Executive Summary

The objective of the National Aeronautics and Space Administration (NASA) Reusable Launch Vehicle (RLV) Program is to develop technology and demonstrations for providing reliable, low cost access to space. Phase I of the RLV program consists of concept definition and technology development leading to a Phase II subscale flight demonstration vehicle, the X-33. Shortly after the NASA Office of Space Access and Technology requested that the National Research Council (NRC) examine the RLV Phase I technology development and test program, decision criteria for this phase were developed by NASA, the Office of Management and Budget (OMB), and the Office of Science and Technology Policy (OSTP); these criteria are cited in the body of the report. The NRC committee took these criteria into consideration when making judgments about whether the Phase I program would provide adequate information to "support a decision no later than December 1996 [whether] to proceed with a subscale launch vehicle flight demonstration which would prove the concept of single-stage-to-orbit (SSTO)." However, it needs to be emphasized that the committee assessed the extent to which the technology development programs represent rational paths (and alternatives) toward RLV goals. The NRC task was limited to the Phase I propulsion and materials technology programs; the NRC was asked not to assess the feasibility of SSTO. However, the technologies required for an SSTO vehicle were considered throughout the study because the Phase I development and test programs are structured to focus on three crucial areas in the development of a cost-effective SSTO vehicle: lightweight materials for the tanks and primary structure, efficient propulsion systems, and multimission reusability and operability.

Materials considerably lighter than those currently used for the tanks and primary structure are required because reaching orbit with an SSTO vehicle (using current technologies) requires that about 90 percent of the vehicle's total mass at launch be propellant. In the propulsion area, a significant improvement in the thrust-to-weight (F/W) ratio (sea-level) of the engines is necessary—compared to the F/W ratio of the two existing large-thrust liquid oxygen/liquid hydrogen engines, the Russian RD-0120 and the U.S. space shuttle main engine (SSME).

Achieving orbit with the required payload is only part of the challenge that has been undertaken in the NASA/industry RLV program. The other, equally important challenge is to demonstrate a system that is capable of achieving a lower cost per launch



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Reusable Launch Vehicle: Technology Development and Test Program Executive Summary The objective of the National Aeronautics and Space Administration (NASA) Reusable Launch Vehicle (RLV) Program is to develop technology and demonstrations for providing reliable, low cost access to space. Phase I of the RLV program consists of concept definition and technology development leading to a Phase II subscale flight demonstration vehicle, the X-33. Shortly after the NASA Office of Space Access and Technology requested that the National Research Council (NRC) examine the RLV Phase I technology development and test program, decision criteria for this phase were developed by NASA, the Office of Management and Budget (OMB), and the Office of Science and Technology Policy (OSTP); these criteria are cited in the body of the report. The NRC committee took these criteria into consideration when making judgments about whether the Phase I program would provide adequate information to "support a decision no later than December 1996 [whether] to proceed with a subscale launch vehicle flight demonstration which would prove the concept of single-stage-to-orbit (SSTO)." However, it needs to be emphasized that the committee assessed the extent to which the technology development programs represent rational paths (and alternatives) toward RLV goals. The NRC task was limited to the Phase I propulsion and materials technology programs; the NRC was asked not to assess the feasibility of SSTO. However, the technologies required for an SSTO vehicle were considered throughout the study because the Phase I development and test programs are structured to focus on three crucial areas in the development of a cost-effective SSTO vehicle: lightweight materials for the tanks and primary structure, efficient propulsion systems, and multimission reusability and operability. Materials considerably lighter than those currently used for the tanks and primary structure are required because reaching orbit with an SSTO vehicle (using current technologies) requires that about 90 percent of the vehicle's total mass at launch be propellant. In the propulsion area, a significant improvement in the thrust-to-weight (F/W) ratio (sea-level) of the engines is necessary—compared to the F/W ratio of the two existing large-thrust liquid oxygen/liquid hydrogen engines, the Russian RD-0120 and the U.S. space shuttle main engine (SSME). Achieving orbit with the required payload is only part of the challenge that has been undertaken in the NASA/industry RLV program. The other, equally important challenge is to demonstrate a system that is capable of achieving a lower cost per launch

