Current Liquid Propellant Systems - Advances in Chemistry (ACS

11 Current Liquid Propellant Systems

Downloaded by NORTH CAROLINA STATE UNIV on May 8, 2013 | http://pubs.acs.org Publication Date: June 1, 1969 | doi: 10.1021/ba-1969-0088.ch011

JACOB SILVERMAN and MARC T. CONSTANTINE Rocketdyne, A Division of North American Rockwell Corp., Canoga Park, Calif. 91304

The requirements for selecting a fuel and oxidizer as a liquid bipropellant system are usually a compromise between the demands of the vehicle system, the propulsion system, and the propellants themselves. The vehicle and propulsion system will determine performance levels, physical property requirements, thermal requirements, auxiliary combustion requirements, degree of storability and packageability, hypergolicity, etc. The final propellant selection must not only satisfy such requirements but is also dictated by thermochemical demands which the fuel and oxidizer make on each other. Frequently, specifically required properties are achieved through the use of chemical additives and/or propellant blending.

" C r o n i the initial application of the liquid rocket propulsion system concept i n the V - 2 rocket to the present development of various systems used to perform the Apollo missions, a variety of elements, compounds, and mixtures have been utilized as liquid rocket propellants. Since there has been no single liquid chemical or combination of liquids suited to all requirements of the present spectrum of rocket propulsion systems, a number of liquid propellants have been developed from various chemical families. Significantly, the hydrocarbons, amines, hydrazines, boranes, nitrogen oxides, nitric acids, halogens, and oxygen have a l l contributed to the growing technology of liquid rocket propellants. The physical and chemical characteristics of these candidate liquid propellants vary widely. However, all of the liquids which have found application as rocket propellants have one common characteristic—they are designed to fit the particular requirements of at least one particular rocket engine and vehicle system. Obviously, few liquids initially fulfill the requirements of a propulsion system designed to perform a particular mission. Thus, various compromises must be undertaken between the 301 In Propellants Manufacture, Hazards, and Testing; Boyars, C., et al.; Advances in Chemistry; American Chemical Society: Washington, DC, 1969.

302

PROPELLANTS

MANUFACTURE,

HAZARDS,

A N D TESTING


requirements of the system and the inherent characteristics of the available candidate propellants. Sometimes the propellant is modified, or some of the desirable design or operational features of the systems are changed, but i n the end the resultant compromise between the system and propellant w i l l perform the originally conceived mission. To establish the relationship between current liquid propellant applications and the available propellant technology, this paper has been divided into three sections. A section on basic propellant considerations describes the normal parameters used to evaluate propellant candidates and their influence on the propulsion system. Although such considerations have been thoroughly discussed in many previous publications ( e.g., Ref. 3 ) , their importance i n establishing the basic criteria for propellant system selection requires a limited review i n this text as a background aid to the reader. Current liquid propellants and propellant candidates are discussed i n a second section i n terms of capabilities and limitations as well as potential application areas ( the compositions of all propellants discussed are defined i n the Nomenclature section at the end of this article). Finally, a section of propellant tailoring illustrates examples of propellant formulation and describes propellant problem-solving techniques. I n conclusion, the results of these considerations are illustrated by the current liquid propellant systems. Basic"PropellantConsiderations Theoretical Performance Requirements. E a c h propulsion system has a minimum performance level that must be achieved to conduct its required mission. After careful analysis of the mission profile, this required performance level is usually related i n terms of required specific thrust or specific impulse (I ). This nomenclature defines the pound of thrust produced per pound per second of propellant flow: 8

where I = specific impulse ( s e c ) , F — engine thrust (lbs.), and W = total weight flow rate of propellant feed (lbs./sec). This propellant system figure of merit represents the work done b y the enthalpy drop i n the rocket engine system: 8

(2) where k = dimensional constant; H = enthalpy of incoming propellants; and H = enthalpy of the combustion products at the exit plane of the rocket engine. c

e

In Propellants Manufacture, Hazards, and Testing; Boyars, C., et al.; Advances in Chemistry; American Chemical Society: Washington, DC, 1969.

11.

