via an exchange ofmomentum with some reaction mass inaccordance with Newton’s Third Law ofMotion. But rockets differ from all otherforms of propulsion since they carry thereaction mass with them (self contained)and are, therefore, independent of theirsurrounding environment.

Other forms ofpropulsion depend on theirenvironment to provide thereaction mass. Cars usethe ground, airplanes usethe air, boats use the waterand sailboats use the wind.The rockets we are mostfamiliar with are chemicalrockets in which thepropellants (reaction mass)are the fuel and oxidizer.With chemical rockets, thepropellants are also theenergy source. A conventional chemicalrocket is a type of internal combustionengine burning fuel and oxidizer in acombustion chamber producing hot, highpressure gases and accelerating themthrough a nozzle. In electric and nuclearrockets, the propellant is essentially aninert mass.

According to Newton’s Second Law,the thrust force is equal to the rate ofchange of momentum of the ejected

matter, which depends on both how muchand how fast propellants are used (massflow rate) and the propellant’s speedwhen it leaves the rocket (effectiveexhaust velocity).

Like other forms of transportation,rockets consist of the same basic elementssuch as a structure providing the vehicleframework, propulsion system providingthe force for motion, energy source forpowering the vehicle systems, guidance

system for direction controland last and most important(indeed the reason forhaving the vehicle at all), thepayload. Examples ofpayloads are passengers,scientific instruments orsupplies. When a rocket isused as a weapon fordestructive purposes, we callit a missile; its payload is awarhead.

ROCKET PHYSICS

Sir Isaac Newton (Fig. 5-1) set forththe basic laws of motion; the means bywhich we analyze the rocket principle.Newton’s three laws of motion apply toall rocket-propelled vehicles. They applyto gas jets used for attitude control, smallrockets used for stage separations or fortrajectory corrections and to large rockets

Fig. 5-1. Sir Isaac Newton


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used to launch a vehicle from the surfaceof the Earth. They apply to nuclear,electric and other advanced types ofrockets as well as to chemical rockets.Newton’s laws of motion are statedbriefly as follows:

Newton’s 1st Law(Inertia)

Every body continues in a state ofuniform motion in a straight line,unless it is compelled to change thatstate by a force imposed upon it.

Newton’s 2nd Law(Momentum)

When a force is applied to a body,the time rate of change ofmomentum is proportional to, andin the direction of, the appliedforce.

Newton’s 3rd Law(Action—Reaction)

For every action there is a reactionthat is equal in magnitude butopposite in direction to the action.

In relating these laws to rocket theoryand propulsion, we can paraphrase andsimplify them. For example, the first lawsays, in effect, that the engines mustdevelop enough thrust force to overcomethe force of gravitational attractionbetween the Earth and the launch vehicle.The engines must be able to start thevehicle moving and accelerate it to thedesired velocity. Another way ofexpressing this for a vertical launch is tosay that the engines must develop morepounds of thrust than the vehicle weighs.

When applying the second law, wemust consider the summation of all theforces acting on the body; theaccelerating force is the net force actingon the vehicle. This means if we launch a200,000-lbf vehicle vertically from theEarth with a 250,000-lbf thrust engine,there is a net force at launch of 50,000-lbfthe difference between engine thrustand vehicle weight. Here the force of

gravity is acting opposite to the directionof the thrust of the engine.

As the rocket operates, the forcesacting on it change. The force of gravitydecreases as the vehicle’s mass decreases,and it also decreases with altitude. As therocket passes through the atmosphere,drag increases with increasing velocityand decreases with altitude (loweratmospheric density).1 As long as thethrust remains constant, the accelerationprofile changes with the changing forceson the vehicle. The predominate effect isthat the acceleration increases at anincreasing rate as the vehicle’s massdecreases.2

Figure 5-2 shows the generalacceleration and velocity profiles duringpowered flight. The acceleration andvelocity are low at launch due to the smallnet force and high vehicle mass at thattime. Both acceleration and velocity

increase rapidly as the engine burnspropellants (reducing vehicle mass andincreasing the net force).

At first stage burnout, the accelerationdrops (the acceleration at this point is dueto the environment: gravity and drag) andis generally opposite the direction of 1The term “Max Q” refers to the highest struc-tural pressure due to atmospheric drag.2As the net force on the vehicle increases and themass decreases, the acceleration increases at anincreasing rate.

Fig. 5-2. Acceleration and Velocity


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motion. With second stage ignition,acceleration and velocity will increaseagain. As the upper stage rocketengine(s) burn more propellants, rapidincreases in acceleration and velocityoccur. When the vehicle reaches thecorrect velocity (speed and direction) andaltitude for the mission, it terminatesthrust. Acceleration drops as the netforce on the vehicle is due to theenvironment, mainly gravity, after thrusttermination, or burnout, and the vehiclebegins free flight. For vehicles with three,four or more stages, similar changesappear in both the acceleration andvelocity each time staging occurs.Staging a vehicle increases the velocity insteps to the high values required for spacemissions.

Once a vehicle is in orbit, we say it is ina “weightless” condition. In fact, thevehicle is continually in free-fall, alwaysaccelerating toward the center of theEarth. The acceleration still depends onthe summation of the forces acting on thevehicle (or the net force).

In a free-fall condition, we don’t haveto continually counter act the force ofgravity, the vehicle’s momentumaccomplishes this task.3 In this“weightless” condition, even a very smallthrust (0.1 pound) operating over a longperiod of time can accelerate a vehicle togreat speeds, escape velocity and morefor interplanetary missions.

To relate Newton’s third law, or“action-reaction law” to rocket theoryand propulsion, consider what happens inthe rocket motor. All rockets developthrust by expelling particles (mass) at highvelocity from their nozzles. The effect ofthe ejected exhaust appears as a reactionforce, called thrust, acting in a directionopposite to the direction of the exhaust.The rocket is exchanging momentum withthe exhaust.

3When the vehicle’s orbit doesn’t intersect theEarth’s surface, we say the gravitational force isbalanced by the inertial force.

It is his Third Law of Motion thatexplains the working principle of allpropulsion systems.

A rocket engine is basically a devicefor expelling small particles of matter athigh speeds producing thrust through theexchange of momentum. When liquid orsolid chemicals are used as propellants,the exhaust consists of gas molecules.Recent scientific advances have involvedexperimental and theoretical work onrocket engines using ions (charged atomicparticles), nuclear particles and evenbeams of light (photons) as “propellants.”

Two items are necessary forpropulsion: matter and energy. Matter isthe reaction mass and is the source ofmomentum exchange. The reaction massbegins with the same momentum as therocket vehicle, but as the rocket expelsthis mass, the rocket and all remainingpropellants receive an equal increase inmomentum in the opposite direction.