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Reusable Launch Vehicle: Technology Development and Test Program and be clearly competitive with other launchers worldwide. In the case of SSTO and maximum reusability, all of the components for the vehicle primary structures, the cryogenic tanks, the thermal protection system (TPS), and the propulsion system must first be developed. Then it must be demonstrated that these components are reusable with minimal inspections or replacements for at least 20 missions and have a lifetime of at least 100 missions. The committee reviewed the RLV program and found the three phase approach to the program to be sound. Phase I of the program includes demonstrations of critical technologies. These demonstrations will be required before proceeding with the more costly, largely subscale flight demonstrations of Phase II. The committee found that the Phase I development, test, and analysis programs are appropriate to support a decision about proceeding with Phase II, subject to implementation of the committee's recommendations. Three prime contractors have proposed three distinct RLV designs and are pursuing different paths in critical technology areas (in some instances a given contractor is pursuing several paths at this stage). NASA centers are providing supporting and complementary research and development in many instances; thus, if there is a failure along one path, alternative paths may be pursued. Phase II must successfully demonstrate that the technical challenges have been met before industry teams can proceed with costly, full-scale RLV development in Phase III. Using this phased approach, NASA can avoid the high development costs and technical risks of previous programs that depended on significant technological advances being concurrent with vehicle development. The committee studied the four major technology areas in Phase I of the RLV program: composite primary structures, aluminum-lithium (Al-Li) and composite cryogenic tanks, TPS, and propulsion systems. However, the committee did not address issues of design integration of component technologies into flight vehicle configurations. In any event, because of the current stage of vehicle design by industry partners and NASA, it was not feasible for the committee to make definitive assessments. The committee's recommendations reflect those aspects of the technology programs believed to require special emphasis. Other important aspects of the programs, even those involving significant challenges, were not addressed in the report if the committee believed that the participating industrial teams and NASA were not only well aware of the challenges but were also paying sufficient attention to meeting them in the program plans. The major findings and recommendations in each of the four technological areas crucial to Phase I are discussed below. COMPOSITE PRIMARY STRUCTURES The technology development program is robust, well organized, and addresses all of the major issues. There are three basic structural approaches: basic composite materials, an isogrid design for the intertank, and a sandwich structure design being developed by a NASA center. Major contractor test articles include an 8-ft-diameter by 38-inch-long DC-XA intertank; an 8-ft-diameter by 10-ft-long ground test intertank; an

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Reusable Launch Vehicle: Technology Development and Test Program 8-ft-diameter filament-wound isogrid, a one-fourth segment of a full-sized intertank (designed to address scaleability concerns); a segment of a full-scale thrust structure; and a full-scale wing box section for one of the RLV configurations. NASA centers are providing considerable analyses, material characterization, and subscale component tests, as well as an intertank/cryotank interface with a joint that is 8 ft in diameter and 6.5 ft long. Under cooperative agreements with industry, NASA also will provide structural test articles for system-level tests. Many of these test articles will be subjected to combined-load testing for life cycle; and some will undergo acoustic and damage-tolerance testing. Integrated health monitoring systems will be attached to many of the full-scale segments during testing. Efforts to validate analysis techniques and to address scaleability to single stage RLVs is progressing satisfactorily. Testing ranges from extensive coupon and other subscale tests, to panel tests, to reasonably large test articles and includes continuous validation of the necessary predictive tools at every stage. Although the approach is sound, the committee is concerned about the 15 percent maximum weight growth margin specified by the program managers; 20 to 25 percent weight growth is typical during the early stage of design development. The need to control weight growth tightly this early in the program places a premium on accurate calculation of structural performance and weight and on early verification that the structure can be built at or below the predicted weight. The committee was unable to cover fully issues such as aging, ease of assembly, and maintenance of structures because of the time constraints. However, the committee considers these issues to be very important over the long-term. "Scaleability" refers to scaling to larger or smaller sizes the physical attributes of a given test article according to scaling laws. If the laws are not known, an iterative process must be used; that is, the predictions based on scaling models must be checked against actual test results at each scaling step. The discrepancies between the model predictions and actual data are used to improve the model for the next step. In most realistic situations, the scaling laws are not known exactly; therefore, extensive testing is required to provide sufficient data to build confidence in the model. It is also important to note that various physical (and chemical) phenomena that directly affect the RLV design and performance, scale in significantly different ways with changes in geometric size (e.g., structural strength and stability under load and thermal stress versus aerodynamic heating rates versus heat conduction through solids—all interacting in the design of launch vehicle structures). Also, with changes in size or other parameters affecting loading, dynamics, or configuration, failure phenomena may be encountered that have not occurred in previous situations.

