5
Propulsion Capabilities for Earth-to-Orbit Vehicles

HISTORY OF LIQUID AND SOLID PROPULSION

The liquid rocket propulsion systems available in the United States stem from three generations of development. First, there are the liquid bipropellant engines developed in the early 1950s and 1960s for the first generation of intermediate-range and intercontinental ballistic missiles. These include the liquid-oxygen/hydrocarbon engines for the Thor and Atlas, and the nitrogen-tetroxide/Aerozine-50 engines for the Titan. Some of these went out of production, but are now available in somewhat improved versions for the Delta II, Atlas II, and Titan II. Titan III, Titan-34D, and Titan IV launch vehicles are major modifications of the Titan II. Titan IVs are still being produced and are expected to continue in production throughout this decade.

In the second generation are the engines developed for the civil space program in the 1960s. These include the very large F-1 liquid-oxygen/hydrocarbon engine for the first stage of the Saturn V, the J-2 liquid-oxygen/liquid-hydrogen engine for the Saturn upper stages, and the relatively small RL10 liquid-oxygen/liquid-hydrogen engine, which is used for the Centaur upper stage. Some other upper-stage engines were also developed in this time period for DoD, such as the nitrogen-tetroxide/Aerozine-50 Transtage engine.

The third generation of liquid engine development produced only the Space Shuttle Main Engine (SSME), a sophisticated, high-pressure, liquid-oxygen/liquid-hydrogen engine tailored specifically to the requirements of the Space Shuttle.

Solid rocket motor developments were initially for jet-assisted takeoff (JATO) and small missiles. The first large high-performance motor design was for the Polaris submarine-launched ballistic missiles, where the requirements for storeability and the logistics of shipboard operations made the solid rocket very attractive. These same characteristics led to their use in the Minuteman silo-deployed intercontinental ballistic missiles (ICBMs). These capabilities provided the base upon which the technology for very large solid rocket boosters, suitable for space launch vehicles, was built. A segmented, 156-inch-diameter solid motor demonstrated the feasibility and practicality of segmented, solid rocket motors, a concept employed as a strap-on booster in the Titan III launch system. When the Space Shuttle program was initiated, solid



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From Earth to Orbit: An Assessment of Transportation Options 5 Propulsion Capabilities for Earth-to-Orbit Vehicles HISTORY OF LIQUID AND SOLID PROPULSION The liquid rocket propulsion systems available in the United States stem from three generations of development. First, there are the liquid bipropellant engines developed in the early 1950s and 1960s for the first generation of intermediate-range and intercontinental ballistic missiles. These include the liquid-oxygen/hydrocarbon engines for the Thor and Atlas, and the nitrogen-tetroxide/Aerozine-50 engines for the Titan. Some of these went out of production, but are now available in somewhat improved versions for the Delta II, Atlas II, and Titan II. Titan III, Titan-34D, and Titan IV launch vehicles are major modifications of the Titan II. Titan IVs are still being produced and are expected to continue in production throughout this decade. In the second generation are the engines developed for the civil space program in the 1960s. These include the very large F-1 liquid-oxygen/hydrocarbon engine for the first stage of the Saturn V, the J-2 liquid-oxygen/liquid-hydrogen engine for the Saturn upper stages, and the relatively small RL10 liquid-oxygen/liquid-hydrogen engine, which is used for the Centaur upper stage. Some other upper-stage engines were also developed in this time period for DoD, such as the nitrogen-tetroxide/Aerozine-50 Transtage engine. The third generation of liquid engine development produced only the Space Shuttle Main Engine (SSME), a sophisticated, high-pressure, liquid-oxygen/liquid-hydrogen engine tailored specifically to the requirements of the Space Shuttle. Solid rocket motor developments were initially for jet-assisted takeoff (JATO) and small missiles. The first large high-performance motor design was for the Polaris submarine-launched ballistic missiles, where the requirements for storeability and the logistics of shipboard operations made the solid rocket very attractive. These same characteristics led to their use in the Minuteman silo-deployed intercontinental ballistic missiles (ICBMs). These capabilities provided the base upon which the technology for very large solid rocket boosters, suitable for space launch vehicles, was built. A segmented, 156-inch-diameter solid motor demonstrated the feasibility and practicality of segmented, solid rocket motors, a concept employed as a strap-on booster in the Titan III launch system. When the Space Shuttle program was initiated, solid

