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3 Reusable Booster System Technology Assessment The statement of task for this study (see Appendix A) specifically calls for addressing the technical maturity of the key elements critical to the reusable booster system (RBS) implementation and the ability of current technology development plans to meet technical readiness milestones. This chapter identifies these key elements, addresses their technical maturity, identifies risks, and addresses potential risk mitigation activities. The technical approach at the system level that the Air Force presented to the committee for eventually devel- oping a flight-ready RBS is based on a three-step process. The combined descriptions of the proposed Air Force RBS program presented by the Air Force Research Laboratory (AFRL), the Air Force Space and Missile Systems Center (SMC), and the Aerospace Corporation involve first building a small-scale Pathfinder vehicle and flight testing it in the 2015 time frame to demonstrate the technical feasibility of performing the rocketback return-to- launch-site (RTLS) maneuver with horizontal landing. If the Pathfinder is deemed successful, then the next step in the RBS development progression would be to scale up to an intermediate-size reusable booster demonstrator (RBD) (approximately 63 ft long and having a dry weight of about 25,400 lb). The RBD would demonstrate the oxygen-rich staged combustion (ORSC)-engine-based main propulsion system (MPS) (using one NK-33 or AJ-26) and the aerodynamic scalability from the much lower cost Pathfinder to the intermediate-sized RBD. Following the RBD demonstration it was advocated by AFRL that the next step in the process should be another intermediate- sized vehicle called RBX; however, the baseline plan as presented by SMC advocated going directly from RBD to the full-scale RBS design, development, test and evaluation (DDT&E) program. 3.1 ASSESSMENT OF TECHNOLOGY MATURITY OF KEY ELEMENTS If it is accepted that significant cost savings will accrue through reuse of the first-stage, then it is necessary to determine whether the technologies needed to develop a reusable first stage for the RBS can be realized in an affordable development and certification program. To answer this question, it is necessary to first identify the new enabling technologies whose development is required, the risks associated with those developments, and the needed risk mitigation plans. As stated in Chapter 2, the RBS will require development of technologies that would enable the successful execution of the rocketback maneuver of its first stage and inspection confirmation of the reusability of the recov- ered first stage before its next scheduled flight. The committee judges that meeting these requirements will involve technology development in four principal risk areas: (1) high-performance hydrocarbon-fueled booster engines; 23
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24 REUSABLE BOOSTER SYSTEM--REVIEW AND ASSESSMENT (2) rocketback RTLS maneuvers; (3) integrated vehicle health management (IVHM); and (4) adaptive guidance and control (AG&C). Secondary technology risk areas include (1) lightweight structures that can handle the expected loads; (2) robust power, fluid, and actuator systems; (3) advanced assembly and manufacturing techniques; and (4) upper-stage LO2/LH2 engines. Note that application of these technologies carries some risk because they have to be applied to a new vehicle configuration with a flight profile that has never been flown before. The identified technology items will require various degrees of risk mitigation effort involving both analysis and testing, to achieve the technology readiness level (TRL) needed to proceed to substantial RBS development. Some of these technologies apply principally to reusable vehicles; others also have application to potential future expendable vehicles. The high-risk technologies and their application areas are summarized in Table 3.1. In the following subsections, the four principal technology risks are discussed, secondary risks are addressed and operational and infrastructure issues, as they relate to the RBS concept, are discussed. The chapter concludes with a summary discussion of RBS risk assessment and mitigation efforts. 3.2 MAIN PROPULSION SYSTEM There are two main propellant options for the RBS MPS, liquid oxygen/liquid hydrogen (LO 2/LH2) or liquid oxygen/rocket propellant (LO2/RP); each has two subchoices for the engine power cycle behind them, open or closed, for a total of four basic options. Because the committee believes that neither a pressure-fed liquid engine nor solid rocket boosters would be tractable options for the RBS, it limits discussion to these four options. The first options for the RBS MPS are a LO2/LH2 propulsion system using either an open-power cycle or a more efficient but higher pressure closed cycle. The open cycle can be either combustion tap-off or gas generator; TABLE 3.1 Highest Technology Risks Risk Area Risk Item Reusable Expendable Hydrocarbon-fueled Combustion instability X X booster engine Oxygen rich, staged combustion X X Power balance X X Physics-based analytical predictive models X X Injector X X Materials/coatings for O2-rich environment X X Turbomachinery X X Long-life bearings X Transients X X Requirements for vehicle integration X X Rocketback return to Sloshing/propellant management X launch site maneuver Plume interactions X Thermal management X Deep throttling X Structural dynamics X Aerodynamics X Kinematics and mass properties management X Integrated Vehicle Reliable/robust sensors X X Health Monitoring Real-time critical decision making: data to action X X (IVHM) Identify and develop nondestructive inspection options and quantify reliability X X System integration into asymmetric vehicle configuration X Adaptive guidance Integration with IVHM X X and controls Real-time control algorithms X X Fast response actuators X Software verification and validation X X
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REUSABLE BOOSTER SYSTEM TECHNOLOGY ASSESSMENT 25 the latter is used for the Delta IV RS-68 or the Saturn V upper-stage J-2 engine. An example of the closed-cycle engine is the staged-combustion design used on the space shuttle main engine (SSME), which is in the process of being simplified to a lower cost expendable version called the RS-25E. (The RS-25E is the baseline engine selected by the Air Force for the RBS expendable upper stage.) Yet another closed-cycle LO 2/LH2 engine could be based on an expander cycle such as is used for the RL-10 family of upper-stage engines. While an LO2/LH2 MPS will result in a much higher specific impulse (Isp) (approximately 390 sec at sea level) than a hydrocarbon-fueled engine, it is a nonoptimal choice for the RBS first stage, owing to the extremely low propellant density of hydrogen and its deep cryogenic properties, which must be maintained at approximately -420°F. The use of these deep cryogens results in large, heavy, and more complicated aerodynamic configurations as well as a relatively poor stage mass fraction. Since mass fraction is just as important as I sp in the basic vehicle "rocket equation" design solution, a higher density, more easily storable fuel such as kerosene is a better choice than liquid hydrogen for the RBS booster. The many operational advantages of using kerosene in the first stage, as compared to cryogenic hydrogen, provide a further rationale for the selection of a hydrocarbon fuel. The selection of a higher density fuel leads to the second set of options for the RBS MPS, which is engines that operate using liquid hydrocarbon fuels, such as rocket propellant (RP)-1, as recommended and advocated by the Air Force in their presentations to the committee. Liquid methane or even ethane or propane (such as the main constituents of liquefied natural gas) might also be a good choice for a higher density, better performing fuel, but there is only limited technology experience with these fuels and LO2 oxidizer in rocket combustion devices and in vehicle flight experience. Thus, while the committee believes that these fuels might be a good choice for future advanced launch systems, rocket engines designed for using liquefied natural gas-type fuels are not currently at a sufficient TRL for serious consideration. The committee therefore agrees with the Air Force baseline selection of LO 2/RP-1 for the RBS MPS. Again there are also two suboptions for this propellant combination: open cycle, either gas generator or tap-off, and closed cycle ORSC. The ORSC engine is physically the highest performance (Isp) design approach for RP-1 fuels for several fundamental reasons, including these: (1) operation at high chamber pressure which provides higher net Isp because of improved combustion efficiency; (2) higher sea level (liftoff conditions) area ratio; and (3) typi- cally much higher engine thrust-to-weight ratio. Key characteristics for open-cycle gas generator or tap-off cycle engines versus ORSC closed-cycle engines are summarized in Table 3.2. Table 3.3 summarizes the advantages of an ORSC LO2/RP-1 engine over an open-cycle gas generator and lists the issues/concerns as well. The turbine drive gas from the fuel-rich gas generator presents a more benign condition for all the materials in the hot-gas flow path but forces the turbine to run at much higher operating temperatures in order to achieve the necessary turbine drive power. As discussed previously, the RBS uses the booster main propulsion system to accelerate the upper stage(s) and payload to the staging velocity and to provide sufficient impulse to allow the first stage to glide back to the launch site. With these dual demands on the main propulsion system, the first-stage sizing becomes very sensitive to the specific impulse and thrust-to-weight ratio of the LO2/RP-1 engine. For this reason, the committee agrees with the Air Force assessment that the ORSC cycle is the preferred engine cycle for the RBS system. A typical ORSC engine is depicted schematically in the very simple diagram in Figure 3.1. Liquid oxygen and fuel are fed into a high-pressure preburner, which initially combusts at greater than stoichiometric mixture ratios (MR = 10-20), and then the initial combustion products are quenched downstream in the preburner cham- ber with large amounts of LO2. The LO2 is usually injected down below in the preburner chamber through some type of tangential slots, such that the effluent gas leaves the preburner at an MR of 60-80 or so and an exhaust TABLE 3.2 LO2/RP-1 Typical Range of Operating Conditions Cycle Nominal Chamber Pressure Range (psia) Vacuum Isp Range (s) Thrust/Weight Ratio Open, gas generator 500-1,000 300-315 70-80 Closed, oxygen-rich staged >2,000-3,500a 325-350 100-120 a More typical for Russian engines such as RD-170 and NK-33.
