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Hypersonic Technology for Military Application (1989)

Chapter: 2 Technologies Relevant to Hypersonic Vehicles and Their Status

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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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Suggested Citation:"2 Technologies Relevant to Hypersonic Vehicles and Their Status." National Research Council. 1989. Hypersonic Technology for Military Application. Washington, DC: The National Academies Press. doi: 10.17226/1747.
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STATUS OF HYPERSONIC TECHNOLOGIES 13 2.0 TECHNOLOGIES RELEVANT TO HYPERSONIC VEHICLES AND THEIR STATUS 2.1 Aerodynamic - Propulsive Integration As usually conceived, the hypersonic air-breathing propulsion system uses a low speed system for operation from standstill to about Mach number 2.5, a ramjet for operation to about Mach number 6.5, and a scramjet for Mach numbers above 6.5. If these three systems are to be combined{, this must be done in a way that does not dlegrade their individual performances when they are active, and with acceptable weight and complexity. This is a major chal- lenge for the designer of hypersonic aircraft, and must be recognized as such. We will not deal with it comprehensively in this discussion, however. Our focus here is primarily on hypersonic propul- sion. Integration of the airframe and propulsion system is a central feature of all conceptual designs for hypersonic flight vehicles mainly because the engine capture area must be a large fraction of the vehicle frontal area. Among the contributing factors are: 1) a low thrust per unit of engine air- flow at hypersonic speeds, which results from the small fractional change in energy of the engine air- flow that can be achieved through combustion; 2) the need to fly as high as possible to minimize the heat load on the structure, which results in a pro- portionately low engine mass flow for any given capture area; and 3 ~ the desire to make efficient use of the compression by the bow shock of the vehicle, which leads to the need to capture, in the engine, most of the flow through this shock. The need to maintain a weak bow shock to minimize losses and for a large capture area require slender configur- ations in which the entire forebody, or at least its lower surface, comprises the engine inlet. (See Figure 2-1.) The same factors dictate that the aft end of the fuselage serve as the expansion sur- face for the propulsive streamtube. While the resulting configurations are conceptually appealing, especially to the propulsion-oriented, they pose prob- lems that, though not entirely new, are certainly more serious than for more conventional designs, in which the pro- pulsive streamtube and fuselage and wing airflows are farther apart. Thus, in most if not all conceptual designs for hypersonic vehicles, the propulsion system is assumed to ingest the boun- dary layer flow that develops on the forebody. Most successful propulsion installations in the past have avoided this. If the propulsion system ingests the boundary layer: l ~ the ramjet, whether operating in the subsonic or supersonic combustion mode, must pass two parallel streams of very different velocities and temperatures, or 2) the boundary layer flow must mix with the free-stream. The former will lead to constraints on the supersonic combustion process, because the pressure must be equalized between the supersonic and subsonic streams, imposing serious performance penalties. The latter may result in losses, or in heating of the supersonic stream, which is counter to the principal rationale for the supersonic combustion ramjet, namely to lower the combustion temperature. These issues are discussed further in the propulsion sections. However, it is

14 proper here to ask whether the engine should ingest the boundary layer from the forebody. One can imagine con- figurations in which the boundary layer flow bypasses the combustor, the inlet ingesting flow primarily from outside the boundary layer, but we have seen little evidence of serious consideration of this possibility. The high degree of integration also poses serious control problems if, as seems probable, the inlet compression and nozzle expansion occur only on the lower surface of the vehicle. Both the inlet and nozzle then contribute strongly to the vehicle's lift, particularly at high flight Mach numbers. A balance be- tween the lift forces on the inlet and nozzle will determine the pitching moment produced by the propulsive streamtube. While the lift and thrust can conceivably be oriented through the center of mass and the pitching moment can be nutted at the design point, it is not yet clear how these balances can be maintained at off-design conditions, without large forces from control sur- faces. Furthermore, even if a design can be evolved to meet these require- meIlts in normal operation, what pitching moment will follow from a sudden loss of heat addition in the combustor, and the resulting modification of the nozzle expansion? Although an inlet upstart probably will be unacceptable in a scramjet operating at very high Mach numbers, how will control deal with such an upstart? A further control difficulty may arise from the ingestion of the ramp boundary layer by the engine. To see this, suppose the airplane pitches upward so as to increase the angle of attack. The ramp boundary layer and shock layer now accumulate more losses, higher static pressure, hotter and lower stagnation pressure air, and these are ingested into the engine. With no change in fuel flow, tile engine thrust and the pressure on the nozzle face HYPERSONIC TECHNOLOGY FOR MILITARY APPLICATION change and ~ strong pitching moment is developed. Although quasi-steady operation of the engine is involved, the coupling with the airframe dynamics is strong and must be dealt with by the control system. This is not only an issue of sensors and control response, it is a question of engine and airframe design. For example, should steps be taken to design or position the engine to be less sensitive to such disturb- ances? In the extreme, should the engine avoid ingesting the boundary layer entirely? The presently proposed configurations for the NASP are subject to such airframe interactions that may pose significant control problems. Most of the current airplane config- urations use a modular propulsion system with several engines side by side under the airframe. The individual modules probably will operate under nearly identical conditions and usually not interact or interfere with each other. However, when changing propulsion mode from low speed engine to ramjet and from ramjet to scramjet, each of the modules will sometimes experience a start-up transition that might induce airflow disturbances that propagate upstream of the inlet. It is unlikely that this phase of operation will occur simultaneously for each of the modules and, indeed, it may not occur symmet- rically with respect to the desired thrust axis. Furthermore, if for some reason the starting transients are severe, the inlet disturbances or inlet malfunction can propagate from one module to the other, perhaps leading to a catastrophic malfunction of the propulsion system. Such difficulties can be closely coupled with yawing disturbances of the airplane which may, in turn, be induced by unsymmetric mode changes of the engine modules. The large nozzle expansion required for efficient hypersonic operation leads to ~ serio us base drag problems at transonic speeds, where the nozzle

