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EXECUTIVE SUMMARY The Committee on Hypersonic Technology for Military Application of the Air Force Studies Board was formed to: · Evaluate the potential military applications of hypersonic aircraft. Assess the status of technologies critical to the feasibility of such vehicles. . MILITARY APPLICATIONS Early in our proceedings, we found that firm military operational concepts do not exist for applications of hyper- sonic aircraft. Determination of oper- ational requirements must await a better understanding of critical technologies. Thus, the focus of this report is an evaluation of the status of these tech- nologies. From this evaluation we concluded that: Hypersonic aircraft technology and ramJet/scramjet propulsion offer potentially large increases in speed, altitude, and range of military aircraft, and may enable or extend important Air Force missions. The simplest (and probably most feasible) hypersonic vehicle would cruise in the range below Mach number 8. The most attractive missions involve flight to orbital or near-orbital speed above the sensible atmosphere. These missions offer flexible recall, en r cute redirection, and return to base. Any potential military advantage in the speed range between Mach num- ber ~ and orbital velocity would be negated by technical difficulties in the areas of surface heating and thrust, and weapons carriage, aim- ing, and release. The global or near-global range 1 EXECUTIVE SUMMARY . coupled with the cryogenic fuels of hypersonic aircraft will require unusual base support requirements that must be considered in any judgment of their operational utility. NATIONAL AEROSPACE PLANE PROGRAM Part of our task is to advise the Commander of AFSC on the: · Research and development strategy of the National Aerospace Plane (NASP) research vehicle designed for single-stage- to-orbit (SSTO) capa- bility. Research vehicle approach for the aerospace plane that will maximize the acquisition of knowledge in the technical areas most critical for hypersonic vehicles. The aerospace plane program is in the technology development phase, and decisions to build an SSTO vehicle will not be made before 1990. Due to the technical uncertainties, we agree with this approach. In particular, substantial technology demonstrations are needed, especially in high temperature materials and structural concepts appropriate to them, before a commitment to build an SSTO research vehicle can be technically justified. We recommend that the NASP pro- gram office retain the ultimate goal of demonstrating the technical feasibility of SSTO capability, but maintain an option of selecting less than SSTO as part of a prudent risk management strategy. Con- sideration should be given to incorporat- ing design features into this research vehicle that allow such modifications as may be found necessary from flight experience.
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2 The research vehicle should also have auxiliary rocket propulsion to enable controlled flight with some inde- pendence from the air-breathing propul- sion system. The auxiliary propulsion will enable greater exploration of the vehicle's flight characteristics and will help ensure flight safety. The drawback is that this auxiliary propulsion system will make the research vehicle heavier, thus making it more difficult to achieve the structural weight fraction required for SSTO capability. Further, the program office might design a series of flight research vehicles, each one able to access an increment of the total flight corridor up to orbital conditions. TECHNOLOGY ASSESSMENT The following is a summary of the status of the major technical areas critical to hypersonic aircraft. Tech- nical reviews such as this one tend to focus on deficiencies and problems. Because our purpose is to help the Air Force support solid approaches and correct deficiencies in its program, this report will follow that pattern. The committee wishes to be on record, how- ever, as strongly favoring a vigorous, vehicle-focused hypersonic technology program, carried on at least at the level of the planned aerospace plane program. Propulsion-Airframe Integration The success of any hypersonic vehicle will depend as much, or perhaps more, on the integration of the various contributing technologies into a complete system, as on the individual technol- ogies. Such integration is quite mission- specific. Integration of the propulsion system and airframe is basic to deter- mining the aerodynamic shape of a hypersonic vehicle. How such integra- tion is achieved becomes increasingly important as the maximum air-breatliing Mach number rises toward orbital values. HYPERSONIC TECHNOLOGY FOR MILITARY APPLICATION Efficient operation at very high Mach number requires configurations that pose serious integration problems at other Mach numbers. Consequently, the low speed propulsion system must be inte- grated with the hypersonic propulsion system and overall aerodynamics in a way that does not degrade performance prohibitively over the entire vehicle speed range. Finally, the amount of hydrogen required for engine cooling exerts an unusually high leverage on the airplane size. Under the most severe conditions, the molar flow rate of hydrogen coolant within the airframe and engine structure is more than double the molar flow rate of air through the engine. Conse- quently, the cooling system must be effectively integrated with airframe and engine system, and coolant plumbing and pumping losses must be minimized. To help solve these problems in the NASP research vehicle program, we recommend that engine-airframe integra- tion receive emphasis in the design definition by teams drawn from both engine and airframe contractors. Propulsion Systems The key technology for hypersonic air-breathing propulsion at higher flight Mach numbers is the supersonic combus- tion ramjet (scramjet). In this device, injection of hydrogen fuel and its mixing with air is the most influential factor affecting engine performance. Current techniques for modeling related phen- omena are inadequate. Conventional one-dimensional or quasi one-dimensional computation of the reacting flow in the combustor is insufficient and often mis- leading. Further, the stability of the scramjet flow with hydrogen reaction is not understood. Flow instability poses the possibility of developing strong shock waves and catastrophic loss of the engine.
