5
Propulsion

INTRODUCTION

Propulsion will play a significant role in determining whether the goals of SSTO will be achieved in the RLV program. Performance of the only high thrust reusable engine in the nation, the SSME, has been excellent, but the SSME does not have the sea-level F/W ratio needed to place an adequately sized SSTO vehicle into the required orbits. Operability of current engines, in terms of the maintenance, parts replacement, and inspections required after each flight, are too costly in time and manpower to meet the low cost-per-flight goals of the RLV. Basic contributing factors to these costs are the lack of significant demonstrated reusability of critical components and adequate, reliable health monitoring instrumentation with automated rapid engine health diagnosis. Production costs of the current engines are also high because of their complexity, including the large number of parts needed and the manufacturing technology that was available when the SSME was developed. Overcoming these shortcomings are the basic objectives of the NASA/industry program in propulsion. The role of propulsion in four specific areas—program performance, producibility, reusability, and maintainability/operability—are as follows:

  • Performance. Rocket engine performance coupled with the vehicle size, weight, and payload characteristics are the key challenges to developing an RLV with SSTO capabilities. Based on presentations from the contractors and engine companies, an engine sea-level F/W of 75 to 80 with a vacuum trajectory average specific impulse (Isp) of at least 440 seconds is needed. The SSME has a F/W of 51 and a vacuum Isp of 453 seconds.

  • Producibility. In order to produce an RLV engine with high-quality equipment at lower cost, simplifications in design and careful verification of manufacturing methods will be required. Simplified but thorough inspection methods, coupled with experienced state-of-the-art engine production facilities, are also needed. During Phase I of this program, key rocket engine components are being evaluated with the primary objectives of reducing the costs of production and operation.



The National Academies | 500 Fifth St. N.W. | Washington, D.C. 20001
Copyright © National Academy of Sciences. All rights reserved.
Terms of Use and Privacy Statement



Below are the first 10 and last 10 pages of uncorrected machine-read text (when available) of this chapter, followed by the top 30 algorithmically extracted key phrases from the chapter as a whole.
Intended to provide our own search engines and external engines with highly rich, chapter-representative searchable text on the opening pages of each chapter. Because it is UNCORRECTED material, please consider the following text as a useful but insufficient proxy for the authoritative book pages.

Do not use for reproduction, copying, pasting, or reading; exclusively for search engines.

OCR for page 55
Reusable Launch Vehicle: Technology Development and Test Program 5 Propulsion INTRODUCTION Propulsion will play a significant role in determining whether the goals of SSTO will be achieved in the RLV program. Performance of the only high thrust reusable engine in the nation, the SSME, has been excellent, but the SSME does not have the sea-level F/W ratio needed to place an adequately sized SSTO vehicle into the required orbits. Operability of current engines, in terms of the maintenance, parts replacement, and inspections required after each flight, are too costly in time and manpower to meet the low cost-per-flight goals of the RLV. Basic contributing factors to these costs are the lack of significant demonstrated reusability of critical components and adequate, reliable health monitoring instrumentation with automated rapid engine health diagnosis. Production costs of the current engines are also high because of their complexity, including the large number of parts needed and the manufacturing technology that was available when the SSME was developed. Overcoming these shortcomings are the basic objectives of the NASA/industry program in propulsion. The role of propulsion in four specific areas—program performance, producibility, reusability, and maintainability/operability—are as follows: Performance. Rocket engine performance coupled with the vehicle size, weight, and payload characteristics are the key challenges to developing an RLV with SSTO capabilities. Based on presentations from the contractors and engine companies, an engine sea-level F/W of 75 to 80 with a vacuum trajectory average specific impulse (Isp) of at least 440 seconds is needed. The SSME has a F/W of 51 and a vacuum Isp of 453 seconds. Producibility. In order to produce an RLV engine with high-quality equipment at lower cost, simplifications in design and careful verification of manufacturing methods will be required. Simplified but thorough inspection methods, coupled with experienced state-of-the-art engine production facilities, are also needed. During Phase I of this program, key rocket engine components are being evaluated with the primary objectives of reducing the costs of production and operation.

OCR for page 55
Reusable Launch Vehicle: Technology Development and Test Program Reusability. A thorough evaluation of all the factors that make current engines and engine-vehicle interfaces less than efficient for reuse is critically needed. Phases I and II technology and test programs are intended to minimize operational delays and hardware failures for the RLV program. Long life and repeatable performance with minimal inspection will be essential. Maintainability/Operability. Engine health monitoring, simplified turnaround procedures, and improved and expedited before- and after-flight procedures are important goals. Several technology programs in Phase I are directed at engine health monitoring and simplifying operational systems. The committee reviewed the engine technology projects established by the prime vehicle contractors and the engine contractors with the intent of determining whether the approach was adequate to support a decision about whether to proceed with an X-33. The engines that will be used to meet the design criteria for an X-33 may differ substantially from the engines contemplated for an eventual RLV. Because of the lengthy development period required for major modifications of existing or new engines, the contractors and engine companies are developing engines to fulfill the X-33 requirements in the engine-vehicle while pursuing the development of the more-capable products that will be needed for an RLV. Therefore, the committee considered the initiation of programs to meet key long-range RLV engine requirements as well as X-33 engine requirements. DECISION CRITERIA The propulsion technology area will be adjusted by August of 1995 to reflect the needs of the X-33 industry partners. Propulsion systems not required by the proposed X-33 or RLV will not be funded by this program. A propulsion concept for the RLV configuration will be selected prior to the Phase II decision which will be required by the preferred RLV configuration. A documented analysis will have been completed prior to the Phase II decision which demonstrates that the selected propulsion subsystems are scaleable to a full-scale RLV and that reuse/operations requirements will be adequately demonstrated by a X-33 vehicle. Estimated requirements for the RLV, which will be supported by this analysis, include a 100 mission life with 20 flights between depot maintenance and a 50 percent reduction in engine inspection time between flights as compared to the shuttle. Results from component work will be documented and provided with the above analysis. Only propulsion technology supporting the X-33 contractors will be pursued within this program.