STATUS OF HYPERSONIC TECHNOLOGIES pressure ratio is far too low to fill the entire base area. This problem has appeared in non-afterburning operation of aircraft where the nozzle area required for afterburning operation has dictated a large base area. Ejector nozzles, which fill this base area with a secondary or tertiary air stream, have provided partial solutions, but this avenue will be much harder to follow for hypersonic aircraft. Base burning is a possible alternative, but one whose heating and fuel consumption implic- ations have not been fully explored. Another class of problems arises from the need to integrate the low speed propulsion system with the high speed ramjets, with acceptable weight and complexity, and without serious interference with the function of the scramjet at high Mach numbers. There should be no extraneous projections into the airflow that would cause strong shocks or excessive heating. All sur- faces of the engine flow path must be actively cooled, and this further argues for a minimum of complexity. While we certainly do not argue that innovative solutions to these design problems are improbable, there is no base of success- ful designs on which to draw to solve them. The importance of aerodynamic- propuisive integration for hypersonic vehicles has been recognized for many years, and has been highlighted over the last two by active participants in the NASP program and by advisory groups. But today there is more enthusiasm than integration. This problem is not being adequately addressed and will remain so as long as responsibility for the pro- pulsion system and vehicle are divided. We urge that an organizational structure be created" where responsibility for the conceptual design of both engine arid vehicle resid e in one organization. This must be done soon, so that these issues are faced in the conceptual design 15 phase, not after a configuration has been selected. 2.2 Propulsion Systems In addition to the several technol- ogical areas that are either unique to the scramjet or are emphasized to an unusual degree, there are important issues of ( 1 ) transition (2) the limit- ations to development and testing above Mach number 8, and (3) the extraordin- arily sensitive interaction between the engines and the airframe. These issues will be discussed separately after we have examined the basic technologies pertinent to the high-speed engine. 2.2.1 Basic Scramjet Engine If the requirements underlying the principles fundamental to the scramjet engine are not satisfactorily met, the engine may perform~no better, or per- haps worse, than a high performance rocket. The main issue is to maintain the static temperature of the air in the combustor at a reasonable value while the aircraft is flying in the Mach num- ber range 10-24. At a very high air temperature the eventual reaction of hydrogen and oxygen to water vapor is very slow or incomplete or both, and the specific impulse of the engine falls from a value in excess of 1000 seconds to well below 500 seconds. This concept breeds two conflicting technological problems. First, due to the high air velocity in the combustor, the combustor would have to be very long to achieve a reasonable residence time. Second, extreme heat transfer rates and wall shear losses require the combustor to be as short as possible. How to balance these issues and whether or not there is an acceptable balance underlie the factors discussed below. The processes of injecting the hydrogen fuel and mixing it with air in

16 the scramjet appear to be the most dif- ficult obstacles to the realization of a successful engine; and they are processes in which our present fundamental and technological base is weakest. Hydrogen fuel must be injected into the engine with very low losses, at local Mach numbers as high as 8, where losses have the greatest effect because of the small fractional heat addition due to combus- tion. It is generally agreed that shear layer mixing rates drop under some conditions of supersonic relative motion between streams. The technological basis for this is not extensive and design experience is lacking. Possible alternatives, such as mixing augmenters, wall injection, and shock enhanced · · · . . . mixing are In an even more primitive technological state. It must be made clear here that satisfactory mixing for a chemical reaction process contrasts sharply with one in which, for example, momentum is being exchanged. The mixing must be complete on the molecular level to allow combustion. Not only is this a more time-consuming process but the exper- imental difficulties of assessing the completeness of molecular mixing are considerable, and therefore the tech- nological basis will be slow to develop. Considerable effort is now being expended in appropriate investigations and the results will be of unusual value not only to the present development but also to future efforts of scramjet development. It is not now clear just how extensive the data will have to be to impact the NASP Program. During the most important periods of scramJet operation the combustor Mach number is in the range of 2.5 to S. This flow field is quite complex due to the heat release, which is controlled by the molecular mixing process, and by the ramp boundary layer and bow shock layer that may be ingested by the engine. The heat release has a pro HYPERSONIC TECHNOLOGY FOR MILITARY APPLICATION pounced effect on the structure of the flow field which, in turn, strongly influences the mixing processes. This coupling introduces complexities that we find very difficult to cope with either experimentally or computationally. Today it is impossible to describe with certainty the best geometry of the com- bustion chamber, for any Mach number or altitude. A serious concern among workers experienced in the field is the stability of the hypersonic flow in the combustion chamber during combustion. Would a small disturbance imposed on a design flow field decay, diverge, or lead to a pulsating combustion process? A central obstacle to understanding the result of such a time-dependent disturbance to the combustion chamber flow is our current incapability to either experimentally measure or to compute this chemically- reacting flow. The ingestion of the ramp boundary layer and shock layer may lead to large variations in the air density and stag- nation pressure over the cross-section of the inlet. The air density from top to bottom of the engine inlet may some- times vary by a factor of four. Exper- ience has shown that such conditions facilitate communication of disturbances through the boundary layer ahead of the engine, which may result in unfavorable inlet conditions and even inlet instabil- ity. Such a disturbance may couple with the combustion process in the chamber with possibly unfortunate results. Our experience with this problem is restric- ted to a much lower Mach number re- gime than that appropriate to the NASP and requires serious experimental atten- tion and perhaps design compromises. Almost exclusively, our ability tO calculate chemically-reacting unsteady flow fields is restricted to one dimen- sion. Such steady one-d imensiona1 calculations are being used in engine performance calculations, yet quasi one

STATUS OF HYPERSONIC TECHNOLOGIES dimensional analysis can cope neither with the mixing-controlled combustion issue nor with the stability problem. But difficulties arise even at the more elementary level of performance calcu- lation. When the engine ingests the ramp boundary layer and bow shock lay- er, the gas entering the combustor has a very non-uniform temperature distribu- tion over its cross-section. As a con- sequence, the chemistry, which has a strong and non-linear temperature dependence, may vary even more vio- lently over the cross-section. In trying to adapt one-dimensional analysis to this problem one is faced with the issue of choosing appropriate average values for each cross-section, which introduces large and unacceptable errors in the results. Also, the output of a one- dimensional analysis can provide only a uniform input to the nozzle calculation that follows. Even an approximate cal- culation of the nozzle expansion process makes it clear that the pressure distrib- ution over the nozzle surface and, con- sequently, the calculated performance of the engine, is very sensitive to the details of the gas state distribution at the start of expansion. It is most important to make at least an approx- imate accounting for the two- and three-dimensional nature of the combus- tor and nozzle gas dynamics. As a result of the relatively high temperature and short residence time in the combustion chamber, the hydrogen- oxygen reaction will not reach an equilibrium water vapor concentration before reaching the nozzle. So it is important that the reaction between OH and H be completed to the maximum degree during expansion in the nozzle. Because it is most unlikely that crucial experiments can be done, a high degree of reliance must be placed upon compu- tations. These must, at the least, be for two-dimensional reactive flow arid pref- erably three-dimensional. Moreover, we must account for the non-uniform state of the gas at the nozzle entrance. At present this magnitude of numerical calculation is impossible and the situation is unlikely to change very soon. The nozzle may be one aspect of the NASP that undergoes significant development during flight test. As we have noted above, the ingestion of the ramp boundary layer and shock layer exacerbates the problems of the intake, combustor, and nozzle. Although it is difficult to study without better computational capabilities, we must nevertheless consider whether or not the boundary layer should be swallowed. Even at the cost of some possible reduction of the airplane performance, the additional design certainty that would accrue from capturing clean air might result in an overall benefit. 2.2.2 Cooling Problems Although active cooling with hydro- gen will be employed over several sections of the airplane, the larger part of the cooling load is connected with the engine. Even the most optimistic estimates show that the hydrogen flow rate required for cooling exceeds the stoichiometric fuel flow requirement above Mach number 15. Less optimistic analyses suggest that above Mach num- ber 20 the hydrogen cooling requirement may exceed! four times that for combus- tion. This means that the molar flow rate of hydrogen in the cooling passages of the airframe is more than twice the total flow rate of air into the engines. And because it is the molar flow rate of gas that determines the pumping power requirement, details of the cooling passages and the active management of the coolant flow to different regions of the airframe and engine are as impor- tant as the external gas dynamics. In fact, because the airframe is largely sized by the hydrogen tankage, the cooling requirement exerts an unusually high leverage on the airplane size and