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EXECUTIVE SUMMARY For the scramjet to operate at peak performance throughout its Mach number range, extensive geometric changes must be made while maintaining minimum internal aerodynamic losses. These changes place additional demands on the design of coolant passages and on cool- ant management. Also, transitions between the various operating modes of the propulsion system present severe design and control problems if extreme loads on the engine and airframe are to be avoided. Since experimental verifica- tion of most transitional modes cannot be accomplished in ground-based facil- ities, the complete engine will undergo much of its high Mach number develop- ment during the flight program. Thus, it will be necessary to incorporate auxiliary rocket propulsion in the research vehicle with fuel supply and controls separate from the scramjet, to be able to vary the engine operating point at fixed flight conditions. Additionally, some rocket propulsion must be incorporated in the propulsion system of vehicles with orbital capabil- ity, for final insertion into orbit and for the de-orbit maneuver. Aerodynamics The aerodynamic problems of hyper- sonic flight can be considered in two categories. Below Mach number 10, the problems are primarily of a fluid mech- anical nature, where one must accurately determine the pressure distribution, skin friction, heat transfer, flow field details, and mixing. Above Mach number 10, the same fluid mechanics problems remain, with the additional complication of the rate kinetics of real gases, the special low density phenomena of high altitude flight, and the effect of small bluntness on slender bodies. It is not yet possible to simulate or compute with any degree of certainty, these phenomena over the entire flight range. 3 There is a critical need for aero- dynamic data, both for input to compu- tational fluid dynamics (CFD) studies, for CFD validation and for engineering design. At low hypersonic speeds no current facility has a high enough Mach number (S to 10) with an adequate Rey- nolds number and size for the necessary aerodynamic and configuration studies. At high hypersonic speeds - above Mach number 10 - it will be necessary to simulate the aerodynamics (viscous effects and flow field), while including real gas effects at conditions ranging from the continuum low altitude regime to the free molecular regime at high altitude. No current facility or facilities can simulate the full range of flow parameters with the needed flow quality. Since complete simulation in ground- based facilities cannot be carried out, it will be necessary to check individual elements of CFD codes. Thus, there will be a need for flight tests for full valid- ation of CFD codes particularly at high Mach numbers. To provide essential data for hyper- sonic vehicle designs, we recommend: · The immediate consideration of a quiet tunnel in the Mach number 10 range and of a size and Reynolds number capability to permit testing of hypersonic configurations at close to full-scale conditions. The Gas Dynamics Facility at the Arnold Engineering Development Center' for example, might be so modified. A combined CFD/experimental pro- gram focused on key hypersonic aerodynamic elements. That a high priority be given to a national integrated program to determine reaction rates at high temperatures as a needed input into CFD computations. That special attention be paid to the problems of low density flows be- cause of the special phenomena associated with high altitude, high Mach number flight. We also sug
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4 HYPERSONIC TECHNOLOGY FOR MILITARY APPLICATION gest that the personnel of the NASA Ames Research Center could provide a key nucleus for such an effort. Control, Guidance, and Information Systems Hypersonic vehicles present signif- icant challenges in -control of the vehicle trajectory and also in the con- trol of vehicle configuration, aero- dynamics, 'dynamics, ' propulsion, and cooling with 'particular emphasis on the need for their complete integration. These demands in turn present difficult challenges in the associated fields of instrumentation and information systems. The net result is that information and control technology joins materials, pro- pulsion, and aerodynamics as an enabling technology. Since the lower surfaces of the pro- jected hypersonic aircraft are part of the engines, angle of attack and sideslip will significantly affect engine perfor- mance. Precise control of the entire integrated vehicle, perhaps as accurately as to within 0.1 degree, is therefore necessary. ' Further, exacting speed control may be needed for engines transition and aerodynamic heatir~g that, in conjunction with the other demands on the control system, may prove dif- ficult to achieve. In the presence of maneuvers and thrust asymmetries, speed control presents a challenge in' controls technology far beyond that posed by any aircraft yet designed.' Offsetting this very severe challenge is the potential offered by advanced control' technology. But to realize these potentially signif- icant returns, it will be important to systematically identify those areas where such benefits are possible and incor- porate