SILVERMAN AND coNSTANTiNE

303

Liquid Propellant Systems

The calculation of a propellant system's theoretical I is a complex process which involves a number of different assumptions and estimations. Most standard computations, which are described thoroughly in a number of texts ( 1 ), involve a model which assumes instantaneous adiabatic combustion at a constant pressure or volume i n the chamber section of a rocket engine, followed by one-dimensional isentropic expansion of the gases in the nozzle section to an assigned pressure or nozzle area. The model further assumes gas ideality, kinetic and thermal equilibrium between condensed and gaseous phases, and a negligible volume of condensed phase. The typically reported I assumes that chemical equilibrium is achieved i n the chamber and maintained throughout the expansion process (shifting equilibrium). s


8

Specific impulse, calculated by this technique, represents a 100% conversion of chemical energy to mechanical energy, and, therefore, is an upper limit to the performance available from a real rocket engine. However, regardless of the technique utilized, the theoretical I of each of the systems are compared on a common basis (e.g., at the same combustion chamber pressure, nozzle geometry, exit pressure, etc.) with the desired performance level dictated by the mission and engine system rquirements. 8

In systems i n which the vehicle configuration is volume limited, theoretical performance comparisons, using density impulse (I d) are also necessary. This nomenclature, which is the product of the I and the bulk density of the propellants, defines the amount of thrust available i n a unit volume of the propellant. The relative importance of I d to I must be defined in the mission analysis. s

s

8

s

Desirable Physical Characteristics. A number of physical characteristics are important i n evaluating and selecting a liquid propellant system. Among these is the normal liquid range of the propellants. This should conform to the operating range of the vehicle system. In most cases, thermal regulation is usually undesirable or impossible; therefore, matching low temperature properties of the propellant to the system's operating environment is more feasible. Conversely, the desirability of a high boiling point and high critical temperature is obvious for those systems exposed to high temperatures. Another desirable physical property is high propellant density. One requirement for high propellant density has already been noted i n the discussion of performance (l d) requirements. The density of the propellant controls the size of the propellant tankage. E v e n i n systems i n which the volume of the vehicle is not critical, the smaller tankage results in a reduction in structural weight and aerodynamic drag of the vehicle. In addition, the change in density with temperature variation should be 8



304

PROPELLANTS M A N U F A C T U R E , HAZARDS, A N D TESTING

low to limit propellant system design and operational problems attendant to expansion and contraction of the liquid during temperature changes. In systems where the propellants are used for thrust chamber cooling, properties related to heat transfer are important. Candidate propellants should have high specific heats, high thermal conductivities, high boiling points, high decomposition temperatures, etc. Mass transfer requirements dictate low propellant viscosity. In addition to the excessive work required to transfer a viscous propellant from the tank to the combustion chamber (either through pumping or pressurization), the injection problems associated with viscous propellants have often compromised efficient combustion. L o w propellant vapor pressure allows a more efficient pump design and avoids one problem area i n fluid pumping. Engineering Properties. Engineering properties usually include storability, thermal stability, materials compatibilities, shock and thermal sensitivity, and toxicity. Failure in any of these areas would eliminate the propellant from further consideration. Storability is the physical and chemical stability of the liquid propellant during storage either i n propellant handling or i n missile systems. This property is always related to various materials of construction and at temperature ranges normally associated with potential storage conditions. The complete absence of decomposition or chemical reaction of the propellants i n the presence of various types of materials, common system contaminants (i.e., moisture), and maximum storage temperatures is normally preferred; however, some minimum rates or levels are permitted i n various applications. Because propellants are constantly subjected to abnormally high temperatures i n various parts of the propulsion system during operation, high thermal stability is desirable. Decomposition of the propellant at temperatures experienced i n the combustion chamber cooling jacket, the injector, and/or the gas film on the combustion chamber wall, can cause undesirable product deposition (resulting i n local "hot spots" and burnout), explosion i n the cooling jacket and/or injector, undesired reaction chains i n the combustion chamber, etc. Materials compatibility is the resistance of materials of system construction to chemical or physical attack b y the propellants or the products of propellant combustion. In addition to destroying the integrity of the pertinent structural member, corrosion of the material by the propellant results i n contamination of the propellant with the corrosion products, which i n turn deteriorates the propellant's physical and chemical properties. As i n storability, materials compatibility is related to usage temperatures, as well as to special materials handling, cleaning, and/or passivation techniques. Although limited materials compatibility of a



11.