OCR for page 55
Reusable Launch Vehicle: Technology Development and Test Program NASA/INDUSTRY PROGRAMS Phase I propulsion technology programs are supported by the three prime vehicle contractors and are being conducted by the engine contractors (i.e., Rocketdyne, Aerojet, and Pratt & Whitney), several material contractors (e.g., Allied Signal, FMI), and Pennsylvania State University, as well as by the NASA MSFC and Lewis Research Center. MSFC provides direct support to the contractors, performs complementary tasks, and funds technology development tasks related to propulsion. The technology development, test, and analysis programs planned for Phase I include: analysis of existing rocket engines that may be considered for X-33; evaluation of modifications to existing engines for upgrades to X-33; analyses of new engine systems both for the X-33 and for advanced RLVs; technology programs to improve performance, producibility, reusability, and maintainability/operability for application to the X-33 and RLV; and advanced concept studies to evaluate candidates to improve future RLVs. Table 5-1 is an assessment of technology programs in terms of performance, producibility, reusability, and maintainability/operability. Engines The propulsion systems for the three vehicle configurations under consideration for Phase II have different requirements. Each contractor is proposing a different propulsion package using the common oxidizer/fuel combination of LOX/LH2. Prior to the first meeting of the committee, the proposed use of the tri-propellant LOX/LH2/kerosene was abandoned, apparently because the complexities of requiring a third propellant in both the vehicles and the engines would offset the marginal improvement in performance. The engine systems being considered by the airframe contractors are grouped below by existing engines, modified and upgraded engines, and new engines, along with the programs to be carried out in both Phase I and Phase II to evaluate alternate ideas and features of these engine types. Various components for each of these engines will be analyzed and, assuming favorable results, will be included in the hardware and tested in cold flow, hot flow, and structural test facilities. Existing Engines Four engines that have undergone extensive ground and flight testing are proposed for application to the X-33, the SSME, RD-0120, RL-10, and VULCAIN. SSME. This is the basic LOX/LH2 engine that has powered the space shuttle for 70 flights. The SSME develops a nominal 470,000 lb vacuum thrust at an Isp of 453 seconds and is capable of throttling from 104 to 65 percent. The SSME has an excellent

OCR for page 55
Reusable Launch Vehicle: Technology Development and Test Program TABLE 5-1 NASA Industry Cooperative Propulsion Technology Programs Program Performance Producibility Maintainability/ Operability Reusability Remarks Engine-specific programs to achieve Phase II SSME Derivatives (RKD) Block II+         Block II+ and Block III engine programs require NASA's implementation of Block II; engine improvements independent of X-33; under the shuttle contract. Nozzle area ratio 57:1 ✓   ✓ ✓ Nozzle modifications will enhance engine F/W. Block III         All programs are pertinent. Light weight nozzle ✓ ✓       Universal main chamber ✓ ✓ ✓ ✓ Schedule details require clarification. Jet boost pumps ✓ ✓ ✓ ✓   Enhanced health monitoring ✓   ✓ ✓   Electro-mechanical actuator valves ✓ ✓ ✓ ✓   Block III controller ✓ ✓ ✓ ✓   RD-0120 (Aerojet) Familiarization   ✓     All programs are highly pertinent to establishing performance of Russian engine to U.S. standards. Lightweight nozzle-analysis & design ✓ ✓     Component changes will increase F/W ratio. Nozzle area ratio: 37.5:1 ✓   ✓ ✓   Altitude compensating nozzle analysis ✓         Component hardware test ✓ ✓ ✓ ✓   Electronic controller breadboard ✓ ✓ ✓ ✓  

OCR for page 55
Reusable Launch Vehicle: Technology Development and Test Program Program Performance Producibility Maintainability/ Operability Reusability Remarks Life test CADB and NASA/MSFC ✓ ✓ ✓ ✓   J-2S Aerospike (RKD)         All programs are pertinent to support Lockheed design. Modular thrust cells cold flow and combustion ✓         Aerospike nozzle flows ✓       Schedule for first two programs may need to be accelerated. SR-71 flight flow evaluation ✓         Programs supporting technology for specific RLV engines RL-400 (P&W)         This is a new engine. Some turbopump work is supported by contract. NASA coop programs are not detailed. Advanced nozzles: two position; dual contour ✓         Milled channel nozzle ✓ ✓ ✓ ✓   Nonmetallic skirts ✓ ✓       Turbopump; hydrostatic bearings, integral rotor ✓ ✓ ✓ ✓   Full-flow preburner ✓   ✓ ✓ Intense materials evaluation is needed for oxygen-rich combustors. Gas-Gas injector ✓   ✓ ✓ Intense materials evaluation is needed for oxygen-rich combustors. RS-2100 (RKD) Engine definition ✓       This is a new engine.