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From Earth to Orbit: An Assessment of Transportation Options rocket boosters similar to the Titan III booster design were selected for the first stage. This, along with the continuing production of both sea-based and land-based missiles through the 1980s, has led to continuous technology development for large solid motors. The result of this process is that the U.S. rocket engine industry is able to produce modern solid propellant motors in a wide range of sizes. However, large liquid rocket engines are based on technologies from the 1950s and 1960s, with the exception of the SSME. Meanwhile foreign entities, mainly Arianespace, the National Space Development Agency of Japan (NASDA), and the former Soviet Union have proceeded with liquid engine developments established on a more recent technology base. FLIGHT-PROVEN U.S. AND INTERNATIONAL ENGINES AND MOTORS In the following sections, flight-proven engines and their international competition are briefly reviewed and assessed. Comprehensive descriptions of them are available from various sources.1,2,3 Descriptive data for flight-proven and proposed U.S. and international engines and motors are presented in Tables 4 through 7. Flight-Proven U.S. Engines and Motors F-1 and J-2 Engines The first stage, liquid-oxygen/hydrocarbon F-1 and the upper stage, liquid-oxygen/liquid-hydrogen J-2 engines were central to the success of the Saturn V Apollo launch vehicle. This capability was essentially abandoned at the conclusion of the SkyLab program with the advent of the Space Shuttle development, and the engines have not been in production since then. Sixty-five F-1 engines were launched successfully on 13 Saturn vehicle flights. The F-1 engine is not presently designed for throttleable thrust or for reusability, but these options are available on the proposed F-1 upgrade, the F-1A. To date, two F-1A engines have been ground tested. Both the F-1A and the J-2 engine could become part of a family of propulsion systems for near 1   Isakowitz, Steven J. 1991. International Reference Guide to Space Launch Systems. American Institute of Aeronautics and Astronautics. Washington, D.C. 2   U.S. Congress, Office of Technology Assessment. 1988. Launch Options for the Future: Buyer's Guide. 07A-ISC-383. U.S. Government Printing Office. Washington, D.C. 3   Wilson, Andrew, ed. 1991. Interavia Space Directory: 1991–92. Jane's Information Group. United Kingdom.

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From Earth to Orbit: An Assessment of Transportation Options TABLE 4 Characteristics of U.S. and International Liquid-Oxygen/Hydrocarbon Engines Parameter US: Delta II US: Atlas II US: Apollo US: Russia: Energia Russia: Vostok Russia: Proton Japan: H-I Engine designation RS 27A MA-5A F-1 (out of production) F-1A (proposed) RD-170 RD 107 Upper Stage Engine MB3 BLK III Thrust (1,000 lb) 200 (SL) 423 (SL) 1,552 (SL) 1,800 (SL) 1,632 (SL) 185 (SL)   170 (SL)   237 (vac)   1,748 (vac) 2,020 (vac) 1,777 (vac) 225 (vac) 19.1 (vac)   Isp (s) 255 (SL) 264 (SL) 264 (SL) 271 (SL) 309 (SL) 257 (SL)   253 (SL)   302 (vac)   305 (vac) 304 (vac) 337 (vac) 314 (vac) 352 (vac)   Chamber pressure (psia) 700 721 982 1,161 3,556 848 1,124 569 Feed system Gas generator Gas generator Gas generator Gas generator Staged combustion Gas generator Staged combustion Gas generator Mixture ratio 2.25 2.25 2.27 2.27 2.43 2.47 2.6 2.15 Throttling capability 100% only 100% only 100% only Yes (Range unknown) 49–102% 100% only 100% only 100% only Expansion ratio 12:1 8:1 16:1 16:1 26:1 Not available 189:1 8:1 Restart capability No No No No No No 1–7 starts No   SOURCES: Isakowitz, Steven J. 1991. International Reference Guide to Space Launch Systems. AIAA; and manufacturers data sheets.

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From Earth to Orbit: An Assessment of Transportation Options TABLE 5 Characteristics of Flight-Proven U.S. and International Liquid-Oxygen/Liquid-Hydrogen Engines Parameter US: Shuttle US: Apollo Japan: H-I US: Atlas US: Atlas Europe: Ariane-4 Russia: Energia-(core stage) Engine designation SSME (in production) J-2 (out of production) LE-5 (in production) RL10 A-3-3A (in production) RL10 A-4 (ground tests only) HM7B (in production) RD-120 (in production) Thrust (1,000 lb) 470 (vac) 230 (vac) 28.1 (vac) 16.5 (vac) 20.8 (vac) 14.1 (vac) 441 (vac) Isp (s)             354 (SL)   453 (vac) 425 (vac) 448 (vac) 444 (vac) 449 (vac) 444 (vac) 453 (vac) Chamber pressure (psia) 3,200 763 526 475 575 508 3,000 Feed system Staged combustion Gas generator Gas generator Expander Expander Gas generator Gas generator Mixture ratio 6.0 5.5 5.5 5.0 5.5 4.8 6.35 Throttling capability 65–104% 100% only 100% only 100% only 100% only 100% only Yes (Range unknown) Expansion ratio 77.5:1 28:1 140:1 61:1 84:1 62:1 85:1 Restart capability No One restart Yes Yes Yes No No   SOURCES: Isakowitz, Steven J. 1991. International Reference Guide to Space Launch Systems. AIAA; and manufacturers data sheets.