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26 REUSABLE BOOSTER SYSTEM--REVIEW AND ASSESSMENT TABLE 3.3 Advantages of Oxidizer-Rich, Staged-Combustion (ORSC) Rocket Engines Over Open-Cycle Gas- Generator LO2/RP Engines Advantages Issues/Concerns Higher Isp (up to 7-10%). Because of its oxidizer-rich hot gas environment, engine components and plumbing (ducts) need to be made from compatible and flame- resistant materials or require the use of special nonburning, resilient protective coatings that do not erode or chip away during handling, testing, and operation (especially important for reusable engines). Use of higher density fuel enables higher overall mass Greater tendency for combustion instability because of the more fraction and more favorable aerodynamics profile rocket difficult to burn hydrocarbon fuel operating at much higher chamber stages/vehicle design. pressure in the preburner and main combustion chamber. Because of higher chamber pressures, the engine design Preburner design and development is more difficult than fuel rich gas results in nozzles with higher sea-level area ratios and generators (GGs). Because of high operating pressures, there will be significantly higher engine thrust/weight ratios. a greater tendency for combustion instabilities and the need for high- temperature oxidizer and flame-resistant materials in the turbines and any associated hot gas ducts and the oxidizer side of the main combustion chamber (MCC) manifolds and injectors. Results in oxygen-rich shutdown, which minimizes carbon Because of the high preburner operating pressures (6,000-9,000 psia), deposits and "coking" of injector orifices with hydrocarbon ORSC engines will require boost pumps and boost pump devices. fuels--therefore, easier to restart multiple times. Fuel-rich GGs typically run at much lower pressures--~1,000-1,500 psia--and are easier to design with fewer components. Enables pumps to generate required power at much lower operating temperatures than with fuel-rich GG powered turbines, which results in increased life and durability (typically about 700°F versus ~1700°F for fuel-rich GG cycle). Eliminates open-cycle GG exhaust plume interactions and interference issues by running all of the turbine drive gas back into the MCC. Usually allows increased engine service life because of generally lower operating temperatures. NOTE: See for example: (1) Oscar Haidn, Advanced Rocket Engines, Lecture Series at the von Karman Institute, Belgium. RTO-EN-AVT-150, ISBN 978-92-837-0085-2, Published March 2007, Open to the public; (2) Robert Sackheim, Overview of United States Space Propulsion Tech nology and Associated Space Transportation Systems, AIAA Journal of Propulsion and Power 22(6), November-December 2006. temperature of 700-800°F. The preburner exhaust gases are then run through the turbopump assembly (TPA) turbine, which typically drives both the fuel and the oxidizer pumps using a gearbox and separation seals. Note that all staged combustion engines must operate with pump discharge pressures significantly higher than the main combustion chamber (MCC) pressure because the main drive turbine operates in series with the MCC. The LO 2 is fed back to the preburner and the oxygen-rich gas from the turbine exhaust is then fed into the MCC injectors, where it is mixed with the liquid RP-1 coming from the fuel pump. Both the hot oxidizer-rich gas and the RP-1 enter the MCC through the chamber injectors. Typically, the injector has hot gas/liquid RP-1 swirling elements. The number of these swirl injector elements depends on the engine thrust level and scales somewhat with engine size and thrust level. Boost pumps are typically used to feed the high-pressure LO2 and RP-1 into the preburner. The engine start and shutdown sequences and methodology vary and tend to be somewhat complex, but they are usually established by the engine timing, calibration of the power balance, detailed operation of the flow control valves, and other fluidic elements. The oxidizer and the fuel pumps are usually mounted on the same common shaft, and dynamic seals and intercavity inert gas purges keep the two liquids well separated. An upstream start turbine is sometimes used to start TPA full operation.
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REUSABLE BOOSTER SYSTEM TECHNOLOGY ASSESSMENT 27 Oxidizer Fuel Pump Turbine Pump Oxidizer Fuel Preburner X Thrust Chamber FIGURE 3.1 Highly simplified schematic of a closed-cycle, staged-combustion rocket engine. Because the preburner, turbine, and thrust chamber operate in series, the required pump pressure is higher compared to open-cycle engines. SOURCE: Air Force Research Laboratory. Figure 3-1 R02338 There are several conventional options for vehicle thrust vector vector control; these would combine one or more editable thrust chambers with some type of non-toxic propellant reaction control subsystem, which will likely be required at a minimum to achieve roll control, but these will not be discussed here in detail. Engine throttling, which is usually required together with control of the mixture ratio, is accomplished with a complex combination of flow bypass valves, throttle valves, and fixed bias and calibration orifices that are inserted during engine and hydraulics system build-up, calibration, and hot fire tests. All liquid RP-1 enters the thrust chamber cooling jacket prior to entering the MCC injectors and is mixed with the oxidizer-rich preburner exhaust gases to achieve the final main chamber combustion process. The liquid cooling jacket (heat exchanger), together with some film cooling of the MCC chamber wall, maintains the MCC at an acceptable operating temperature while allowing the necessary engine combustion efficiency and associated I sp. 3.2.1 Hydrocarbon-Fueled Booster Engine Risk Assessment In considering the hydrocarbon-fueled ORSC rocket engine that will serve at the MPS for the reusable booster, the committee identified 12 risk areas:
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28 REUSABLE BOOSTER SYSTEM--REVIEW AND ASSESSMENT ˇ Combustion stability. Combustion stability physics for high-pressure liquid-liquid preburners and the MCC gas-liquid injectors for LO2/RP-1 ORSC engines are not well understood in the U.S. rocket industry. Combustion instability issues have plagued many rocket engine development programs during the past 60 years. Both physics- based modeling and a well-defined test program will be required to achieve and demonstrate the required stability margins for both combustion devices and thereby retire these risks. The Air Force Office of Scientific Research, in close cooperation with AFRL's rocket propulsion laboratory at Edwards Air Force Base, has been making signifi- cant investments to analyze, predict, and defeat the combustion instability problem in liquid oxygen, hydrocarbon (LO2/HC) engines. ˇ Injectors. The new ORSC injector will most likely be based on a co-axial swirl design that will have to be tuned to MCC frequencies and set up with either acoustic cavities and/or baffles to ensure stable engine operation. Also, thrust-scaling relationships will have to be established through a combination of analysis and empirical data. While this is not presently seen as a major risk, there will have to be a dedicated experimental testing effort to tune the injector and chamber at the subscale and then the full-scale levels. ˇ Operation in high-pressure, oxygen-rich environments. High-pressure, oxygen-rich environments can be very hard on inert materials, and dangerous conditions can result following initial failures that are difficult to contain when trying to recover to a safe operating state. This type of environment is unique to the ORSC engine. As described above, in an ORSC engine all oxidizer consumed in the engine combustion process is first used to drive the TPA turbine, resulting in high-power margin and a relatively low operating temperature. The resulting oxygen-rich environment is also relatively clean, such that no soot or other residual combustion product deposits are generated during normal operation. Because of these unique conditions, the ORSC cycle generates high-pressure oxygen rich environments that present compatibility challenges for traditional turbine and ducting materials. ˇ Physics-based analytical models. Another risk for the RBS MPS is a fundamental lack of fully anchored physics-based analytical models for ORSC engines. In the anticipated fiscal environment of limited budgets and short development schedules available for RBS development, reliable and accurate analytical models and tools are critical for the successful and cost-effective completion of the planned RBS DDT&E program. Accurately anchored models enable a reduction in the number of design fabrication cycles and, most significantly, expensive test cycles that were required for past rocket engine development programs. The analytical model development needs to be multiscale, with modeling ranging from subscale components to subassembly levels such as the power head (TPA, preburner, and flow control valves) as well as at the full-scale ORSC engine level. This approach will ensure that validated analytical models are developed logically and accurately and that they are well anchored so that validated models at each step of the design process will be available for the ORSC engine when scaling to higher thrust levels. These same models will also benefit other future ORSC engine development work by accel- erating and reducing the total number of expensive tests required. ˇ Valves/sensors/actuators. Some development effort will be required for the necessary reliable advanced fluid valves and control elements to enable a wide range of throttling, mixture-ratio control for proper and efficient propellant utilization, engine balance and calibration, thrust vector control (TVC), IVHM, and various other engine controls. These control elements must have modern, accurate sensors and be fully integrated with an automated digital MPS and engine control system, with the vehicle adaptive guidance and control system, and with the IVHM. ˇ Thrust level. The RBS ORSC engine thrust level requirement described in Air Force presentations to this committee is highly ambiguous. It was stated that the thrust requirement for RBS was anywhere from 350,000 to 500,000 lbf, which is inherently a problem from the standpoint of engine size and injector scaling and of possible high-pressure instabilities. The requirement became more complicated when Air Force briefers said several times that they might be interested in using the new American ORSC engines to replace the Russian RD-180 for the EELV program, which operates at thrust levels between 800,000 and 1.1 million lb f; this would pose a significant engine scaling problem and a much higher level of risk. The thrust level requirement should therefore be established early in the RBS development program and maintained at that level to avoid additional complications. If the Air Force wants to make the new engine compatible with the thrust requirement of NASA's Space Launch System (SLS), which is stated to be 1.1 to 1.2 million lbf, this ambiguity would also have to be resolved to enable a single, focused ORSC engine development program without any additional complications that could lead to large program cost and schedule overruns.