OCR for page 55
Reusable Launch Vehicle: Technology Development and Test Program Program Performance Producibility Maintainability/ Operability Reusability Remarks Supercooled propellants ✓       Effect on engine restart needs to be evaluated. Deep throttling ✓       Preliminary tests on SSME. Oxygen-rich turbine drive ✓   ✓ ✓ Required materials evaluation for oxygen-rich combustors initiated. Gas-Gas 40K injector ✓         Restart ✓       Emphasis is required. Dual bell nozzle ✓         Jet pump ✓ ✓ ✓ ✓ All pump work is pertinent. Schedule details need firming. Revolutionary reusable technology turbopump (RRTT) ✓ ✓ ✓ ✓   RS-2200 (RKD) (RLV Aerospike)         Schedule detail lacking, but acceleration appears to be required to validate engine concept for X-33/RLV. Injector concepts ✓         Single-thrust cell (round/rectangular) ✓         Multicell technology ✓         Alternate nozzle shapes ✓       Altitude compensating. RRTT or IPD pump ✓ ✓ ✓ ✓   Restart ✓         Additional support programs DC-XA Project. (McDonnell) ✓   ✓ ✓ A test vehicle to demonstrate vertical takeoff and landing.

OCR for page 55
Reusable Launch Vehicle: Technology Development and Test Program THROTTLING Throttling is the ability of an engine to operate at various thrust levels different from, and usually lower than, the designated nominal value. Throttling is required both for ascent and landing to control acceleration or deceleration of the vehicle to the appropriate level. Some throttling is necessary regardless of the vehicle design or configuration; however, particular configurations or designs may require less severe throttling. For instance, if the vehicle is designed so that aerodynamic lift can be used in landing, then the throttling requirements are less severe in landing than for a comparable vehicle that relies solely on the engine to decelerate it to the appropriate level. Throttling levels often are quantified as percentages of the optimum nominal thrust of the engine. For example, if the output thrust of a nominally 100,000 lb engine can be controlled to provide a thrust of 10,000 lb, then throttling to 10 percent has been achieved. The capability of an engine to produce a small fraction of its nominal thrust during flight (as given in the previous example) is sometimes referred to as deep throttling. Several methods of throttling have been used in engine systems. These methods all reduce the flow of propellants to the main thrust producing chamber, which results in (1) a reduction in combustion pressure, and (2) a reduction in thrust. performance record and outstanding flight reliability on the space shuttle. However, because of its weight and inadequate F/W, as well as extensive turnaround time and requirements for checking out and refurbishing components, the basic SSME does not meet the long-term requirements for a less costly, easily maintained RLV with rapid turnaround capabilities. Proposed use of the SSME in the X-33 vehicle is marginal. Technology programs are in place to improve the basic engine. The current technology programs will result in the modified SSME Block II, which is described in the section on upgraded engines. RD-0120. This Russian LOX/LH2 engine, designed and tested by the Chemical Automatics Design Bureau, was first used in 1976 and has completed 793 ground tests and two successful flight tests (four engines each) on the Energia vehicles. The engine develops a 441,000 lb vacuum thrust and, with the existing design, has a vacuum Isp of 455 seconds. The RD-0120 reportedly has a throttling capability of 114 percent down to 25 percent. Without major modifications the RD-0120 cannot meet RLV sea-level F/W requirements. Improvements in F/W similar to those for the SSME are required for X-33. Plans are underway by one contractor to upgrade this engine for increased operability and modest increases in sea-level F/W. RL-10s (RL-10A-3-3A and RL-10A-4). These engines have extensive ground and flight experience and develop a 20,000 to 25,000 lb thrust and an Isp of 390 seconds. It is proposed that these engines be used in clusters for various X-33 roll control and descent modes. VULCAIN. A French/SEP (Société Europeenne de Propulsion) engine that develops a 200,000 lb thrust. VULCAIN is currently used in the Ariane vehicle program. It is proposed for cluster installation in specific X-33 applications.

OCR for page 55
Reusable Launch Vehicle: Technology Development and Test Program Table 5-2 shows the characteristics of the existing flight-proven 400,000 lb thrust LOX/LH2 engines under consideration for primary propulsion. Upgraded Engines Optional upgrades are available for all of the existing engines to improve specific qualities. Improvements will be evaluated in the proposed technology programs. Application of the advanced features is intended for the X-33, with later development for the RLV. SSME Block II. Modifications of the present SSME engine that will be used in the space shuttle program include improved LOX and LH2 turbopumps, a single-tube heat exchanger, a two-duct powerhead, a simplified low-pressure oxidizer turbopump, a large throat main chamber, and an improved engine controller. This engine will be certified for use in the space shuttle by 1997 and will be available for the X-33. Unlike the current SSME, which requires that engines be pulled between flights to replace turbopumps, the Block II engine should withstand at least 10 launches before engines must be pulled. TABLE 5-2 Characteristics of Flight-Proven 400,000 lb Thrust LOX/LH2 Engines Parameter U.S. Space Shuttle Russia: Energia Engines Engine designation SSME, Block II (in production) RD-0120 (in production)a Thrust (1,000 lb) 395 (SL) 330 (SL) Isp (s) 453 (vac) 341 (SL) 455 (vac) Chamber pressure (psia) 3,200 3,170 Feed system Staged combustion Staged combustion Mixture ratio 6.0 6.0 Throttling capability 65–104% 25–114% Expansion ratio 77.5:1 85.7:1 Restart capability Yesb Yes Dry weight (lbs) 7,675c 7,606 FN/WE 51 43 a Data source: Aerojet briefing presented to the committee on August 4, 1995. b Pending proper propellant conditioning. c Rocketdyne briefing presented to the committee on August 2, 1995.