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From Earth to Orbit: An Assessment of Transportation Options TABLE 6 Characteristics of Flight-Proven U.S. and International Nitrogen-Tetroxide/Hydrazine-Based Engines Parameter US: Titan III Europe: Ariane-4 Russia: Proton Russia: Kosmos China: Long March Engine designation LR87-AJ11 Viking RD-253 RD 216 YF-20 Thrust (1,000 lb)   152 (SL) 331 (SL) 165 (SL) 166 (SL)   548 (vac) 171 (vac) 368 (vac) 194 (vac)   Isp (s)   248 (SL) 285 (SL) 248 (SL) 259 (SL)   302 (vac) 278 (vac) 316 (vac) 291 (vac) 289 (vac) Chamber pressure (psia) 829 848 2,130 1,066 Unknown Feed system Gas generator Gas generator Gas generator Gas generator Gas generator Mixture ratio 1.9 1.7 2.69 Not available Not available Throttling capability 100% only 100% only 100% only 100% only 100% only Expansion ratio 15:1 10.5:1 26.1:1 Not available Not available Restart capability No No No No No   SOURCES: Isakowitz, Steven J. 1991. International Reference Guide to Space Launch Systems. AIAA; and manufacturers data sheets.

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From Earth to Orbit: An Assessment of Transportation Options TABLE 7 Characteristics of the Flight-Proven U.S. RSRM and the Proposed ASRM Solid Rocket Motors Characteristics/Motor RSRM ASRM Diameter/length (in) 146/1,513 150/1,513 Thrust (1,000 lb,) 2,590 (vac) 2,636.6 (vac) Isp (s) 267.9 (vac) 269.18 (vac) Expansion ratio 7.72 7.54 Motor weight (lb) 1,255,978 1,350,381 Propellant mass fraction 0.882 0.895 Inert weight (lb) 148,809 142,313 Metal case weight/segments 98,740/4 99,442/3 Single nozzle weight (lb) 23,965 18,217 Solid propellant type PBAN HTPB Thrust vector control Flexible bearing Flexible bearing Recovery/reuse Yes Yes   SOURCE: Bardos, Russell. November 14, 1991. (National Aeronautics and Space Administration Headquarters). ''Advanced Solid Rocket Motor and a Comparison of the RSRM and ASRM.'' term application but would require new tooling, updating to modern materials and manufacturing procedures, and requalification to today's standards. The Committee recommends that the F-1A and J-2 engines be evaluated for future applications. MA-5 and RS-27 Engines The Rocketdyne MA-5 and RS-27 are engines used on the Atlas and Delta launch vehicles. Both engines were originally developed for the Atlas and Thor missiles, respectively. They use liquid oxygen and a hydrocarbon fuel, a relatively low chamber pressure, and a gas generator cycle that results in a relatively low specific impulse (i.e., performance) for this propellant combination. Having been designed originally for missile applications, these engines

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From Earth to Orbit: An Assessment of Transportation Options do not incorporate the elements of redundancy or the design margins desired for maximum reliability on a new class of vehicles. LR-87 and LR-91 Engines The Aerojet LR-87 and LR-91 engines power the core of the Titan launch vehicle family and have had a history of successful flights. They operate on storable propellants, nitrogen tetroxide and Aerozine-50, both now considered toxic materials. The use of these propellants has raised some concern regarding transport safety and environmental impact associated with launch due to their hazardous nature, and the Committee could identify no distinct advantages that are offered by these engines or their fuels. RL10 Engine The Pratt & Whitney RL1O is a high-performance liquid-oxygen/liquid-hydrogen upper-stage propulsion system that has been in production for well over 20 years. Its expander cycle leads to performance characteristics that are still attractive for modern upper-stage propulsion application. It is used in the Centaur stages for the Atlas and Titan launch vehicles. The design is amenable to upgrading to meet specific stage needs. A version of the RL1O is currently being modified for the one-third scale model (DC-X) of the SDIO Single-Stage Rocket Technology Program's Delta Clipper (DC-Y) presently scheduled for flight test in 1993. Space Shuttle Main Engine The Rocketdyne Space Shuttle Main Engine (SSME) development started more than 20 years ago as a reusable, liquid-oxygen/liquid-hydrogen engine for use on the Shuttle. To maximize the specific impulse and minimize its size and weight, the SSME was designed to develop extremely high chamber pressures and uses a staged combustion cycle that requires even higher turbopump delivery pressures. These characteristics led to severe problems in SSME development and pose continuing problems with the turbopumps. The SSME has been subject to ongoing improvements such as improved oxygen and hydrogen turbopumps under the Advanced Turbopump Program. It is the understanding of the Committee that funding for turbopump development has been partially eliminated in the proposed FY 1993 budget. However, the Committee considers development of the improved turbopumps necessary and the Advanced Turbopump Program of high priority for the future reliability of the Space Shuttle Main Engine.

