Engineering Handbook - Propultion Systems

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Monk, J. C. Propulsion Systems

The Engineering Handbook.

Ed. Richard C. Dorf

Boca Raton: CRC Press LLC, 2000

1998 by CRC PRESS LLC


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176Propulsion Systems

176.1 Performance Characteristics

176.2 Liquid Rocket Engine CyclesPressure-fed Expander Gas Generator Staged Combustion

176.3 Major ComponentsMain Injector Thrust Chamber Turbomachinery

176.4 System Preliminary Design Process

176.5 Conclusion

Jan C. MonkNational Aeronautics and Space Administration

Rocket propulsion is an application of Newton's first, second, and third laws of motion. Newton's

first law of motion states that a particle not subjected to external forces remains at rest or moves

with constant velocity in a straight line. A rocket lifting off the launch pad goes from a state of rest

to a state of motion. Newton's second law of motion states that force equals mass times

acceleration. Force in the equation is the rocket thrust, where mass is the amount of rocket fuel

being burned and converted into gas, which expands and then escapes from the rocket. As the gas

exits the combustion chamber through a nozzle, it picks up speed. Newton's third law of motion

states for every action, there is an equal and opposite reaction. With rockets, the action is theexpelling of gas out of the engine; the reaction is the force or thrust of the rocket in the opposite

direction.

176.1 Performance Characteristics

In the process of producing thrust, rocket engines generate more power per unit weight than any

other engine. To enable a rocket to climb into low-earth orbit, it is necessary to achieve velocities

in excess of 28 000 km per hour. Escape velocity is a speed of about 40 250 km per hour. To

achieve these velocities, the rocket engine must burn a large amount of fuel and push the resulting

gas out of the engine as rapidly as possible. Containing and controlling this power is the basicchallenge in the development of these devices. For example, the power density produced by liquid

hydrogen (LH2) turbomachinery utilized by the space shuttle main engine (SSME) is

approximately 83 horsepower per pound of turbopump weight.

Rocket propulsion system design solutions are quite varied: thrust levels from ounces to millions

of pounds force; liquid and solid propellants; and liquid systems with pressures that are

maintained by turbopumps or pressurized tanks. Liquid system applications vary from small



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pressure-fed, storable monopropellant thrusters for keeping satellites stationary, to large

turbopump-fed, cryogenic bipropellant engines for boost propulsion. Combustion chamber

pressures vary from a few pounds per square inch (psi) to several thousand psi. Generally, liquid

propulsion systems consist of a propellant feed system, an injector, a combustion chamber, and a

nozzle. The propellant feed system includes ducting and valves for controlling flows and, in thecase of pump-fed systems, turbomachinery that draws propellants from lightweight propellant

tanks and increases the pressure to the level necessary to support the desired combustion chamber

pressure.

The ideal rocket propulsion equation is

Videal = g0 Isp lnM0

M1(176:1)

where Videal is the ideal delta velocity imparted on a vehicle, g0 is the gravitational constant, Ispis the propulsion system's specific impulse, M0 is the initial mass of the vehicle, and M1 is the

final or burnout mass of the vehicle. This equation provides two important performanceparameters: specific impulse, which is a measure of propulsion system efficiency expressed in

seconds, and vehicle burnout mass, which includes all structures (tankage, thrust structure, etc.),

residual propellants, engine systems, feed systems, pressurization systems, auxiliary systems,

electronic systems, upper stages, payload supporting structures, and the payload itself.

One of the more important internal rocket engine parameters is characteristic exhaust velocity,

commonly referred to as C-star (C) , which relates combustion chamber pressure, chamber throat

area, and propellant flow rate. Theoretical characteristic exhaust velocity C is computed as

follows:

C

=

Pns Atg0

_wtc (176:2)

where Pns is nozzle stagnation pressure in psi, At is throat area in square inches, and _wtc is

chamber propellant mass flow rate in pounds-mass per second. A number of losses will reduce the

actual C realized. These losses are generally a function of injector design and are related to

mixture ratio maldistribution, mixing, etc. The actual C realized by a given design is,

Cact = c C (176:3)

where c is C efficiency, typically between 0.80 and 0.99.