It takes energy to accelerate thereaction mass (impart momentum). Thefaster propellants are accelerated, themore propulsive force achieved; however,it also takes more energy.

ROCKET PERFORMANCE

There are several rocket performanceparameters that, when taken together,describe a rocket’s overall performance:1) Thrust, 2) Specific Impulse, and 3)Mass Ratio.

Thrust (T)

The thrust is the amount of force anengine produces on the rocket (and on theexhaust stream leaving the rocket,conservation of momentum). The amountof thrust, along with the rocket mass,determines the acceleration. The missionprofile will determine the required andacceptable accelerations and thus, therequired thrust. Launching from theEarth typically requires a thrust to weightratio of at least 1.5 to 1.75. Once thevehicle is in orbit and the vehicle’smomentum balances the gravitational


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force, smaller thrust forces are usuallysufficient for any maneuvering.

Specific Impulse (Isp )

Specific impulse is a measure ofpropellant efficiency, and numerically isthe thrust produced divided by the weightof propellant consumed per second(ending up with units of seconds).

So, Isp is really another measure of arocket’s exhaust velocity. Specificimpulse is the common measure ofpropellant and propulsion systemperformance, and is somewhat analogousto the reciprocal of the specific fuelconsumption used with conventionalautomobile or aircraft engines. The largerthe value of specific impulse, the better arocket’s performance.

We can improve specific impulse byimparting more energy to the propellants(increasing the exhaust velocity), whichmeans that more thrust will be obtainedfor each pound of propellant consumed.We can think of specific impulse as thenumber of seconds for which one poundof propellant will produce one pound ofthrust. Or, we can think of it as theamount of thrust one pound of propellantwill produce for one second.

Mass Ratio (MR)

Since the rocket engine is continuallyconsuming propellants, the rocket’s massis decreasing with time. If the thrustremains constant, the vehicle’sacceleration increases reaching its highestvalue at engine cut-off; for example, thespace shuttle reaches 3 Gs just beforemain engine cut-off.

The purpose of a rocket is to place apayload at specified position with aspecific velocity. This position andvelocity depends on the mission. We canequate the energy needed to do this to thechange in velocity (or delta-v, ∆∆v) therocket imparts to the satellite. For arocket, the ideal ∆∆v gain depends on theIsp (exhaust velocity, ve ) and the massratio.

The more propellant the vehicle cancarry with respect to its “dry” weight, orweight without propellant aboard, thefaster it will be able to go. Mass ratio isan expression relating the propellant massto vehicle mass; the higher the mass ratio,the higher the final speed of the rocket.Therefore, a rocket vehicle is made toweigh as little as possible in its “dry”state. Increasing the weight of the vehiclepayload results in decreasing the massratio, and therefore cutting down themaximum altitude or range. For example,the addition of one pound of payload to ahigh-altitude sounding rocket may reduceits peak altitude by as much as 10,000feet.

PROPULSION TECHNIQUES

From our previous discussion of rocketperformance parameters, we see that wewould like to be as efficient as possible indeveloping thrust. To develop thrust, wehave to exchange momentum with somereaction mass (propellant). Any way thatwe can do this is a valid propulsionoption. We would like to choose theoption that decreased the overall missioncost while still providing for missionsuccess.

We are most familiar with chemicalrocket systems, however, there are otherways we can produce rocket propulsion.The two main ways of accelerating apropellant to provide thrust are:thermodynamic expansion and electro-static/ magnetic acceleration. Themethods for providing the thermal energyfor thermodynamic expansion, orelectricity for electrostatic acceleration,can come from chemical, nuclear, or solarsources.

Thermodynamic Expansion

Thermodynamic expansion is themechanism we are most familiar with. Allof our chemical systems use this methodto accelerate the propellants. However,we can also use nuclear or electricalenergy to heat the propellant.


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In thermodynamic expansion, we heatthe propellant to turn it into a highpressure, high temperature gas. We thenallow that gas to expand in a controlledway to turn the thermal potential energyinto directed kinetic energy, whichproduces thrust. The basic device used tocreate these large volumes of gas and toharness their heat energy is extremelysimple and often contains no movingparts.

The rocket engine usingthermodynamic expansion creates apressure difference between the thrustchamber (combustion chamber) and thesurrounding environment. It is thispressure difference that accelerates thegases.

A rocket engine usually operates atwhat the gas dynamist calls supercriticalconditionshigh chamber pressureexhausting to low external pressure. TheSwedish engineer Carl G.P. De Lavalshowed that for supercritical conditionsgases should be ducted through a nozzlethat converges to a throat (section ofsmallest area) and then diverges totransform as much of the gases’ thermalenergy into kinetic energy.

Nozzles

There are a number of nozzle types;Figure 5-3 depicts four of them. Theconical nozzle is simple and easy tofabricate and provides adequateperformance for most applications;however. it also has off axis exhaustvelocity components which reduces theefficiency. The radial velocitycomponents cancel and don’t contributeto the overall thrust, therefore the energy

going into the radial velocity is wasted.The contoured or bell-shaped nozzleprovides for rapid early expansionproducing shorter (less massive) nozzles,and redirects the exhaust toward the axialdirection near the nozzle exit. The plugand expansion-deflection type nozzles aremuch shorter than a conventional conicalnozzle with the same expansion ratio.

These nozzles have a center body andan annular chamber. The plug changesthe direction of the gas flow from thethroat during expansion from radial to anaxial direction. The expansion of exhaustgas is determined by ambient pressure. Avariation of the plug nozzle is theaerospike, which uses radial auxiliarycombustion chambers around the exit tothe main combustion chamber. Theexhaust plumes from the auxiliarychambers expand to form a "nozzle" forthe gases escaping from the engine. Overexpansion and under expansion can belargely compensated for by increasing ordecreasing the thrust of the auxiliarychambers.

Chemical Rockets

Chemical rockets are unique in that theenergy required to accelerate thepropellant comes from the propellantitself, and in this sense, are consideredenergy limited. Thus, the attainablekinetic energy per unit mass of propellantis limited primarily by the energy releasedin chemical reaction; the attainment ofhigh exhaust velocity requires the use ofhigh-energy propellant combinations thatproduce low molecular weight exhaustproducts. Currently, propellants with thebest combinations of high energy contentand low molecular weight seem capableof producing specific impulses in therange of 400 to 500 seconds or exhaustvelocities of 13,000 to 14,500 ft/sec.

Chemical rockets may use liquid orsolid propellants or, in some schemes,combinations of both. Liquid rocketsmay use one (monopropellant), two(bipropellant) or more propellants.Bipropellants consist of a combination of

Fig. 5-3. Nozzle Types


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a fuel (kerosene, alcohol, hydrogen) andan oxidizer (oxygen, nitric acid, fluorine).The liquids are held in tanks and fed intothe combustion chamber where they reactand then expand through the nozzle.