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REUSABLE BOOSTER SYSTEM TECHNOLOGY ASSESSMENT 29 ˇ Systems engineering. The Air Force presenters offered little description of the RBS needs and/or require- ments for system engineering and integration of the MPS and the other major systems into the RBS vehicle. These requirements must be firmly established early in the development process to avoid serious conflicts, unnecessary complications, and, worst of all, requirements creep, before program critical design review and flight certification. If these design requirements are not firmly established and maintained early in the program, there will be a high probability of serious cost overruns and schedule slippage. ˇ Power balance. Overall engine power balance and flow calibration and control with robust margins and tolerances for wide operating variances must be established early and verified by testing. Otherwise there will cer- tainly be problems with off-nominal operating conditions, which often occurs later in the program or, worst of all, during actual flight. There must be demonstrated anomaly and off-nominal operating capability and robust margins designed into the more complicated ORSC engine to avoid failures that would be catastrophic during flight. ˇ Turbomachinery. There are always risks with new engine turbomachinery. Because of the high-pressure preburner operation, boost pumps need to be used in the engine and as part of the overall cycle. Engine start-up transients and shutdown sequences will all have to be established and fully characterized for this more complicated ORSC, multicomponent engine to ensure safe and repeatable operation. This is not considered to be a major risk, but it is a moderate risk that will have to be addressed during the design and development process with a focused, dedicated effort that will have associated costs. ˇ Long-life bearings. High-speed, long-life reusable turbomachinery shaft bearings will have to be developed and verified for all ORSC engine rotating machinery. This is a low-to-moderate-risk concern but one that must be explicitly addressed because bad bearing choices and subsequent integration into the engine can lead to serious problems later in the DDT&E program. This happened more than once in past programs such as the SSME, when bearing issues were discovered after a number of space shuttle flights perceived to be successful. ˇ Reusability: Other than the SSME, there has been almost no requirement for multiple reuses of U.S. rocket engines. Typically, additional life margin is designed in and demonstrated on all rocket engines to allow for hot- fire testing and even retest before flight. The question of how many engine reuses will be required prior to routine maintenance and scheduled engine overhauls must be answered. This concern is not a major risk, but it is rather a moderate risk that must be mitigated because it is a new requirement for an ORSC engine that has not been previously developed by U.S. industry. ˇ Materials. The last risk, a moderate one, is the mechanical and dynamic design approach and especially the materials to be used for hot, oxidizer-rich, hot-gas ducts from the preburner, which may require flexjoints or axial joints and other complications. As mentioned above, this risk area will have to be empirically evaluated before committing to a design. Thus, the principal technical risks associated with the development of a LO 2/RP ORSC rocket engine concern combustion stability and operation in the high-pressure oxygen-rich environment. Significant addi- tional basic and applied research on combustion stability will be necessary before analytical tools are available that allow confident prediction of combustion instabilities. This work is under way at AFRL, but significant improvement in prediction capabilities cannot be anticipated in the near term. Fortunately, empirical techniques are available to "de-tune" the combustion system if instabilities arise, so this risk is principally one that will need additional development time and resources will be required if instabilities arise as the engine is scaled up during its development phase. The risk associated with the oxygen-rich operation is fundamental and potentially more difficult to overcome. It is well known that Russian engine designs have overcome this material incompatibility challenge by using inert enamel coatings on traditional high-strength turbine alloys and hot-gas ducting. The alloys provide the structural load support, while the enamel coating provides the requisite hot-oxygen-compatible and/or protected surfaces. This type of solution has been used on Russian ORSC rocket engines for over 50 years and is used around the world for gas-turbine applications for jet engines and domestic power generators. There are at least three flight- certified Russian ORSC engine designs (RD-170, RD-180, and NK-33) that are well known to the U.S. rocket engine industry but not produced in the United States. Through various business arrangements with the Russian engine manufacturers, the basic design for these engines is well known and understood by the U.S. manufacturers--
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30 REUSABLE BOOSTER SYSTEM--REVIEW AND ASSESSMENT namely, Pratt and Whitney Rocketdyne (PWR)1 and Aerojet.2 Each company sells a version of different Russian engines (e.g., PWR sells the RD-180 and Aerojet sells the NK-33) to U.S. launch vehicle suppliers. Another oxygen-rich compatibility approach is to develop and use a hot-oxygen inert parent material. 3 This approach is being attempted by the AFRL and large engine contractors (PWR and Aerojet) with some successful results already having been reported. PWR has developed a new hot-oxygen-compatible material called Mondaloy and has been evaluating its applicability and durability in hot oxygen-rich environments. To date, only small coupon pieces of Mondaloy have been manufactured, and the statistical basis for the thermal and mechanical properties needed for engine development is currently lacking. As such, basic issues with weldability, fatigue, and fracture remain to be investigated. Currently, the Russian-developed enamel coatings are a far more mature and proven technology. However, application of these special coatings and/or advanced materials to U.S. ORSC engine designs has not been fully proven, so a comprehensive risk reduction program will be required. This risk mitigation effort must be focused on developing and proving a new hot-oxygen-compatible parent metal alloy or on verification of a coated mate- rial system capable of multiple reuse for the turbine and hot-gas flow-path components similar to the Russian solutions. Thus, if the known coatings, which are inert and will not react with hot oxygen, become the preferred approach, the risk mitigation effort will have to ensure that the coated engine elements are sufficiently durable for multiple reuse under all the necessary environmental exposures. In short, the program will have to verify coating durability under all relevant conditions and demonstrate traditional hot oxygen compatibility to be certified for a flight RBS-ORSC engine that launches an RBS. The committee believes that in spite of the ORSC engine risks and concerns discussed above, there already exists an extensive database and success with this type of engine around the world. Table 3.4 lists all LO 2/HC engines using either open- or closed-cycle designs that have been flown or are flight qualified throughout the world. As can be seen from this table, ORSC engines have been or are about to be flown on launch vehicles in the United States (all Russian designed and built), Russia, Ukraine, India, China, and South Korea. This extensive successful flight history and development experience should provide confidence for the development of a LO 2/HC engine for the RBS booster. Many of these LO2/HC-powered launch vehicles have successfully placed large payloads of all kinds into their required Earth orbits or onto space-science trajectories throughout the solar system. This long history of success also includes putting many human beings in space as well as on the moon. So, these past suc- cessful experiences on both kinds of LO2/HC engines certainly provide assurance that a completely new ORSC engine can be developed here if adequate resources, time, and planned reserves are devoted to an affordable and reasonable program. 3.2.2 Hydrocarbon-Fueled Booster Engine Risk Mitigation In addition to its extensive and successful EELV flight experience with hydrocarbon booster engines, the Air Force described to the committee the types of hardware technology demonstrations and risk mitigation programs that AFRL has been pursuing. These technology programs are briefly described below together with a short over- view of NASA's hydrocarbon engine development activities. The AFRL has been conducting a joint rocket technology program with industry, the Integrated High Perfor- mance Rocket Propulsion Technology (IHPRPT), to advance all forms of rocket propulsion technology, including hydrocarbon boosters, for about the last 20 years. The goals for the hydrocarbon booster portion of IHPRPT are summarized in Table 3.5.4 1 A. Weiss, Pratt and Whitney Rocketdyne, "Reusable Hydrocarbon Rocket Engine Maturity for USAF RBS," presentation to the Committee for the Reusable Booster System: Review and Assessment, February 16, 2012. Approved for Public Release. 2 J. Long, Aerojet, "Reusability and Hydrocarbon Rocket Engines Relevant US Industry Experience," presentation to the Committee for the Reusable Booster System: Review and Assessment, February 16, 2012. Approved for Public Release. 3 R. Cohn, Air Force Research Laboratory, "Hydrocarbon Boost Technology for Future Spacelift," presentation to the Committee for the Reusable Booster System: Review and Assessment, February 16, 2012. Distribution AApproved for Public Release. 4 R. Cohn, Air Force Research Laboratory, "Hydrocarbon Boost Technology for Future Spacelift," presentation to the Committee for the Reusable Booster System: Review and Assessment, February 16, 2012. Distribution AApproved for Public Release.