them into the designs. Both the performance and safety of the vehicle will depend fully on onboard processing. Projections for throughput requirements for the onboard information system range up to 25-50 million instructions per second. This high level of throughput will cause reliability reuqirements to be in the range of 10 million to one billion hours between failures. This level of reliability and throughput has not yet been achieved in a flight qualified installation. Other agencies such as NASA and DARPA have sponsored high reliability, high through- put laboratory level programs that may help assure the availability of an adequate onboard processing system. The thermal environment, even inside a hypersonic vehicle, is likely to be quite hostile to electronic, sensor, and hydraulic systems. Active cooling undoubtedly will be necessary. Advances in high temperature electronics, cabling, and hydraulics also will be highly desir- able to offer more design flexibility and reduce the weight requirements for active cooling. In addition, the combin- ation of speed, low density, and high temperature impose conditions for sensors beyond anything required to date. The problem becomes even more complex when the need for redundancy is added. A program to identify and develop the necessary sensors is imperative. A control system designed to accom- modate all the uncertainties that can be anticipated at this stage of the design would have to be very-robust and com- plex. Considerable savings in controls expenditures are likely if a phased flight development approach is adopted to reduce uncertainty as the flight envelope is expanded from subsonic to hypersonic speeds. To make this possible the infor- mation processing system installed in the prototype aircraft must be exceptionally reliable and be able to expand gracefully to the necessary size and throughpu t required for the final configuration, without need for redesign or recon ~. . ~ ~gurat~on.
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EXECUTIVE SUMMARY Materials, Structures, and Cooling The system challenge posed by a hypersonic vehicle to the materials and structural community is reflected in the structural weight fraction, which is the weight of all the structure divided by the take-off gross weight (TOGW). With presently envisioned air-breathing propulsion systems using cryogenic fuels, a fuel fraction of approximately 0.75 is required for SSTO performance. Design studies show that the consequent struc- tural weight fraction of an SSTO vehicle must be approximately 0. 15. Present military aircraft that operate mainly in the slightly supersonic regime have structural weight fractions of approx- imately 0.30. An increase in either the fuel fraction or the structural weight fraction means either a multi-fold increase in TOGW (if the payload is not to decrease) or a vanishingly small payload (if TOGW is not to increase). Thus, from the viewpoint of structural weight fraction, we consider the SSTO objective very demanding. It can be met only with new materials and new structural concepts. The structures will face severe design conditions arising from aerody- namic heating. The aircraft will require new materials with elevated temperature properties comparable to those of the high strength aluminum alloys at room temperature. A necessary consort to the "strength" materials will be coatings to prevent oxidation at the elevated tem- peratures at which they will have to operate. The aircraft will also require materials for insulation and for both passive and active cooling. Experimental material cannot be readily used in a flight test vehicle. The scale-up of a material system that has been proven in the laboratory to the production quantities demanded by even a research vehicle requires the develop- ment of a materials fabrication technol- ogy. Also, the properties of these new 5 materials are very process-dependent, so that extensive processing experience must be accumulated before they can be reliably employed. Even after the materials are available with nominally adequate properties, the structural designer must confront the effects of the high temperatures, including: stresses induced by the differential expansion of the structure (the so-called thermal stresses), degradation of the material properties, and distortion of the shape of the structure, which will lead to interactions with the flight loads. Designing for thermal stresses requires temperatures to be predicted with great precision. Also required are an accurate knowledge of the thermal conductivity, thermal coefficient of expansion, heat capacity as a function of temperature, and the thermal impedances of joints and connections. Design considerations are further complicated by the additional stresses caused by loads due to maneu- vers, gust, and landing. In sum, vehicles with orbital or near-orbital capability in a single stage will require new materials and novel structural concepts to meet the struc- tural integrity requirements of strength, rigidity, longevity, damage tolerance, and efficiency; all of these at a cost com- mensurate with the importance of the mission. An aggressive and sustained materials and structures program will be required to develop these technologies to meet the presently envisioned objectives of the aerospace plane. Considerable progress toward such a program has been made in the last year, but we still regard materials and structures as a probable limiting factor in meeting the objective of SSTO performance. The Role of Computational Fluid Dynamics Computational fluid dynamics has become a principal tool for aerodynamic and propulsive-flow design of hypersonic