Another useful parameter is thrust coefficient, which relates thrust F, chamber pressure, andthroat area as follows:

F = CF Pns At (176:4)

where CF is the thrust coefficient, Pns is nozzle stagnation pressure, and At is throat area. Once

again, additional parameters must be added to reflect actual values. This yields the following:



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F = CF CF Pns At PaAe (176:5)

where C F is thrust coefficient efficiency, typically between 0.90 and 0.97, Pa is local atmospheric

pressure in psi, and Ae is exit area in square inches. This equation yields thrust at any point

between sea level and vacuum conditions.Specific impulse Isp is an overall efficiency term and is defined as

Isp =F

_wt(176:6)

where Fis thrust level in pounds-force and _wt is the total mass flow rate in pounds-mass persecond. Specific impulse can be computed for the engine or thrust chamber by utilizing either

engine thrust and flow rate or thrust chamber thrust and flow rate, as appropriate. Specific impulse

can also be computed ifC and the thrust coefficient are known. This relationship is expressed as

Isp =CCF

g0(176:7)

Again, one must maintain consistency between theoretical values and actual values.

Thrust and specific impulse are commonly calculated at either sea level or vacuum conditions for

reference or comparative purposes. Later discussions will refer to sea level thrust (Fsl ) , vacuum

thrust (Fvac ) , sea level specific impulse (Ispsl ) , and vacuum specific impulse (Isp vac ) .

Mixture ratio is the ratio between the oxidizer and fuel flow rates, and is expressed in equation

form as

MR =_wo

_wF(176:8)

where _wo is oxidizer flow rate in pounds per second and _wF is fuel flow rate in pounds per second.Mixture ratio can be computed for the engine or thrust chamber by utilizing either engine flow

rates or thrust chamber flow rates, as appropriate.

Expansion ratio " is a ratio of the thrust chamber nozzle exit area, Ae ; and the thrust chamber

throat area, At :

" =Ae

At(176:9)

A more complete definition of these and other rocket engine equations, including solid propellant

systems, can be found inRocket Propulsion Elements [Sutton, 1992].

176.2 Liquid Rocket Engine Cycles

A number of power cycles are available for liquid propellant systems. These include pressure-fed,



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expander, gas generator, and staged combustion cycles. Each cycle has advantages and

disadvantages; the one selected for a given application is determined after a series of system trade

studies. A description of a number of engine systems is given in Table 176.1. A brief description of

each of these power cycles follows.

Table 176.1Liquid Rocket Engine Characteristics

Pressure-fed

This system consists of a thrust chamber assembly, associated ducting and valves necessary for

control, pressurized tankage, and the pressurization system for the tankage. This system is widely

utilized for satellite attitude control, orbital transfer, and as auxiliary propulsion for most major

launch vehicles. Pressure-fed systems are perhaps the simplest of all propulsion systems, but are

performance limited because of the weight penalty associated with increasing chamber pressures.

As pressures increase, tank wall thickness and the mass of the gases needed to maintain tank

pressures increase. Tank pressures are set by chamber pressure plus pressure losses in the cooling

circuit (if any), injector, valves, and ducting. In most pressure-fed applications, combustion

chambers are passively cooled (i.e., film-cooled or radiative/ablative). The space shuttle utilizes

pressure-fed systems for the orbital maneuvering system and the reaction control system. A

schematic of a simple pressure-fed system is given inFig. 176.1.



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Figure 176.1 Pressure-fed propulsion system schematic.

Expander

This is the simplest of the turbopump-fed systems primarily because the power source for the

turbines is the thrust chamber cooling circuit. Only the thrust chamber requires an ignition system.