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From Earth to Orbit: An Assessment of Transportation Options Ballistic Missile Solid Motors Various types of solid motors are currently used in intercontinental ballistic missile (ICBM) systems. These motors are suitable for use in several space launch vehicles with light payloads to low-Earth orbit (approximately 2,000 pounds). For example, the first stage motor developed for the Peacekeeper is also to be used as the first stage of the Taurus launch vehicle. Some of these solid motors may be available as a result of the recent military deactivation of missiles, but their availability is uncertain. In addition, solid motors are subject to aging, and some have been deployed in the field for as long as 30 years. Their condition and performance are not known with complete surety. Redesigned Solid Rocket Motor (RSRM) The Redesigned Solid Rocket Motor (RSRM) is the very large, segmented, solid booster currently used on the Space Shuttle. Two RSRMs provide 80 percent of the total liftoff thrust. The RSRM was derived from and replaced the Solid Rocket Motor (SRM) previously used on the Shuttle, and the main difference is the improved design of the field joint and O-ring seal between the case segments of the RSRM. Since the RSRM rocket case is reusable, it is retrieved, refurbished, refilled with propellant, and returned to Kennedy Space Center for future launches. Table 7 shows the characteristics of the RSRM. The Shuttle is currently operating satisfactorily with the RSRM. According to present planning, the RSRM is capable of meeting all operating requirements of the Shuttle, including the assembly and operational maintenance of Space Station Freedom. With continued meticulous attention to processing and improvements in nondestructive evaluation, there is every reason to believe that the Shuttle can be operated without failure of the RSRM. It should be noted, however, that even though the costs of inspection and operational precautions are high, they must not be allowed to become routine. The RSRM is now a well-understood propulsion system with established characteristics and, as such, the Committee believes that it is capable of safe operation with careful management. Flight-Proven International Engines and Motors The following is an evaluation of some specific foreign rocket engines and motors that the Committee believes to be especially relevant to the considerations in its charge. In overview, the foreign launch vehicles use propulsion systems incorporating modularity to various degrees, which brings a variety of benefits. Based on presentations made to the Committee, the Russians appear to be using noteworthy innovations in materials, manufacturing, and test methods that are worthy of pursuit by technology transfer to the United States.

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From Earth to Orbit: An Assessment of Transportation Options RD-170 Engine The Russian RD-170 is the high thrust liquid-oxygen/hydrocarbon modular engine used in the first stage of the heavy lift Energia launch vehicle. It is technologically comparable to the SSME, having very high chamber pressure, and is in the thrust range to compete with the F-1 engine used during Apollo and the F-1A (F-1 upgrade) proposed by Rocketdyne. The RD-170 incorporates the desirable features of throttling and reusability. It is currently in production, and more than 200 engines have been produced to date. The RD-170 is now used in the Zenit and Energia/Buran launch vehicles. At least twenty-five flights have been successfully completed. The qualification and flight status of this engine makes it attractive for direct application in a heavy lift launch vehicle liquid booster stage. Aerojet, Pratt & Whitney, and Rocketdyne are currently engaged in negotiations to test and/or acquire the RD-170 from the engine manufacturer, NPO Energomash. The Committee recommends that the performance of the RD-170 engine be evaluated thoroughly for application to U.S. systems. One way to hasten the evaluation could be as part of an international cooperative effort in which the National Aeronautics and Space Administration (NASA) might purchase several RD-170 engines and send personnel to monitor engine testing at Kiminsky, Russia. If the results of the performance evaluation are positive, the engines might be manufactured in Russia or under license in the United States. However, the Committee believes that no critical component, such as an engine, should be solely supplied from abroad. RD-253 Engine The Russian RD-253 is the first stage engine for the Proton launch vehicle. It has demonstrated excellent reliability with no known failures in more than 900 flight engines. However, its use of storable liquid propellants (nitrogen tetroxide and unsymmetrical dimethylhydrazine (UDMH)) could raise environmental concerns and pose logistic problems equivalent to those of the Titan launch vehicle. Viking and Other Engines The Viking engine (used on the Ariane-4 first and second stages) and the Chinese Long March engines use nitrogen tetroxide and hydrazine propellants. They do not appear to offer enough advantages to the United States in terms of the technology or the basic capability to merit further investigation.

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From Earth to Orbit: An Assessment of Transportation Options PROPOSED U.S. AND INTERNATIONAL ENGINES AND MOTORS Proposed U.S. Engines and Motors Space Transportation Main Engine (STME) The Space Transportation Main Engine (STME), Figure 2, utilizes liquid-oxygen/liquid-hydrogen propellants and is currently at the technology demonstration stage. Thus, the performance parameters shown in Table 8 are preliminary. The STME is envisioned as a low-cost, highly reliable engine. These goals are to be met through a philosophy of engine development that places reliability at a higher priority than performance. This engine is regarded as the central propulsion capability for the National Launch System (NLS). The STME builds on the technology obtained from development of the SSME, but proposes to use the technology in a less demanding way, emphasizing reliability and low cost rather than ultimate performance, as noted above. Because the STME engine builds on experience gained with the SSME, it is helpful to use the SSME as a basis for comparison. Figure 2 Characteristics of the proposed liquid-oxygen/liquid-hydrogen Space Transportation Main Engine (STME). The physical differences between the STME and the SSME are (1) lower STME chamber pressure, (2) utilization of a gas generator cycle in the STME, and (3) higher STME thrust. As stated previously, the main difference, however, is a dramatic change in design philosophy. The

