Below is the uncorrected machine-read text of this chapter, intended to provide our own search engines and external engines with highly rich, chapter-representative searchable text of each book. Because it is UNCORRECTED material, please consider the following text as a useful but insufficient proxy for the authoritative book pages.
ED 1 lo - x of x a o o ·~ oh Hi of ~ ~ Go en of To ~ cat lo lo lo lo lo lo lo lo lo ~ Lo ~ lo Hi Go ~ ~ lo _1 LO ~) ~ 1 - ~ o ~ Hi Hi 1 1 Lo, · ~ ~ Hi a: 1 1 L~) o ~c ~ n, Q ~< ~) J J 3 3 ct: cc 3 ct ~z Z ' ~LL C ~=: ~=: Z Z 0 ~oO a (V S~ 3 O O ·_ ·_ 1 1 00 ·_ ~ O O ~ a, · ~ Q ~ ~) =: I ~ a C~ ~n ~_ a_ - r~ LL C CO ~: ~5 cn ~ ·_·_ 1 LO ~- ~_ E o S~ ~: ~3 ~_
STATUS OF HYPERSONIC TECHNOLOGIES 53 structural reasons, such schemes as MHD acceleration of a hypersonic stream generated by a wind tunnel, are being investigated as a way to achieve high Mach numbers. The committee has reviewed one such proposal for steady state crossed-field acceleration of air to 25,000 feet per second, and we find the proposal does not reflect understanding of the large body of knowledge devel- oped by the ~~~ I, last 20 years. The analysis on which the proposal is based is limited to "one- dimensional" or channel flow, without consideration of wall effects. These effects limit the feasibility of such devices and have received a great deal of study. The overall conclusion of these detailed studies is that the steady crossed-field accelerator is an ineffec- tive device for producing large gas velocities, because too large a fraction of the input energy goes into heating the walls, as well as the gas to be accelerated. MHD community Over the 2.~.3 Materials and Structures Test Facilities The Air Force recently conducted a study of the high temperature test tech- nology needed for hypersonic vehicle applications. They found that heating capability above 1400° C. will be dif- ficult to achieve and that instrumen- tation is not available for use above 800° C. Also, high temperature strain gauges are not available for temper- atures over 800° C. When testing must combine flowing air, and mechanical and thermal cycling to obtain the necessary data for structural design, one must conclude that additional testing capabil- ities are needed to get the data in an expeditious manner. 2.8.3.1 Structures Testing Hypersonic vehicles will require structures that are ultra-light, temper ature resistant, inspectable, durable, and safe. The structural design concepts depend on vehicle configurations chosen for the flight profiles that will be flown, and from the material choices available. Structural optimization can be done once all the loads are known, the stresses determined, heat transfer known, the aeroelastic behavior of the vehicle determined, and once the strength, stiff- ness, and fracture toughness of the materials selected are known in consid- erable detail for all conditions of vehicle operation. The design criteria for a hypersonic vehicle will determine the amount and type of structural development testing required. This should include the speci- fication of time, temperature, load synthesis, cumulative creep criteria, oxidations criteria which can then influence coating criteria, fracture mechanics and fatigue, sonic fatigue and panel flutter. A typical structural component is shown in Figure 2-12. 2.X.3.2 Facilities The Air Force study mentioned above indicated that there are existing facilities adaptable for major component testing. However, a major structural component test facility was estimated to cost $90 million dollars. A full-scale test facility that would be required for structural certification was estimated to cost $462 million dollars. In the National Space Transportation and Support Study 1995-2010, prepared by the loins DOD/NASA Transportation Technology Team, total structures/- materials funding included facilities. The facilities funding totaled $554 million dollars that included a structural certification facility. In fact, these figures add up to about the same amount as estimated by the Air Force for hypersonic vehicle structural testing and certification.