OCR for page 55
Reusable Launch Vehicle: Technology Development and Test Program Technology programs are also planned to improve the SSME engine. Two improved versions, Block II+ and Block III, are planned. Block II+ incorporates a 57:1 shortened nozzle; the Block III engine features a lightweight nozzle, jet boost pumps, electrical valves, a new combustion chamber, and a new controller. The Block II+ engine could support the X-33, but Block III would be available only for the RLV. The goal for Block II+ is a sea-level F/W of 58; the target for Block III is a F/W value near 70. RD-0120 AD-1. This modification of the basic RD-0120 design features a reduced nozzle expansion ratio of 37.5:1 and an electronic controller. Planned development includes certification of rapid reuse operations and 10 turnarounds without refurbishment. This modification will allow the RD-0120 to support the X-33 program. RL-10A-5-1. This improvement will increase thrust to 25,000 lb and increase the nozzle expansion ratio (59.5:1) for increased altitude performance. VULCAIN Mk II-1. The improved VULCAIN will have an increased MR (from 5.3 to 6.2), a larger nozzle expansion ratio (62:1), and deeper throttling. New Engines There are three new engines projected for use in RLV planning. One of these, the Aerospike, is being funded and developed by an industry contractor for an X-33 demonstration. The other two, RL-400 and RS-2100, are not applicable to the X-33 because they require funding as well as new long-term development programs. Aerospike. This concept, which evolved from considerable development work in the 1960s, begins with development of an engine for the X-33 and is followed by development of a larger, gas generator cycle engine (RS-2200) for use in a future RLV. The X-33 Aerospike is a 200,000-lb thrust unit that uses 2 banks of 12 thrust cells each in a linear array that will be installed in the boattail of a lifting body X-33 using Saturn V J-2S turbomachinery. The technology programs for the evaluation of this concept are integrated through Phase I and Phase II. Single-thrust cell testing will be completed in the third quarter of FY96, and multi-cell preliminary testing will be completed in the fourth quarter of 1996. Engine design, fabrication, hot firing, and flight tests are scheduled for Phase II. A scaled Aerospike system will be flown (piggy backed) atop an SR71 aircraft and hot fired during the altitude portion of the flight to determine rocket exhaust effects. These test results, combined with wind tunnel tests and computational fluid dynamics, will be used as design criteria for one contractor's X-33 flight design. If selected, the advanced Aerospike systems are projected for RLV implementation. RL-400. This is a projected full-flow, bipropellant, preburner engine that uses oxygen-rich gases to drive the oxygen turbopump and increase engine efficiency. The development of this engine is proposed to support an RLV vehicle after the year 2000. Technology programs during Phase I and Phase II are aimed at analysis and evaluation of engine components. RS-2100. This is a projected full-flow, staged-combustion engine that uses a fuel rich preburner on the fuel side and an oxidizer-rich preburner on the oxidizer side

OCR for page 55
Reusable Launch Vehicle: Technology Development and Test Program operating at reduced turbopump temperatures for greater pump operating margins. Many other improvements will be implemented. This engine is also projected to be developed for flight after the year 2000. The two proposed engines now being analyzed and evaluated are the RL-400 and RS-2100. Both use a full-flow, staged-combustion cycle, which is the preferred cycle for conventional nozzle engines. This cycle uses the engine propellant flow, LOX-rich propellant to drive the LOX pump turbine and fuel-rich propellant to drive the fuel pump turbine and mixes the two in the thrust chamber. This method provides substantial power to the turbopumps while maintaining relatively cool turbine inlet temperatures. By passing the entire flow through the turbines, this cycle keeps low turbine operating temperatures, which results in longer life and fewer inspections. Seal leakage concerns are also reduced because the turbine gas is compatible with the pumped fluids on both the fuel and oxidizer sides of the engine. Full-flow, staged-combustion engines may have distinct advantages over other engine types, and several NASA/industry programs have been initiated to investigate this potential. For example, materials evaluation programs for oxygen-rich combustion, new element designs for gas-gas injection, and full-flow preburner development are included in several Phase I and Phase II technology programs. The complementary U.S. Air Force (USAF) Integrated Powerhead Program is directed toward building components for these engines, albeit on a smaller scale than those required for the RLV (250,000 lb thrust). This USAF program and other technology programs being conducted at Air Force facilities (Phillips Labs, and others) are structured to allow transfer of all the results and findings to the contractors on the X-33 and RLV programs. The data will be shared to benefit the entire propulsion community. An engine based upon this cycle will not be available for the X-33. There appear to be no significant "show-stoppers" in the development of full-flow combustion engines although development of the hardware may be complex. One challenge is that preburners must be built of materials and coatings that are resistant to hot oxidizer-rich gases. Russia has demonstrated this technology, but it has not been used in U.S. rocket engine development. Developing new, simplified pumps with fewer parts that are enabled by hydrostatic bearings will also be a challenge. Depending on the rate of progress and available funding, this type of engine may or may not be available in time for use in the RLV. The committee sensed that all of the contractors, especially the engine builders, may be waiting for the government to initiate development of a new engine. Compared with the costs of improving an existing engine, developing a new engine is expensive and may raise questions regarding cost effectiveness. Engine Applicability to X-33 and RLV Both existing LOX/LH2 engines (SSME & RD-0120) can be modified for an X-33 flight vehicle. Technology programs are underway to modify them by reducing weight and permitting longer intervals between servicing. Tests are underway on the SSME testbed at MSFC to demonstrate a 25 percent extension of throttling capability for the