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REUSABLE BOOSTER SYSTEM TECHNOLOGY ASSESSMENT 31 TABLE 3.4 American Heritage or Other Available Source Hydrocarbon Rocket Engines: Previously or Currently in Use, or in Development Thrust Level (lbf) / Rocket Manufacturer/ Vacuum Specific Engine Supplier Cycle Impulse (s) Status Application Comment RS-27 Pratt and Whitney GG 200K S.L., 237K Flown hundreds of Delta II and A2L MD-1 Rocketdyne ALT / 303 flights (~800) Previous Thors, Thor MA-7 Delta, and Delta III MA-5A Pratt and Whitney GG 430K (booster)+60 Flown hundreds of Atlas family up LR-89/ Rocketdyne (sustained) / 297 times on Atlases; to satellite launch LR-105 Project Mercury vehicles many flights (1,404) H-1 Pratt and Whitney GG 200K S.L., 205K Flown many flights Saturn 1B, Jupiter, and Rocketdyne ALT / 301 (152) early Thor Deltas F-1 Pratt and Whitney GG 1,522K S.L. / 307 ~65 flights Saturn V/Apollo 5 used in Rocketdyne first stage RD-180 Pratt and Whitney ORSC ~860K S.L., ~10 flights Atlas III and V Two TCAs, Rocketdyne, 933.4K ALT / 337 one pump Russian-derived NPO Energomash RD-170 NPO Energomash ORSC ~1700K / 337 Flown many times Buran, Zenit, Proton Four TCAs S-3D Pratt and Whitney GG 80K / 310 46 flights Jupiter/Juno Rocketdyne AJ-87 Aerojet GG 300K / 249 Flown many times Titan-I first stage AJ-1 on ICBM test flight AJ-3 AJ-91 Aerojet GG 80K / 310 Flow many times on Titan-I second stage ICBM test flights NK-33 Aerojet ORSC 340K S.L., 380K Intended for use on Developed for Russian PC = 2,109 (AJ-26) ALT / 330 Russian N-1 Moon N1, to be used on psia launcher first stage Taurus II (Antares) NK-39 Khrunichev/ ORSC N-1 second stage Developed for Russian Aerojet N1, intended for use on K1 RS-84 Pratt and Whitney ORSC 1,050K S.L., Finished PDR; Intended for 100 Incorporates Rocketdyne 1,123K ALT / 338 cancelled by NASA missions, reusable advanced launch vehicle technology items, advanced materials, enhanced water-cooled nozzle Merlin SpaceX GG ~80K / ~302 Flown Falcon I, 9 and 27 Privately family funded development Other NPO Energomash ORSC Various / ~330 Some flown Various Russian and RD-190 miscellaneous Ukrainian Rockets Russian continued
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32 REUSABLE BOOSTER SYSTEM--REVIEW AND ASSESSMENT TABLE 3.4 Continued Thrust Level (lbf) / Rocket Manufacturer/ Vacuum Specific Engine Supplier Cycle Impulse (s) Status Application Comment MC-1 NASA/Marshall GG ~75K / ~280 Advanced Looking for one, (Fastrak) Space Flight development, many almost for X-34 Center in-house tests YF-100 China, Inc. ORSC 260K S.L., 301 Flown on Long China's Long March ALT / 336 March launch vehicle 5, 6, and 7 NOTE: ALT, at altitude; GG, gas generator; ICBM, intercontinental ballistic missile; lbf, pounds force; ORSC, oxygen rich, staged combustion; PC, combustion chamber pressure; S.L., sea level; TCA, thrust chamber assembly, K, thousand. TABLE 3.5 Integrated High Performance Rocket Propulsion Technology (IHPRPT) Hydrocarbon Boost Technology/Performance Advancement Goals Goals IHPRPT Goals Related to Reusable Booster System Specific impulse (s) sea level/vacuum +15% Thrust to weight sea level/vacuum +82% Production cost -50% Failure rate -75% Mean time between replacement (cycles) Defined Mean time between overhaul (cycles) Defined Turnaround time (h) Defined Throttle range Defined Sustainability Must derive from sustainable materials and processes SOURCE: Richard Cohn, Air Force Research Laboratory, "Hydrocarbon Boost Technology for Future Spacelift," presentation to the Com- mittee for the Reusable Booster System: Review and Assessment, February 16, 2012. Distribution A: Approved for Public Release. Critical RBS technologies being studied and developed under ongoing AFRL propulsion R&D programs are summarized in Figure 3.2. The hydrocarbon boost (HCB) Phase II demonstration program aims to develop technologies to support the ORSC LO2/RP engine capability. Conducted by both Aerojet and PWR, this program aims to mature advanced hot-oxygen-rich-compatible materials and coatings as well as engine components such as pumps, advanced hydrostatic bearings, valves, actuators, preburners, igniters, main thrust chambers and new engine controllers, IVHM systems and associated sensors. As seen in Figure 3.2, the technology development associated with the HCB Phase II demonstration program and the Advanced Liquid Rocket Engine Stability Technology (ALREST) program are scheduled to run through the year 2020, which would limit technology contributions from these programs to any near-term RBS development activities. The ALREST combustion stability effort is focused on development of a fundamental understanding of com- bustion instabilities in high-pressure LO2/RP combustion systems. These two efforts are planned for completion in 2020. Additional efforts under way include assessment of the characteristics of Mondaloy and development of IVMH diagnostic techniques applied to booster engines. NASA has also been conducting advanced ORSC engine development programs over the last 15 years, working to solve many of the same advanced technology problems as those that the Air Force has been addressing. 5 Several years ago NASA Marshall Space Flight Center (MSFC) funded a new ORSC engine program known as the RS-84. 5D. Lyles, "NASA's Reusable Stages and Liquid Oxygen/Hydrocarbon (LOX/HC) Engines," presentation to the Committee for the Reusable Booster System: Review and Assessment, February 17, 2012. Approved for Public Release.
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REUSABLE BOOSTER SYSTEM TECHNOLOGY ASSESSMENT 33 2009 2010 2011 2012 2013 2014 2015 2016 2017 2018 2019 2020 2021 2022 2023 2024 2025 Near Term Implementaon USET Mid Term Implementaon HCB Phase II Demonstraon Program ALREST Combuson Stability MondaloyTM IVHM/EHM Far Term Implementaon 3GRB Risk Reducon 3GRB Demo (Subscale) Concept RP21 Upper Stage (RBCC) RP21 Boost or Upper Stage Nozzle Themis and Follow-on in-house FIGURE 3.2 Air Force Research Laboratory Liquid Rocket Engine Roadmap as of fiscal year 2012. SOURCE: Richard Cohn, Air Force Research Laboratory, "Hydrocarbon Boost Technology for Future Spacelift," presentation on February 15, 2012. Distribution AApproved for Public Release. This engine program proceeded through the preliminary design phase and was able to conduct some advanced prototype component design, manufacture, and test before it was canceled because it had no firm mission require- ment and the NASA budget did not support sufficientFigure 3-2 technology funds. Nevertheless, some successful advanced component-level results were achieved that are now R02338 directly applicable to a new Air Force ORSC engine program. vector NASA also worked jointly with the Air Force on editable an advanced technology program known as the Integrated Powerhead Demonstration (IPD). The IPD program was initiated by the Air Force but soon evolved into an effort jointly funded by AFRL and NASA MSFC with all hot-fire testing conducted at the NASA Stennis Space Center. The power head was to be integrated into a full-flow staged-combustion engine, where all propellants (in this case LO2 and LH2) flow through oxidizer-rich and fuel-rich preburners that power separate fuel-pump turbine and oxidizer-pump turbine, respectively. The hot-gas exhausts from each preburner flow into the MCC through a gas-gas injection system, making 100 percent of the propellant energy available to produce thrust. The IPD was designed as a ground-based demonstrator for an engine that would produce 250,000 lb f thrust. It was designed, built, and successfully tested, thereby demonstrating compatibility for high-performance and long-life components, materials, and technologies for reusable booster engine applications and (for the first time) a gas-gas injection MCC. After the successful tests at NASA Stennis, the program was terminated by the government, also for a lack of funding and a lack of a well-defined mission requirement. The successful IPD demonstration led to improved understanding of advanced engine components, and many of these results will be directly applicable to advanced component development associated with any type of new ORSC engine. NASA has recently announced a requirement for a new ORSC engine in the million-pound thrust class for its new SLS Advanced Liquid Strap-on Booster (ALSB). The procurement and conceptual design activities are currently under way for a program that will demonstrate significant risk mitigation results for this new ORSC