In contrast, solid propellants are anintimate mixture containing all thematerial necessary for reaction. Theentire block of solid propellant, called thegrain, is stored within the combustionchamber. Combustion proceeds from thesurface of the propellant.

A chemical rocket engine is little morethan a gas generator. The rapidcombination (combustion) of certainchemicals results in the release of energyand large volumes of gaseous products.The gas molecules generated haveconsiderable energy in the form of heat.In ordinary chemical rocket engines, thetemperature of the resulting gases can risehigher than 5,500 degrees Fahrenheit.

For chemical systems in general, liquidpropellants provide higher specificimpulses than solid propellants. We callliquid Hydrogen (LH) and liquid Oxygen(LOX) high energy propellants because ofthe large energy release duringcombustion and the high transfer ofthermal energy into directed kineticenergy of the exhaust stream.

An efficient LH/LOX burning engineproduces around Isp = 390-430 sec. onaverage.4 Solid propellant motorsproduce around Isp = 265-295 sec.

The total impulse of a rocket is theproduct of thrust and the effective firingduration. A typical shoulder launchedshort-range rocket may have an averagethrust of 660 pounds for an effectiveduration of 0.2 seconds, giving a totalimpulse of 132 lbf-sec. In contrast, theSaturn rocket had a total impulse of 1.14billion lbf-sec.

Nuclear Rockets

4The Isp of any particular engine depends upon itsdesign altitude. The Space Shuttle Main Engines(SSME) produce Isp = 363.2 @ sea level, and455.2 @ vacuum.

The nuclear rocket is an attempt toincrease specific impulse by using nuclearreaction to replace chemical reaction asthe energy source. The nuclear reactorgenerates thermal energy and heats thepropellant which is then expandedthrough a conventional nozzle.

Compared to the chemical rocket, thenuclear rocket has some advantages. Theenergy released in a nuclear reaction isvery much larger than that of a chemicalreaction (on the order of a million timeslarger), and since the energy source isseparate from the propellant, we have alarger latitude for propellant choice.Thus, hydrogen would be a goodpropellant because it has the lowestatomic weight, and would provide thehighest exhaust velocities for a givenchamber pressure and temperature.

We might think that the abundantenergy in nuclear rockets would meanthat we could employ indefinitely highchamber temperatures. This is definitelynot the case, however, since the heat istransferred from a solid reactor to thepropellant. Thus the structuralcomponents within the nuclear rocket,unlike those in a chemical rocket, must behotter than the propellant, and thetemperature cannot exceed the limitingtemperature of the structure or thereactor material. The attainabletemperatures in nuclear rockets to dateare considerably below the temperaturesattained in some chemical rockets, but theuse of hydrogen as the propellant morethan offsets this temperaturedisadvantage. Thus, as far as specificimpulse is concerned, the increasedperformance of nuclear rockets is entirelydue to the use of a propellant with a lowatomic weight. The nuclear fission rocketoffers roughly twice the specific impulseof the best chemical rocket (about 800-1,000 seconds), while delivering fairlyhigh thrusts for long periods of time.

One theoretical improvement is a high-density reactor using fast neutrons. Thistype of reactor is expected to producehigher performance levels in a smaller


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package than the thermal (or slow)reactors. Another improvement is a gascore reactor in which the operatingtemperature could be much higher. Thisincrease in temperature would occurbecause of the elimination of the solidcore of fuel elements used in slow andfast reactors. These structural elementsare temperature limited.

NASA’s Lewis Research Center ispursuing a concept for a reusable vehiclepropelled by a nuclear thermal rocket(NTR) to take astronauts to the Moonand back (Fig. 5-4). With the addition ofmodular hardware elements, the lunartransit vehicle would become the core ofa spacecraft to land astronauts on Marsearly in the 21st century.

Specific impulse has reached about 850seconds in nuclear engines, while the best

liquid oxygen/liquid hydrogen combustionengines only approach 475 seconds (in avacuum). Such a system could decreasetransit times to Mars from 9-15 monthsdown to 4-6 months, leaving more timefor exploration. Of course nuclear rocketshave drawbacks. Nuclear reactors are notonly heavy, but while in operation,produce large amounts of radiation. Themass and radiation hazard prohibit its useas a launch vehicle. However, once inspace the benefits on long range missionswould more than offset the extra mass.

Electrothermal Rockets

Another method using thermodynamicexpansion is the arcjet. The arcjet is anelectrothermal rocket because it useselectrical energy to heat a propellant. Inthis method, an annular arc is created in

the chamber and the propellant is heatedto high temperatures as it interacts withthe arc. After the heating, the propellantis expanded through a conventionalnozzle (Fig. 5-5).

This type of propulsion takesadvantage of using hydrogen as apropellant, and, like nuclear rockets,experiences a similar performance gain in

specific impulse (up to 1,200 seconds).Unlike nuclear rockets, arcjets are small,producing little more than several poundsof thrust.

Electrical Propulsion

Electrical and electromagnetic rocketsfundamentally differ from chemicalrockets with respect to their performancelimitations. Chemical rockets are energy-limited, since the quantity of energy islimited by the chemical behavior of thepropellants. If a separate energy source isused, much higher propellant energy ispossible. Further, if the temperaturelimitations of solid walls could be madeunimportant by direct electrostatic orelectromagnetic propellant accelerationwithout necessarily raising the fluidtemperature, there would be no limit tothe kinetic energy we could add to thepropellant. However, the rate ofconversion from nuclear or solar toelectrical energy and then to propellantkinetic energy is limited by the mass ofthe conversion equipment. Since thismass is likely to be a large portion of thetotal mass of the vehicle, the electricalrocket (including electrothermal/static/magnetic) is essentially power-limited.

Fig. 5-4. Reusable Rocket

Fig. 5-5. Conventional Nozzle


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Electrostatic/magnetic rockets convertelectrical energy directly to propellantkinetic energy without necessarily raisingthe temperature of the working fluid. Forthis reason the specific impulse is notlimited by the temperature limitations ofthe wall materials, and it is possible toachieve very high exhaust velocities,although at the cost of high powerconsumption.

Because of the massive energyconversion equipment, electrical rocketshave low thrust, perhaps only one-thousandth of vehicle weight in theEarth’s gravitational field. For thisreason, they are mainly restricted to spacemissions during which the gravitationalforces are very nearly balanced by inertialforces. Low accelerations are quiteacceptable, since the journeys are of longduration.

The propellant of an electrical rocketconsists of either discrete chargedparticles accelerated by electrostaticforces, or a stream of electricallyconducting fluid (plasma) accelerated byelectromagnetic forces.

