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Reusable Launch Vehicle: Technology Development and Test Program 6 Conclusions and Observations In terms of organization and content, the Phase I technology and test program is generally sound. With the participation of three industry partners and three NASA centers, the program for developing and testing materials-related areas (cryogenic tanks, primary structures, and thermal protection systems) as well as the propulsion system are robust in that more than one approach is being pursued in almost all critical areas. The program is rationally directed toward eliminating as many unknowns as possible early in the technology development and test program—before they are on the critical path in a vehicle development program. In several areas of materials and propulsion, the program is pushing the envelope beyond proven technology, and some failures are to be expected before the required knowledge is obtained. However, the committee found that the Phase I development, test, and analysis programs are appropriate to support a decision regarding whether to proceed with Phase II, subject to implementation of the recommendations herein. It should be noted that, owing to time constraints on the study, the committee could not review several important areas, which are listed below: important aspects of the propulsion system, such as plumbing, leak sensors, lines, valves, and joints upstream of the engine, purge systems, pressurization systems, and the small reaction control system and orbital maneuvering system the integrated health monitoring system for all components and NDE technologies ground support equipment for the propulsion system, such as propellant quick disconnect, automation, automated fluid and electrical connections, and safe, operationally efficient ground and flight/vent purge systems operations issues that were explicitly excluded from the committee's charge to make the committee's task feasible within the allotted time
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Reusable Launch Vehicle: Technology Development and Test Program MATERIALS The committee anticipates that when the Phase II exit criteria for the development and testing of materials-related areas have been met, the following results may be achieved: A first estimate of the mass fraction achievable for these components will be available, especially from the larger scale test articles (e.g., 8-ft to 14-ft-diameter cryotanks and larger primary structures). However, for the estimate to be reasonable, each of the various sized test articles must be designed, built, and tested to RLV-scaled conditions using the design codes that are being validated. All of the joints and fittings for the larger test articles should be properly scaled to the RLV flight configuration. Reusability will be partially demonstrated by subscale components and cryostat/pressure box integrated structures, insulation, and TPS panel tests. In all probability, thermostructural cycle testing of the larger test articles will not be adequate to answer the question of whether these components will achieve their cost per flight goals. This will be demonstrated only with adequate cycle testing of the totally integrated large test articles in flight configuration (structure, insulation, and TPS) and in appropriate flight-equivalent environments. Flight tests on the X-33 technology demonstration vehicle, with the individual components integrated in vehicle configuration and with health monitoring systems in place (which have been tested pre-flight with applicable NDE techniques) will verify these essential qualities. For the materials-related technologies, the question of scaleability of test results from subscale test articles, including the X-33 vehicle, to the RLV will be based on systematic validation of the analytical codes, such as NASTRAN and Mechanica, that were used to design the RLV. For this validation to be effective, all test articles, from small bottles of composite material for the LH2 tank (as an example) through intermediate-sized test articles, and to the 8-ft-diameter tanks, must be designed by the codes being validated, properly scaled to RLV conditions, and tested with scaled forces and in appropriate thermal environments. The validated code should then be used to design the X-33 vehicle tanks (scaled to RLV conditions) and the flight results compared to pre-flight predictions of the design codes. The other requirement for proper scaleability is that the test articles must be fabricated by the exact same process to be used for the RLV article. If the program is carried out in this manner, as anticipated, it should validate design codes for the RLV. However, compelling reasons for resolving the scaleability issue for the propellant tanks and primary structures include: The super lightweight external Al-Li tank using 2195 alloy is being developed for the shuttle program. Because this tank is 28 ft in diameter, the design codes will be validated to that size, so much less information