SILVERMAN A N D CONSTANTINE


305

propellant may not necessarily eliminate the propellant from consideration, it imposes restrictions on the ultimate mission. The propellants should be stable with respect to initiation and propagation of deflagration or detonation. Because of their energetic nature and high reactivity, most propellants are sensitive to compression or thermal shock under certain conditions (i.e., contamination, high temperature, incompatible materials); however, many of these propellants can be utilized if the pertinent conditions are well characterized and avoided. O f course, there are some chemicals (i.e., ozone) whose sensitivities are beyond present handling technology; thus, their use as propellants is presently precluded. Although toxicity of a propellant candidate ( or its combustion products) is important i n a l l propellant selections, emphasis varies on this factor. Toxicity is usually of passing interest i n upper stage or space vehicles, which can utilize specialized handling techniques and controlled launch sites to eliminate the possibility of contamination during launch. However, where the application involves the handling of the propellant system by a large number of personnel in the field or a potential launch over a populated area, high toxicity eliminates a propellant from further considerations. Economic Factors. Economic factors are related to the availability and cost of the propellant as well as the cost of the equipment required to transport, store, and supply the propellant. Generally, low cost is a prime requisite for a propellant which w i l l be utilized in large quantities and/or i n multiunits (i.e., booster stages of launch vehicles and i n military weapons ). However, where utilization of a high-cost propellant may be required to complete the mission, the cost factor can be of secondary importance. This situation is usually associated with upper stages of a space launch vehicle. Combustion Characteristics. Ignition, combustion efficiency, and combustion stability are the principal combustion-related considerations i n liquid propellants. Hypergolic ignition, defined as the spontaneous ignition upon contact of fuel and oxidizer, is a distinctly advantageous property. Although hypergolicity is not a prime requisite of all propulsion systems, this desirable feature eliminates the complexity and weight of auxiliary ignition systems. In systems requiring pulsing operation and/or several restarts, hypergolicity is a necessity. Combustion efficiency is usually described i n terms of specific impulse efficiency (percentage of theoretical specific impulse achievable). Specific impulse efficiencies depend greatly on both chemical composition of the propellant and the physical design of the injector, combustion chamber, and nozzle configuration. Efficiencies can thus vary from 90-98%.


306


Under certain conditions, propellants may exhibit high frequency vibratory combustion. Such vibration can cause extensive hardware damage and/or a mission abort. In most instances, however, combustion instability is related principally to the physical design of the combustion chamber rather than the chemical properties of the propellants.


Current Liquid Propellants and Propellant Candidates Current liquid propellants are reviewed from two points of view. First, the current application areas are described in terms of their requirements. Secondly, the capabilities and limitations of various liquid propellant candidates are indicated with a few descriptive illustrations. Application Areas. Applications can be divided into two general groupings—space exploration and military applications. Space exploration may range from a sounding rocket into the upper atmosphere to manned exploration of the moon. Military applications range from small air-to-air missiles to I C B M requirements. Although many of the requirements between the two may be similar, the primary difference is that the space systems are designed to be handled by a relatively few highly specialized and trained personnel i n well-controlled environments, while military systems are designed to be handled by a large number of personnel under various situations i n the field. Space exploration application areas can be further s u b d i v i d e d — namely, booster propulsion, upper stage propulsion, and spacecraft control propulsion. The requirements i n each of these general areas are different. Propellants that are comparatively easy to handle i n large quantities, have unlimited availability, and are relatively cheap are normally the choice for booster application. Although only moderate performance levels are required, the bulk density of the propellant should be fairly large to preclude unreasonably sized vehicles. The propellant system need not be hypergolic, although this is obviously desirable. The propellants can have almost any liquidus range as long as they can be handled within the available technology. Because these systems are usually pump fed and regeneratively cooled, both propellants should have good mass transfer properties, and at least one of the propellants (usually the fuel) should have good heat transfer properties. Upper stage propellant applications are usually based primarily on performance—i.e., high specific impulse. If the upper stage is a multistart vehicle, hypergolicity is usually required. Careful consideration is also given to the propellant physical and chemical stability as well as to the matching of the propellant's liquidus range to the space environmental temperature if the propellant system is to remain operational i n space



11.