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Reusable Launch Vehicle: Technology Development and Test Program MAJOR RECOMMENDATIONS REGARDING PRIMARY STRUCTURES Test articles of each size must be designed, built, and tested to RLV-scaled conditions using design codes that are being validated. Furthermore, all of the joints and fittings for the larger test articles should be properly scaled to the RLV flight configuration. This may require full-scale testing of some joints. The planned combined-loads tests, which simulate the appropriate thermal and acoustic environments integrated with flight vehicle interfaces (e.g., TPS on the cryogenic tank or the intertank), should be conducted with as many cycles as possible. Many health monitoring systems and nondestructive evaluation (NDE) techniques were mentioned in the briefings, but there does not seem to be a well-ordered program to identify which measurements will be made and where or how NDE will be used to penetrate multiple layers of material. The committee is aware of the extreme difficulty of this task and strongly recommends the development of a clear plan for certifying readiness for launch of flight-critical hardware. Weight requirements (not only the 4lb/sq ft given in the decision criteria) must be defined for each vehicle concept. REUSABLE CRYOGENIC TANKS Another key component of the RLV program is the development of reusable cryogenic tanks. Both Al-Li and organic-matrix composite tanks are under development by NASA and industry partners. The development programs are generally robust in that most critical areas are addressed by more than one approach. Both the Al-Li cryogenic tanks and the organic-matrix composite tanks are discussed below. Al-Li Cryogenic Tanks Two alloys with different properties are being investigated for use in the cryogenic tanks. In addition, three fabrication techniques—net shape extrusion, net shape spin forming, and net roll forging—are being considered, and two welding techniques—variable polarity plasma arc and friction-stir welding—are being developed and tested. To address scaleability issues, the RLV program will fabricate and test both 8-ft and 14-ft diameter tanks; data from the space shuttle super lightweight tank, with a 28-ft-diameter tank, are being added to the RLV Al-Li database. Several groups are conducting material properties characterization tests, including tests to assess reusability (e.g., fatigue and crack growth rate) and liquid oxygen (LOX) compatibility for each alloy.

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Reusable Launch Vehicle: Technology Development and Test Program Several areas of concern affect producibility, operability, and reusability of the tanks. These include assuring proper microstructure and texture for all product forms and addressing the issues of weldability and weld repair, lot-to-lot variations, and anisotrophy in Al-Li alloys. NASA and the industry partners are aware of most of these concerns; they are identified here because they are crucial to the success of the program. Major Recommendations Regarding Al-Li Cryogenic Tanks The committee's recommendations for Al-Li cryogenic tanks are as follows: Recent microstructure and texture analyses have shown that current processing methods produce excellent products. However, each casting and product form must be characterized extensively to assure that the required microstructure properties are obtained. Other product forms for which such characterization is required include all welded, weld repaired, and extruded near-net-shape formed products. Weldability and weld repairs are major issues of concern. Although the 2195 alloy can be welded, repair or second-pass welding is a major problem. Marshall Space Flight Center (MSFC) has been experimenting with an aluminum-silicon (Al-Si) filler material. The committee recommends caution in the application of this material because it forms an AlLiSi phase that attracts and absorbs moisture, which introduces the possibility of stress-corrosion cracking. Tests for stress-corrosion in the 2195 weld zones should be rigorous. Because of the limited database on Al-Li alloy 1460 and the possible lot variations in Al-Li alloys, extensive testing is needed on small samples of all product forms. The fatigue, crack growth, and stress-corrosion behavior of welds and weld repairs should be determined. Because Al-Li alloys have been shown to be more anisotropic than conventional aluminum alloys, texture, strengths, and elastic moduli should be characterized at various orientations. Weight predictions (other than 0.7 lb/ft3 for an oxidizer tank or 0.5 lb/ft3 for a hydrogen tank, as given in the decision criteria) must be clarified for each vehicle concept. Achievement of these predictions must be verified using properly designed and scaled articles. Organic-Matrix Composite Cryogenic Tanks There are three approaches to development of organic-matrix composite cryogenic tanks. Carbon cloth layups impregnated with epoxy formed to shape and subsequently cured in an autoclave (an oven capable of raising the pressure to desired levels) have traditionally been used for the heat shield on reentry vehicles and similar applications.