~4 HYPERSONIC TECHNOLOGY FOR MILITARY APPLICATION There has been some discussion of activating the NASA Plumbrook high temperature test facility in Sandusky, Ohio, and the NASA Dryden facility at Edwards Air Force Base, but to date there seems to be no funding to activate these facilities. In 1985, the Aerospace Industries Associations Aerospace Technical Council established a High Temperature Test Facility (HTTF) Col- laborative R&D Ad Hoc Group to deter- mine whether the Members of A.I.A. should form a partnership to develop such an HTTF. In February of 1987, the Ad Hoc Group, after visiting the NASP program office, reported that DOD and NASA were doing a "good job assessing and developing the necessary test facil- ities" that would cover most of the needs identified by the HTTF ad hoc group and that the group should be dis- banded with no further action. As of March 198S, there seems to be no posi- tive action with funding to proceed to define these facility needs. The major test facility, now being refurbished to carry out testing in aerothermal loads and high temperature structures, is the Langley S-Foot high temperature tunnel, a Mach number 7 blowdown type of facility in which methane is burned in air under pressure and the resulting combustion products are used as the test medium with a maximum stagnation tem- perature near 3800° R to reach the required energy level of flight simula- tion. This facility will, however, not be ready for testing until the late fall of 1988. There is an urgent requirement for the development of major high temperature materials and structural component test facilities. The develop- ment of such facilities is mandatory to insure that an adequate data base of material properties is developed and that structural design concepts can be evaluated to support hypersonic vehicle design. In the end, flight experiments may be necessary as an adjunct to evaluate the structural concepts being considered because of the inability to adequately simulate the combined environments of temperature and flow. Experimental data can be obtained in several ways. Rocket or free-flight tests lack adequate communication between the flying vehicle and the test engineer. Only a small amount of data are obtained from each flight, and each flight represents a very large expendi- ture of money and engineering time. To get a maximum return from flight tests, simulated flight tests first should be performed in the laboratory. 2.X.3.3 High-Speed Wind Tunnel Tests To obtain high temperatures in a wind tunnel the air must be heated to the desired temperatures, and this poses new problems in wind tunnel design. Such a venture entails its own host of difficulties and results undoubtedly will not be forthcoming for some time. 2.~.3.4 Laboratory Tests with Heating Devices Laboratory heating devices that produce thermal energy can be devel- oped. Such devices should be inves- tigated, and the possibilities are many. Radiant devices, such as those in ordinary household cooking ranges or in refractory ovens, can be arranged in a dense pattern over a broad area to obtain a distributed source of high radiant energy. The transient heating phase could be controlled by changing the distance between model and heater. It could also be done by a system of shutters on the heater. Chordwise and spanwise variation of temperatures could be obtained by painting or otherwise preparing the exposed surface of the model to achieve different adsorptions. The design of the heater could also be arranged so that its thermal output varies across its face. But the scarcity
- ~ AN k ~ \ m to0~// /~//A:\ Lot ~ 3 V o Hi/ S\\\ \ ~?: ~ cn o m ~- / N/~/: / I: o E o ~ C_) Cot ~ _. 1 ~
STATUS OF HYPERSONIC TECHNOLOGIES 55 . . . Of experimental data and the complexity of the problem would indicate that much can be learned and perhaps should be first learned from experiments on simple models using simple experimental appara- tus. 1. Test section 1 ft. in diameter or more. 2. In contrast, there is evidence that the Soviet Union continued to build hypersonic facilities through the 1970s and 1980s. 3. Chord length defined as the square root of the test section area.