307

for an extended period. Such factors as toxicity, cost, density, etc., are of much lesser importance. Selection of propellants for spacecraft control propulsion usually depends on factors related to long term reliability and ease of operation. Spacecraft control propulsion systems include reaction control systems, attitude control systems, orbital altitude, maneuvering systems, etc. A l l require hypergolic propellants for multirestart and/or pulsing modes of operation. Although specific impulse is not of primary importance, reasonably high bulk density (weighted average of fuel and oxidizer) is usually required because of the limited volume configuration. These systems are usually pressure fed and utilize ablative, film, or radiation cooling for the combustion chamber. Military applications can be subdivided into two distinct types— fixed launch site systems and mobile systems. Both types place emphasis on specific impulse, density impulse, excellent storability and stability, and instant readiness. Fixed-site systems, which include most of the long range ballistic missiles, are systems whose launch facilities are well established and are usually environmentally controlled. These vehicles are usually pump fed and regeneratively cooled. Hypergolicity is preferred, but not necessarily required. Mobile weapon systems include the ordnance weapon systems, such as the air-to-air, air-to-surface, surface-to-air, and short-range, surfaceto-surface missile. Because of the limited volume requirements, heavy emphasis is placed on density impulse. These systems are completely prepackaged (including propellants) and require complete storability of the propellant as a liquid over a temperature range — 6 5 ° - 1 6 0 ° F . Because of personnel proximity at all times, toxicity is sometimes an important factor. The large number of units which are utilized usually imposes restrictions on propellant cost. Current Propellant Capabilities and Limitations. In selecting a propellant system for any application, the various considerations and factors briefly outlined above are weighted for each of the propellant candidates against the requirements of the proposed system. Initially, the potential candidates are screened by evaluating their performance potential (I and/or I d, as required). Some insight into the performance of different combinations can be obtained by utilizing a generalized performance criteria chart such as shown in Table I. This table illustrates the performances that may be achieved through combination of various oxidizer and fuel types. The numbers in the table represent theoretical calculations of I (seconds) at standard conditions of 1000 p.s.i.a. and shifting equilibrium with optimum expansion to sea level conditions (using the "typical procedure" previously described). The thermodynamic properties of the propellants and the potential reacs

s

s


308


tion products, which are required for these calculations, were established by the J A N A F Thermochemical Working Group (2). The propellants i n the top row of Table I represent the highest performing members of the various halogen, oxyhalogen, nitrohalogen, hydroxyl, nitrogen oxide, and interhalogen oxidizer families, respectively. The left-hand column represents various fuels of the hydrogen, beryllium, borane, hydrocarbon, hydrazine (and amine), aluminum, and lithium families.


Table I.

Generalized Performance Criteria

Specific Impulse at 100 p.s.i.a. Chamber Pressure, Shifting Equilibrium, Optimum Expansion to Sea Level Oxidizer Fuel H BeH

OF

0

N F,

411 378 363 344 364 348 365

401 360 365 347 339 322 332

391 351 343 311 313 309 269

361 350 340 314 334 327 333

2

2

2

2

62^6

CH N H AIH3 LiH 4

2

F

4

C/F

H0 2

2

2

2

343 329 317 293 312 304 313

341 332 321 283 292 301 253

314 357 332 281 282 318 268

5

Such a table is indicative of the performance potential of these propellant groupings, although the performances w i l l vary slightly from member to member in each propellant family. Table I shows that the highest performance is realized from systems containing F ( in particular, the F / H system). In general, maximum performance is achieved with various combinations of the cryogenic propellants, H , B H , C H , F , O F , 0 , N F . One exception are those combinations involving B e H . Typical performances of earth-storable propellants are noted in the combinations of H 0 , N 0 , and C1F with B e H , N H , A 1 H , and L i H . 2

2

2

2

2

2

2

2

6

4

2

4

2

2

2

2

4

5

2

2

4

3

A n expansion of Table I is shown i n Table II to illustrate the performances available in combinations of other members of the various chemical families. These differences are noted through the examples of two different propellant system groupings. The upper section of Table II utilizes two different members of the nitrogen oxide oxidizer family, N 0 and I R F N A (inhibited red fuming nitric acid), with several members and mixtures of the hydrazine and amine fuel family, N H , N H , CH N H (monomethylhydrazine, M M H ) , ( C H ) N H (unsym-dimethylhydrazine, U D M H ) , N H - ( C H ) N H (50-50), and M A F - 4 ( UDMH-diethylenetriamine, 60-40). As noted, there is an 8-10-second difference in I between the N 0 - and IRFNA-oxidized systems. The I performance range between various fuels is 23 seconds. 2

2

3

2

3

3

2

s

2

4

3

2

2

2

2

4

4

3

2

2

4


s

11.