OCR for page 55
Reusable Launch Vehicle: Technology Development and Test Program current SSME. This improved capability may be adequate for X-33 requirements. However, in view of the approach being taken, this demonstration is of questionable value for two reasons. First, it does not reproduce the throttling dynamics associated with going from full thrust down to 25 percent; and second, the SSME uses Rocketdyne pumps instead of the Pratt & Whitney ATD pumps planned for the RLV. The current SSME must demonstrate deep throttling to satisfy one prime contractor's RLV requirement; another requires only current throttling capability of 65 percent. The RD-0120 reportedly has throttling capabilities to 25 percent and, even with the weight-reduction potential, will require an increase in F/W before it can be used as a cluster engine in an RLV. When upgraded by the technology programs described here, both the SSME and RD-0120 engines are viable options. Both could be ready in time, and both are directly applicable to the RLV. Work on reducing the cost of these engines was not apparent at the time of this study. The Aerospike engine is viable for the lifting body X-33, as has been demonstrated in the Aerospike technology programs. But engine characteristics of the required shape, size, performance parameters, and interaction must still be demonstrated (as scheduled). To meet both X-33 and RLV requirements, the research and development of Aerospike should include new manufacturing and chamber configuration modifications. To meet RLV specifications, an even greater effort will be necessary. The risks involved in timely development and modification of this engine system for the X-33 are higher than the risks for more conventional, proven engines. The smaller engines (RL-10 and VULCAIN) are being considered as auxiliary control systems and orbital maneuvering systems. Technology programs for these applications are notably absent from the Phase I and II programs. Performance With some improvements in performance, current engines can be used for X-33 demonstrators; however, RLVs will require very high rocket engine performance (thrust/weight and specific impulse). A series of technology programs are planned to improve these performance characteristics. Sea-Level F/W Improvement Rocket engines for the X-33 and RLV must show a considerable increase in sea-level F/W over current levels. The SSME Block I, which flew on the shuttle in 1995, and RD-0120 have F/W values of 51 and 43, respectively, whereas the requirement for the RLV is between 75 and 80, depending upon vehicle configuration, vehicle weight, and engine specific impulse. An increase of this magnitude will require some increase in sea-level thrust and major weight reduction in many engine components. At the same time, high reliability and low cost must be maintained, and reusability must be improved. This is not a simple task.

OCR for page 55
Reusable Launch Vehicle: Technology Development and Test Program For the X-33, modifications for both the SSME and the RD-0120 have been defined. These modifications could achieve a F/W approaching 60. The SSME Block II+ configuration with a truncated rocket nozzle (higher thrust and lower weight) is expected to come close to the desired performance. The RD-0120 with the same nozzle truncation and additional changes to other hardware would achieve somewhat lower performance levels. Historically, changes in sea-level F/W values have required long periods of development. Both the SSME and RD-0120 will require extensive modifications to achieve sea-level F/W values of 75 to 80. SSME modifications have been developed in detail, and the contributions of each modification to improved sea-level F/W performance have been quantified. The RD-0120 team also has identified required modifications, which would require long development times. (Details of these modifications are not included in this report to protect proprietary information.) The development of a new engine, RS-2100, for example, would not be constrained by existing envelopes and components. But a new engine would not benefit from the accumulated knowledge and experience of the existing engines. In either case, the task will be difficult to complete successfully. Throttling Whichever vehicle concept is selected, significant throttling capability will be needed. The Aerospike will throttle by keeping pumps at full rpm for pitch changes and diverting the flow from one thruster bank to another through a differential throttling valve; for yaw conditions, pump speed changes will be needed. Work on the differential throttling valve is included in the Phase I development and test program. Conventional bell engines will throttle by changing pump speed. The first approach (Aerospike) requires development of a sensitive valve and flexibility in the injection system to accommodate variable flows. Conventional bell engines rely primarily on the preburner(s) and pump. Programs are in place to demonstrate both conventional engine and Aerospike configuration throttling, but these programs may need to be expedited to meet RLV goals. Concerns about deep throttling during pump speed changes include controlling pump speed with turbine inlet temperatures that do not drop rapidly and are low enough to freeze moisture in the pumps; controlling the preburner injector dynamics, and ensuring sufficient flow to cool the combustion chambers and nozzle; overall system dynamics; and sustained preburner combustion at the mixture ratios required for deep throttling. Any of these engines that can successfully and reliably deep throttle would be viable for X-33. Turbopumps and Pumps Reducing the weight of the turbomachinery is a crucial factor in improving engine F/W. All three engine companies are establishing benchmarks for new turbomachinery

OCR for page 55
Reusable Launch Vehicle: Technology Development and Test Program with precision casting, hydrodynamic bearings, single crystal turbine blades, stout rotors, improved casings and shrouds, and other alternate designs. The proposed new pump designs are applicable only to RLV and will not be ready in time for X-33. The schedules presented to the committee are tight. Combustion and Mixture Ratio Several programs are planned to improve combustion performance and overall engine balance. In support of potential RS-2100 engine concept, a full-flow mixed preburner cycle featuring a LOX-rich preburner, which reduces engine power requirements, will be tested. Testing will include an analysis of material sensitivity to oxygen-rich products. Techniques for running preburner oxygen-rich gaseous combustion products into a main injector along with heated gaseous hydrogen fuel to achieve complete high-efficiency (gas-gas) combustion are also being evaluated. This approach does not apply to the Aerospike engine, which uses a conventional gas generator cycle. A main combustion mixture ratio of 7:1 is also being tested to evaluate performance loss versus vehicle mass fraction gain. Gas generators Oxygen-rich combustion, single versus dual units, and full-flow preburner development and test programs are scheduled. However, the programs do not yet appear to be detailed enough to verify the expected performance results. Injectors Gas-gas injection (for RLV), new element designs, and full-flow, high-pressure injection technology programs are proposed for improving engine performance; however, the advantages of these technologies must be evaluated in more detail to determine the feasibility of using them for the X-33. Combustion Chambers Programs to develop a large throat chamber, alternate materials, milled slot design modifications, and shape reduction were discussed with the committee. All of these concepts are viable and should result in performance gains.