18 weight. The actual cooling requirement is now one of the least certain elements in the entire vehicle, so the airframe probably will be over-sized to accom- modate this uncertainty. The design of cooling passages in the engine and assurance of their effectiveness are made much more dif- ficult, and may perhaps be compromised, by the extensive geometric changes required of the engine over its Mach number range. It is in the regions of the hinged joints, and the scramjet may require several of them, that meticulous design and careful coolant management must be exercised. The emergence of coolant effective- ness, active coolant management, and a high degree of integration of the coolant system with the structural design as a significant and novel aspect of the NASP is just beginning to be taken seriously. Furthermore, this issue will command attention for any hypersonic airplane, and the technology base that is acquired here will be of permanent value. 2.2.3 Propulsion System Transition The acceleration of the NASP from take-off to orbital velocities requires not only widely different operating conditions for each engine but the transition between the two or three different engines, each of which is designed to operate in a specific Mach number range. In the "conventional" arrangement, one engine will operate from zero to low supersonic Mach num- bers, a ramjet with subsonic combustion up to the lower limit of the scramjet range and, finally, the scramjet to near orbital velocities. It may be anticipated that these changes of propulsion mode .. . . . . wit he sensitive processes, requiring careful control to assure that each newly fired engine starts properly and that the transition induces no extreme and unusual loads. Though we have HYPERSONIC TECHNOLOGY FOR MILITARY APPLICATION some experience with this sort of problem, two important factors make the present situation unique. First, the transition to the scramjet mode lies in a Mach number range far outside our experience and where an error may be costly. Second, in contrast with our experience, experimental verification and development will not be possible. The chance for serious trouble here is so likely and the potential damage so great that the issue is significant. The development problems of the individual propulsion systems are so severe in themselves that considerations of this problem might be postponed too far into the program. 2.2.4 Auxiliary Rocket Propulsion Because the ground-based test facilities for the NASP will be unable to accommodate the Mach number and enthalpy range appropriate to the high speed engine, much of the engine testing and development must be planned as part of the flight program. Engine develop- ment and testing will occur as the flight envelope of the airplane is expanded into the higher Mach number range. During this process, the airplane will require a separate propulsion system to allow a range of conditions under which to test the engine. This probably would take the form of several hydrogen- oxygen rockets distributed appropriately over the airframe but physically separate and with an independent control system. These rockets should be capable of taking over the propulsion and trim of the airplane during testing of the high- speed engine. This will allow much greater flexibility in exploring the operating envelope of the engine at nearly constant inlet conditions. This requirement will hold not only for the NASP but for any hypersonic research vehicle using scramjet propulsion. It appears that there is a quite separate requirement for some degree of

STATUS OF HYPERSONIC TECHNOLOGIES rocket propulsion as a component of the final NASP propulsion system itself. Rocket propulsion will be needed to insert and stabilize the vehicle in orbit, and possibly to maneuver in orbit and de-orbit. It is very likely that the rockets will prove invaluable in main- taining thrust and trim during change of propulsion mode. Part of the long-term growth pro- cess for a conventional aircraft and propulsion system is the gradual improvement of the thrust level and efficiency of the engine, and a cor- responding improvement of the overall aircraft performance. The NASP will be no exception to this pattern and will, in fact, undergo even more striking changes than the conventional aircraft. In its initial form, the scramjet probably will become ineffective at velocities con- siderably below orbital velocity and will require a component of rocket propulsion during some final portion of its accel- eration. But the long-term growth of the engine will reduce, though probably not eliminate, this requirement, and the propulsion system must accommodate this growth without severe design changes to the entire airplane. Rocket thrust should be an integral part of the final operational NASP propulsion, not just a temporary feature of the flight test program. 2.3 Aerodynamics The aerodynamics of a hypersonic vehicle must be closely integrated with the power plant. Presentations to this committee arbitrarily focused on the aerodynamics of the entire vehicle including the inlet, but excluded the propulsion system element, consisting of the fuel injectors, combustion chambers, and expansion nozzle, despite the fact that many items included in the aero- dynamic discussion are important in the propulsion system itself. The aero- dynamic discussions must address flight 19 conditions from take-off to orbit, conditions of high dynamic pressure (2000 psf seems to be an approximate, practical, upper limit), low Reynolds number conditions at high altitude, and conditions in the non-continuum region approaching free molecular flow. The aerodynamic considerations require the treatment of real gases with full viscous effects applied to configurations that consist of blunt noses and leading edges on slim bodies with complex configur- ations that generate three-dimensional gradients and shock waves. The requirement is to predict, with reason- able accuracy, the following parameters, assuming similar inputs will be provided by the propulsion system: · Local pressures, heat transfer, and skin friction, on the surface of the entire vehicle, including the flow through the inlet. a) To provide the forces (lift, drag, moment, etc.) over the full flight envelope b) To provide the detailed thermal loads for cooling and structural design, and c) To provide dynamic inputs to the aerodynamic and thermal . control systems. The three-dimensional flowfield detailed information on the conditions and the state of the with local gas, for configuration and control design, for inlet positioning conditions, and optimum inlet design for flow to the combustion chamber. The statuses of the various aero- dynamic technology elements have been evaluated on the basis of the experi- mental data base from wind tunnels and flight, and from the examination of computations that have been validated or calibrated under the appropriate condi- tions. Current computational capability enables us to calculate complex flows, but is limited in dealing with the details of viscous arid real gas effects. Current technology provides the capability to