S I L V E R M A N A N D CONSTANTINE

Table II.

309


Theoretical Specific Impulse of Selected Propellant Systems

1000 p.s.i.a. Shifting Equilibrium, Optimum Expansion to Sea Level Fuel Oxidizer N H^ 292 283 2

N 0 IRFNA 2

4

NH 269 260

CH N H 289 279

3

3

2

3

(CH ) N H 286 277 3 2

2

2

50% NH 50% (CH ) N H 289 280 3 2

2

r

2

2

MAF-4 282 274


Fuel Oxidizer

o

2

Flox-30 Flox-70 OF F 2

2

RP-1 311 325 350 347 344

300 317 344 341 318

The lower half of Table II demonstrates the performance differences between two members of the hydrocarbon family, C H and RP-1 ( a kerosene cut), with various oxidizer selections. The oxidizer selections are designed to demonstrate the change of performance between the oxygen and fluoride groups with various intermediate mixtures, flox 30 ( F - 0 , 30-70), flox 70 ( F - 0 , 70-30), and an oxyfluoride compound, O F . The difference between C H and RP-1 performance, as F is added to 0 , is particularly noteworthy. This difference is fairly consistent between the two as F is added; however, after reaching a maximum in each system (345 seconds at 7 1 % F with RP-1, and 353 seconds at 83% F with C H ) , the performance dropoff with additional F content occurs at lower F concentrations and is more rapid with the RP-1 fuel. This difference is a result of the RP-1, which has a hydrogen/carbon ratio ( H / C ) of —2, requiring more oxygen than C H ( H / C = 4 ) . In addition to the use of tables such as these to establish the theoretical performance potential (a like comparison can be made of I d) of various propellant systems under the desired operating conditions, the individual propellants are compared with respect to the other considerations. A n example of a comparison of liquid range of various propellants to a specific set of required conditions is shown in Table III. In this example, the operating temperature environment of the particular vehicle is —65° to + 1 6 0 ° F . W i t h this information, the liquidus ranges of various oxidizers and fuels are plotted under these conditions. (In the table, the solid line represents the range between the normal freezing and boiling points, with the end of the broken line representing the boiling point at 50 p.s.i.a.) Those propellants with liquid ranges entirely within the shaded area fulfill the initial liquid range requirements. 4

2

2

2

2

2

4

2

2

2

2

2

4

2

2

4

s


310


Table III.

N o r m a l L i q u i d Range


Oxidizers

Propellant

Tailoring

In many situations, the propulsion and vehicle system requirements cannot be met by the available propellant combinations. Such problems are often solved b y tailoring. This involves the formulation of a desirable set of characteristics by mixing selected ingredients. Several examples of new propellants that have been developed i n this manner are noted below. A n example of propellant tailoring is the fuel used to launch the first U . S. satellite into orbit. The original fuel for the launch vehicle was ethyl alcohol. M A F - 4 (also known as hydyne or U - D E T A ) , a mixture of 6 0 % U D M H and 4 0 % diethylenetriamine ( D E T A ) , was formulated to simulate the physical properties of C H O H but provide the increased propellant performance (using liquid oxyen as the oxidizer) requirements of the mission. Other M A F fuels, M A F - 1 ( D E T A - U D M H - a c e t o n i t r i l e , 50-40-10), and M A F - 3 ( U D M H - D E T A , 20-80) were formulated primarily to i n crease density. These formulations were based on density impulse requirements of various prepackaged propulsion systems as w e l l as to maintain the freezing point and viscosity characteristics of the U D M H . Formulation of the N H - U D M H (50-50) blend provided a relatively high performance but thermally stable hydrazine-type fuel suitable for regenerative cooling of large thrust chambers. 2

2

5

4


11.