Electrostatic Rockets

These are commonly called ionrockets. Neutral propellant is convertedto ions and electrons and withdrawn inseparate streams. The ions pass through astrong electrostatic field producedbetween acceleration electrodes. Theions accelerate to high speeds, and thethrust of the rocket is in reaction to theion acceleration (Fig. 5-6).

It is also necessary to expel theelectrons in order to prevent the vehiclefrom acquiring a net negative charge.Otherwise, ions would be attracted backto the vehicle and the thrust would vanish.They remove these excess electrons by re-injecting them back into the exhaust ionbeam.

Ion rockets offer very high specificimpulses (a typical figure being 10,000seconds with values ranging up to 20,000seconds), but very low thrust, one-halfpound being high. It has been estimated

that an ion rocket employing cesiumpropellant would require over 2,000 kWof electrical power per pound of thrust.

The propellant for ion engines may beany substance that ionizes easily. Unlikethermodynamic expansion, the size of themolecules is not a primary factor. Themost efficient elements are mercury,cesium or the noble gases.

Electromagnetic Rockets

There are three major types ofelectromagnetic rockets: magnetogas-dynamic, pulsed-plasma and traveling-wave. All methods use a plasma withcrossed electric and magnetic fields toaccelerate the plasma.

A plasma is an electrically conductinggas. It consists of a collection of neutralatoms, molecules, ions, and electrons.The number of ions and the number ofelectrons are equal so that, on the whole,the plasma is electrically neutral. Becauseof its ability to conduct electrons, theplasma can be subjected toelectromagnetic forces in much the sameway as solid conductors in electricmotors.

Magneto-gas-dynamic Drive. Strong external electric and magneticfields direct and accelerate the plasmastream, imparting high exhaust velocity.The performance is limited due to non-perpendicular currents flowing in theplasma at high field strengths. Thespecific impulse is lower than ion rocketsbut still very high (around 10,000

Fig. 5-6. Ion Acceleration


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seconds). The mass flow rate is restrictedso the thrusts remain low.

Pulsed-Plasma Accelerators.One of the disadvantages of the

steady crossed-field accelerators is thatthey require a substantial external fieldand therefore, a massive electromagnet.It is possible to make an accelerator forwhich an electromagnet is unnecessary byusing the plasma current itself to generatethe magnetic field, which gives rise to theaccelerating force. Whereas the crossed-field accelerator is analogous to a shuntmotor (which has separate current circuitsfor the electric and magnetic fields), theanalog of this type of accelerator is theseries motor in which the magnetic field isestablished by the same current whichinteracts to establish the crossed fieldforce.

Traveling-Wave A third type of plasma accelerator,

sometimes called the magnetic-inductionplasma motor, offers potential advantagesover both the foregoing accelerators. Itrequires neither magnets or electrodes,and relies on currents being induced in theplasma by a traveling magnetic wave.

If the current in a conductorsurrounding a tube containing a plasmaincreases, the magnetic field strength inthe plane of the conductor will increase.Then an electromotive force will beinduced in any loop in this plane. If theconductor current increases rapidlyenough, the induced electric field willestablish a substantial plasma current.The induced magnetic field and plasmacurrent then interact to cause a body forcenormal to both, which tends to compressthe plasma toward the axis of the tube andexpel it axially.

A traveling-wave accelerator makesuse of a number of sequentially energizedexternal conductors along the tube. Asthe switches are fired in turn, themagnetic field lines move axially along thetube, interacting with induced currentsand imparting axial motion to the plasma.

The inward radial force on the plasmain this accelerator appears to offer anadvantage in keeping the hightemperature plasma away from the solidwalls of the tube. The fact that noelectrodes are needed is also an attractivefeature.

STAGING

Currently, the only practical methodwe have for launching satellites is withchemical systems. As we found out in therocket performance section, specificimpulse and mass ratio limit our chemicalsystems’ performance.

What does this mean in terms ofsatellites and space probes? A rocket hasto provide enough energy, essentially25,000 ft/sec (17,500 mph), to orbit theEarth as a satellite and 36,700 ft/sec(25,000 mph) to escape the Earth’sgravitational field and become a planetoidcircling the Sun.

A body must attain a velocity of nearly35,000 ft/sec to hit the Moon. Nopractical rocket of one stage can reach thecritical velocities for satellites or spaceprobes.

A solution to this problem is to mountone or more rockets on top of oneanother and to fire them in succession atthe moment the previous stage burns out.For example, if each stage provides about9,000 ft/sec in velocity when fired asabove, it would take three stages to put asatellite in orbit, or four stages to reachthe moon or go beyond it into space as adeep space probe orbiting the sun.

Staging reduces the launch size andweight of the vehicle required for aspecific mission and aids in achieving thehigh velocities necessary for specificmissions.

Multistage rockets allow improvedpayload capability for vehicles with a high∆∆v requirement, such as launch vehiclesor interplanetary spacecraft. In amultistage rocket, propellant is stored insmaller, separate tanks rather than a largersingle tank as in a single-stage rocket.Since each tank is discarded when empty,


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energy is not expended to accelerate theempty tanks, thereby achieving a highertotal ∆∆v. Alternatively, a larger payloadmass can be accelerated to the same total∆∆v. The separate tanks are usuallybundled with their own engines, with eachdiscardable unit called a stage.

The same rocket equation describesmultistage and single-stage rocketperformance, but it must be applied on astage-by-stage basis. It is important torealize that the payload mass for any stageconsists of the mass of all subsequentstages plus the ultimate payload itself.The velocity of the multistage vehicle atthe end of powered flight is the sum ofvelocity increases produced by each of thevarious stages. We add the increasesbecause the upper stages start withvelocities imparted to them by the lowerstages.

A multistage vehicle with identicalspecific impulse, payload fraction andstructure fraction for each stage is said tohave similar stages. For such a vehicle,the payload fraction is maximized byhaving each stage provide the samevelocity increment. For a multistagevehicle with dissimilar stages, the overallvehicle payload fraction depends on howthe ∆∆v requirement is partitioned amongstages. Payload fractions will be reducedif the ∆∆v is partitioned suboptimally.

ROCKET PROPELLANTS

The type of rocket engine determinesthe corresponding type of propellantstorage and delivery systems. In the caseof chemical rocket engines, thepropellants may be either liquid or solid.

Rocket engines can operate oncommon fuels such as gasoline, alcohol,kerosene, asphalt or synthetic rubber, plusa suitable oxidizer. Engine designersconsider fuel and oxidizer combinationshaving the energy release and the physicaland handling properties needed fordesired performance. Selectingpropellants for a given mission requires acomplete analysis of mission, propellantperformance, density, storability, toxicity,

corrosiveness, availability and cost; sizeand structural weight of the vehicle; andpayload weight.