56 Below are the principal findings we have adduced from our review of hyper- sonic technology for military application, and the recommendations we offer to further these technologies and their applications. 3.1 Potential Military Hypersonic Applications ( 1 ~ Hypersonic aircraft technology, in association with air-breathing propulsion, offers potentially large increases in speed, height, and range of military aircraft, and may enable or extend important Air _ · . force missions. (2) Operational hypersonic aircraft will necessarily have very large turning radii, and useful missions therefore require global or near-global range. (3) Cryogenic fuels are necessary, and any studies of hypersonic aircraft missions should therefore include a careful examination of the base support requirements which they imply. (4) The simplest class of hypersonic cruise vehicle would fly up to Mach number 8. This class can signif- icantly advance the reconnaissance and strike missions now done by the SR-71. (5) The most attractive potential Air Force missions involve flight to orbital or near-orbital speeds above the sensible atmosphere. In con- trast to ballistic missiles and satellites, these offer flexible recall, en route redirection, and return to base. HYPERSONIC TECHNOLOGY FOR MILITARY APPLICATION 3.0 FINDINGS AND RECOMMENDATIONS (6) Sustained hypersonic flight in the atmosphere between the two ex- tremes of (4) and (5) above pre- sents major technical difficulties. Problems of surface heating, thrust, vehicle stability and control, infrared signature, aiming, and weapon release could make any potential military advantage in this speed range unlikely. 3.2 Propulsion-Airframe Integration Engine-airframe integration is a key aspect of configuration def- inition for hypersonic vehicles- increasingly so as the maximum air- breathing Mach number increases. The combination of long forebody and low Reynolds number produce a thick entropy layer that must be ingested by the engine or diverted. Its thickness is sensitive to Mach number and Reynolds number, and will vary significantly over the flight corridor. The Low Reynolds number is dic- tated at high Mach number by the need to reduce heat transfer rates and pressure loadings, to transition to rocket propulsion for orbital insertion, or both. 4) A very large ratio of capture area to frontal area results from low Reynolds number (high altitude) and small fractional energy addition due to combustion. 5) Efficient operation at very high Mach numbers require configur- ations that pose serious integration problems at off-design Mach num
FINDINGS AND RECOMMENDATIONS her. The large nozzle expansion leads to very large base drag at transonic speeds. The interaction of the nozzle expansion plume with the slipstream, and with the reaction control system, will influence both the net thrust and moments at near orbital speeds. 6) Integration of the low speed pro- pulsion system with the hypersonic propulsion system, in a way that does not degrade the performance at hypersonic speeds, is a major concern. 7) The variation of engine-inlet boundary layer conditions with flight conditions (Mach number, Reynolds number, and altitude) must be quantitatively predictable, or an engine concept must be devised that is insensitive to the boundary layer thickness. 8) Items 5, 6 and 7 above are un- solved problems. Engine-airframe integration should receive more emphasis, by teams drawn from both engine and airframe contrac- tors. 3.3 Propulsion Systems I ~ Injection of hydrogen fuel and rapid mixing with air with minimum loss is the most influential factor affecting the engine length and heat load. (2) The heat release pattern in the engine is determined by the rate of molecular mixing, and the super- sonic flow in the engine is extremely sensitive to the heat release pattern. The ingestion of ramp boundary layer and bow shock layer by the engine poses difficult problems of engine design and penalizes engine 57 performance. (4) The stability of the scramjet flow with hydrogen reaction is not understood; instability poses the possibility of developing strong shock waves and catastrophic loss of engine. Short term design studies and long term research studies of hydrogen injection and mixing should be increased as soon as possible to assure that this issue does not become an obstacle to high-speed engine development. Measurement of molecular mixing should be emphasized and the exploration of novel techniques of mixing augmen- tation must be encouraged. Because the combustor heat release pattern is mixing controlled and, further, because the state of the air entering the combustor may be extremely non-uniform, the mixing process must be understood to the extent that it can be controlled" as well as accelerated. (6) One-dimensional or quasi one- dimensional computation of reacting flow in the combustor is inadequate and often misleading. (7) The H-OH reaction must be com- pleted for the scramjet to perform well. Much of this reaction will happen during expansion in the nozzle and this reaction may "freeze out" early in the expansion process. (8) Under the most severe conditions of operation the molar flow rate of hydrogen in the cooling passages is more than double the total molar flow rate of air through the engines. Effective use of coolant and minimization of pumping losses is imperative and an unusual degree of integration with structural design is required.