OCR for page 55
Reusable Launch Vehicle: Technology Development and Test Program Nozzles On an SSTO vehicle, optimizing nozzle performance on Isp, from sea-level takeoff to the vacuum space environment, will be critical. Several nozzle technology programs are planned for reducing the expansion ratio operation at low altitudes and expanding to high ratios as altitude increases. Development of a two-position, moveable bell nozzle, a dual-inner-contour nozzle to induce flow separation, and an Aerospike altitude compensating nozzle are directed toward this objective. A series of wind tunnel flow tests are planned to evaluate various nozzle shapes upon expansion, with and without a double contour. New lightweight nozzles, alternate fabrications of nozzle shapes, variable nozzle expansion area ratios by means of inserts, dual-step nozzles for sea-level/altitude performance, and milled slot nozzles are also being investigated. These investigations are aimed at: (1) improving sea-level performance of a high nozzle expansion ratio; and (2) reducing the weight of the units. The programs have high merit, and there is much to be learned from them. Producibility, Reusability, Maintainability and Operability Manufacturing Operations In general, the engine companies did not highlight overall engine manufacturing procedures and materials in their presentations to the committee. No product improvement programs were presented that addressed the question of manufacturing hardware exactly to print, although this is a common problem in existing hardware programs and decreases producibility. Because there is only one major U.S. development program for reusable engines, attention should be paid to the issues of producibility. One of the existing flight-proven engines that may be applicable to the X-33 (i.e., the RD-0120) was designed, developed, and qualified in Russia, where the Voronezh plant manufactures as many as 20 engines a year. Manufacturing time for each engine is reported to be one year. The time required to manufacture one SSME is four to five years, leading to the conclusion that it would be prudent to examine carefully the manufacturing processes and controls used by Russia's Chemical Automatics Design Bureau (CADB). Under a strategic business partnership between Aerojet and the CADB, Aerojet plans to test the RD-0120 in the United States and, pending contract award, will create U.S. production facilities for the RD-0120. A NASA/industry program will facilitate this technology transfer and provide benchmark life testing of the RD-0120 both in Russia and at MSFC. The life tests will provide answers to questions about the maturity of the X-33 and questions on long life engine design. Life tests will be duplicated at NASA MSFC.

OCR for page 55
Reusable Launch Vehicle: Technology Development and Test Program Turbomachinery (turbines, pumps, assemblies) Programs are in progress to enhance the producibility and maintainability/operability of the high speed turbomachinery required for reusable high-performance engines. For example, the producibility of the alternate high-pressure LOX and hydrogen turbopumps of the SSME Block II engine has been improved substantially through development of nonmetallic bearing balls, integral turbine tip seals, precision castings, and single crystal turbine blades. Precision castings have eliminated the sheet metal housings, which required many welds, in the SSME Phase II engine turbopump designs, thus reducing the need for tedious crack inspections and crack weld repairs during fabrication. The LOX pump is already in use and should be easily available for X-33. The fuel pump is still being tested, but it should also be ready in ample time for X-33. The increased weight of these pumps can be a great drawback, however. Improved turbomachinery is vital to the RLV. In this context, improved means reducing weight substantially while maintaining or exceeding the reusability demonstrated by the Block II SSME turbopumps. If Block II pumps live up to expectations, they will approach the minimum reusability requirements for the RLV. Engine F/W is critical to the SSTO, and engine weight is particularly important because of the aft location. Applicability of the best available jet engine technology and experience must be emphasized. This may be expensive, but payoffs will be high. Revolutionary Reusable Technology Turbopump (RRTT) and Other Advanced Turbopumps All of the engine manufacturers, in cooperation with NASA, are evaluating advanced turbopumps for future engines. The current programs are applicable to new engine concepts in the 400,000-lb-thrust class, such as the Aerospike, the RL-400, and RS2100. Although the approaches differ in detail, all advanced concepts involve reducing the number of parts, using hydrostatic bearings, investigating advanced manufacturing approaches, and utilizing new materials. The record of the SSME pumps clearly indicates that this is a fruitful area for improvement. Development of advanced concepts is essential to achieving the ultimate goal of a highly operable vehicle. Advanced turbopumps incorporating new materials and substantially fewer parts, as exemplified by the RRTT concept, will greatly reduce required maintenance and enhance operability by reducing failure modes and eliminating currently required inspections. Hydrostatic (Hydrodynamic) Bearings This technology is fairly mature but has not been used in rocket engine turbopumps, although some tests were conducted in connection with the alternate SSME pumps. Hydrostatic bearings are being considered for use in the RRTT and other advanced turbopump designs. These bearings may eliminate many of the failure modes that are a major source of problems in current engines and make designing long-lived