Pump discharge pressures are set by chamber pressure plus pressure losses in the cooling circuit,turbine, injector, valves, and ducting. The combustion chamber is regeneratively cooled. In some

applications, extensible radiation-cooled nozzle extensions are used to increase area ratio while

maintaining a short stowed length. Expander cycles are limited in the combustion chamber

pressure that can be attained because the energy available to drive the turbine(s) is obtained from

the combustion chamber cooling circuit. For applications that require operation at sea level, this

reduces the area ratio that can be achieved without side loads. Nozzle flow separation is discussed



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generators are operated at relatively low mixture ratios because turbine temperatures must be

maintained in the 1000 to 2000 degrees Rankine range. The main combustor mixture ratio is biased

higher to offset this parasitic flow. The combination of poor thrust efficiency of the GG gases and

main chamber mixture ratio bias results in a specific impulse penalty. The gas generator cycle was

utilized on the F-1 and J-2 engines of the Saturn V launch vehicle and is currently in use on theDelta, Atlas, and Titan launch vehicles. A schematic of a simple gas generator system is given in

Fig. 176.3.

Figure 176.3 Gas generator engine system schematic.



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Staged Combustion

The staged combustion cycle provides the highest performance of conventional chemical rocket

engines. Turbine power is derived from a separate combustor or preburner which also utilizes the

same propellants as the main system. In bipropellant systems, the hot gas is routed through the

turbopump turbines to the main injector where it is mixed with the other propellant and is

combusted in the main chamber. Pump discharge pressures are set by chamber pressure plus

pressure losses in the cooling circuit, turbine, injector, valves, and ducting. Thrust chambers are

regeneratively cooled. Staged combustion cycle engines developed in the U.S. have utilized a

fuel-rich preburner. Several rocket engine systems developed in Russia have utilized an

oxidizer-rich preburner. In the former case, the fuel-rich hot gases are mixed with oxidizer in the

main chamber. In the latter, oxidizer-rich hot gases are mixed with fuel in the main chamber. The

staged combustion cycle utilizes all propellants in the main combustion chamber, which provides

maximum performance. A schematic of a simple staged combustion system is given inFig. 176.4.

Figure 176.4 Staged combustion cycle engine system schematic.



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A variant of the staged combustion cycle is the full flow cycle, in which the oxidizer pump

turbine is driven with an oxidizer-rich preburner and the fuel pump is driven with a fuel-rich

preburner. This cycle offers some simplification in turbomachinery design because of the

simplified seal design between the pump end and the turbine end, and a significant reduction in

turbine temperatures because all the propellants can be utilized in the turbine drive circuits. This

concept is currently under study as a candidate engine system for the next generation launch

vehicle. A schematic of a simple full flow staged combustion system is given in Fig. 176.5.

Figure 176.5 Full flow staged combustion cycle engine system schematic.



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combustion chambers, a film of fuel is injected at the outer circumference of the injector. In order

to produce a stable combustion process, baffle elements which are cooled with fuel are commonly

used. The most common injector concepts are coaxial, showerhead, and impinging. These are

illustrated inFig. 176.6.

Figure 176.6 Injector concepts.

176.3 Major Components

Main Injector

The purpose of the injector is to introduce propellants into the combustion chamber in a controlled

manner, to atomize the propellants, and to mix the propellants at the proper mixture ratio in a

homogenous manner. Mixture ratio variations across the injector face are one of the most commonproblems that the designer will encounter, and these maldistributions lead to combustion efficiency

losses. In some cases, maldistributions are deliberately introduced. To enhance the durability of



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The coaxial injector consists of a series of concentric tubes into which the oxidizer is introduced

through a center tube and the fuel introduced through the annular area formed by a second tube.

This type of injector is commonly used in oxygen-hydrogen engines. Impinging and showerhead

injectors, which are commonly used in oxygen-kerosene and storable propellant engines, consist of

a series of two sets of orifices. One injects the oxidizer, and the other the fuel. The number oforifices, injection velocities, and injection angles are selected to provide consistent atomization and

mixing of propellants. Impinging injectors slant the orifices to impinge the two propellant streams

against each, enhancing mixing. The design utilized by the conventional impinging injector

consists of a series of concentric copper rings containing the injection orifices. The rings alternate

between oxidizer and fuel. The outer ring is generally a fuel ring and contains a set of smaller

orifices that control film coolant for the combustion chamber. Separate manifolding routes

propellants to each set of orifices.