Liquid Propellants

The term “liquid propellant” refers toany of the liquid working fluids used in arocket engine. Normally, they are anoxidizer and a fuel, but may includecatalysts or additives that improveburning or thrust. Generally, liquidpropellants permit longer burning timethan solid propellants. In some cases, theypermit intermittent operations. That is,combustion can be stopped and started bycontrolling propellant flow.

Many combinations of liquidpropellants have been investigated.However, no combination has all thesedesirable characteristics:

• Large availability of raw materialsand ease of manufacture

• High heat of combustion per unit ofpropellant mixture

• Low freezing point (wide range ofoperation)

• High density before combustion(smaller tanks)

• Low density after combustion(higher γ)

• Low toxicity and corrosiveness(easier handling and storage)

• Low vapor pressure, good chemicalstability (simplified storage)

Liquid-propellant units can beclassified as monopropellant, bipropellantor tripropellant in nature (Fig. 5-7). Amonopropellant is a single liquidpossessing the qualities of both anoxidizer and a fuel. It may be a singlechemical compound, such asnitromethane, or a mixture of severalchemical compounds, such as hydrogenperoxide and alcohol. The compoundsare stable at ordinary temperatures andpressures, but decompose when heatedand pressurized, or when a catalyst startsthe reaction.


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Monopropellant rockets are simple,since they only need one propellant tankand the associated equipment. The mostcommon monopropellant systems usehydrazine. Bipropellant units carry fueland oxidizer in separate tanks and bringthem together in the combustion chamber.At present, most liquid rockets usebipropellants. In addition to a fuel andoxidizer, a liquid bipropellant may includea catalyst to increase the speed ofreaction, or other additives to improve thephysical, handling or storage properties.

A tripropellant has three compounds.The third compound improves the specificimpulse of the basic propellant.

Liquid propellants are commonlyclassified as either cryogenic or storablepropellants. A cryogenic propellant isone that has a very low boiling point andmust be kept very cold. For example,liquid oxygen boils at -297° F, liquidfluorine at -306° F and liquid hydrogen at-423° F. Personnel at the launch site loadthese propellants into a rocket as nearlaunch time as possible to reduce lossesfrom vaporization and to minimizeproblems caused by their lowtemperatures.

A storable propellant is one that isliquid at normal temperatures andpressures and may be left in a rocket fordays, months, or even years. Forexample, nitrogen tetroxide boils at 70° F,

unsymmetrical dimethyldrazine (UDMH)at 146° F and hydrazine at 236° F.However, the term storage refers tostoring propellants on Earth. It does notconsider the problem of storage in space.

As described earlier, in order to storethe liquid propellants within the rocketvehicle until such time as they areintroduced into the combustion chamberof the rocket engine, large tanks arerequired. Once combustion starts andpressure is built up inside the combustionchamber, the propellants will not flowinto the combustion chamber of their ownaccord. A method of forcing thepropellants into the combustion chamberagainst the combustion pressure isrequired. Two methods presently used toaccomplish this are shown in Figure 5-8.The simplest of these provides a gaspressure, usually helium, in the propellanttanks sufficient to force the propellantsout of the tanks through the deliverypiping and into the combustion chamber.

The pressurization method requirespropellant tanks that are strong enough towithstand the pressure and this, in turn,means thick tank walls and increasedtankage weight. This decreases the massratio. Therefore, there is a definite limit

to the size of the rocket vehicle that canuse the pressurization method.

The second method, as previouslydescribed, utilizes pumps to drain thepropellants from the tanks and force theminto the combustion chamber. Thisrequires a pump for each propellant aswell as some method of driving the

Fig. 5-7: Liquid Propellants

Fig. 5-8. Propellant Feed Types


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pumps. These pumps are usually thecentrifugal type. They are generallydriven by a turbine mounted on the samedrive shaft. The turbine, in turn, ispowered by a small gas generator thatmay use the decomposition of high-strength (highly concentrated) hydrogenperoxide to produce steam. Othersources of turbine power may be the tworocket propellants, burned in a smallauxiliary combustion chamber, or a smallsolid-propellant grain burned to producedriving gas. A novel method involvesbleeding some of the combustion gasfrom the rocket engine back to theturbine. This is a system which essentially“bootstraps” itself into operation. Pumpdelivery systems allow the use ofextremely thin-walled propellant tanks,which increases the possible mass ratio.

With liquid propellants, the combustionprocess starts when the propellants areinjected into the rocket engine. Thepropellants are driven into the combustionchamber through an “injector,” whichoften looks like an overgrown showerhead. The injector serves to break up thepropellants into atomized spray, thuspromoting mixing and completecombustion. Injectors are extremelydifficult to design, as there are nodefinitive mathematical equations thatanalyze their operation. Modern injectorsare built as a single unit that forms theforward end of the combustion chamber.They are perforated with hundreds of tinyholes, the number, size, and angle ofwhich are critical.

Propellants may be chosen so that theyreact spontaneously upon contact witheach other. Such propellants are knownas hypergolic and do not require a meansof ignition in order to get combustionstarted. Ignition for non-hypergolicpropellants requires an igniter. Ignitersare usually pyrotechnic in nature,although some engines have used sparkplugs.

Typical non-hypergolic combinationsare alcohol/LOX, gasoline/LOX, liquidhydrogen/LOX, alcohol and nitric acid,and kerosene (RP-1)/LOX. Typical

hypergolic combinations are aniline andnitric acid, fluorine/hydrazine, fluorineand hydrogen, hydrazine/hydrogenperoxide, and aniline and nitrogentetroxide.

Monopropellants are chemicals whichdecompose in the presence of a suitablecatalyst or at a suitable temperaturereleasing energy in the process.Hydrogen peroxide (75 percent pure orbetter) is a common monopropellant usedin many vehicles for small adjustment orvernier rockets. Such strong peroxidemixtures, however, must be handled withgreat care because they decompose withexplosive suddenness in the presence ofimpurities. Other monopropellants arenitro-methane (CH3NO2), ethylene oxide(C2H4O) and hydrazine (N2H4). Many ofthese propellants are highly unstable,many are highly toxic and some are both.

Liquid propellant engines are extremelyversatile, can be throttled, and can beused again by simply reprovisioning thepropellant tanks. They provide highspecific impulses, but are more complexand therefore, less reliable than a solidmotor.

While it is possible to argue endlesslyover the merits of both types, it is safe tosay that both solid-propellant motors andliquid-propellant engines will continue tobe used in the future for specificapplications where their respectiveadvantages outweigh their disadvantages.