S I L V E R M A N A N D CONSTANTiNE

311


of Selected Propellants


Fuels

The current RP-1 hydrocarbon fuel used i n high thrust boosters is an example of a special kind of tailoring. This hydrocarbon blend or distillation cut was selected to meet a series of special property and combustion requirements for liquid oxygen-oxidized high thrust systems. Propellant tailoring has also been studied as a means of upgrading the performance of present liquid oxygen systems by adding F . This potential performance improvement ( indicated i n Table II ) would apply to both hydrocarbon- and hydrogen-fueled systems. Experimental studies with both systems have verified the propellant system performance increases which can be realized. 2

Conclusions As a result of this constant evaluation and compromise between the demands of the vehicle and propulsion systems and the current propellant technology, various liquid propellant systems have been developed and are being applied i n current vehicle systems (Table I V ) . In Table I V thrust level is used to demonstrate the size of the propulsion system. Some of the systems i n this table have been phased out, while others are still i n development. However, Table I V does represent the current status of operational liquid bipropellant systems.


312


Table IV.

Propellant System, L0 /RP-1 2

Current Liquid Bipropellant Applications Engine System, Common Designation


2

2

Application Area

Thor

170,000

Space booster, I R B M (Delta 1st stage)

H-l

200,000 300,000 300,000 (1st stage) 80,000 (2nd stage)

Uprated Saturn I Space booster ICBM

Atlas

370,000-390,000

F-l

1,500,000

Space booster, ICBM Saturn V , S-IC

Centaur

15,000

Blue Streak Titan I

L0 /H

Thrust Level, lbs.

Uprated Saturn I, SIII (2 engines) Saturn I, SII (6 engines)

L0 /NH 2

3

LO /C H 0H H 0 /JP-4, JP-5 2

2

2

5

2

IRFNA/UDMH

J-2

200,000

Saturn V , SII & SIII

X-15

15,000-58,000

Experimental rocket plane

Redstone AR-2, 3

75,000 3300-6600

SRBM Auxiliary rocket engine for aircraft

Warrior

3500-10,200

Auxiliary rocket engine for aircraft

AJ10-118

7500

Space booster (Delta 2nd stage) Surface-to-surface missile

Lance Agena IRFNA/MAF-1

Bullpup

IRFNA/MAF-3 IRFNA/MAF-4

TD-174 P4

16,000

550 (booster) 106 (sustainer)

Space booster (2nd stage) Air-to-surface missile Air-to-air missile Target drone


11.

SILVERMAN A N D CONSTANTINE


Table I V .


Propellant System,

Engine System, Common Designation

313

Continued Thrust Level, lbs.

Application Area

IRFNA/JP-4

Aerobee 100 sustainer

2600 (total)

Sounding rocket

IRFNA/anilinefurfuryl alcohol

Aerobee 150 and 150A sustainer

4100 (total)

Sounding rocket

IWFNA/turpentine N 0 /N H U D M H (50-50)

Emeraude Apollo service module RCS Ullage rocket

62,700 100

Space booster Reaction control

1750

AJ10-131

2200

F750 L 2 . 2 K

2200

Saturn V , S-IVB Ullage control General purpose space engine Multiple restart space engine

Lunar module ascent engine

3500

Lunar module liftoff (moon)

F720 L 8 . 0 K

8000

Multiple restart space engine

Transtage

8000

Lunar module descent engine

1050-10,500

Upper stage propulsion Lunar landing engine

Apollo service module

21,900

2

4

2

4

N 0 /N H U D M H (50-50) 2

4

2

4

YLR113-AJ-1 Titan II

N 0 /MMH 2

4

Advanced Syncom RCS Gemini RCS Transtage ACS Apollo command module

N 0 /MMH 2

4

Gemini OAMS Radiomic

50,000-150,000 430,000 (1st stage) 100,000 (2nd stage)

Space propulsion Rocket sled Space booster, ICBM

Reaction control 25 (two 8-engine sets) 25 (4 engines) 45 (4 engines) 93 (two 6-engine sets) 25 (8 engines) 85 (2 engines) 100 (6 engines) 85-100

Reaction control Attitude control Attitude control Orbital change General purpose attitude control


314


Table I V . Continued Thrust Level, lbs.