OCR for page 55
Reusable Launch Vehicle: Technology Development and Test Program pumps easier because they would not have critical speed problems. Although highly reliable pumps can be designed without this technology, hydrostatic bearings show great promise and could be available for use in an RLV engine with further development. High Performance Low Maintenance Powerheads This unit features a higher performing injector and a new, single-tube heat exchanger. Manufacturing heat exchangers for converting liquid oxygen to gaseous oxygen (to pressurize the oxygen tanks) has been complicated because the exchanger consists of a primary tube, a bifurcation joint, and two secondary tubes that are assembled by welding. Advanced technology to produce the very long jointless tube of the appropriate material has recently been developed, and single-tube heat exchangers can now be fabricated. The new heat exchanger eliminates a potential Category I failure (i.e., a failure involving loss of life or mission) that might have occurred as a result of leakage in one of the many welds in the original heat exchanger. Eliminating the welds also enhances producibility. The single-tube heat exchanger also improves maintainability and operability, and eliminating welds reduces concern about leakage and failure of tubes. This, in turn, reduces the need for inspections and checking for leaks. Combustors and Nozzles The programs for improving the producibility of the main combustion chamber (MCC) are based on eliminating welds and developing new fabrication processes. The potential payoff of this approach is illustrated in the case of the SSME, for example. Rocketdyne reported that the SSME Phase II engine MCC requires 40 months to manufacture. MCC manufacture of the SSME Block II engine will be reduced to 24 months by using precision castings of the combustion manifolds rather than welding and by improvements in the plating and assembly process. Changing to a proposed milled channel combustor allowed the manufacturer to demonstrate production of a universal MCC in 12 months. This universal MCC will need to be certified prior to use in the SSME, but these changes in fabrication and material processes have yielded dramatic improvements in producibility. The goal is to extend the life of the SSME thrust chamber to at least 100 flights. Although the universal MCC would greatly enhance engine producibility, it may not be available in time for the X-33. Programs for simplifying the fabrication of nozzles have also been proposed. One of these combines Russian technology and advanced manufacturing technologies to produce a lightweight, milled, channel nozzle.

OCR for page 55
Reusable Launch Vehicle: Technology Development and Test Program SSME Block III Controller/New RD-0120 Controller All of the engine manufacturers propose programs to use electromechanical actuated valves and simplified electronic controllers. Although the current SSME controller has not caused major problems, a newer controller with modern electronics will weigh less and be more reliable; it will also enhance operability and maintainability. The additional capacity of the controller for health monitoring should be even more beneficial. A new U.S. technology controller for the RD-0120 will offer similar improvements. The new SSME and RD-0120 controllers should have applicability to the X-33. Valve Actuation If electrical actuators are used to open and close valves, pneumatic and hydraulic systems, along with their numerous parts and potential for leaks, could be eliminated. The higher cost of electrical actuators derives from the need for intermittent high electric power. In general, electric valve actuation is a mature, available technology that can improve reusability. And the overall simplification of this method will reduce failure modes. The primary concern seems to be whether a sufficient variety of valve/actuator combinations is available to fit all needs. It was not clear to the committee how thoroughly this question is being addressed. Health Monitoring Onboard or built-in health monitoring for rocket engines involves installing instrumentation to measure critical temperature, pressure, vibration, and rotation speed. By monitoring these parameters during operation, and particularly by noting trends, engine health may be assessed, which can eliminate many aspects of ground inspection (e.g., torque checks on pumps) and expedite maintenance by allowing the scheduled replacement of components. The software to monitor and analyze the data is as critical as the instrumentation. An overarching architecture to determine which measurements are necessary is essential for vehicle health monitoring. This architecture will define the software instrumentation to be developed and the ways subsystems will interact with the vehicle controller. Although built-in health monitoring will not directly enhance near-term reusability, this capability will greatly increase confidence by increasing knowledge of the condition of hardware. Onboard health monitoring can pinpoint failures before they happen, so the affected components can be replaced in a timely manner. By establishing trends and identifying weak points, onboard health monitoring will encourage product improvements that will substantially enhance reusability. There is general agreement about the importance of engine health monitoring; however, there is less agreement about exactly what to measure and how to measure it.

OCR for page 55
Reusable Launch Vehicle: Technology Development and Test Program The latter is particularly important because spurious data or the failure of sensors could cause major problems. The development of software to analyze data must go hand in hand with the development of sensors. This is a complex problem with a high payoff and should be pursued vigorously because test and evaluation on the X-33 are essential. The committee received relatively little information about health monitoring of the rocket engine from NASA/industry partners. For the proposed RLV and, for reasons of traceability, the X-33, enhanced engine health monitoring is essential. Relevant work is underway to develop and test new nonintrusive sensors, and/or more rugged sensors; analysis software is also being developed. It is not clear if the program will reach the needed level of maturity in a timely fashion. High Reliability Sensors Erratic readings and the mechanical failure of sensors have long been a problem in rocket engine work. Development of more rugged, nonintrusive sensors to measure temperature and pressure is being pursued with an eye toward improving reliability and reducing failures that might cause engine damage. Such failures have caused premature shutdown of shuttle engines in flight and have raised concerns that mechanical failures could cause catastrophic damage. Given the importance of health monitoring and the continuing concern about engine damage from sensor failure, newer more robust and reliable sensors are vital. The development of sensors should receive special emphasis in the RLV program. New Fabrication Techniques New fabrication methods, such as friction-stir welding and near-net-shape forming show promise for reducing fabrication-induced stress and cracking that lead to potential failures. Friction-stir welding is a developmental process that will require significant maturation, whereas near-net-shape forming is a fairly well-developed technology. The potential reduction of failure modes should greatly enhance reusability. A new welding process, for example, could reduce many of the concerns about Al-Li welding. However, it was not clear how rapidly the new processes are being pursued or whether they will be available in time for the RLV. New Materials Composite materials show promise for reducing weight and increasing the robustness of some engine parts, such as lines and valves. Experimental work is in progress on composite cryogenic valves and lines. Ground tests are also in progress, and these components may be flight tested in 1996. If successful, these components could be