20 compute real gas effects if the signif- icant reaction rates are known. In general, aerodynamic characteristics for three-dimensional configurations up to about Mach number 10 are reasonably known or can be predicted, if transition location and the extent of transition can be provided. Beyond that, where real gas effects become important and viscous effects more complex, very little three- dimensional aerodynamic information is available beyond the blunt body con- figuration for realistic design procedures. The following comments address the inclusion of detailed viscous effects, real gas effects, and real gas complex flow conditions. 2.3.1 Viscous Effects The primary viscous effects are experienced in the boundary layer where the hypersonic conditions require a considerable extension of low speed experience and computations to high Mach numbers and "cold wall" conditions. The main problem is to define the condition of the boundary layer at any point on the vehicle during any cond- ition of flight. This definition is required to calculate heat transfer and skin friction (critical elements in the design of the structure and cooling system), to predict and avoid unaccept- able separations, and to compute per- formance. 2.3.1.1 Transition Point Determination Determination of the transition point is a critical element in the design and performance of a hypersonic vehicle. Unfortunately, there is no well-founded theory for this determination. Conven- tional low Mach number use of the para- meter Rex/M at values as low as 150 to about 300 has not been validated under the conditions of hypersonic flight or for complex flows. New attempts to use the parameter en are being studied by HYPERSONIC TECHNOLOGY FOR MILITARY APPLICATION NASAiLangley. The problem is that wind tunnel data are biased by wind tunnel disturbances and free flight results are primarily for simple bodies, axisymmetric, with zero gradients. Considerable work is being done in this field, including extensions of laminar, linear stability theory, and attempts to evaluate cold wall effects in the lower hypersonic region. Still missing are high Mach number data with 3-D gradients and shock waves. Proposals are being considered for extended wind tunnel tests under the appropriate conditions and for flight tests of bodies that might include more complex geometries. The problem is made even more difficult by the lack of knowledge of the disturbance field in the stratosphere, which could be a factor in triggering transition. 2.3.1.2 Extent of the Transition Region Although it is critical to determine the transition point, the condition of the boundary layer downstream of that point is also important. There is some indi- cation that the transition region between fully laminar and fully turbulent flow gets longer as Mach number increases. It is possible, therefore, that a con- siderable region of transitional flow will occur over a hypersonic vehicle or in a hypersonic inlet. The characteristics in this region are unknown. It is believed to be partially laminar and partially turbulent, in a temporal sense, but detailed information in this region must await solution of the transition point problem of sub-section 2.3.1.1. 2.3.1.3 Turbulent Boundary Layer Somewhere downstream of the tran- sition point, the flow will be fully turbulent. At higher Reynolds number conditions, low altitude flights ~ considerable part of the hypersonic vehicle may be turbulent. Depending on the trajectory chosen, this turbulent

STATUS OF HYPERSONIC TECHNOLOGIES flow could occur under Mach number conditions that have not been exten- sively studied. The effects of high external Mach number and cold wall, with gradients have not been well- val~dated for the higher Mach number regimes. 2.3.1.4 Test Conditions It is possible to do considerable work on items 2.3.1.1 through 2.3.1.3 in ground facilities below Mach number 10, although the results to date are only preliminary and effects of wind tunnel conditions on such phenomena as the transition point have not been fully explored. The effects of cold walls in 3-D complex flows, with the gradients and shock waves that are required for optimized configurations in this lower Mach number range, along with inlet studies, can be studied, but have not. These same problems at high Mach numbers are critical issues because facilities are limited, and all of the parameters for boundary layer validation must be applied, including wind tunnel disturbance effects (now including entropy and constituent effects), wall cooling and catalysis, and wall rough- ness. 2.3.2 Real Gas Effects Blunt body flows have been calcula- ted and validated to some considerable detail. There is some question of the results for very high altitudes where the constituents of the atmosphere and the local conditions are not well known. Accurate measurement of the conditions ahead of the vehicle in flight would provide an important input into the analysis of the results, and considerable effort is being expended in this area' although solutions to the problem are not yet in hand. 21 Slim, complex, three-dimensional bodies with blunt roses and leading edges provide a special problem in real gas effects. Although the blunt nose solutions are well known, the streamlines downstream of the blunt nose carry different histories; mixing and recombin- ation rates depend on local conditions. A similar problem is experienced in the expansion of the combustion chamber flows through the nozzle. Although the chemical kinetics for the blunt nose are well known, not all of the data needed for recombination rates for the down- stream flows are in hard. These data should be obtainable, however. The effects of real gas kinetics on boundary layer characteristics are only beginning to be explored. The effects are probably not important below Mach number 10. Above 10, surface catalysis and chemical reactions in the boundary layer flow, with cold walls, could affect boundary layer characteristics and the transition point and transition region. Studies are only in their elementary phase and some reasonable solutions for the boundary layer characteristics will be required before the addition of real gas effects can be attempted in detail. 2.3.3 Real Gas Effects in Complex Flows Inlets, bodies' fins, and wings, and their interactions, cause a combination of viscous effects, shock waves, and strong gradients with real gas effects. These must be understood to give detailed flowfields, such as for the combustor, where the local conditions and the constituents (state of the gas) must be known at each point. Local, configuration-specific hot spots must be identified. Inlets have been tested at low Mach number. At high Mach num- bers, the inlet tests nest be closely associated with ~ ~ en -defined initial flow determined by the details of the forebody. The lack of adequate testing facilities and the ~or~plexity of the flow

22 HYPERSONIC TECHNOLOGY FOR MILITARY APPLICATION makes this an important problem that will require considerable further work, both for computational validation and for inlet optimization at high Mach numbers. 2.4 Control, Guidance, Instrumentation, and Information Systems As with propulsion, aerodynamics, and materials, further developments in controls are required for hypersonic flight. Challenges in controls-related areas include vehicle trajectory, con- figuration, and effective dynamics, and the related fields of instrumentation and information systems. The Wright Bro- thers, in 1901, realized that approaches to problems in propulsion, structures, and aerodynamics were more or less understood, but that an "inability to balance and steer still confronts students of the flying problem...Eand] when this one feature has been worked out, the age of flying machines will have arrived...." Today for hypersonic fight, we recognize a host of aerodynamic, materials, propulsion, etc., problems clearly enough that we can, to some degree, define and even assess them. But the means to achieve directed and governed propelled hypersonic flight can be seen, at best, only dimly. The problem is how to exert simultaneous control over aircraft flight path, aerodynamic attitude (local flows at critical locations), and propulsion. All guidance and control system aspects are somewhat critical, but the most impor- tant "missing means" are probably the fundamental system architecture and control structure, special sensors, effecters, and the information processors that implement the control structure. In this section, we discuss various issues central to controlled hypersonic flight: physical factors, vehicle dynam- ics, vehicle-flight control systems, sensors, and information systems. Although some of the key problems in these areas can be alleviated by adopting the system design and test philosophy of gradual and punctuated buildup described below, an aggressive technology develop- ment program in controls, sensors, and information systems is still necessary. The uniqueness of vehicle dynamics and flight control disciplines to hyper- sonics is sometimes a matter of degree rather than of kind. Control of sideslip, for example, is a ubiquitous general problem in aircraft, but the control precision needed for some hypersonic aircraft configurations can be an order of magnitude greater than for other craft. Further, the ways to achieve this control on a multiply-redundant system basis are, at best, obscure and possibly beyond the present state of the art. This ancient problem could become a design driver at best and a potential show stopper at worst. 2.4.1 Physical Factors Hypersonic vehicles will demand sophisticated control for the engines, inlet, and guidance and control of the aircraft. Subsystems probably will be more complex than even an entire air- craft automatic flight control system of today. The demands on the information processing equipment will be very sig- nificant. This will lead to hierarchical distributed fault-tolerant information systems. Embedded in these systems will be many sensors and effecters at diverse locations around the aircraft. This, in turn, will place unique requirements on the subsystems and sensors. All the aircraft designs will have a large cryo- genic tank at the center of the fuselage. The many sensors, wires, fiber cables, and hydraulic lines will have to pass through these tanks or be routed around them. Either way, there is a technology issue. At a minimum, the components will have to operate in the environment peculiar to their location. Some sensors and effecters together with their con- nectors and cabling will be near the