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42 REUSABLE BOOSTER SYSTEM--REVIEW AND ASSESSMENT When the dynamic pressure and aerodynamic loads are sufficiently reduced during upper-stage ascent, the payload fairing will be released and jettisoned. The environmental loads generated by this event will have to be accommodated but they will probably pose no unusual material or structural technology risk. The only risk that might be encountered would be one from some new requirement to save launch weight by introducing new and lighter materials, which could challenge the structural design and the design of the release mechanism. From a structural design standpoint, the RBS LES, except for the attachment and release geometry and mecha- nisms, will likely be very similar to the upper stages now in use with EELVs. The propellant tanks will have to accommodate loads from the nonlinear attachment and subsequent separation; however, the loads analysis and analytical models should be readily adaptable from experience with the space shuttle's unique geometry and the Delta IV Heavy side-mounted liquid boosters. If a decision is made to use lightweight composite tanks, there will be additional technical risks having to do with the lightweight composites used to contain cryogenic propellants. There is still significant concern over the very negative experience with NASA's X-33 cryogenic composite tanks. The TRL for using these composite materials with cryogens is still low, and much more technology development would be required. There is also the very important question of how to best mount large composite tanks on the vehicle's primary structure and what type of fasteners or bonding techniques to use. The answer will depend heavily on the materials of construction and the design approaches. Then too, there is always the question of designing for distribution of flight loads. Should the loads be distributed through the tanks or around them or in some load-sharing combination? How will the design best be accomplished? Other design solutions such as engine structural mounts and thrust load accom- modation are likely to be readily extractible from standard analytical techniques and past launch vehicle flight experiences. An additional complicating factor for the reusable booster is that RTLS maneuver functions require additional hardware elements (wings, control surfaces, landing gear, and so on) that cause the RBS to suffer appreciable intrinsic mass penalties compared to conventional expendable boosters. The basic rocket equation, V = Isp g ln[1 / (1 - )] where is the mass fraction (the ratio of the fuel mass to the total mass), shows us that for a given DV capability, significant reductions in RBS mass require significant improvements in specific impulse or total-system mass frac- tion (fuel mass/total RBS mass). For ORSC LO2/HC engines, specific impulse gains greater than approximately 10 percent are unlikely. However, the hardware elements needed for the RTLS maneuver invite both invention and the consideration of new materials. For example, Elias, in his presentation to the committee, 16 projected significant potential gains in structural material strength-to-weight ratio based on nanomaterial technologies such as large- scale nanotube structures. He stressed that the mass reductions stemming from such new materials are necessary for RBS-type systems to be cost competitive with next-generation expendable systems, especially in the emerging commercial arena. While large-scale nanotube structures are unlikely in the near future, significant reductions may be achieved by incorporating new fiber composite materials for the load-bearing structures. While challenges exist in meeting inert mass fraction requirements for RBS, many of the structural design concerns about the best selection of construction materials should be readily resolved following flight testing of the subscale Pathfinder and the much lower cost RBD flight test vehicles. Thus, the principal technology risks for achieving the needed inert mass fraction of the RBS structure are associated with the structural materials and process selection and the accurate determination of the structural loads. Obtaining accurate flight profile aerodynamic, aerothermodynamic, and other environmental loads from subscale tests to enable validated and accurate CFD loads for design of the full-scale RBS is therefore essential. 16 A. Elias, Orbital Sciences Corporation, "Orbital Sciences Corporation Opinions on Launch vehicle Reusability," presentation to the Committee for the Reusable Booster System: Review and Assessment, March 28, 2012. Approved for Public Release.
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REUSABLE BOOSTER SYSTEM TECHNOLOGY ASSESSMENT 43 3.6.2 Power, Fluid Thermal, and Actuation R&D While not considered a major technical risk, the RBS still must have its own power source during the full flight profile. Once the ground power umbilical disconnect separates from the booster at liftoff, the RBS will need a power source that is either accessible and replaceable or some type of rechargeable/regenerative power supply. There are four technically mature candidates to perform this function: ˇ Primary thermal batteries. These batteries are inexpensive, reliable and easy to activate, but relative to other power source listed here are heavy and will have to be replaced after every flight and mission abort. ˇ Chemical energy to power an Auxiliary Power Unit (APU). APUs have a long history of success on aircraft, Space Shuttle, etc., but their operations can be complicated and the APU expendables must be replenished after each flight. ˇ Fuel cells. Fuel cells offer higher efficiency with a compact footprint, but they are expensive, safety con- cerns remain, and they require recharging of expendables after each flight. ˇ Rechargeable/reusable batteries. Li-ion, NiMH, and other such batteries generally have a long life with proven high reliability and low maintenance without additional expendables, but they are expensive and require a readily available power source for recharging and regular conditioning. The selection of the appropriate internal power source will be decided by selection of the optimum approach for meeting the concept of operations, reliability, and performance criteria. If RBS operability, including fast ground turnaround and responsiveness, are an Air Force requirement, rechargeable/reusable batteries for the short-duration booster flight are likely the best choice. 3.6.3 Assembly and Manufacturing Modern capabilities for manufacturing and assembly of all types of vehicles throughout the world have greatly advanced in recent times. Incorporation of fully automated and robotic operations has greatly increased productivity for airplanes, automobiles, ships, mobile tool systems (e.g., tractors, lawn mowers, and the like) and even for the few expendable rockets being produced in the United States--Delta IV, Atlas V, and Falcon 9. The successful use of these advanced manufacturing methods for almost all modern vehicles has improved productivity and decreased cost and production time. The benefits of all the advanced, automated methods now in use have been widely recognized and are being employed by an increasing number of manufacturing companies the world over. The current EELVs have adopted friction stir welding in the manufacture of their aluminum propellant tanks and adapters. This automated process generates less heat than conventional welding processes, results in a smaller heat-affected zone with its attendant poorer mechanical properties, and eliminates most defects and the need for weld repairs. If the RBS propellant tanks are made of aluminum, friction stir welding should be explored as the primary metal joining method. The RBS baseline plan, while only planning to manufacture eight complete flight vehicles, may have real potential for cost savings using modern manufacturing technology approaches. In addition, it may be prudent to build these vehicles over a stretched-out span to permit incorporation of improvements as vehicles are built and to have an ongoing manufacturing capability if replacement boosters are needed. While a highly automated plant for booster manufacturing may initially be more expensive than the conventional option, it better accommodates personnel turnover and improves manufacturing quality. The ongoing production of upper stages and payload fairings for RBS can clearly benefit from a highly automated, modern manufacturing facility. Accordingly, the committee believes that even with the initial requirement for only eight flight vehicles, producing the vehicles using the most modern, automated manufacturing and assembly methods would likely provide high payoffs. This approach will unquestionably improve the probability of meeting the stated RBS LCC goals, even including the high initial capital investment required to implement these advanced highly automated modern techniques.
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44 REUSABLE BOOSTER SYSTEM--REVIEW AND ASSESSMENT Over the years the major space providers have been evolving their satellite and launch vehicle capabilities, incorporating new technology into existing designs. By way of contrast, other providers, such as Iridium and SpaceX have performed clean-sheet exercises by combining the use of state-of-the-art materials and state-of-the-art manufacturing techniques in their product lines to help reduce costs and increase reliability. Clean-sheet exercises, when done effectively, allow for incorporation of a holistic vision of the end-use system into the early concept design. Additionally, simplification concepts such as standardization and modularization as well as early systems engineering to incorporate IVHM diagnostic sensor systems can be designed into the system from the beginning and validated during testing. The manufacturing process can also be designed to run in parallel with the postflight maintenance and servicing. 