OCR for page 55
Reusable Launch Vehicle: Technology Development and Test Program ready for the RLV engine. This will probably require some acceleration of the effort to ensure that the engine companies are comfortable with the new components. Such technology improvements, at relatively low cost, address the critical question of engine weight and will contribute substantially to reducing the turnaround time by reducing the need for inspections and maintenance between flights. FINDINGS AND RECOMMENDATIONS Findings The most significant finding is that the prime contractors believe an engine sea-level F/W ratio of greater than 75 is required for the RLV. The SSME Block II and the RD-0120 engines provide sea-level F/Ws of 51 and 43 respectively, and the SSME Block II+ (with a short nozzle) may achieve F/W of 58. An increase of approximately 30 percent in sea-level F/W presents developers with a difficult challenge. Developments to achieve this increase have been identified by the contractors; however, the committee believes that achieving greater than 75 F/W will be very difficult, even with a totally new engine. Upgrading an existing engine to meet this challenge, although less costly than developing a new engine, will be even more difficult. The shortened nozzle modifications can facilitate an increase in the sea-level F/W performance of the SSME or RD-0120 engine. Two of the three prime contractors appear to have selected the SSME or RD-0120 for the X-33 technology testbed. At least one, and possibly both engines, will require throttling to a level deeper than has been currently demonstrated by SSME. Tests are underway on the SSME testbed engine to demonstrate throttling to 25 percent. However, it is not clear to the committee that there is a fall back position if this demonstration is unsuccessful. In addition to development of the engine for the X-33 in Phase II, there are plans for an engine development and ground test program leading directly to engine technology for an RLV. However, these plans are not well defined at this point. The X-33 engine will only partially demonstrate scaleability to the RLV, so engine development and testing on the ground will have to be relied upon to demonstrate scaleability to the RLV. The schedule for development and qualification of the Aerospike engine for flight on the X-33 may be difficult to meet because development of a new engine typically takes as long as a decade. Some of the concepts being evaluated may require restart of the main engines, including both engine/propellant conditioning and ignition mode. Although this has been done in the past with a variety of engines, restarting the main engines during flight can be difficult. In addition, engine/propellant conditioning may constitute a significant mass penalty in a mass-critical vehicle. Development efforts to reduce engine turnaround time significantly after each flight do not adequately reflect the desire for a rocket engine that can be handled the way an operational jet engine is handled. The proposed turnaround procedures for the SSTO

OCR for page 55
Reusable Launch Vehicle: Technology Development and Test Program were arrived at by reducing the number of similar procedures currently used for the space shuttle rather than by initiating procedures tailored to the RLV. Recommendations RLV engine sea-level F/W requirements to achieve SSTO should be revalidated independently by the prime contractors and by NASA's vehicle design and performance groups. Current F/W goals of greater than 75 will be very difficult to achieve with existing engines and even new ones, without compromising the structural margins required to satisfy reusability goals. If the requirement of high (sea-level) F/W is revalidated, the committee recommends that development of the selected RLV engine be initiated at the beginning of Phase II and pursued vigorously. Because of the different requirements for the X-33 vehicle engine, it will not advance the F/W goal by much. Therefore, the decision to proceed with Phase II will have to be based largely on data from the development program. There should be ongoing trade studies to assess whether larger, but still viable, vehicles will satisfy the F/W requirement. The decision criteria for progressing from Phase II to Phase III for the propulsion system should reflect the required RLV engine performance targets (such as F/W of greater than 75 and vacuum Isp of 440 seconds or higher). NASA should evaluate the contractors' detailed analyses of projected methods and component improvements for achieving a sea-level F/W greater than 75. The practicality of each required component design should be documented by the engine contractors and evaluated by NASA and, perhaps, by an independent group of propulsion experts. The RLV engine ground program for Phase II should be thoroughly defined and executed to provide a high level of confidence that RLV engine requirements will be met. If the prime contractors considering SSME or RD-0120 engines for the X-33 demonstrator determine that higher sea-level F/W performance is needed, development of the short (truncated) nozzle should begin soon. Dual contour nozzles specifically designed for optimum expansion conditions through the flight trajectory must be verified by hot firing tests in a flight-like environment to assure smooth flow transition without excessive side loads or abrupt skewed shock conditions. A plan for developing and qualifying the required throttling must be proposed. Specific details and experimental verification of deep throttling must be demonstrated on the engines proposed for the X-33 and RLV vehicles requiring this capability. Requirements for restarting the main engine should be evaluated. A plan showing how this requirement will be met should be developed.

OCR for page 55
Reusable Launch Vehicle: Technology Development and Test Program The detailed configuration of combustor body and throat shape, nozzle shape and expansion ratio, and vehicle integration for the X-33 and RLV Aerospike engine should be completed before the Phase II decision date. Throttling and thrust vector methods, including interaction effects between adjacent engines should also be evaluated. Overall engine health monitoring requirements should be better defined. More robust and reliable health monitoring methods and instrumentation than are currently used should be developed and thoroughly tested. NASA should evaluate the program and engine changes required to meet the goals of rapid turnaround times. In general, operability and engine reliability requirements should be developed for the X-33 and RLV. Producing an RLV engine that does not have to be touched between flights unless problems are indicated by on-board health monitoring or visual inspection should be a design goal. NASA and industry should consider funding high risk/high payoff technology efforts after the X-33 is selected.