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Solid Propellants

The solid-propellant motor (Fig. 5-9)is the oldest of all types and is by far thesimplest in construction. Since thepropellants are in solid form, usually

mixed together, and since a solid-propellant charge undergoes combustiononly on its surface, there is no need toinject it continuously into the combustionchamber from storage tanks. Solidpropellants are therefore, placed right inthe combustion chamber itself. A solidpropellant rocket motor combines boththe combustion chamber and thepropellant storage facilities in one unit. Asolid-propellant charge, or “grain,” isignited and burns until it is exhausted,changing the effective size and shapeduring its operation.

Since a solid-propellant grain burnsonly on its surface, the shape of the grainmay be designed to regulate the amountof grain area undergoing combustion.

Since the thrust is dependent upon themass flow rate, which is in turn dependentupon the amount of propellant beingconsumed per second, the thrust outputof a solid-propellant rocket motor can bedetermined in advance, or “programmed.”A grain that burns with constant areaduring the thrust period yields constantthrust and is known as a restricted orneutral-burning grain (It might, forexample, burn from the aft end to theforward end in the manner of a cigarette)(Fig. 5-10).

In addition, a grain may be designed toburn with increasing area and thrust(progressive) or with decreasing area andthrust (regressive). Choice of grain styledepends on the motor’s use.

There are many chemical combinationsthat make good solid propellants. Asidefrom gunpowder and metal-powdermixtures (such as zinc and sulfur) whichhave erratic burning rates and poorphysical properties, there are two classesof solid propellants which were originallydeveloped for rockets during and afterWorld War II and are in wide use today:double-base (homogeneous) andcomposite (heterogeneous) propellants.Double-based propellants consist chieflyof a blend of nitrocellulose andnitroglycerin with small quantities of salts,wax, coloring and organic compounds tocontrol burning rates and physicalproperties. The double-based propellantsmay be regarded as complex colloids withunstable molecular structure.Homogeneous propellants have oxidizerand fuel in a single molecule. The blastfrom a small chemical igniter easily startsthe rapid recombination of this structurein the process of burning. Aging allows aslower rearrangement of the molecules,and thus often significantly changes the

burning properties of the propellant.Double-based propellants can be formedefficiently in many shapes by eithercasting or extrusion through dies.

Fig. 5-10. Grain Configurations

Fig. 5-9. Solid Propellant Motor


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Composite propellants, as the nameimplies, are mixtures of an oxidizer,usually an inorganic salt such asammonium perchlorate, in a hydrocarbonfuel matrix, such as an asphalt likematerial. The fuel contains small particlesof oxidizer dispersed throughout. Thefuel is called a binder because the oxidizerhas no mechanical strength. Usually incrystalline form, finely ground oxidizer isapproximately 70 to 80 percent of thetotal propellant weight. Composites areusually cast to shape. Current work withcomposites and double-based propellantsincorporates light metals (such as boron,aluminum, and lithium), which yield veryhigh energies.

Although less energetic than goodliquid propellants (lower specificimpulse), solids have the advantages offast ignition (0.025 seconds is common)and good storability in the rocket.Making them, however, is costly,complex and dangerous.

An ideal solid propellant would possessthese characteristics:

• High release of chemical energy• Low molecular weight of

combustion products• High density before combustion• Readily manufactured from easily

obtainable substances by simpleprocesses

• Insensitive to shock andtemperature changes and nochemical or physical deteriorationwhile in storage

• Safe and easy to handle.• Ability to ignite and burn uniformly

over a wide range of operatingtemperatures

• Nonhygroscopic (nonabsorbent ofmoisture)

• Smokeless and flashless

It is improbable that any propellant willhave all of these characteristics.Propellants used today possess some ofthese features at the expense of others,depending upon the application and thedesired performance.

Propellant Tanks

The function of the propellant tanks issimply the storage of one or twopropellants until needed in the combustionchamber. Depending upon the kind ofpropellants used, the tank may be nothingmore than a low pressure envelope or itmay be a pressure vessel for containinghigh pressure propellants. In the case ofcryogenic propellants (described later),the tank has to be an exceptionally wellinsulated structure to keep propellantsfrom boiling away.

As with all rocket parts, weight of thepropellant tanks is an important factor intheir design. Many liquid propellant tanksare made out of very thin metal or are thinmetal sheaths wrapped with high-strengthfibers. These tanks are stabilized by theinternal pressure of their contents, muchthe same way balloon walls gain strengthfrom the gas inside. Very large tanks andtanks that contain cryogenic propellantsrequire additional strengthening or layers.Structural rings and ribs are used tostrengthen tank walls, giving the tanks theappearance of an aircraft frame. Withcryogenic propellants, extensive insulationis needed to keep the propellants in theirliquefied form. Even with the bestinsulation, cryogenic propellants aredifficult to keep for long periods of timeand will boil away. For this reason,cryogenic propellants are usually not usedwith military rockets/ missiles.

The propellant tanks of the shuttle canbe used as an example of the complexitiesinvolved in propellant tank design. Theexternal tank (ET) consists of two smallertanks and an intertank. The ET is thestructural back bone of the shuttle andduring launch it must bear the entirethrust produced by the solid rocketboosters and the Orbiter main engine.

The forward or nose tank containsLOX. Antislosh and antivortex bafflesare installed inside the LOX tank as wellas inside the other tank to prevent gasbubbles inside the tank from beingpumped to the engines along with the


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propellants. Many rings and ribsstrengthen this tank.

The second tank contains LH. Thistank is two and a half times the size of theLOX tank. However, the LH tank weighsonly one third as much as the LOX tankbecause LOX is 16 times denser than LH.

Between the two tanks is an intertankstructure. The intertank is not actually atank but a mechanical connection betweenthe LOX and LH tanks. Its primaryfunction is to join the two tanks togetherand distribute thrust loads from the solidrocket boosters. The intertank alsohouses a variety of instruments.

Turbopumps

Turbopumps provide the required flowof propellants from the low-pressurepropellant tanks to the high-pressurerocket chamber. Power for the pumps isproduced by combusting a fraction of thepropellants in a preburner. Expandinggases from the burning propellants driveone or more turbines which, in turn, drivethe turbopumps. After passing throughthe turbines, exhaust gases are eitherdirected out of the rocket through anozzle or are injected, along with liquidoxygen into the chamber for morecomplete burning.

Combustion Chamber and Nozzle

The combustion chamber of a liquidpropellant rocket is a bottle-shapedcontainer with openings at opposite ends.The openings at the top inject thepropellants into the chamber. Eachopening consists of a small nozzle thatinjects either fuel or oxidizer. The mainpurpose of the injectors is to mix thepropellants to ensure smooth and completecombustion with no detonations.Combustion chamber injectors come inmany designs.