Thrust Chamber

A number of design solutions have been utilized in thrust chambers, varying from passively cooledablatives to a number of regeneratively cooled concepts. In some applications, the thrust chamber

is composed of two separate components. The upper portionincluding the throat region and a

portion of the expansion regionis commonly called a combustion chamber. The lower

portionconsisting of the remainder of the expansion regionis called a nozzle. Regeneratively

cooled thrust chamber designs include brazed tube bundles, copper with milled channels, and steel

with milled channels. The bundled tube concept utilizes: steel tubes, pressed to vary the shape

necessary for formation of the overall thrust chamber shape; a structural shell in the combustion

chamber region with a number of straps spaced along the thrust chamber length for additional

strength; and necessary manifolding for inlet and discharge coolant flow. For higher pressure

applications (greater than approximately 1800 psia), the heat load produced by the combustion

process exceeds the capability of brazed tube designs. For these applications, a copper liner is

required in the high heat flux region. This configuration consists of a slotted copper liner, structural

jacket, and manifolding. Figure 176.7 illustrates these two thrust chamber concepts.



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Figure 176.7 Two thrust chamber concepts.



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MANNED MANEUVERING UNIT

Astronaut Bruce McCandless II is a few meters away from the cabin of the earth-orbiting Space

Shuttle Challengerin this 70 mm photograph taken on February 7, 1984. McCandless is one of the

two 41-B mission specialists who participated in this historical extravehicular activity (EVA). This

spacewalk represented the first use of a nitrogen-propelled, hand-controlled device called the



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Manned Maneuvering Unit (MMU), which allows for much greater mobility than that afforded

previous spacewalkers, who had to use restrictive tethers.The MMU is a self-contained backpack with nitrogen gas propulsion that allows orbiter crews to

move outside the payload bay to other parts of the orbiter or to other spacecraft. The MMU latches

to the spacesuit (Extravehicular Mobility Unit, EMU) backpack and can be donned and doffed by

an astronaut unassisted.MMU controls follow the layout familiar to spacecraft crews: the left-hand controller governs

foreaft, rightleft, and updown translations, while the right-hand controller handles roll, pitch,

and yaw motions. The controllers may be used singly or in combination to give a full range of

movement within the operating logic of 729 command combinations, including attitude hold.Thrust impulses are from 24 dry nitrogen gas thrusters each with 7.56 newtons thrust. Two

25-by-76 centimeter (9.8-by-30 inch) Kevlar filament-wrapped aluminum nitrogen tanks each hold

5.9 kilograms (13 pounds) of nitrogen when fully charged. Two 16.8 volt, 752 watt-hour silver

zinc batteries supply MMU electrical power, enough for one six-hour EVA. The nitrogen tanks

could be recharged in less than 20 minutes at the payload bay MMU service rack.Built by Martin Marietta, Denver, CO, the MMU is 1.2 m (49.4 in.) high, 81 cm (32.5 in.) wide,

and 1.1 m (44.2 in.) deep with control arms extended. The MMU weighs 136 kg (300 lb) when

charged with nitrogen. With a spacesuited crewman and consumables added, on-orbit mass isabout 335 kg (740 lb). (Photo courtesy of National Aeronautics and Space Administration.)

Turbomachinery

The turbomachinery design process of liquid rocket engines is very similar to a normal

pump/turbine design, except for two critical areas. The first is the critical need to minimize weight.

This is perhaps the greatest difference. As stated earlier, the power density of the space shuttle

main engine turbopump is 83 horsepower per pound of turbopump weight. The second difference

is the dynamic and steady state environments that rocket engines require. Although a number of

turbojet engines operate at turbine temperatures significantly higher than most rocket engines, theyattain the steady state operating point in a matter of minutes, not in one to four seconds as do

rocket engine turbines. This produces severe thermal strains that tax the ability of materials to

sustain. Other environments that provide problems in some materials are oxygen and hydrogen.

Particle impact, fretting, and rubbing in an oxygen environment can lead to disastrous fires.