N 0 /JP-X MON/MMH Downloaded by NORTH CAROLINA STATE UNIV on May 8, 2013 | http://pubs.acs.org Publication Date: June 1, 1969 | doi: 10.1021/ba-1969-0088.ch011

2

4

Air turbo rocket Surveyor vernier engine

MON/UDMH

Application Area

22 48 (2 engines)

Reaction control Agena secondary propulsion

500 30-104

Drone Attitude and velocity control

16 (2 engines) 200 (2 engines)

Agena secondary propulsion

Table I V shows that the L 0 / R P - 1 and N 0 / N H - U D M H (50-50) propellant combinations have emerged as the current workhorse propellant systems for first stage or booster applications. These systems are relatively inexpensive, have good physical and engineering properties, and reflect a high degree of development. Both have been and are systems being used i n I C B M applications, although the former is a nonhypergolic cryogenic system, while the latter is a hypergolic storable combination. Current upper stage systems are based on propellant combinations with performances ranging from those of the I R F N A / U D M H and the N 0 / N H - U D M H (50-50) systems, to the high levels achieved with the L 0 / H system. The development of the nitrogen oxide-type oxidizer/ hydrazine-(or amine)-type fuel propellant system for second-stage application was based on the need for wide liquid ranges and hypergolicity, whereas the development of the L 0 / H system for upper stage application responded to the need for high performance. The spacecraft control propulsion systems utilize primarily moderate performing nitrogen oxide/hydrazine systems. These systems are reliably hypergolic with good storability and stability under the programmed environmental conditions. The present I C B M system requirements are identical to those of the large boosters. The present performance and prepackageable requirements of the ordnance systems are well suited to the combination of I R F N A with various prepackageable hydrazines and amines. 2

2

4

2

2

2

4

2

4

4

2

2

2

Nomenclature Fuels D E T A : Diethylenetriamine, ( N H C H ) N H M M H : Monomethylhydrazine, C H N H U D M H : wnsym-Dimethylhydrazine, ( C H ) N H 50-50 F u e l blend: 5 0 % N H - 5 0 % U D M H 2

2

3

4

2

2

3

3

2

2

2

2

4


11.

315



Hydradyne V : 75% N . H - 2 5 % M M H M A F - 1 : 50.5% D E T A - 4 0 . 5 % U D M H - 9 . 0 % C H C N M A F - 3 : 80% D E T A - 2 0 % U D M H M A F - 4 (Hydyne, U - D E T A ) : 60% U D M H - 4 0 % D E T A JP-4, JP-5: Hydrocarbon fuels (kerosene cut) developed for jet pro pulsion applications J P - X : 60% TP-4-40% U D M H R P - 1 : A hydrocarbon fuel blend (kerosene cut) developed for rocket propulsion applications 2

4

3


Oxidizers Flox: Various mixtures of liquid fluorine and liquid oxygen (e.g., flox 30 is 30% F - 1 0 % 0 ; flox 70 is 70% F - 3 0 % 0 ) I R F N A (inhibited red fuming nitric a c i d ) : 84.4% H N 0 - 1 4 % N 0 1% H O-0.6% H F I W F N A (inhibited white fuming nitric a c i d ) : 97.5% H N O - 1 . 5 % H O-0.3% N O -0.7% H F L o x ( L 0 ) : L i q u i d oxygen, 0 M O N : Mixtures of N O and N 0 (e.g., M O N 10 is 10% N O - 9 0 % N 0 ; M O N 30 is 30% N O - 7 0 % N 0 ) 2

2

2

2

3

2

4

2

a

2

2

4

2

2

2

2

4

4

2

4

Literature Cited (1) (2) (3)

Huff, V. N., Gordon, S., Morrell, V. Ε., ΝACA Rept. 1037 (1951). Stull, D. R. et al., "JANAF Thermochemical Tables," Dow Chemical Co., Midland, Mich., Aug. 1965. Sutton, G. P., "Rocket Propulsion Elements, An Introduction to the Engi neering of Rockets," Wiley, New York, 1963.

RECEIVED April 14,

1967.


Current Liquid Propellant Systems - Advances in Chemistry (ACS

Recommend Documents