STATUS OF HYPERSONIC TECHNOLOGIES skin, in the engine, and in the cryogenic fuel. They must accommodate large temperature gradients while guaranteeing electrical, optical, and hydraulic integrity. If routed near the aircraft subskin, they must withstand the very high temperatures expected in that area, yet exposure to cryogenic temperature will be a challenge to hydraulics, if used. Typical maximum operating temper- ature for today's fiber cables is 75-85° C. New fiber for aircraft applications to 200° C is being tested. Beyond 200° C, current organic jacket material be- comes inappropriate. Questions concern- ing the optical properties of specific fiber types exist for very high temper- atures. The waveguide itself is made of fused silica, the material used in space shuttle windows. However, without a protective jacket, this fiber is subject to serious mechanical and chemical damage, particularly in the vibratory environment likely in a hypersonic vehicle. Two program paths are needed to minimize risks. One is a technology program focused on the thermal capabilities of sensors, cables, and connectors and the other is careful attention to the details of the thermal environment to which this type of equipment will be subjected. Hypersonic vehicles with their very high temperature will present an unusual requirement on control system sensors. Even in conventional vehicles, it is a challenge to be able to locate the sen- sors involved in stabilization of flexible body modes at useful locations. Nodes and antipodes are sometimes preferred by the flight control system designer, but such locations may be counter- indicated by other demands. This prob- lem will be compounded by the likely significant variation in mass distribution as fuel is consumed. Because these vehicles will usually exhibit both lateral and longitudinal instabilities (requiring a high-bandwidth controller to correct) and also are projected to have low 23 frequency body bending modes, active control of body bending will be needed. This will require a set of sensors distributed around the body at particular locations selected to expedite control. Just as the cables may be exposed to very high temperatures, so will these sensors. Most accelerometers and rate gyros available for such applications are limited to approximately 85° C maximum temperature. Piezo-electric accelero- meters are available that can operate up to around 250° C. Clearly, either cool- ing must be provided or a new family of high temperature body motion sensors will be needed. The problem is even more complex because the body bending parameters could be a function of tem- perature. In a normal situation, when a sensor cannot be properly placed, a model of the structure can be used to estimate the signal that would be obtained from properly located sensors. This is not a preferred solution. Where bending parameters are temperature- dependent, it is even less desirable since the model required here will be even more complex. This will place consid- erable robustness demands on the control system. Almost all modern aircraft use hydraulics to provide the necessary forces for control. Hypersonic aircraft likewise will need high bandwidth, high strength controls. Either hydraulics or newer technology high strength elec- trical actuators will be required. Either way, high pressure hydraulics (or elec- trical power cables) must be either routed under the skin where they will suffer a totally new environment due to the high temperatures that could be encountered or passed through or near the cryogenic tanks. A key factor will be the chemical stability of the hydraul- ic fluid which limits its lubrication properties. Carburization and viscosity losses usually limit petroleum- based fluids to applications around 200° C.

24 Synthetic emulsions may work up to approximately 400° C. Fluids would need fire retarding capability in the event of a leak. Servo valves present additional problems due to magnetic degradation. Their performance in most cases begins to be affected at about 200° C. Again, this important technol- ogy development issue must not be overlooked and the design must consider protection of this type of equipment also. 2.4.2 Vehicle Dynamics and Controls 2.4.2.1 "Aerodynamic-Attitude" (Angle of Attack and Sideslip) Control Projected hypersonic "aircraft" are basically engines with attachments. The local flow directions at key locations near the engines are central to engine operation. The basic airplane perfor- mance is extremely sensitive to the angle of attack and sideslip. On large vehicles, the lower frequency structural and slosh modes may be significant contributors to the local effective angle of attack and sideslip as seen by the engine. It follows from this litany that the establishment and maintenance of the local angle of attack and sideslip within very narrow bounds by the inner loops of the attitude control/structural mode control systems are an enabling, lim- iting, and driving technical requirement. This amounts to precision control of the vehicle as an engine mount! The precise requirements are a bit obscure, but angle of attack established and held to 0.1 degree even while maneuvering is sup- portable for some proposed configur- ations. Just how this is to be done as part of the flight control system (which simultaneously controls the flight path vector) is a question of the first order. The maintenance of near-zero aero- dynamic sideslip in the presence of HYPERSONIC TECHNOLOGY FOR MILITARY APPLICATION thrust asymmetries, during maneuvers, etc., in the supersonic and hypersonic regimes is a critical problem in almost all its aspects - control power and effectiveness, sensing, and actuation. Precise sideslip control is needed in essentially all hypersonic designs. 2.4.2.2 Integrated Propulsion/Flight Control and Guidance Speed control is sometimes essential to the trajectory adjustment control for hypersonic vehicles. Ramjet and scram- jet engines have some unknown dynamic features that can seriously affect a tightly integrated propulsion/flight control system. Thus, the very feasibil- ity of speed control may be questionable. Even without speed control, ways to cope with upstart, minimizing thrust asymmetries (e.g., through center of gravity control, offsetting effecters, etc.) are not certain. 2.4.2.3 Aerothermoelastic and Slosh Mode Characteristics and Control The nature of the structural dynamic and slosh mode characteristics, how these are affected by temperature, and the implications for structural and slosh mode control are a first magnitude problem, especially for manned vehicles. Because flight control systems have been required to cope with structural and slosh modes for many years there is a tendency to treat this as a problem that will yield to available flight control technology. Things will be different with hypersonic aircraft. In the past, most flight control systems have relied on proper sensor placements and sup- pression of the lower frequency struc- tural and slosh modes picked up by the sensors using notch or low-pass filters. For large flexible vehicles that are flown statically unstable to maximize performance, and for which active con