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From Earth to Orbit: An Assessment of Transportation Options TABLE 8 Comparison of the STME and SSME Characteristics/Engine STME SSME Propellants LOX/LH2 LOX/LH2 Feed system Gas generator Staged combustion Thrust (1,000 lb) 650 (vac) 470 (vac) Throttling 70–100% 65–104% Isp (s) 428.5 (vac) 453 (vac) Chamber pressure (psia) 2,250 3,200 Mixture ratio 6.0 6.0 Expansion ratio 45:1 77.5:1   SOURCE: NASA/George C. Marshall Space Flight Center; and manufacturers data sheets. primary drivers of the SSME development were high delivered performance (Isp) and low weight. The stated objective of the STME design is to give priority to reliability, manufacturability, and reduced cost—at the expense of performance. This design philosophy influenced the choice of lower chamber pressure and the use of the gas generator cycle, design factors that result in a lower delivered specific impulse for the STME. Attendant with these choices is a significant reduction in the turbopump outlet pressure that, in turn, is expected to result in a more robust design. Further, the reduced chamber pressure decreases the heat transfer load to the thrust chamber walls, which also contributes to increased design margins. The stated design philosophy is also geared to producibility of the engine in large numbers at low cost. This has resulted in a closer look at engine materials and manufacturing methods to be employed. The use of low-cost castings in lieu of machined forgings, minimization of weld joints, and relaxed tolerances where possible all are elements of the manufacturing technology aspects of the STME design. The Committee enthusiastically supports the STME design philosophy. Once the design thrust level has been finalized, the Committee trusts that the urge to relax this philosophy and move toward higher performance will be avoided diligently, even if initial development and tests are highly successful. Escalating the performance requirements during development would negate the positive features of the design.

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From Earth to Orbit: An Assessment of Transportation Options The Committee believes the STME development should proceed immediately and vigorously and should be tailored to optimum application in a launch vehicle with a payload capacity of approximately 20,000 pounds to low-Earth orbit (LEO). The proposed thrust level of the STME is 650,000 pounds. The Committee endorses this design thrust level, with the assumption that the vehicle will be launched at less than maximum STME thrust level for improved reliability and margin. The Committee assumes that the initial sizing of the STME not only will accommodate the requirements of a 20,000-pound payload vehicle, but also will allow STME utilization as a core stage engine in our future launch vehicle fleet, along with strap-on booster propulsion modules to augment the payload capabilities of larger vehicles. This approach is consistent with highly successful launch vehicle experience in the international space community, as evidenced by the extensive Russian inventory of strap-on propulsion systems, the Ariane family of launch vehicles, and the Delta launch vehicles. Advanced Solid Rocket Motor (ASRM) The Advanced Solid Rocket Motor (ASRM) is a solid rocket booster intended to replace the RSRM for the Space Shuttle vehicle, providing a proposed 12,000-pound Shuttle payload increase. It incorporates a larger diameter case; advanced, light-weight nozzle; different field joint design; and continuous processing of propellant. The ASRM is intended to be produced in a government-owned, contractor-operated facility, currently under construction in Yellow Creek, Mississippi. The characteristics of the ASRM are shown in Table 7. The projected benefit of the ASRM is an anticipated 12,000-pound payload increase. The significant savings due to reduction in the number of Shuttle missions to support SSF could warrant the use of the ASRM, provided program and technical risks can be fully contained. The Committee is aware of the large investment that has been made in ASRM development and in related facilities; however, there is considerable risk remaining. In particular, the use in the case segments of welded circumferential joints in place of mechanical joints, and a steel alloy that has not previously been used in large rocket cases, introduce significant risk, as pointed out by the 1991 NRC Committee on Advanced Solid Rocket Motor Quality Control and Test Program. Continuous casting of the propellant also is unproven for the large sized motors of the ASRM. These and other new features, such as the lighter-weight nozzle design, could lead to program extensions or failures requiring expensive corrective action. The current RSRM is capable of meeting all operational requirements of the Space Shuttle. The Committee believes that the balance of costs, technical and programmatic risks, and potential benefits tips in favor of avoiding integration of the ASRM into the Space Shuttle system at this time. Regarding the utility of the ASRM for other future space launch systems, the Committee understands its potential as a strap-on for the heavy payload end of NLS, (i.e., NLS-1); however, the Committee has found no compelling rationale for such use other than the fact that

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From Earth to Orbit: An Assessment of Transportation Options it might be introduced in a reasonably short time. The Committee believes that NASA should rely on the current RSRMs and that the ASRM program should be reconsidered. It believes that NASA and the nation would be well served if the development of the NLS were directed toward strap-on boosters that have pad hold-down with engine shutdown capability and throttleability as means toward increased reliability. In its discussions, the Committee found a number of considerations that favor liquid propulsion systems as compared to solid propulsion systems. Liquid rocket engines permit a more flexible approach to modular clustering and are amenable to verification before launch. However, the most compelling characteristics favoring liquids are throttleability and thrust termination capability, which can enable first-stage booster designs to incorporate an engine redundancy capability. A recent study4 of the causes of liquid propulsion system failures has shown that a large fraction of such failures were noncatastrophic in nature; that is, they resulted in loss of thrust from the engine, but not in a failure that would destroy the vehicle or other engines. This is termed a benign failure. As mentioned earlier, the historical ground and flight data have shown the benign failure ratio (failures/engine flights) for liquid propulsion systems is greater than the catastrophic failure ratio. For solids, there have been no benign failures for flight motors.5 When benign failures dominate, reliability enhancement through incorporation of active redundancy becomes feasible and often economically attractive. Indeed, as far as the propulsion system is concerned, active redundancy is a means of improving reliability. It is a proven approach for aircraft but, in the past, has not been employed on expendable launch vehicles. However, the very high cost of failures, particularly when expensive payloads are involved, makes the use of these techniques worthwhile even for expendable vehicles. Solid Rocket Motor Upgrade (SRMU) The Solid Rocket Motor Upgrade (SRMU) is a segmented rocket motor design being developed to provide more lift capability for the Titan IV launch vehicle. It incorporates a composite case, but no other noteworthy technological advantages over the RSRM and the ASRM. 4   Leonard, B.P. 1991. Projected Launch Vehicle Failure Probabilities With and Without Engine Segment-out Capabilities. L-Systems, Inc. El Segundo, Calif. 5   Smith, O. Glenn. 1991. Reliability in Manned Transportation Systems. Johnson Space Center. Houston, Tex.