3.6.4 Upper-Stage Development The RBS high-performance expendable second stage will need to be developed to be operationally compat- ible with the newly developed reusable first stage in a manner that would allow meeting overall mission goals as described by the Air Force. The baseline second stage uses LO 2/LH2 propellants owing to the inherent need for very high performance upper stages for heavy payload orbit injection (as was the case, for instance, for the Saturn S-IVB, the Delta IV, and the Centaur upper stage on Atlas I, II, III, and V). The need for high performance in the second-stage engine is especially true for the RBS concept, where the staging velocity is very low compared to an optimized expendable system. With the large DV requirement for the RBS second-stage, the projected overall launch mass of the RBS system would need to be much larger if a high-performance LO 2/LH2 engine was not used. The Air Force baseline approach for this high-performance upper-stage engine is to use the existing, flight- proven RD-25E--that is, the nonreusable version of the SSME that is being developed by NASA. Backup options for a high-performance upper-stage engine (such as the Russian-produced RD-0120) are also potentially available. Given this baseline approach and the potential availability of alternatives, the committee believes that development of the RBS upper-stage engine does not pose a significant technology risk. However, for the range of payloads initially described in the Air Force RBS launch model, the upper stage will be quite large and will require large, lightweight cryogenic propellant tanks. The chosen materials (metal or composites) may introduce additional tech- nology risks. Tank technology concerns as well as questions of propellant management and the possible benefits of CBTs for LO2 and LH2 to save mass are discussed elsewhere in this report. AFRL has been developing improved analytical prediction techniques for application to upper- stage engines in its Upper Stage Engine Technology program. In addition, the Air Force is now supporting an advanced upper- stage engine technology development program, the Affordable Upper Stage Engine Program, which is developing the technology base for a new high-performance, higher-thrust-level LO 2/LH2 engine to ultimately replace the RL-10 family, which has been around and slowly evolving since 1962. Other than the J-2 engine used on Saturn's upper stages and the SSME, the RL-10 and its various derivatives have served as the only high-performance, upper-stage engine for most U.S. launch vehicles (Delta IV, all Atlas/Centaur versions, and some Titan missions with a Centaur upper stage). In addition to the baseline engine options of the RS-25E or RD-0120, AFRL has the option to develop a new scaled-up version (much higher thrust level) of an RL-10 class (high performance expendable) LH 2/LO2 engine if neither the RS-25E or RD-0120 engines are available or suitable for the RBS upper stage. However, if a new upper-stage high-performance LH2/LO2 engine needed to be developed for any reason, this would add significant cost to the RBS development program, which was not briefed to the committee by any Air Force, Aerospace Cor- poration, or contractor presentations (or by any other entity). 3.7 OPERATIONS AND INFRASTRUCTURE The baseline RBS development plan includes three launch sites: two at Cape Canaveral Air Force Station (CCAFS) in Florida and one at Vandenberg Air Force Base (VAFB) in California. The first RBS launch complex will be all new and built at CCAFS and will accommodate the RBS-Y full scale development vehicle. It will initially operate in parallel with the current CCAFS EELV launch sites. Once the Atlas component of the EELV program
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REUSABLE BOOSTER SYSTEM TECHNOLOGY ASSESSMENT 45 is phased out, Atlas pads can be converted to accommodate the RBS, and the initial RBS-Y pad at CCAFS could also be used for RBS. There will be many design decisions at RBS facilities to manage cost but few technology decisions. Most operability enhancements envisioned for RBS infrastructure are being used or are currently in development for existing launch vehicles or other processing/manufacturing applications. In the opinion of the committee, the RBS facilities will face a design and cost control challenge but probably not be a significant tech- nology development challenge. Applicability of existing EELV infrastructure to RBS will be somewhat limited because the RBS booster is so large, especially its wing and vertical tails and its LES. Table 3.8 compares the medium-lift RBS with the Atlas V-551 EELV compiled from data provided by the Air Force, the Aerospace Corporation, United Launch Alliance, and contractors, whose representatives briefed the committee. This comparison can be used to identify which cur- rent Atlas infrastructure may be applicable for RBS use. Some existing Atlas EELV horizontal processing facilities may be usable as is or with reasonable modification. Existing launch control centers can probably be used for RBS with upgraded equipment. EELV vertical processing and some launchpad facilities may require major rework if the winged RBS booster with a piggyback-mounted LES will not fit. However, it still makes economic sense to modify current launchpad facilities to avoid lengthy environmental approvals for new sites and to take advantage of existing exhaust ducts, vehicle transportation equipment, propellant storage and transfer systems, fiber-optic communications, and payload air conditioning capabilities. Processing operations at the launch sites will most likely be based on current EELV practices, since the RBS manifest for the Air Force is "launch on schedule." Payload and expendable upper stage operations will follow current practices with enhanced automation, while those for the reusable booster will need to be modified to accom- TABLE 3.8 Comparison of Reusable Booster System (RBS) (medium lift) and Atlas V-551 Evolved Expendable Launch Vehicle RBS (approx.) Atlas V-551 Booster Diameter, Length (ft) 17, 110 12.5, 106.5 Inert and Propellant Weight (klb) 105, 900 41.7, 626.3 Number of Engine(s); Thrust (klbf) 5 AJ-26; 1,655 1 RD-180 = 860 Wing, Tail Span (ft) 60, 36 Solid Rocket Booster Diameter, Length (ft) 5.2, 65.5 Number of SRBs, Weight (klb) 5, 514.7 Number of SRBs, Thrust (klb) 5, 1897.8 Second Stage Diameter, Length (ft) 16, 130 10, 41.6 Inert, Propellant Weight (klb) 38, 340 4.9, 45.9 Number of Engines; Thrust (klbf) 1 RS-25E; 500 1 RL-10; 22.3 Interstage Adapter Diameter, Length (ft) 12.6, 13.6 Weight (klb) 8.8 Payload Fairing Diameter, Length (ft) 18, 77 17.8, 76.8 Weight (klb) 9 8.8 Spacecraft (to low Earth orbit) Weight (klb) 50 41.5 Total Dry Weight (klb) 200 626 Gross Lift-off Weight (klb) 1,340 1,298 Total Sea Level Thrust (klbf) 1,655 2,548 Stack Height (ft) 180 225 Booster Diameter (ft) 17 24 NOTE: The RBS data is approximate and conservative, to be used for comparative purposes only. SRB, solid rocket booster.
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46 REUSABLE BOOSTER SYSTEM--REVIEW AND ASSESSMENT modate landing, "safing" of the vehicle (making it safe for ground crews to carry out needed handling activities) and turnaround for relaunch. The following subsections are arranged to reflect a launch system nominal (historical) processing flow, with RBS influences and deviations noted. RBS operations may require changes to existing infrastructure and launch procedures. A review of the current practices and discussion of potential RBS impacts follows. 3.7.1 Range Safety Current practice for an established launch vehicle requires submittal of a preliminary flight data package to Range Safety at L-60 days and a final data package at L-30 days. For each upcoming unmanned vehicle launch, the Range reviews the nominal planned vehicle trajectory and establishes limit lines. If the actual flight trajectory exceeds those limits, the Range Safety Officer destroys the vehicle. RBS use of AG&C may allow these limits to be relaxed to account for RBS's ability to respond automatically to off-nominal conditions and still perform its mission. Range tracking must be performed by two independent sources. The RBS will interact with the Range Safety and the Airspace Control systems during both launch and return. In neither regime is it likely that the booster will present new challenges, even to the Range Safety systems in place today, which are certain to evolve between now and when an RBS system would fly. From the perspective of Range Safety, there is no significant difference between an uncrewed reusable system and an expendable booster. The same requirements for information about location, velocity, direction, and booster status, and need for a flight termination system will apply. Flight termination systems are well understood, and the RBS would present no new challenges. The Federal Aviation Administration Office of Commercial Space Transportation (FAA/AST) is facing the challenge of handling uncrewed vehicles in the airspace today, and the committee believes that the challenges to doing so will be resolved long before the RBS is flying. FAA/AST has limited experience in supporting a crewed vehicle returning from space to a landing in the United States, but that experience will grow substantially over the next decade. The experience with autonomous entry into the airspace and landing is far more limited today (e.g., the X-37B). However, the ability to manage this kind of operation has been demonstrated and will likely not present any major barriers to RBS operations. The various test and demonstration vehicles involved with the RBS, e.g., the Pathfinder, demonstrator and Y-vehicles, will probably be flown on traditional ranges. The Range Safety and Airspace Management requirements must be addressed during these phases of the program or alternatives found to the operational ranges. Finally, the launch range will now have to support operations associated with the booster's return to the launch base airstrip for landing. These operations are expected to be further complicated for the RBS Heavy with two returning boosters. Their return has to be staggered to permit sequential landing/exit from the runway. The Eastern Range has dealt with the space shuttle orbiter's return for landing, and similar considerations apply for RBS. The orbiter returns with some residual Orbital Maneuvering System propellants, and the RBS booster will have remaining LO2/RP-1 and attitude control system (ACS) propellants. These must be off-loaded and the propellant systems made safe prior to towing the booster(s) to a ground maintenance facility. 3.7.2 Launch Readiness Reviews Current expendable launch vehicles are subjected to a number of readiness reviews starting several weeks prior to launch. This process has been institutionalized over the past 60 years and with slight differences, is practiced by all government agencies and current launch system providers. The reviews are held with the launch service provider's management, payload provider, launch customer, Range Safety, launch base management, and the engi- neering tiger team. To achieve responsive launch with RBS, this process will need to be significantly revised and shortened while maintaining launch system reliability. A major cultural change by all participants in the prelaunch process will be required to accomplish a more streamlined and less costly review process.