STATUS OF HYPERSONIC TECHNOLOGIES trot of the lower frequency structural modes is needed for structural system optimization, the traditional approaches will no longer apply. Active control of the instability and of lower frequency structural/slosh modes will demand very large controller bandwidth. Lags appropriate for traditional "gain-stabil- ized" structural mode filters are inconsistent with the controller band- width needs and cannot be tolerated by the control system. Even if they were, such lags would also lead to piloting difficulties and poor flying qualities. (For example, relatively large effective time delays, partially due to structural mode filters, played a major role in the ALT 5 Shuttle PIO.) The net result for structural/slosh mode control is that the system may have to be "phase stabilized" in such a way that the net effective vehicle dynamics are appropriate for piloted control. This has never been done for manually controlled aircraft- existing specifications don't even permit it to be tried! Thus, the problem is completely different in both kind and degree from what one would casually expect as an extrapolation of existing technology. 2.4.2.4 Engine/Inlets/Diffuser, etc., Control The vehicle-flight and propulsion system controls have been mentioned above as a major overall integration problem. And even in cases where the engine can be looked at as an entity that is somewhat separate from the vehicle, the maintenance of reasonable operating conditions by the engine control system will not yield to a simple extrapolation of existing technology, and perhaps not even of existing control concepts. For example, while the control of thermal states throughout the vehicle is difficult, within the engine it is awe- inspiring! Close coordination of thermal 25 cooling and engine flows is essential, with even closer integration of other engine and thermal variables being a likely requirement. Since these vehicles will be so critically dependent on controls, it is important to consider some issues relat- ing to control system integrity. While highly reliable fly-by-wire control sys- tems are now standard, the level of integration of flight, aerothermoelastic, thermal, and propulsion controls required here poses a significant new dimension to the problem. Control system archi- tectures for these vehicles will be driven by the simultaneous requirements for tight local control of high frequency phenomena and coordinated action across the entire system to assure appropriate control integration. Hence the need for distributed hierarchical systems with life critical integrity. The basic principles for such systems are not established, proven architectures are not available, and only rudimentary design, evaluation, and validation tools are in place. As one would expect in the interdis- ciplinary arena of controls/aerodynam- ics/propulsion/thermal/structural dynam- ics interactions, the quantitative dimen- sions of many facets of these problems will be configuration specific. The problems themselves, however, are gen- eral, and the fundamental control system architectures available to solve them are relatively limited. For example, one can list perhaps a dozen specific overall vehicle dynamic phenomena stemming from interacting aerodynamic, structural, and propulsion sources that must be off- set or corrected by the controls in some flight regimes. These phenomena then define the fundamental control system architectural possibilities as well as the key sensing and effecting means. Yet, as we have noted, the definition, feasi- bility, and detailed consideration of technology needs pertinent to these architectural possibilities, associated sensing and effecting means, etc., appear

26 to have received insufficient attention. 2.4.3 Sensors Hypersonic vehicles place new requirements on sensors required for feedback control. In addition to the thermal requirements outlined above, entirely new classes of sensors are needed. The aircraft is, in effect, a flying inlet and outlet for the engine. Control of the flow is essential. This will require air data, pressure, and local flow sensing on a scale never before required. Unfortunately, such sensors do not exist. Probably the finest aero- dynamic data sensing system yet tested was the "Q Ball", a servo-driven spherical nose installation on the X- 15. It delivered angle of attack, side-slip, dynamic pressure, etc., over a very wide range of conditions - up to a maximum q of 2,200 psf (at about Mach number 5) and worked well up to a maximum achieved Mach number of 6.7 (at a q of about 780 psf). The Shuttle Entry Air Data System (SEADS) has been operated on one flight up to Mach number 27 but at very low density. When used as basic sensors for high bandwidth flight con- trol, these sensors normally will have to be multi-redundant. Even worse, heat transfer rates will increase as leading edge radius decreases. Considering the shapes of projected hypersonic vehicles, a compound problem of very high tem- perature and no room for equipment such as the "Q ball" will further complicate the sensor problem even before redundancy is considered. Measurements of mass flow, density, and three-dimensional velocity are needed at very high data rates. Species (H2O, O. OH, etc) measurements using various experimental techniques may be fruitful in these areas, but are yet to be proven in ground tests, much less on board. Even temperature sensing is a problem. Thermocouples are useful up to perhaps 1700° C, but installation on HYPERSONIC TECHNOLOGY FOR MILITARY APPLICATION the skin with wiring is a severe challenge. We have a long way to go in the sensor area and an ambitious R&D program will be needed in this area also. 2.4.4 Information System Requirements For one hypersonic vehicle concept, three airframe and engine contractors have made preliminary projections of throughput for the onboard information system that range up to 25-50 million instructions per second (MIPS). A reli- ability requirement in the 10-7 to 10-9 failures per hour range is likely to be necessary. Whereas the reliability requirement can be met with todays technology, obtaining the necessary throughput at that reliability in an aircraft-qualified installation is a considerable technical challenge. Figure 2-2 indicates how well some current systems can do. While some programs are aimed at very high throughputs, such as DARPA's strategic computing initia- tive, very little work has been done to provide that level of computing capabil- ity at the reliabilities necessary for this application. A system of this capability that can also grow by accretion of functions and equipment will have to be designed in parallel with the aircraft to assure that necessary system architec- ture can be built into the aircraft. Obviously, it will be necessary to develop the technology in time to do this. If a phased test program is employed, full availability of this system can be deferred. Its basic design, however, and laboratory proof of its viability must be established in time to make a commitment for defined hyper- sonic vehicle programs. 2.4.5 System Design Philosophy The uncertainties in the vehicle and propulsion system dynamics constitute a significant, perhaps even a governing, challenge for the development of hyper