The purpose of the nozzle is to providefor gas expansion to achieve themaximum transfer of thermal energy intodirected kinetic energy.

HYBRID ROCKETS

Another rocket engine should bementioned. Composite (hybrid) enginesare combinations of solid and liquidpropellant engines. In a compositeengine, the fuel may be in solid forminside the combustion chamber with theoxidizer in a liquid form that is injectedinto the chamber.

Though not in widespread use, they dooffer some advantages in rocketpropulsion. Figure 5-11 depicts asimplified structure of the hybrid system.

Theoretical work on hybrid propulsion

dates back to the 1930s in both the U.S.and Germany. In the 1940s, a hybridmotor was built that burned Douglas Firwood loaded with carbon black and waxin 10% liquid oxygen. Germany’swartime experiments tried powdered andre-formed coal fuel cores, but even cleancoal contained too many impurities to bea good rocket fuel. Work continued intothe 1960s with both the Navy and AirForce funding research.

The hybrid fuel burns only on contactwith the oxidizer, and cracks in the fuelgrain do not admit enough oxidizer tosupport catastrophic failures common tosolids. Also, unlike conventional solids,the flow of oxidizer makes the hybridthrottleable and restartable. Even thoughhybrids cannot match the density-impulseof solid rocket motors loaded withaluminum, motors with thrusts rangingfrom 60,000 to 75,000 pounds have beentested. Future tests expect thrustsreaching 225,000 pounds.

Fig. 5-11. Hybrid Rocket System


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Safety is an inherent advantage, claimmakers of hybrid systems. As notedabove, cracks in the fuel, because they arenot exposed to the oxidizer, do not causean explosion. Hybrid propulsion makeslaunch vehicles safer in flight. Enginethrust can be verified on the pad beforereleasing the vehicle for flight. And,unlike solids, hybrids can be shut down onthe pad if something goes wrong.

Environmental concerns are lessenedusing hybrid systems specially designed tominimize pollution effects. The hydrogenchloride in solid fuel exhaust has alreadybecome an environmental concern for theacid it dumps on the surface of the Earth,and the damage it does to the protectiveozone. Aluminum oxide, an exhaustcomponent of traditional solid rockets, isalso environmentally suspect. A hybridlaunch vehicle using polybutadiene fueland liquid oxygen produces an exhaust ofcarbon dioxide, carbon monoxide andwater vapor similar to that ofkerosene/liquid oxygen engines.

COOLING TECHNIQUES

The very high temperatures generatedin the combustion chamber transfer agreat deal of heat energy to thecombustion chamber and nozzle walls.This heat, if not dissipated, will causemost materials to lose strength. Withoutcooling the chamber and nozzle walls, thecombustion chamber pressures will causestructural failure. There are manymethods of cooling, all with the objectiveof removing heat from the highly stressedcombustion chamber and nozzle.

Radiation Cooling

This is probably the simplest method ofcooling a rocket engine or motor. Themethod is usually used formonopropellant thrusters, gas generators,and lower nozzle sections. The interior ofthe combustion chamber is covered with arefractory material (graphite,pyrographite, tungsten, tantalum ormolybdenum) or is simply made thick

enough to absorb a lot of heat. Coolingoccurs by heat loss through radiation intothe exhaust plume. Radiation cooling canset an upper limit on the temperatureattained by the walls of the thrustchamber. The rate of heat loss varieswith the fourth power of the absolutetemperature and becomes more significantas the temperature rises.

Ceramic Linings

In relatively small (low temperature)rockets, the interior walls of thecombustion chamber and nozzle may belined with a heat-resistant (refractory)ceramic material. The ceramic gets hot,but because it is a poor conductor of heat,it prevents the metal walls of themotor/engine from becoming overheatedduring the short operating period. Thismethod is not adequate for large rocketsin which the more intense heat must betransferred rapidly from the walls of thethrust chamber. Ceramic linings are alsotoo heavy for use in large rockets.

Ablation Cooling

As mentioned earlier, in the ablationcooling method, the interior of the thrustchamber is lined with an ablative material,usually some form of fabric reinforcedplastic. This material chars, melts andvaporizes in the intense heat of thenozzle. In this type of “heat sink cooling,”the heat absorbed in the melting andburning (the energy alters the chemicalform instead of raising its temperature) ofthe ablative material prevents thetemperature from becoming excessivelyhigh. The charred material also serves asan insulator and protects the rocket casefrom overheating. The gas produced byburning the ablative material provides anarea of “cooler” gas next to the nozzlewalls. The synthetic organic plasticbinder material is reinforced with glassfiber or a synthetic substance. Solidrocket motors use ablative coolingalmost exclusively, as there are no otherfluids to use to cool the nozzle throat.


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Film Cooling

With this method of cooling, liquidpropellant is forced through small holes atthe periphery of the injector forming afilm of liquid on the interior surface of thecombustion chamber. The film has a lowthermal (or heat) conductivity since itreadily vaporizes and protects the wallmaterial from the hot combustion gases.Cooling results from the vaporization ofthe liquid which absorbs considerableheat. Film cooling is especially useful inregions where the walls becomeexceptionally hot, e.g., the nozzle throatarea.

Transpirational Cooling

This technique is very similar to filmcooling. The combustion chamber has adouble-walled construction in which theinside wall is made of a porous material.Propellant is circulated through the spacebetween the walls and seeps continuouslythrough inner wall pores into thecombustion chamber. There it forms afilm which rapidly vaporizes. The coolingaction is much the same as film cooling,but has the additional advantage ofallowing considerable heat to be absorbedby the propellant within the walls of thechamber. This method is also referred toas evaporative or sweat cooling. Majordrawbacks to transpirational cooling arethat it is difficult to manufacture this typeof chamber, and also difficult to maintaina steady liquid flow through the pores.

Regenerative Cooling

This is the most common method ofcooling for cryogenic propellant rockets.It involves circulating one of the super-cooled propellants through a coolingjacket around the combustion chamberand nozzle before it enters the injector.The propellant removes heat from thewalls, keeping temperatures at acceptablelevels. At the same time, the temperature

of the propellant rises, causing it tovaporize faster upon injection. Thiscooling method is often used with gasgenerator systems as a way to driveturbopumps (Fig. 5-12).

Solid Rocket Motor Cooling

In solid propellant motors, the nozzle

serves the same purpose as in the liquidengine. Because there is no super-cooledpropellant available to provide cooling,we use other methods for thermalprotection. If not properly constructed,the walls of the combustion chamber willbecome excessively hot. This could causecase failure under the high operatingpressures existing in the interior. Toprevent this, the inner wall of the motorcase is coated with a liner or inhibitor.This liner provides a bond between thepropellant grain and the case preventingcombustion from spreading along thewalls, and acts as a thermal insulator,protecting the case from heat in areaswhere there is no propellant. Theunburned propellant provides additionalthermal protection as it must be vaporizedbefore it will burn.