58 HYPERSONIC TECHNOLOGY FOR MILITARY APPLICATION (9) The hydrogen requirement to cool the engine exerts an unusually high leverage on the airplane size and weight. It is essential to refine the accuracy of and confidence in estimates of cooling requirements before final selection of airplane size. ( 10) High priority and additional emphasis must be given to the research and design studies con- cerned with the utilization and management of hydrogen coolant flow. This is of particular importance in the portions of the engine that experience geometric changes during the acceleration. (1 1) Film cooling and sweat cooling with hydrogen have very attractive fea- tures and both technological and research efforts must be augmented. The gas dynamic peculiarities of using hydrogen as the coolant should be emphasized in these studies. This work must be accel- erated because coolant requirements have such a powerful impact upon the airframe design. (12) The scramjet must operate at peak performance throughout its entire Mach number range during accel- eration. The configuration and geometric changes required over this range are very extensive and must be done with the minimum introduction of shocks and other losses. (13) The geometric changes required of the scramjet over its Mach number range place demands upon design of cooling passages, coolant flow management and seals that are of unprecedented difficulty. (14) Transition between the three oper- ating modes of the propulsion system, subsonic to ramjet, ramjet to scramjet, and the re-start and reverse transitions upon re-entry, present extremely sensitive and difficult problems. These must be solved to avoid placing unaccept- able structural and thermal loads on the airframe and engine, which may lead to failure. ( 15) The transition from one engine mode to another, especially from the ramjet to scramjet, might produce large unsteady loads and unsatisfactory starting. To insure against these problems, sufficient rocket propulsion should be incor- porated into the powerplant com- plex to suppress any severe problems during transition. ( 16) Some rocket propulsion must be incorporated into the final propul- sion system a) to reach orbit from the scramjet Mach number limit, b) to facilitate the gradual introduc- tion of advanced scramjet technol- ogy over the life of the airplane, and c) for de-orbit maneuver. (17) The high-speed engine development should be predicated on the prob- ability that most of the develop- ment will be done in flight test. To develop an engine in flight, an auxiliary rocket propulsion system, separate from the NASP engine package, will be needed to augment thrust and assure airplane trim during high-speed engine tests. ( 18) Complete scramjet engines will undergo development and testing during the flight program, not in ground-based facilities. Con- sequently, it is necessary to incorporate some rocket propulsion - separate from any that may be integrated with the scramjet - for setting desired engine test condi- tions and extending the flight envelope of the airplane.
FINDINGS AND RECOMMENDATIONS (19) With boundary layer ingestion, the scramjet engine is quite sensitive to angle of attack. This variation in engine operation will be reflected in changes of pressure distribution over the discharge nozzle, resulting in a large pitching moment. These difficulties may be avoided only through unusually careful integration during airframe and control system design and development. (20) The modular design of the scramjet engine allows interaction between the inlets of adjacent modules during ramjet start-up and transi- tion from ramjet to scramjet oper- ation. This interaction can propagate inlet malfunctions from one module to adjacent modules. This behavior appears likely to be particularly sensitive to yawing motions of the airplane. (21) The strong interaction between the high speed engine, the forebody ramp, and the external nozzle, and the very powerful dependence of this interaction upon pitch and yaw of the entire airplane has important implications upon the control system and the coupling of engine and airplane. Consideration should be given to design compromises that would reduce this potential problem, even at the expense of reduced performance of the earliest NASP configurations. 3.4 Aerodynamics This committee has identified two main aerodynamic problem areas. At low hypersonic Mach numbers (below about 10), the problem is mainly one of fluid mechanics. The prediction of the boun- dary layer and flow field characteristics are required to permit the detailed determination of the pressure distribu- tion, skin friction, heat transfer, and 59 the flow field condition around the body and through the inlet to the combustion chamber. Above Mach number 10, the aerodynamic problems involve the factors identified in the lower Mach range, with the additional complication of the rate kinetics of real gas effects and the special problems of low density flows and small bluntness dimensions. Neither the low nor high Mach number areas are currently amenable to detailed wind tunnel exploration or validated compu- tation to provide a well-grounded base for design, although some results are available from facilities that partially simulate the real flows. Therefore, progress must rely on a fragmented approach, where limited experiments and computation will in time provide an adequate base for design. Validation of this base will require flight tests that include many elements simultaneously, a situation not amenable to full simulation on the ground or by validated compu- tation. 3.4.1 Low Hypersonic Speeds (Mach Numbers 6 to 10) The prime requirements are to identify the location and details of transition initiation, the transition region, the mixed flowfield and boundary layer characteristics around and through complex geometries with cold walls, and the mixing phenomena in the combus- tion chamber. These problems are recognized and are within reach of current technology but have not been solved. 2) A unique Mach number 3.5 "quiet" research facility at NASA Langley is beginning to provide results indicating the necessity for such flow characteristics. A Mach number 6 research facility has been approved, but not built. These research facilities are inadequate for the requirements of Mach