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STATUS OF HYPERSONIC TECHNOLOGIES sonic vehicle control systems. These lead to a requirement for a very robust control system and perhaps sophisticated outer loop guidance capability. Through careful planning, including early deci- sions in the areas of systems architec- ture and on some of the subsystem physical factors, we believe some of the demands in these areas can be amelio- rated by the exploitation of a carefully phased flight test program. The full-up controls capability will not be necessary during early flight test in a program that slowly evolves toward hypersonic speeds, the full-up controls capability could evolve slowly during the flight test program. Most important among the attributes required for this approach is a control configuration which is at first as simple as possible but that can grow by accre- tion as the incremental test program proceeds. The most significant implica- tion of this is for the digital informa- tion systems architecture. From the very beginning, it must provide the necessary levels of safety, fault toler- ance, and reliability. But it must also be able to evolve into a much larger system without rewiring the aircraft, making any significant physical changes, or reworking the software. This means the system architecture must from the start be designed to be expandable while simultaneously providing the necessary levels of reliability. When such a system is in place, it can be used for fundamental flight control purposes at the beginning and then evolve toward a fully integrated engine, airflow, struc- ture control system. In the early stages of the flight program, for example, nominal trajectories and minimally compensated inner loop control functions would be used. More elaborate compen- sation and optimal control and guidance laws would follow when needed. As the program begins to reach the middle velocities where auxiliary constraints such as temperatures become important, additional control loops will be added as 27 needed. They would be added with good knowledge of the vehicle parameters and dynamics from previous flight tests. This would reduce the need for overly robust control systems that will be a challenging technical problem and should simplify the controls requirements for the vehicle. If during any flight, a parameter, such as a temperature limit, was reached, the flight speed would be reduced and perhaps that particular test terminated until a full understanding of why the limit was exceeded was obtained and the control law modified. This punctuated buildup and elaboration of controls consonant with an orderly phased flight test program leads to a full capability that probably will not be as elaborate as would have to be planned for if, the full capability was to be installed in the vehicle from the beginning of the flight test program. This advantage will have to be carefully balanced against the implications of cost and schedules to the test program should a problem arise that requires significant redesign and validation of the control system. 2.5 Materials for Hypersonic Vehicles 2.5.1 Introduction Structures for hypersonic vehicles must be lightweight, high-temperature resistant, inspectable, durable, and reliable. These requirements demand materials and concepts for structural design of hypersonic vehicles that are still in the emerging technology phase of development. Since different portions of the vehicle will encounter different pressures and atmospheric and temper- ature regimes, depending on the vehicle configuration and on the trajectory flown, we must consider a wide variation of potential structural concepts. These will be influenced by mission require- ments, desired performance, and the required payload.

28 Many considerations affect materials choices for hypersonic vehicles. Some of the more obvious: load-temperature- time cycle, vehicle reusability require- ments, interior environment require- ments, ant! the type of fuel and pro- pulsion system chosen. Materials- related factors that affect the lifetime of the structural elements are: resistance to chemical attack oxidation and corrosion resistance chemical stability as affected by alloy depletion at the surface and by diffusion at elevated temperatures interphase stability at high temper- atures between fibers and matrix, between coatings and contacted phases, etc. metallurgical stability, i.e., grain growth, over-aging, or precipitation during use and in thermal cycling creep high fatigue resistance low crack propagation rates, and designs that are inspectable, refurbishable, and damage tolerant. Temperature regimes range from cryogenic temperatures (-268° C) for fuel containment up to 1100-2200° C for leading edges, control surfaces, and stagnation points. Also, dynamic pressures up to 2000 psf may be exper- ienced, posing very stringent require- ments on both structural concepts and materials choices. For example, the wing leading edge is one of the most critical design areas because of the severity of the aerothermal environment. We must overcome such problems as localized heating, thermal gradients, connection to the airframe structure, aerodynamic loads, thermal stresses and distortions, and interface heating. A number of leading edge concepts exist such as radiatively-cooled, heat-pipe cooled, and actively cooled designs. Similar considerations apply to nose caps and control surfaces as well as to the fuselage. HYPERSONIC TECHNOLOGY FOR MILITARY APPLICATION At high altitude, atmospheric oxygen begins to dissociate above Mach number 7, with both oxygen and nitrogen almost fully dissociated above about Mach num- ber 15, and ionization above about Mach number 20. These real gas effects at the stagnation regions will require coatings to protect against oxidation, to form high emittance surfaces to lower surface temperatures, and to protect against erosion. Many factors will determine which materials are best for the various structural elements of a hypersonic vehicle. In briefings to this committee on vehicle concepts and configurations, large portions of the vehicle structure were described as stiffness critical, other areas were judged to be buckling or crippling critical, other areas strength critical. Some areas were designed by fracture toughness considerations, ther- mal fatigue, creep, etc. In examining these requirements, it becomes evident that a great deal of materials testing will be required to create an adequate materials properties data base. Many materials of interest are anisotropic and process dependent; these factors compli- cate and expand the amount of testing needed to develop a reliable data base. Finally, many of these properties must be obtained for temperatures from -268° C to 2200° C. The committee was briefed by the Air Force, NASA Langley personnel, and the airframe and propulsion contractor teams on the NASP program concerning material choices and the status of their developments. Centralization of the Materials Technology Maturation Program according to the type of materials under consideration is an important correct step to obtain the required data in a cost-effective and expeditious manner. It is important to realize that getting property data meaningful for design requires material specimens in the proper size, thickness, and finish that have been processed with a well-defined ped

STATUS OF HYPERSONIC TECHNOLOGIES igree. Most of the available data consist of tensile strength properties at room temperature and higher obtained on small coupon-type specimens and by three-point bend tests. Much of the data are not yet of sufficient quantity to be statistically significant for design purposes and emanate only from a single source. In some vehicle configurations described to us, the structural weight fraction necessary to achieve the per- formance goals set for the NASP re- search vehicle dictates that the vehicle be made largely of metal-matrix compos- ites, carbon-carbon composites, thermo- plastics, titanium, and titanium alumin- ides. A materials development program must include several related vehicle design inputs, such as stiffness vs. strength, minimum gauge requirements, acceptable creep elongation and fatigue life, and behavior in oxidizing and other chemical environments. Compared to the materials data required for subsonic aircraft, hypersonic aircraft materials data base requirements are much more extensive because of the need for thermal, chemical, oxidation and other data at elevated temperatures over specified time periods. 2.5.2 High Temperature Materials Potential system requirements dictate that the structural weight fraction of proposed hypersonic vehicles be much smaller than that of conventional sub- sonic aircraft. To satisfy the need for minimal structural weight fraction, one is forced to consider material concepts on the edge of emerging technology. High temperature materials are needed that maintain useful properties at the necessarily elevated working temper- atures; they must be high in stiffness, high in strength, low in density, oxi- dation resistant, have high fracture to ugliness, and creep resistance. In addition they must have several other desirable properties, including fabricabil- ity, joinability, ease of assembly, reasonable cost, reproducibility, and they must be available in sizes and shapes suitable for fabrication into final product sections. High temperature, lightweight structural materials of interest include the new RST high temperature aluminum alloys, high temperature RST and oxide dispersion strengthened titanium alloys, the titanium aluminide and nickel aluminide intermetallic alloys (TiAl, Ti3Al, NiAl, and Ni3Al), and metal matrix component alloys with reinforcing particulates, or fibers, of SiC and other refractory materials. The high temper- ature materials that are available in quantity and well characterized are the successfully applied nickel and cobalt base superalloys dispersion-strengthened with fine refractory carbides and oxides. The requirements for cryogenic fuel containment pose material needs that include light-weight super-insulation, organic-based graphite composite mater- ials for propellant tanks and internal structural applications and a variety of non-structural materials. Temperatures of interest go down to -240° C. The conventionally processed titan- ium alloy, Ti-6Al-4V, has strength to 300° C, and the RST-processed Ti (dis- persion-strengthened) alloys may sustain temperatures of at least 800° C. How- ever, titanium and its alloys oxidize above 550° C in long time applications, and are also susceptible to hydrogen embrittlement. This reactivity poses significant applications problems because hydrogen will be the primary fuel and will be used for regenerative cooling of parts of the structure. There is a need for the development of high thermal conductivity, low density materials (that are not embrittled by hydrogen and maintain adequate high temperature strength) for piping for regenerative cooling of the structure and parts of the