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Reusable Launch Vehicle: Technology Development and Test Program A second approach is to use winding machines to apply graphite filaments coated with epoxy to a mandrel in the desired shape; this is followed by curing in an autoclave. The third approach is to use a sandwich construction of honeycomb or foam core between sheets of graphite epoxy. At least two sizable tanks will be fabricated using each method: one 8-ft-diameter by 16-ft-long tank; and one 8-ft-diameter by 9-ft-long tank. Fabrication of a third, slightly larger tank is under review. Several organizations will conduct material properties characterization and subscale tank/bottle and panel tests to address the issues of basic weight, strength, and reusability. The committee has two concerns about these tests. Although the sizes selected for the test tanks are reasonable, producing full-scale tanks five times larger than the test article while maintaining the required material properties may be difficult. Second, if autoclaving is necessary, it is unclear that there will be an autoclave large enough to accommodate the full-scale tanks or primary structures; and the cost and time for building one must be evaluated. There are multiple approaches to evaluating the critical issue of joining the tank to the intertank structure, both in design and tests. These evaluations may reduce the risks in this important area. Major Recommendations Regarding Organic-Matrix Composite Cryogenic Tanks The committee's recommendations on organic-matrix composite cryogenic tanks are as follows: A detailed plan addressing producibility of full-scale organic-matrix composite tanks should be developed, and the advisability of demonstrating fabrication techniques should be evaluated. The necessity for autoclaving must be evaluated and the availability of large-capacity autoclaves should be resolved as soon as possible. Thermal/load-cycle testing should be conducted on all tanks 8 ft in diameter (or larger) that have cryogenic insulation interfaces with neighboring components and TPS affixed to demonstrate that the integrated system will satisfy reusability requirements. These tests will provide a database for the RLV comparable to the database for the tank to be tested and flown on the DC-XA. Weight predictions (aside from the ones specified in the decision criteria) must be clarified for each vehicle concept, and satisfaction of predicted weight requirements must be verified using properly designed and scaled test articles. THERMAL PROTECTION SYSTEMS The current well-balanced program for developing advanced thermal protection materials addresses the key issues of significantly improved operability and reusability,

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Reusable Launch Vehicle: Technology Development and Test Program without exceeding the weight requirements allotted to that system. Two NASA centers have proceeded along two distinct but complementary development tracks. One approach takes advantage of the long heritage of the Shuttle TPS with significant improvements in the robustness of reusable blankets and ceramic tiles; the second pursues the use of metallic panels to improve robustness. A third concept, the use of ceramic-matrix composites, is being developed for application to the highest-temperature areas of the vehicle during reentry (i.e., the nose and leading edges of the wings and control surfaces). Each of these approaches raises some concerns, and the program appears to be addressing them. Producibility does not appear to be a major issue for new TPSs. But there are important concerns about the resistance of the tiles (refractory or metallic) to particle impact at liftoff and landing and, especially, in orbit at Space Station altitude, where it is predicted that penetration of a tank may occur at least once in a 100-mission cycle. It is clear that both the Shuttle-improved and metallic TPS are more resistant than the earlier Shuttle TPS, but the performance of both systems has yet to be fully quantified for various operational conditions. Tests for this type of resistance to damage are in progress. A major workshop was conducted at Ames Research Center (ARC) to define experimental programs for evaluating environmental and vibroacoustic effects on the TPS. Environmental effects being evaluated include rain/particle erosion, lightning, and pad ice/frost. A valuable, comprehensive "robustness test matrix" was devised by the community of experts. Other questions, such as the levels of rain that can be absorbed safely and issues related to waterproofing, still need to be resolved. The time and effort required after each Shuttle flight to return the vehicle to service are unacceptable in terms of turnaround time for an RLV. Neither NASA nor the industry partners has successfully developed either agents or coatings that provide permanent waterproofing or rapid techniques of applying waterproofing to the TPS materials. It remains to be seen if existing technology and ongoing research can solve the problem of waterproofing. Significant development in methods of attaching the TPS to the tanks, insulation, and structure is needed for both of the generic TPS systems. This is an important problem area. Concepts that satisfy the requirements for structural integrity in flight and, at the same time are easily replaceable, will require innovative development and testing. The Shuttle-improved TPS system includes adhesive bonding, which is safe for flight but difficult to replace. The metallic panels depend on mechanical attachments that are still under development. A permissible heat leakage rate into the propellants has been specified for the space shuttle external tank. However, a corresponding rate has not been determined for an X-33 or RLV. Neither the sensitivity of propulsion efficiency to propellant temperature nor the resulting permissible heat leakage rate is known for subcooled propellants that have not been previously used in operational vehicles.