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From Earth to Orbit: An Assessment of Transportation Options Proposed International Engines and Motors The advanced propulsion systems under development in Europe (Vulcain HM-60) and Japan (LE-7) are similar to existing technologies in the United States and are not discussed in detail in this report. However, the characteristics of some of the international engines are contained in Tables 9 and 10. TABLE 9 Characteristics of Proposed U.S. and International Liquid-Oxygen/Liquid-Hydrogen Engines Parameter US: NLS Europe: Ariane-5 Japan: H-II China: CZ2E/HO Engine designation STME Vulcain HM-60 LE-7 YF-75 Thrust (1,000 lb)   180 (SL) 190 (SL) 17.6 (SL)   650 (vac) 225 (vac) 243 (vac)   Isp (s) 428.5 (vac) 430 (vac) 445 (vac) 440 (vac) Chamber pressure (psia) 2,250 1,450 2,090 Not available Feed system Gas generator Gas generator Staged combustion Gas generator Oxidizer/fuel mixture ratio 6.0 5.1 6.0 Not available Throttling capability 70–100% 100% only 100% only Not available Expansion ratio 45:1 45:1 60:1 Not available Restart capability Multiple No No One restart   SOURCES: Isakowitz, Steven J. 1991. International Reference Guide to Space Launch Systems. AIAA; and manufacturers data sheets.

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From Earth to Orbit: An Assessment of Transportation Options TABLE 10 Characteristics of Proposed International Nitrogen-Tetroxide/Hydrazine-Based Engines Parameter Europe: Ariane-5 India: GSLV Engine designation L-7 VIKAS Thrust (1,000 lb) 6.1 (vac) 165 (vac) Isp (s) 316 (vac) 293 (vac) Chamber pressure (psia) 218 763 Feed system Pressure fed Gas generator Mixture ratio 2.05 1.7 Throttling capability 50–100% 100% only Expansion ratio 83:1 Not available Restart capability Multiple No   SOURCES: Isakowitz, Steven J. 1991. International Reference Guide to Space Launch Systems. AIAA; and manufacturers data sheets. ENGINE AND MOTOR TESTING System Reliability and Tests One point that is not always clear in discussions of the reliability of liquid rocket engines versus solid rocket motors is that the solid rocket motor test includes all facets of the propellant delivery system, whereas the liquid rocket engine may not. Specifically, the liquid rocket engine will include the turbopump (if any), which accepts the propellant flow from a facility source that

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From Earth to Orbit: An Assessment of Transportation Options is not flight weight or flight type in its design and operation. The analysis of propulsion system reliability mentioned previously indicated that significant portions of liquid propulsion system unreliability are associated with the system upstream of the turbopump (tanks, low-pressure feed lines, valves, etc). It is the opinion of the Committee that increased emphasis on propulsion system tests, including the propellant feed system, should be a major aspect of any new launch system program. Increased emphasis is also required in the design phase to include innovative methods to monitor propulsion system health and implement any required shutdowns at appropriate locations. The Committee found no evidence that this design approach is currently being addressed in the NLS program. Propulsion Test Facilities The nation's capability for testing liquid rocket engines and solid rocket motors is becoming constrained by the pressures of residential and commercial development in the vicinity of rocket manufacturers' facilities and the increasingly stringent environmental standards being imposed. The following is a summary of the most likely test sites for future engine and motor development activities, including a brief assessment of potential restrictions and constraints. Government-Owned Test Facilities The major government-owned engine and motor test facilities are: Edwards Air Force Base (AFB), California—The F-1 development test stands (6 in number) are still in place and could be reactivated in approximately 6 months' time. For instance, stand C-1 was converted from a liquid engine test capability to accommodate a Titan 34 solid strap-on test capability (nozzle down, vertical firing) within six months. It is presently being used to test the Solid Rocket Motor Upgrade (SRMU) for the Titan IV. Test stand 2-A is being modified to accommodate the STME combustion device development program. The area of Edwards AFB is substantial and, to date, the emissions have been able to meet state and local county standards by the time the exhaust cloud reaches the base boundaries. Although the area to the west of Edwards AFB is growing, the prevailing winds are from the southwest. Only the town of Boron is in the area of likely coverage by exhaust clouds, and it is far enough away to ensure that no danger is presented. Arnold Engineering Development Center, Tullahoma, Tennessee—The Air Force has established a number of simulated altitude test facilities that can accommodate both liquid