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REUSABLE BOOSTER SYSTEM TECHNOLOGY ASSESSMENT 47 3.7.3 Spacecraft Processing Current facilities at CCAFS and VAFB are suitable for RBS payload processing and fairing encapsulation. At CCAFS these include the Astrotech commercial processing facility, the Air Force defense satellite commu- nications system processing facility (4-m diameter only) and shuttle payload integration facility; and the NASA vertical processing facility, spacecraft assembly and encapsulation facility, multipayload processing facility (4-m diameter only), and payload hazardous processing facility. These facilities support spacecraft storage, preparation, propellant loading, solid rocket motor and ordnance preparations, dynamic balancing, adapter mate and encapsula- tion. The VAFB facilities include the Astrotech Payload processing facility and the Space Systems International integrated processing facility, which provide these same services. The facilities are all capable of processing the large 4- and 5-m diameter fairings currently used on the Atlas V and Delta IV EELVs. Transporters exist to move the encapsulated spacecraft to their launch sites. In the United States, large spacecraft processing, encapsulation, and transport has historically been performed vertically. Upon arrival at the launch complex, the encapsulated spacecraft is mated vertically to the stacked vehicle upper stage. Horizontal spacecraft processing is also feasible, and could be used in RBS processing, as demonstrated by the Russians on Soyuz, Proton, and Zenit, and by the United States on the Orbital Pegasus and SpaceX Falcon-9 vehicles. 3.7.4 Launch Vehicle Processing Options The current EELVs are processed using three different scenarios. Understanding these scenarios and their associated infrastructure requirements will be helpful when discussing potential RBS processing and facility requirements. EELV processing and infrastructure information was obtained from United Launch Alliance, the manufacturers of Atlas and Delta IV. ˇ Atlas V at CCAFS. For relatively frequent launches, it is efficient to integrate and check out as much of the vehicle as possible away from the launch pad. This approach is used for Atlas V at CCAFS. After vertical integration on a mobile launch platform (MLP) and checkout in the vertical integration facility (VIF), the launch vehicle, minus its payload, is transported vertically on the MLP to a clean pad and placed over the exhaust duct. The vehicle is tanked and a wet dress rehearsal (WDR) is performed. The vehicle is then moved back to the VIF on its MLP for integration of the encapsulated spacecraft and then transported back to the pad for launch. Higher launch rates off the single pad could be accommodated by constructing a second VIF and MLP. ˇ Atlas V at VAFB. Because of infrequent launch requirements at VAFB, Atlas uses a conventional mobile service tower (MST), with vehicle integration occurring on the fixed launcher inside the MST. Vehicle stages are individually transported horizontally to the launch site, rotated to vertical, and stacked in their launch configura- tion with the MST hoist. WDR is performed and then the encapsulated spacecraft is transported to the launch site vertically and stacked onto the launch vehicle with the MST hoist. The MST is then pulled back for propellant loading, countdown, and launch. ˇ Delta IV at CCAFS and VAFB. Delta IV facilities integrate its booster(s) and upper stage in a horizontal integration facility at both CCAFS and VAFB. The horizontally integrated booster(s) and upper stage is then moved horizontally to the launch pad using a transporter/erector. Rotation onto the launch mount is accomplished with hydraulic pistons located in the concrete pad surface in front of the launch mount. A MST is used for access to the erected vehicle. The MST is rolled back to its launch position and a WDR is performed. Following WDR, the MST returns to its service position, and the encapsulated spacecraft is transported separately (vertically) and hoisted onto the stacked launch vehicle by the MST hoist. The MST is then pulled back for propellant loading, countdown, and launch.
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48 REUSABLE BOOSTER SYSTEM--REVIEW AND ASSESSMENT 3.7.5 Booster and Upper Stage(s) Processing Solid upper-stage processing is relatively insensitive to orientation for checkout, and existing infrastructure will be adequate. The current EELVs do initial checkout of their stages horizontally. For Atlas, the liquid core booster, solid rocket boosters (SRBs) and upper stage are checked out separately. In processing the Delta IV, vehicle checkout occurs after the booster(s) and upper stages are mated. Both the RBS solid and large expendable stages can be processed either vertically or horizontally, but because of the large booster and LES dimensions, the committee believes that separate horizontal processing will likely result in significantly lower facility cost. RBS booster and LES processing involves extensive electrical checkout of nonhazardous components, similar to that currently performed on EELVs. This automated computer-controlled checkout includes electrical continuity and isolation testing of all electrical harnesses and uses test squibs to validate all pyrotechnic events. 3.7.6 Booster/Upper Stage Integration and Checkout For RBS at CCAFS, off-pad integration, either horizontally or vertically, would be very desirable for accom- modating potential mission model growth and/or responsiveness requirements. At VAFB, unless there are major changes in RBS launch requirements such as increased responsiveness, either off- or on-pad integration should be acceptable. The extent to which current Atlas launch facilities at CCFS and VAFB must be modified for RBS will help determine which processing technique is most cost effective. However, it is likely that most of the existing major processing facilities for the Atlas (VIF at CCAFS and MST and SLC-3E launcher at VAFB) will require extensive modification or complete replacement. For a new launch complex, such as the initial RBD launch site at CCAFS, a trade-off of facility versus ground support equipment (GSE) requirements will be most cost effective. If maximum use of the current Atlas launch facilities is desirable, then different processing approaches may be implemented at space launch complex (SLC)-41 and SLC-3E. However, if identical off-pad processing is mandated at all three RBS launch sites, at the very least, the Atlas pad at VAFB will require major reconfiguration. 3.7.7 RBS Transport and Pad Installation Mated booster(s) and upper stage(s) can be transported either horizontally or vertically from an integration/ checkout facility to the launchpad. At SLC-41, the Atlas booster and Centaur upper stage are vertically integrated with the MLP in the VIF. The Atlas booster to MLP interface includes rise-off disconnects for all booster fluid and electrical services. The MLP has an umbilical mast for the Centaur and payload services. The MLP also includes autocouplers that mate to the GSE launch mount service connections. Since the Atlas V-551 dry weight is much greater than the RBS dry weight (see Table 3.8), the MLP that currently transports a fully assembled Atlas V-551 from the VIF to the pad can probably be used without major structural modification to transport the RBS. However, the existing MLP umbilical mast will have to be revised (removed or articulated) to accommo- date the winged RBS booster, and MLP launch hold-down/release and rise-off umbilical interfaces will require modification, at the least. EELV upper-stage services are provided via an umbilical mast, as are spacecraft services (air-conditioning, ground power, and communications). The piggyback mounting of the LES onto the RBS booster provides an opportunity to also use rise-off disconnects for LES services at the base of that stage and perhaps for some payload services routed through the upper stage. Accommodating spacecraft services adds weight to the upper stage but could eliminate the need for an umbilical mast. 3.7.8 Wet Dress Rehearsal The wet dress rehearsal is accomplished with a complete launch vehicle but without the encapsulated payload installed. WDR includes tanking the vehicle and performing a complete vehicle countdown, stopping just short of engine start functions. Atlas RP-1 leak testing is accomplished following RP-1 loading by taking the propellant tank to liftoff pressure for 10 to 15 minutes. Cryogenic leakage checks are accomplished during WDR (and on
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REUSABLE BOOSTER SYSTEM TECHNOLOGY ASSESSMENT 49 launch day) by cycling the fill/drain valves and remotely monitoring the compartment temperatures and perform- ing mass spectrometer analysis of compartment gas percentages using a hazardous materials gas detection system. All umbilicals remain connected. WDR has been accomplished on liquid propellant launch vehicles prior to launch for the last 60 years. In the early days, a captive hot fire test was conducted in conjunction with WDRs. Except for one or two tests with new vehicles, captive hot fire testing on an operational launch pad has ceased since it is deemed high risk. For RBS, the continuation of WDR as a regular prelaunch activity should be reevaluated. WDR is employed, especially on vehicles with cryogenic propellants, to check for component operability at low temperature and propellant leaks. The risk of finding these problems later, on launch day, is a launch delay. For Atlas at SLC-41, on-pad access does not exist (there is no MST), so depending on the problem, the MLP/launch vehicle may need to return to the VIF for corrective action. This 3- or 4-day span necessitates waiting for another range launch slot. RBS will need to be a robust design as compared to current expendable vehicles in order to meet the planned reductions in launch operation costs, so the need to perform routine WDRs may be significantly reduced. If an RBS booster is processed rapidly for its next mission, perhaps the previous launch can be considered a success- ful WDR for the booster. It is important to note that if WDR for RBS can be eliminated, then complete vehicle vertical integration in a VIF (booster, upper stage and encapsulated spacecraft) would be a very efficient process- ing technique. Current Atlas launch site propellant storage capacity can be used for RBS but will need to be significantly expanded to accommodate the propellant requirements of the much larger RBS upper stage. Existing propellant transfer infrastructure can probably be used as is for RBS, except propellant loading time will be extended. Payload air conditioning capability is likely adequate without any modification, assuming EELV payload conditioning requirements remain unchanged for RBS. 3.7.9 Payload Integration Following successful completion of Atlas WDR, the launch vehicle/MLP returns to the VIF for integration of the encapsulated payload. For Delta, the encapsulated payload is transported vertically to the launch pad and lifted by the MST crane to mate with the upper stage. If a complete vertically integrated launch vehicle with payload is transported to the pad, then additional on-pad checkout can be minimized. If the encapsulated spacecraft's integra- tion occurs on-pad, then additional checkout is required there. After Atlas payload integration and checkout in the VIF, the complete vehicle stack is returned to the launch pad on the MLP. 3.7.10 Propellant Loading and Launch Countdown This process is highly automated for the current EELVs. Automatic propellant loading is controlled by com- puters located at the launch pad. It is monitored by engineers in the remotely located Launch Control Center who have override capability if something goes wrong. Fully automated countdown is conducted by computers that sequence events and monitor data for out-of-tolerance conditions. Health and status data are also available via an automated system to responsible system engineers. These activities occur hands-off unless some anomaly is detected. Events can be halted at any time during the count. Fairly rapid countdown recycle is possible (depend- ing on the problem) until the last few seconds before liftoff. Atlas currently has the tools with its automated data management system to make a sound launch decision within the 126 minutes of cryogenic tanking operations prior to the final 4 minutes to launch a safe and successful mission. The current launch control processes and facilities used for Atlas V and Delta EELVs at both CCAFS and VAFB could be adapted for RBS. 3.7.11 Exhaust Ducts and Acoustic Suppression System Existing exhaust duct and acoustic suppression water system capacity at both SLC-41 and SLC-3E are likely to be adequate for RBS. Liftoff rocket exhaust mass flow is critical in exhaust duct and suppression system sizing. The mass flow is proportional to total engine thrust level and the propellants used. Both launch sites are designed