Susceptibility of materials to hydrogen embrittlement reduces the variety of materials available for

the designer or requires platings to protect materials. Another environment to which rocket engine

turbomachinery is susceptible is rotor dynamics, which is considerably more critical than in

conventional rotating machinery because of the reduced weight of rocket turbopumps. Structural

design considerations, including explanation of the processes utilized in the SSME, can be found in

Structural Design/Margin Assessment[Ryan, 1993].

176.4 System Preliminary Design Process

A number of theoretical thrust chamber performance computer models are readily available that

provide the basic performance parameters needed to support a conceptual design. The most

common is a Finite Area Combustor Theoretical Rocket Performance Program, commonly referred

to as the One-Dimensional Equilibrium (ODE) Program referenced in Computer Program forCalculation of Complex Chemical Equilibrium Compositions and Applications, Supplement

ITransport Properties [Gordon et al., 1984]. This model generates performance data for various



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propellant combinations as a function of mixture ratios, combustion chamber pressures, and nozzle

expansion ratios. A sample set of data for liquid oxygen (LO2 )/liquid hydrogen propellants isgiven inTable 176.2.

Table 176.2Theoretical Performance of Oxygen-Hydrogen Combustor



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follows:

Propellants: Liquid Oxygen/Liquid Hydrogen

Sea-Level Thrust: 400000 pounds force

Mixture Ratio: 6.0:1.0

In some instances, a minimum specific impulse value is specified. However, in most cases, the

design value should be selected as the result of vehicle-engine trade studies, along with engine

weight, recurring costs, and nonrecurring costs.

One of the first choices to be made by the engine designer is the value of combustion chamber

pressure. This choice is also the result of a series of trades. For this exercise, 2000 psia has been

chosen and is a good first approximation for the optimum value when recurring costs are one of the

more important parameters. This value will provide good performance, while simplifying

turbomachinery to two pump stages with moderate turbine temperatures. An initial assumption

needs to be made as to engine cycle. For a booster application, either a gas generator cycle or astaged combustion cycle usually provides optimum performance. For this exercise, the staged

combustion cycle is selected. This simplifies the initial set of calculations in that the engine flow

rate and thrust chamber flow rate are approximately identical.Table 176.3provides a typical ODE

output for a thrust chamber operating at a chamber pressure of 2000 psia and a mixture ratio of 6.0.

The first parameter to select is area ratio. From previous vehicle trade studies, the optimum nozzle

exit pressure (Pe) for a first stage or booster vehicle is approximately 6.5 psia, the optimum for a

single-stage-to-orbit vehicle is approximately 4.0 psia, and the optimum for a parallel burn core

stage (booster and core stages ignite at sea level) is 2.5 psia. Standard atmospheric conditions for

sea level and various altitudes can be found in Terrestrial Environment (Climatic) Criteria

Guidelines for Use in Aerospace Vehicle Development, 1993 Revision [Johnson, 1993].

Table 176.3Theoretical Performance of Oxygen-Hydrogen Combustor at a Chamber Pressure of2000 psia and a Mixture Ratio of 6.0

The following process outlines a methodology of determining the initial set of overall system

requirements for a liquid rocket engine. A preliminary set of top-level engine requirements must be

established by the vehicle systems designer for a booster engine. For this exercise, they are as



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Another consideration is side loads on the nozzle. For an overexpanded nozzle (Pe < Pa ) ,

unsymmetrical flow separation results if significant dynamic loads are applied to the nozzle

(pressure times surface area forces). These side loads can ultimately destroy a nozzle and, at a

minimum, result in significant weight increases. Side loads are computed using both empirical

techniques and detailed nozzle fluid flow with computational fluid dynamics techniques.Calculating the locations within the nozzle where flow separation will occur is very difficult and

side loads can usually be quantified accurately with test data.