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From Earth to Orbit: An Assessment of Transportation Options rocket engines and solid rocket motors of very large size. For example, the Peacekeeper second-stage motor was tested there. NASA/Marshall Space Flight Center, Huntsville, Alabama, has a number of large liquid rocket test facilities capable of handling SSME engines as well as larger engines. The buffer zone is quite large, but the issues of noise, air, and groundwater pollution have created questions leading to interaction of NASA/Marshall with various state and local authorities and residents. NASA/Stennis Space Center, Mississippi, has 4 very high thrust test stands available for liquid-oxygen/liquid-hydrogen rocket engine testing. Three stands will be in active use until 1994 for SSME testing. NASA is presently evaluating reactivation of the other stand. Currently, a horizontal, solid rocket test stand for the ASRM is being built. Some of the local population is opposed to solid rocket testing in the area. The concern is the combination of high humidity and the presence of hydrogen chloride (HCl) in the solid rocket exhaust. This situation must be monitored closely to see how it develops. Industrial Test Facilities The following are the major industrial engine and motor test facilities: The Thiokol Corporation in Wasatch, Utah; Hercules, Incorporated in Magna, Utah; and United Technologies in Coyote, California, have test facilities suitable for large solid rocket development efforts. There are presently no known environmental constraints that may affect the ability to continue such testing. Rocketdyne, Santa Susanna, California, has extensive liquid rocket test facilities in the Santa Susanna Mountains. Since initial construction, residences have gradually been built closer and closer to the gates of the facility. Strict Los Angeles County air quality and groundwater standards must be met. The impact on the cost of testing and the ability to conduct tests at will is unknown, but believed to be restrictive. Aerojet Propulsion Division, Sacramento, California, has extensive liquid and solid rocket test facilities in the Folsom area. There has been substantial commercial development in the area. As a result, Aerojet must restrict test operations to certain atmospheric (wind) conditions to ensure that no effluents pass over occupied buildings outside test facility boundaries.

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From Earth to Orbit: An Assessment of Transportation Options Each of the above testing sites is unique and must be examined individually with regard to the environmental impact. ENVIRONMENTAL EFFECTS OF CHEMICAL PROPULSION Because it has received much discussion and is a source of current controversy, the question of environmental pollution by chemical propulsion commands detailed discussion. The Committee, upon review of the American Institute of Aeronautics and Astronautics (AIAA) Workshop Report on the ''Atmospheric Effects of Chemical Rocket Propulsion,''6 accepts the conclusions contained therein. Specifically, the effects of currently projected launches of the Shuttle and Titan IV annually will result in the depletion of ozone globally on the order of 0.0065 to 0.024 percent; ozone depletion in the 60–90 N latitudes will be about 0.2 percent; ozone depletion near the exhaust plume is greater than 80 percent during ascent but returns to normal within three hours; acid rain is less than 0.01 percent of global anthropogenic sources; local launch effects are confined to within 2,500 feet of the launch pad; effects on global warming are less than one-millionth of all carbon dioxide (CO2) produced; and for the current projected Shuttle and Titan IV launch rates, local toxicity hazards and air quality are considered manageable. These facts, which are hardly in dispute, have not allayed concern in the community of environmentalists regarding the effects of space launch vehicle emissions on the atmosphere. These concerns, often expressed in language that suggests dire global consequences of continuing to launch payloads into space, are not supported by the findings of the broadly based group of scientists and engineers encompassing the propulsion and launch vehicle industries. Nevertheless, the protests do pose a possible problem in terms of public reaction to the continued or expanded use of large solid boosters. Because of concern over the potential detrimental environmental effects of some launchers, the Committee endorses continuing research to better identify and understand these effects. Data suggest that pollution due to combustion products from launch vehicles, at the frequency and scale that is anticipated, is not significant in comparison with other 6   American Institute of Aeronautics and Astronautics. Atmospheric Effects of Chemical Rocket Propulsion. October 1, 1991. Washington, D.C.

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From Earth to Orbit: An Assessment of Transportation Options anthropogenic pollution on a global scale. It is, however, a serious local concern in the vicinity of launch test sites and deserves further investigation. PRIORITIES FOR INVESTMENT A plan is needed to provide an array of engines with a range of thrust levels and propulsion system capabilities for all stages of future launch vehicles. The proposed STME can be used for first stages of future launch vehicles. In addition, the Rocketdyne F-1A and the current Russian RD-170 engines should be evaluated for liquid booster applications. Hybrid engine technology should be investigated to determine its suitability with new launch systems. The propulsion system technology for second-and third-stage applications should emphasize low system cost and high reliability as in the initial booster stage. Revival of the Rocketdyne J-2 or upgrading of the Pratt & Whitney RL10 liquid-oxygen/liquid-hydrogen engines could provide upper-stage engines fitting the range of desired propulsion capabilities. Space Transportation Main Engine The modern Space Transportation Main Engine (STME) is the beneficiary of a significant investment in the underlying technologies, including fabrication methodologies, advanced materials, and revised design procedures aimed at improving the reliability and cost of a production engine. Development and qualification of the Space Transportation Main Engine (STME) should proceed immediately and vigorously. The initial use of the STME is to propel the NLS vehicles, and its development is crucial. In addition, it is the Committee's recommendation that the STME development team invest in demonstrating a prototype of the 650,000-pound thrust engine concept for the 20,000-pound payload vehicle. The design philosophy should reinforce the robustness aspects of the engine, and should include health monitoring and control technology. Space Shuttle Propulsion Systems The Space Shuttle will remain the centerpiece of the U.S. manned presence in space well into the first decade of the next century. The Committee believes that efforts underway to improve reliability, reduce the cost, and simplify production and refurbishment of the SSME should be continued and that investment should be made in maintaining the operating integrity and effectiveness of all aspects of the propulsion system.