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6 HYPERSONIC TECHNOLOGY FOR MILITARY APPLICATION vehicles. Computers can simulate simul- taneously the hypersonic flight para- meters of velocity, free stream density, physical scales, and free stream thermo- chemical state, while the present ground-based test facilities cannot. The role of CFD is especially significant for high altitude hypersonic flight conditions involving reaction-rate dependent air chemistry. Current supercomputers and numeri- cal methods are able to simulate three dimensional flows using the Reynolds- averaged (time-averaged) Navier-Stokes equations. However, the lack of capa- bility for modeling of turbulent stresses, heat flux, and transition locations, form the principal current limitations of CFD techniques. Although the validation of three-dimensional hypersonic CFD codes is now relatively limited, we expect this situation to improve considerably in the near future as various results from suet codes and hypersonic experiments become available. The most intense local heating rates on vehicles such as the projected NASP research vehicle are expected to be on cowl lips, caused by shock-on-shock heating. Present CFD simulations of this phenomenon agree well with experi- ments for the lower portion of the hypersonic flight corridor where shock wave thickness is negligible. However, neither computations nor experiments have yet been made for these high altitude hypersonic flight conditions. The limitations of current CFD simulations are not inherent to CFD itself, but are due largely to the present state of supercomputer development. Advanced CFD technology will require more powerful supercomputers than those expected to be available in the near future. Our recommendations are that: · an effort be made to significantly improve the modeling of turbulent . mixing flows relevant to hypersonic flight and of turbulent boundary layer flows in the presence of heat transfer and pressure gradient. high-altitude shock-on-shock cowl lip heating be investigated by direct simulation Monte Carlo methods, appropriate new experiments, and continuum equations more advanced than Navier-Stokes. the direct numerical simulation of turbulence in compressible flows should be an integral part of the long range technology development program of the Air Force. Test Requirements Most U.S. hypersonic facilities were built when the major national effort was on ballistic missiles and space vehicles. A careful examination is required of their utility for the design and testing of the new generation of hypersonic vehicles. The nation must consider not only current, but also long term needs for hypersonic testing. Since ten years or more are required to design, build, and commission new facilities, it must be accepted that only currently available facilities will con- tribute to the testing needs of the NASP program during the technology develop- ment and design phases. It is very important, therefore, to ensure that all existing facilities are used as effectively as possible. We should note that many of these are 20 to 40 years old and flow conditions in them are not sufficiently well known. Research must be done in such facilities, to identify their flow characteristics. Components of hypersonic vehicles will have to be tested at unpreced- entedly high temperatures, for which existing but inactive facilities such as those at Plum Brook, Ohio, and the NASA Dryden Test Facility should be reactivated. Serious consideration should
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EXECUTIVE SUMMARY also be given to the construction of facilities for component testing at temperatures above the 1200° C limit of these existing facilities. We do not know the local composi- tion, temperature, and turbulence of the upper atmosphere well enough to infer high altitude hypersonic flight conditions from ground test data. Measurements and analyses should be made of these characteristics of the atmosphere to provide a firmer base for use of test data in vehicle design and in flight test planning. It will not be possible to test hypersonic vehicles with orbital or near- orbital capability over the full range of 7 their flight conditions prior to flight. Thus, hypersonic aircraft research above Mach number ~ will require a philosophy of flight research somewhat more akin to that practiced for missiles and launch vehicles. A clear recognition of this will be essential to success of the program. The NASP flight research program will collect more basic information on aerodynamics, propulsion, control, and structural behavior than any previous flight research program. We recommend that very careful consideration be given to the flight test scenario during the design of the research vehicle, not afterward.
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