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Reusable Launch Vehicle: Technology Development and Test Program will have to be extrapolated to the 40-ft-diameter tank of the RLV than from the 14-ft-diameter test article in the X-33 program. Another important factor will be the scaleability of the analysis of the welding of the Al-Li tank. If using dissimilar materials becomes necessary, welding will induce stresses because of coefficient of thermal expansion mismatches, and scaling will have to be correlated with subcomponent (full-size) testing. For the LH2 composite tank, the committee believes that the program should "demonstrate fabricability and structural capability of a full-scale cryotank during Phase II." This statement reflects a concern about fabricating such a large graphite-epoxy tank with the desired properties (and the lack of a large autoclave, if needed), and the scaling accuracy of the design codes over such a large extrapolation range (from 8-ft-diameter test tanks to the 40-ft-diameter RLV tank). Correlation of analysis codes to full-scale sizing will be difficult because (based on contractors' tests of 32-inch bottles to 8-ft tank design) areas that can not be stressed to critical levels because of scaling issues will not correlate correctly into the mathematical models; thus the test results may not indicate critical stresses in the full-scale RLV design in complex geometric areas, such as Y-joints, skirt, and conic and internal tank structures. Full-scale test articles for these more difficult areas will probably be required. For the composite primary structures, one of the contractors will fabricate test articles that are segments of full-scale RLV components to address the scaleability issue. These will be segments of the intertank, thrust structure, and composite wing structure. Other contractors will use 8-ft-diameter intertanks. Scaleability of load peaking, that is, edge effects at geometric discontinuities, will be important for correlating from testing of scaled designs. These data will be useful for designing structural interfaces for the RLV. The global effect of TPS on the RLV primary structure must be understood. Increased loads and thermal environments caused directly or indirectly from the TPS must also be correlated into mathematical models for structural influences. Many of the test articles of the TPSs will be full-scale. In cases where subscale test articles are used, or where the thermal or aerothermal environment is not simulated, appropriate analytical codes developed and validated over the years, including the large shuttle database, will be used to properly scale the test results to full-scale RLV conditions. PROPULSION By the end of Phase I, the propulsion program will have identified X-33 and RLV versions of existing engines (SSME and RD-0120) for two RLV contractors, and the Aerospike engine for the X-33 vehicle will be defined. Critical component technologies
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Reusable Launch Vehicle: Technology Development and Test Program will be identified and/or under development for all candidate engines. Hot fire tests are planned for three different engines: a three-cell X-33 Aerospike configuration current SSME demonstration of rapid turnaround time, high mixture ratio, and throttling extended-life RD-0120 engine benchmark tests in Russia to demonstrate reusability, followed by life tests in the United States with operability enhanced electronics added to the benchmark engine. OBSERVATIONS The following suggestions may fall outside the charter of this committee, but were deemed important. The committee would like to emphasize that satisfying the Phase II decision criteria does not guarantee that the two major objectives of SSTO, performance and low cost per launch, will be achieved. Some of the reasons have been discussed in the preceding section. Other concerns have to do with the considerable extrapolation in size of test articles to full-scale and issues still to be resolved with propulsion development. But satisfying the Phase II criteria will be a good benchmark indicating that risks have been lowered enough to allow proceeding to the next phase. The committee believes this is the same rationale NASA, OMB, and OSTP used to formulate the decision criteria. It is important to realize that the RLV requires engine technology beyond the technology demonstrated in the X-33 engine. The X-33 should test as many critical components for the SSTO and TPS as possible. The cryotanks and composite primary structures should be designed to RLV-scaled conditions. The X-33 is a test program and, as such, should be prepared to take prudent risks. It would be advisable to demonstrate fabrication and structural capability of a full-scale composite cryotank during Phase II to avoid undue risks during Phase III when heavy expenditures will be necessary. For a suborbital X-33, the committee's queries about launch sites and, much more important, landing sites, revealed that this issue requires considerably more attention because of technical constraints imposed by various potential orbits. As described in the introduction to this report, the study focused on specific technologies that must be advanced to achieve an SSTO RLV. The ASEB study committee and NASA identified the four categories of technology to be studied: PVS, RCTS, TPS, and propulsion. These categories are discussed individually in the preceding chapters, but little
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Reusable Launch Vehicle: Technology Development and Test Program attention has been paid to the interaction and interdependence of these technologies for SSTO/RLV. Each of the three design concepts has unique interdependencies that must be dealt with by the contractors in the iterative design integration and optimization process. It is important to recognize that creative and innovative integration of designs and vehicle shaping can provide benefits and create risks as significant as the benefits and risks of developing individual technologies. One design concept may involve higher risks but potentially greater payoffs in terms of size, cost, and ultimate commercial value. Another may be heavier or more costly to operate but create lower technical and commercial risks. The committee simply points out that these factors must be evaluated before the crucial decisions planned for late 1996 can be made.
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