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From Earth to Orbit: An Assessment of Transportation Options Because the engine has been plagued with problems, there is no question that Shuttle operations can be enhanced by continuing to upgrade the SSME. As noted earlier, a specific area of concern involves the turbopumps that deliver liquid-hydrogen and liquid-oxygen propellants to the engine main combustion chamber. Pratt & Whitney, under NASA contract, has delivered the first set of 16 improved oxygen turbopumps for the SSME. Budget cuts in NASA's Alternate Turbopump Program (ATP) made by Congress in the hydrogen fuel pump improvement program will delay the flight qualification required to show that the turbopumps are interchangeable with the current Rocketdyne turbopumps and that they have the required mission lifetime. It is important to pursue both the alternate oxygen and hydrogen pumps under the ATP because Shuttle operations are likely to continue well into the first decade of the next century, and it is believed that both turbopumps can offer greatly increased safety margins and reduced requirements for pump replacement/changeout. The Redesigned Solid Rocket Motor also deserves future investment to ensure the continued success of Space Shuttle operations. In particular, the RSRM could benefit from continued investment in improved manufacturing technology to ensure the repeatability of manufacture and nondestructive inspectability of the motor. PRIORITIES FOR LONGER-TERM PAYOFF Engines Major investment priorities have been listed in the foregoing paragraphs. The Committee believes that there are also other promising engines in early stages of development that require investment in experimental development and testing to validate their engineering practicality. These are the modular plug engine, the modular bell engine, dual-fuel engines, and variable mixture ratio liquid-oxygen/liquid-hydrogen engines. The technology for these engines is discussed further in Chapter 6. Booster Stages Although not required for the 20,000-pound payload class vehicle lift capability recommended by the Committee as the critical U.S. need, a modular approach to providing booster strap-ons is required for 135,000-pound (and greater) payload class vehicles. Propulsion candidates for booster rockets include liquid, solid, and hybrid rockets. Some characteristics of solids, liquids, and hybrids are shown in Table 11. Hybrid rocket boosters as shown in Figure 3, use a liquid oxidizer and a solid fuel, and have been studied and tested in a variety of thrust ranges since the 1960s. The Committee

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From Earth to Orbit: An Assessment of Transportation Options TABLE 11 Characteristics of Solid, Liquid, and Hybrid Boosters Characteristics Solid Boosters Liquid Boosters Hybrid Boosters Specific impulse Relatively low specific impulse Relatively high specific impulse Relatively low specific impulse Cost Lower up-front development and production costs Higher up-front development and production costs Not available Failure mode Failures tend to be catastrophic Failures tend to be noncatastrophic Failures tend to be noncatastrophic Shutdown/restart No shutdown capability in event of malfunction Shutdown possible in event of malfunction Shutdown possible in event of malfunction Active redundancy No capability Yes Yes Functional test capability after assembly No functional test of production article Functional test possible both after assembly and before liftoff Functional test of production articles possible before liftoff Environmental Emission of chlorides and aluminum oxide particulates Possible environmental effects are currently not well understood Environmental effects are comparable to those of a liquid/hydrocarbon booster Handling and storeability Fire hazard in handling, storage, and vehicle assembly Fully storable, but materials and bond are temperature sensitive and subject to aging Use of cryogenic liquids requires careful handling, storage, and launch preparations Use of cryogenic liquids requires careful handling, storage, and launch preparations Fuel grain design is more robust than that of a solid booster Complexity Operational simplicity Turbomachinery results in complexity Greater operational simplicity Reusability Reusability of cases Tankage probably not reusable Reusability of cases Propellant density High propellant density Relatively low density leads to larger volume High propellant density

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From Earth to Orbit: An Assessment of Transportation Options recognizes that the hybrid rocket has not achieved the flight-proven status of solid and liquid boosters. However, the attractive features of hybrid rockets such as the ability to control liquid oxidizer flow that permits throttling and engine shutdown make them viable candidates for thrust augmentation booster applications on future vehicles. Figure 3 Hybrid motor that uses liquid oxygen and a solid fuel. The liquid oxidizer flow can be controlled to permit throttling and engine shutdown.  Courtesy of Thiokol Corporation. The time scale for the development of a 135,000-pound payload class vehicle, would permit investment in hybrid rocket motor development now, so that its status is advanced to the point at which it can be quantitatively evaluated in competition with solid and liquid bipropellant systems designed to directly comparable criteria. Two liquid rocket engines that the Committee also believes are candidates for booster applications are the F-1A and the RD-170 engines as discussed earlier.