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Reusable Launch Vehicle: Technology Development and Test Program Major Recommendations Regarding TPS The committee's recommendations on TPS are as follows: NASA should evaluate the probability that particles in space will penetrate not only the TPS but also the propellant tanks during a 100-mission life cycle. NASA also should assess the impact of penetration on the reentry survivability of the RLV. The "robustness test matrix" evaluations should be carried out as soon as possible, with early emphasis on determining hypervelocity impacts and the resistance of new TPS candidate materials to environments known to cause the most problems for the Shuttle orbiter. Activities related to metallic panel attachments should be enhanced, and more-operable attachment mechanisms for the Shuttle-improved TPS should be investigated to assure easy replacement. Metallic and ceramic-matrix composite standoff panels should be tested in arc jets to demonstrate that there is no overheating at the attachment points. Methods of waterproofing need to be pursued vigorously if reasonable ground-processing times are to be achieved. Permissible heat leakage rates into liquid hydrogen (LH2) and LOX propellants should be established for normal and subcooled propellants. PROPULSION SYSTEMS The prime contractors have indicated a requirement for an engine sea-level F/W ratio greater than 75 for the RLV. The SSME Block II and RD-0120 ratios are 51 and 43, respectively, with a projection that SSME Block II+ (with a short nozzle) may achieve an F/W ratio of 58. Shortening the nozzles of the SSME or RD-0120 engines will increase sea-level F/W performance. Therefore, an increase of 30 percent or more will be required, which presents developers with an extremely difficult challenge. Methods of achieving this increase have been identified by the contractors, and, although in the opinion of the committee achieving a F/W ratio greater than 75 will be very difficult, it is by no means impossible. In addition to developing the X-33 engine in Phase II, an engine development and ground test program is planned that will lead directly to the engine technology for an RLV. However, these plans are not well defined. Because the characteristics of the X-33 engine are only partially scaleable to the RLV, it is through engine development and testing on the ground that the scaleability to the RLV will be demonstrated. Efforts to significantly reduce engine turnaround time after each flight have not yet achieved the objective of a rocket engine that can be handled much the way an operational jet engine is handled.

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Reusable Launch Vehicle: Technology Development and Test Program Major Recommendations Regarding Propulsion Systems The committee's major recommendations on propulsion systems are as follows: RLV engine sea-level F/W requirements to achieve SSTO should be revalidated by the prime contractors and independently using NASA's vehicle design/performance groups. Current goals of greater than 75 F/W will be difficult to achieve in the SSME or RD-0120 derived engines, as well as in new engines, without raising concerns about the structural margins required to satisfy reusability goals. If the requirement of high sea-level F/W is re-validated, the committee recommends that development of the selected RLV engine be initiated at the beginning of Phase II and vigorously pursued. Because the X-33 vehicle engine will make only reasonably small contributions to the F/W goal, the development program will be the major source of data for a decision about proceeding with Phase III. Concurrent trade studies should be conducted to assess whether larger, but viable, vehicles will alleviate the F/W requirement. The decision criteria for proceeding from Phase II to Phase III for the propulsion system should reflect the required RLV engine performance targets (such as a sea-level F/W of greater than 75 and vacuum Isp of 440 or higher). NASA should evaluate the contractor's 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 an independent group of propulsion experts. The ground RLV engine 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 a short (truncated) nozzle should begin soon. The X-33 and RLV Aerospike engine configuration details of combustor body, throat shape, nozzle shape, expansion ratio, and vehicle integration should be completed before the Phase II decision date. The throttling and thrust vector methods proposed, including interaction effects between adjacent engines, should be evaluated. More robust and reliable health monitoring instrumentation than is currently used should be developed and thoroughly tested. The overall approach to health monitoring and assuring flight readiness of the propulsion system within turnaround goals should be defined. NASA should evaluate the program and engine changes required to meet the rapid turnaround goals. In general, operability and engine reliability

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Reusable Launch Vehicle: Technology Development and Test Program requirements should be developed for X-33 and RLV. The fact that the RLV engine will not be subjected to major inspection or maintenance between each flight unless problems are indicated by on-board health monitoring or visual inspection should be considered in the design. GENERAL OBSERVATIONS Although NASA and its industry partners have adopted reasonable approaches to advance the state of the art in both space materials and propulsion during Phase I, formidable challenges remain. The committee did not address issues related to time or money constraints. In addition, because of time constraints on the study, the committee could not review several important areas: 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