In solid-propellant rockets, thenozzle’s form is often achieved with ashaped insert which keeps the nozzlethroat cool to prevent significant damageduring the operation of the motor.Common insert materials include bothrefractory substances, like pyrographite

Fig. 5-12. Regenerative Cooling


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and tungsten or ablative substances. Theablative materials are fabric reinforcedhigh temperature plastics as previouslydiscussed. There is usually no significantchange in motor performance due todeterioration of nozzle throat ablatives.

Another method of keeping the nozzlethroat cool is the use of a cooler burningpropellant located near the throat areawhich will burn and form a thin layer ofcooler gas next to the nozzle walls. Thisthin film of gas protects the nozzle fromthe high temperature gas created by themain propellant.

THRUST VECTOR CONTROL

In a rocket, the rocket engine or motornot only provides the propulsive force butalso the means of controlling its flightpath by redirecting the thrust vector toprovide directional control for thevehicle’s flight path. This is known asthrust vector control (TVC). TVC can bedivided into those systems for use withliquid engines and those for solid motors.

When choosing a TVC method, weneed to consider the characteristics of theengine/motor and its flight application andduration. Also, the maximum angularaccelerations required or acceptable, theenvironment, the number ofengines/motors on the rocket, availableactuating power, and the weight andspace limitations are all weighed againsteach other to produce a cost effective, yetappropriate, system of control. Theeffective loss of engine performance dueto the use of a particular TVC methodand the maximum thrust vector deflectionrequired are major design considerations

Liquid Rocket TVC Methods

Gimbaled Engines

Some liquid propellant rockets use anengine swivel or gimbal arrangement topoint the entire engine assembly. Thisarrangement requires flexible propellantlines, but produces negligible thrust losses

for small deflection angles. This methodis relatively common (Fig. 5-13).

Vernier Rockets

Vernier rockets are small auxiliaryrocket engines. These engines canprovide all attitude control, or just rollcontrol for single engine stages during themain engine burn, and a means ofcontrolling the rocket after the mainengine has shut off (Fig. 5-14).

Jet VanesJet vanes are small airfoils located in

the exhaust flow behind the nozzle exitplane. They act like ailerons or elevatorson an aircraft and cause thevehicle to change direction byredirecting the rocket. Jet vanesare made of heat-resistantmaterials like carbon-carbon andother refractory substances.Unfortunately, this controlsystem causes a two to threepercent loss of thrust, anderosion of the vanes is also amajor problem (Fig. 5-15).

Fig. 5-13. Gimbaled Engine

Fig. 5-14. Vernier Rocket

Fig. 5-15.Jet Vanes


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Solid Rocket TVC Methods

Rotating Nozzle

The rotating nozzle has no throatmovement. These nozzles work in pairsand are slant-cut to create an area ofunder expansion of exhaust gases on oneside of the nozzle. This creates anunbalanced side load and the inner wall ofthe longer side of the nozzle. Rotation ofthe nozzles moves this side load to anypoint desired and provides roll, yaw andpitch control. This system is simple butproduces slow changes in the velocityvector. Rotating nozzles are usuallysupplemented with some other form ofTVC.

Swiveled Nozzle

The swiveled nozzle changes thedirection of the throat and nozzle. It issimilar to gambaling in liquid propellantengines. The main drawback in using thismethod is the difficulty in fabricating theseal joint of the swivel since this joint isexposed to extremely high pressures andtemperatures (Fig. 5-16).

Movable ControlSurfaces

Movable ControlSurfaces physicallydeflect the exhaust orcreate voids in theexhaust plume todivert the thrustvector. This methodincludes jet vanes, jettabs, and mechanicalprobes. These TVCapproaches are all based on proventechnology with low actuator powerrequired. They suffer from erosion andcause thrust loss with any deflection.

A similar system is the jetavator, a slip-ring or collar at the nozzle exit whichcreates an under expansion region (asdiscussed in conjunction with rotatingnozzles). The jetavator is a movable

surface which allows the under expandedregion to be moved 360 degrees aroundthe rocket nozzle to produce pitch andyaw control. This system was developedfor the Polaris SLBM.

Secondary Fluid Injection

A secondary fluid is injected into theexhaust stream to deflect it, therebychanging the thrust vector (Fig. 5-17).Fluid injection creates unbalanced shockwaves in the exhaust nozzle whichdeflects the exhaust stream. There aretwo types of fluid injection systems.

The Liquid Injection TVC uses bothinert (water) and reactive fluids (rocketpropellants) for the TVC. Reactive fluidcombustion in the exhaust plume createsthe greater effect. Hydrazine, water,nitrogen tetroxide, bromine, hydrogenperoxide, and Freon have all been used.

The Hot Gas Injection TVC uses gaseither vented from the main combustionchamber, or from an auxiliary gasgenerator. These gases are “dumped”into the nozzle to cause the unbalancedshock wave.

Fig. 5-17. Thrust Vectoring

Fig. 5-16.Swiveled Nozzle


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SUMMARY

Table 5-1 summarizes the capabilities of the different types of rocket engines andpropellants. Each has its own advantages and disadvantages. Specific use of a particulartype depends upon the mission.

Type Thrust(1000 lbs)

Isp Missions

ChemicalLiquid

Solid

1500

2000-3000

260-455200-300

Manned missions near Earth andMoon. Instrumented probes to Venusand Mars.

Nuclear 250 600-1000 Heavy payload manned missions toMoon, Venus and Mars.

Arc-Jet .01 400-2500 Very heavy payloads from Earth orbit.Plasma .005 2000-10,000 To other planets and stationkeepingIon .001 7500- 30,000 For deep space missions

Table 5-1. Rocket Engines and Propellants

TOC


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REFERENCES

Asker, James R., “Moon/Mars Prospects May Hinge on Nuclear Propulsion,” AviationWeek & Space Technology, December 2, 1991, pp. 38-44.

Hill, Philip G., Peterson, Carl R., Mechanics and Thermodynamics of Propulsion.Addison-Wesley Publishing Company, MA, 1970.

Jane’s Spaceflight Directory, Jane’s, London, 1987.

Space Handbook, Air University Press, Maxwell Air Force Base, AL, January 1985.

Sutton, George P., Rocket Propulsion Elements, John Wiley & Sons, New York, 1986.

Wertz, James R., and Wiley J. Larson, ed., Space Mission Analysis and Design, KluwerAcademic Publishers, Boston, MA, 1991.

Documents

ROCKET THEORY - United States Air Force