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50 REUSABLE BOOSTER SYSTEM--REVIEW AND ASSESSMENT to accommodate the Atlas V-551. The single booster RBS thrust level at liftoff is approximately 65 percent that for an Atlas V-551 with five ground ignited SRBs. In addition, the rocket exhaust area for an Atlas with five SRBs is two times larger than for RBS (see Table 3.8). 3.7.12 Flight Including Abort Facility requirements are associated with telemetry and Range Safety tracking and, if necessary, command destruct. These requirements are likely to be similar to those currently available to EELVs. 3.7.13 Booster Landing and Safing The unmanned booster returns to the launch base for landing at its air strip. Facility requirements will be similar to those used to monitor the space shuttle orbiter's return for landing, except without the concern for residual toxic and/or hypergolic reaction control system propellants. Guidance and navigation will be a fully automated onboard booster avionics function. Once the booster has safely landed, it must be towed to a protected or remote location for safing. The main LO2 and RP-1 tanks will be drained of residuals and purged. This can be accomplished by providing horizontal tank sumps (preferred) or by rotating the booster to a vertical position. Main engines will be purged. ACS propellants must be off-loaded and tanks purged. An important operability goal is to avoid the use of hazardous or toxic RCS propellants that would require use of Self Contained Atmospheric Protection Ensemble suits. Any unused pyrotechnic devices must be safed. 3.7.14 Postflight Booster Checkout, Maintenance, and Storage The booster is towed on its landing gear using a standard airliner tug from the airstrip to a hanger for routine ground turnaround checkout and maintenance. The booster's IVHM system will have captured any anomalies or out-of-tolerance conditions that occurred during flight and is used after landing to assess postflight vehicle health. This IVHM data will likely preclude the need for most of the technician-performed, hands-on checkout. These data identify specific maintenance requirements for returning the booster to flight readiness status. After mechanics and technicians complete required maintenance, vehicle IVHM, augmented as required by automated ground checkout equipment, will be used to validate turnaround booster maintenance. Some limited nondestructive evaluation of sensitive areas or components, such as inside the rocket engine nozzle, will also be required. 3.7.15 Booster Depot Maintenance At the proposed depot maintenance interval of 10 flights or whenever a major anomaly occurs that cannot be handled during regular ground turnaround maintenance, the booster will be undergo depot-level maintenance. Items requiring removal and replacement, such as landing gear tires, will be determined by equipment improvement upgrades and component life qualification results. Booster postflight turnaround maintenance, depot maintenance, and storage are best accommodated in a common facility to minimize facility cost and eliminate transfer of the booster from one location to another. 3.8 SUMMARY OF RBS RISK ASSESSMENT AND MITIGATION EFFORTS The development plan described in the presentations to the committee lacked the kind of details that would enable it to evaluate the technical merits, their creditably and risks, or even the accuracy of cost estimates. For example, other than achieving the top-level goals of demonstrating the rocketback maneuver at simulated staging
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REUSABLE BOOSTER SYSTEM TECHNOLOGY ASSESSMENT 51 # KEY TECHNICAL RISK AREAS & MITIGATION FAST PF RBD RBS 1 Rocketback Return To Base Method Flight Dynamics 2 Reusable & Throttleable LOx/Hydrocarbon Main Engine Hydrocarbon Boost 3 Advanced Low Maintenance Structures 4 Integrated Systems Health Management / Autonomic Logistics 5 Adaptive GN&C Flight Software 6 Propellant Management & Slosh Dynamics 7 Rapid Mission Planning 8 Agile, Responsive Ranges 9 Green/Operable RCS Propellants 10 Operable Thermal Protection Systems 11 Non-pyro Separation Systems 12 Payload and Stage Separation Dynamics 14 Operable Power Systems (Batteries & APU) 15 Advanced Ground Processing Automation FIGURE 3.4 Sample risk mitigation strategy. SOURCE: Slater Voorhees, Lockheed Martin Corporation, "Reusable Booster System (RBS)," presentation to the committee, March 28, 2012. Approved for Public Release. conditions above Mach 3 and subsequent return toFigure 3-4site, it was not clear what else the Pathfinder dem- the launch R02338 onstration would achieve. Also, there was no description or discussion of the Pathfinder MPS design. For example, vector is it intended to use one ORSC engine, such as the editable existing NK-33? Or, would it just use an existing, off-the-shelf, available open-cycle engine such as Fastrac17 or a Falcon 1 Merlin? There was no discussion of which, if any, propulsion elements and associated information would be directly applicable to the RBS design. The committee believes that developing a better definition of the goals and objectives of the Pathfinder pro- gram is important for two reasons. First, it is important to understand whether the cost is worth the investment. As stated above, there are important technical questions that can be answered in a Pathfinder program, but these questions will be answered only if the program is properly structured. Second, an understanding is needed of how the results of the Pathfinder program can be used to decide whether it is worthwhile to proceed to the next step, the RBD. So, decision gate criteria should be developed to enable a decision on what to do next after a successful Pathfinder demonstration. The development of RBS capabilities will challenge many of the limits of state-of-the-art capabilities in such disciplines as propulsion, aerothermodynamics, controls, structures, health monitoring and sensing. The Air Force plans to develop the technologies needed to achieve the goals for the RBS program in phases. An example of a program that follows this approach is provided in Figure 3.4, taken from Lockheed Martin's presentation 18 to the committee. In the figure, the technology development is shown to occur under the following programs: the 17 Fastrac was a LO /RP-1 60,000 lbf gas-generator cycle engine that delivers a vacuum I of approximately 280 s. It was developed by 2 sp NASA Marshall Space Flight Center (MSFC) and intended for flight application on the X-34 program by Orbital Sciences, but the program was cancelled and the engine inventory was placed in storage at MSFC. 18 S. Voorhees, Lockheed Martin, "Reusable Booster System (RBS)," presentation to the Committee for the Reusable Booster System: Review and Assessment, February 28, 2012. Approved for Public Release.
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52 REUSABLE BOOSTER SYSTEM--REVIEW AND ASSESSMENT Future Responsive Access to Space Technologies (FAST), Pathfinder, the RBD, and RBS full-scale development. (Note that the AFRL-funded FAST program has been completed.) This phased technology development suggests that the maturation of the HCB, Advanced Low Maintenance Structures, IVHM, adaptive AG&C flight software and operable thermal protection systems, which was initiated under the FAST program and will need to continue under subsequent programs (e.g., Pathfinder, RBD, and RBS). The Pathfinder program will also initiate devel- opment of capabilities for executing the rocketback RTLS maneuver and for handling propellants management and slosh dynamics, agile responsive ranges and payload separation dynamics. The RBD program will continue the development of the above listed technologies and initiate development of capabilities for rapid mission plan- ning, green/operable RCS propellants, operable power systems for batteries and APU, non-pyro separation, and advanced ground processing automation. The cost benefits of these technologies will be quantified during testing of the full-scale RBS Y-vehicles. Attaining the goals of the RBS program critically depends on successful demonstration of the relatively large number of widely varying technologies and capabilities. However, it is often difficult to assess the challenges associated with overcoming certain technical barriers until these are studied and characterized over a period of time. Having said this, it seems at the outset that executing the RTLS rocketback maneuver without damaging the booster and ORSC engines will be critical to the success of this program because such a maneuver has never been demonstrated at the scale of the RBS. Executing the RTLS rocketback maneuver will be challenging because it will depend on developing capabilities for considerable hydrocarbon rocket thrust throttling, handling the sloshing and dynamics of the propellants, controlling complex aerodynamics that may involve interactions of the rocket exhaust plume with the system's aerodynamics, and providing adequate thermal protection. Additionally, the attainment of "reusability," which is critical to achieving the expected cost savings, will require development of a reliable IVHM system. Such an IVHM system has to be capable of monitoring the health of critical components of the reusable part of the system. To be effective, the development of the IVHM would logically be undertaken in concert with the development of other system components such as the structure and the propulsion system. However, it is not clear at this point how this will be accomplished if the RBS is to employ legacy propulsion systems such as proposed with the use of the AJ-26. Finally, development of an adaptive guidance and control system will be critical to the operation of the unmanned, autonomous reusable booster that will need to operate in a complex environment with inherent uncer- tainties in vehicle performance.