As the application for this exercise is a booster, an exit pressure of approximately 6.5 psia is

desired. Using Table 176.3, an expansion ratio of 30 yields an exit pressure of 6.2 psia, which is

close enough for a first approximation of the engine characteristics. Using the initial vehicle thrust

and mixture ratio requirements, theoretical values ofC and CF obtained fromTable 176.3, and

assumed values ofC efficiency of 0.99 (readily obtainable with oxygen/hydrogen coaxial tube

injectors) and CF efficiency of 0.97, various engine parameters can be computed using Eqs.

(176.3), (176.5), (176.6), (176.8), and (176.9) and are shown below in order of

computation:

Fvac 400 000 lb

Propellan

ts

Oxygen/hydrogen

MR 6.0

Cycle Staged combustion

Pns 2000 psia" 30:1

CF 1.8311

C* 7539.8 ft/s

Ai 109.22 in2

Ae 3276.7 in2

_wt 932.27 lbm/s_wf 133.18 lbm/s_wo 799.09 lbm/s

Ispvac 429.1 s

Fsl 351 846 lbf

IspSl 377.41 s

176.5 Conclusion

The science of rocketry has enabled some of humankind's greatest achievements, ranging from

instantaneous global communications and accurate weather forecasting via geostationary satellites

to trips to the moon. NASA's space shuttle is one of the most complex flying machines ever built

and is the only partially reusable launch vehicle. While several countries have rockets capable of



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carrying a variety of payloads, the U.S. and Russia are the only countries with spacecraft that can

transport a crew to and from orbit. France and China have expendable launch vehicles now in use,

and Japan is developing another.

The U.S. is on the threshold of a next generation launch system. NASA aerospace engineers and

industry experts are exploring new concepts, including the first fully reusable launch vehicle.Developing rockets for 21st century missions promotes enhanced technologies to meet new

challenges, including balancing design requirements between operability, performance, weight,

and cost. The next generation of spaceship will open new doors to the space frontier.

Defining Terms

Ablation: A passive cooling technique in which heat is carried away from a vital part by

absorption into a nonvital part, which may melt or vaporize and then fall away, taking the

heat with it.

Combustion chamber: A devicewhich includes a throat regionto mix, burn, and control

propellants.Injector: A device to distribute and inject propellants into the combustion

chamber.

Nozzle: A device used to accelerate the combusted gases.

Propellant: Fuel [the chemical(s) the rocket burns] and an oxidizer (oxygen compounds) to ignite

the fuel.

Sea level: Standard atmospheric conditions at an altitude of zero feet.

Side loads: Unsymmetrical loads put on a nozzle because of internal flow separation of

overexpanded gases.

Regeneratively cooled: A cooling technique in which propellants, usually fuel, are utilized to

remove heat from the inner wall of a combustor in a heat exchange

process.

Thrust chamber assembly: An assembly consisting of the main injector, combustion chamber,

and nozzle. Depending upon fabrication techniques, the combustion chamber and nozzle can

be separate components or combined into a single component.

Vacuum: Conditions where atmospheric pressure can be considered to be 0.0

psia.

References

Gordon, S., McBride, B., and Zeleznik, F. 1984. Computer Program for Calculation of Complex

Chemical Equilibrium Compositions and Applications, Supplement ITransport

Properties. NASA Technical Memorandum 86885, National Aeronautics and Space

Administration, Office of Management, Scientific and Technical Information

Program.

Johnson, D. 1993. Terrestrial Environment (Climatic) Criteria Guidelines for Use in Aerospace

Vehicle Development, 1993 Revision. NASA Technical Memorandum 4511, National

Aeronautics and Space Administration, Office of Management, Scientific and Technical



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Information Program.

Ryan, R. 1993. Structural Design/Margin Assessment. NASA Technical Paper 3410, National

Aeronautics and Space Administration, Office of Management, Scientific and Technical

Information Program.

Sutton, G. P. 1992.Rocket Propulsion Elements, 6th ed. John Wiley & Sons, New York.

Further Information

American Institute of Aeronautics and Astronautics (AIAA)

American Society of Mechanical Engineers (ASME)

National Space Society

Marshall Space Flight CenterCentral Technical Library (Phone: 205-544-4524)

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Engineering Handbook - Propultion Systems