30 HYPERSONIC TECHNOLOGY FOR MILITARY APPLICATION propulsion system. A portion of the Materials Maturation program is devoted to this task. Materials under consider- ation include the titanium aluminides and copper based metal matrix composites. 2~5~2el The Aluminides For service above about 700-1000° C for relatively long times, there is much interest in the intermetallic compounds: Ni3Al, NiAl, Ti3Al, and TiAl. These basic materials offer significantly lower density (the lightest being the TiA1 phase), improved stiffness, high strength-to-weight, and improved oxi- dation resistance due to the high aluminum content, to operating temper- ature levels significantly higher than those of their metallic base alloy counterparts. However, these are intermetallic compounds and have poor ductility at room temperature and below, except for the modified Ni3A1 phase-alloyed with boron. They have relatively low strength at temperatures above about 800° C in the unalloyed state and have relatively low melting temperatures compared to those of the refractory metals, and the ceramic and carbon- based composite materials. For structural application at 700- 1000° C, the titanium aluminides show much potential. They are up to 50% lighter and have higher specific stiffness than the superalloys (see Table 2-A). Titanium aluminides have been proposed as matrix materials for SiC and other fiber-reinforced metal matrix composites, where the high temperature strength of the fibers offsets the rapid fall-off in strength of the titanium aluminide above 800° C. All of these materials are products of newly emerging - technologies with a limited data base for design, development and testing to ascertain the resultant properties. Although the high specific strength and low density evoke great interest for structural applications in the 700-1000° C range, the materials are inherently brittle at low temperature and efforts are underway to increase their ductility. Such improvements in low temperature ductility are mandatory if these materials are to be useful. Titan- ium aluminide honeycomb panels have been successfully produced by brazing. The key to titanium aluminide metal matrix composites is the fabrication of foil, which can be diffusion bonded around fibers such as SiC. Foil 0.003" thick has recently been produced with sufficient tolerance for this application. However, reproducibility of results requires considerable additional work, and higher temperature oxidization testing is urgently needed. Little is known about the interac- tions of the matrix and fibers and particulate reinforcements at elevated temperatures. It is likely that a barrier coating will be needed on the fibers to prevent degradation. Since these mater- ials are desired in minimum gauge struc- tures and in honeycomb panels, the problem of internal oxidation at elevated temperatures must also be addressed. Titanium aluminides and carbon- carbon composites are often suggested for higher temperature applications. RST titanium is specified but no infor- mation about properties j availability of shapes and structures, etc. seems avail- able. Finally, composite metallic materials, again largely based on titanium alloys, are specified as major components of the vehicle structure; mainly in SiC fiber-reinforced titarl~um and titanium aluminizes. Data are extremely limited; manufacturing methods are not well knowll; size and shape potential are in question. Vehicle designers appear to be relying on material producers to supply these materials and data. RSR titanium aluminide (RSR powder metallurgy) Is of

Superal 1 oys Ti Al 1 oy Ti 3A1 Ti A1 Density (lusting) 0.3 0.16 0.15 0.14 Modul us (x106 psi ~28 16 18 25 Max Temp: (° C) Creep 1000* 540 815 1000 Oxi Cation 1000 600 650 1040 . . *1095 °C for-ODS (or MA) alloys TABLE 2-A Ti tan i up Al umi n i des Property Compari son

STATUS OF HYPERSONIC TECHNOLOGIES unknown merit, especially for large structures. Few details are available on the powder metallurgy processes being used especially regarding oxygen con- tamination, powder characteristics, and the subsequent consolidation methods and conversion to structural shapes. Titanium aluminide foil material and honeycomb structures are specified but supporting data on product yield, sizes, methods of joining, shapes, and proper- ties are not available other than those quoted above. Two specialty alloys, MA 956 and MA6000E, mechanically alloyed, nickel base, oxide dispersion-strength- ened alloys are in production and used in aircraft. They have a fair data base, but we have limited data on fabrication, performance, oxidation, resistance, etc. A general criticism of the status of most composite materials, and particu- larly of the aluminides is that mechan- ical test data are extremely limited, and, in many specific cases, unavailable. Most of the data are short-time tensile data; few data are available for lives of 100 to 1000 hours at high temperatures in appropriate atmospheres. Crack growth data are generally unavailable. Longtime tests at maximum temperature are unavailable. Choices of materials are based largely on inadequate data and optimistic projections. Although titanium aluminide appears to be the choice of some airframe designers, others depend heavily on carbon-carbon composites. Most test work has been with carbon-carbon (including small panels). There has been very little work with the titanium aluminides and there has been very little material to work with. In this respect the use of carbon- carbon may be somewhat more realistic, particularly at the higher temperatures. The joining methods with carbon-carbon composites and metal matrix composites are a serious concern and as yet not much work has been done. The NASP technology maturation program is aimed at improving materials 31 through conventional alloying, by apply- ing RST principles, through oxide disper- sion strengthening and by application of the aluminides as the matrix for com- posite structures. Since the basic aluminides are relatively new structural materials, their properties and those of the alloys that may be developed from them are largely unknown. The aim is to produce alloys capable of operation up to about 1000- 1 100° C while main- taining fabricability, formability, and while retaining oxidation resistance for thousands of hours, and exhibiting suitable strength, toughness, etc. Some properties of titanium alumin- ide materials are compared to conven- tional nickel base superalloys and titanium alloys in Table 2-A. Improve- ments in high temperature strength using metal matrix composites are shown in Figure 2-3. 2.5.2.2 Superalloys The superalloy compositions and structures are based on extensive and excellent performance records for numerous applications where the maximum operating temperatures are about 950° C and life is in the tens of thousands of hours. Current nickel base superalloys have been developed for high strength and long time service in turbine combustion atmospheres with performance far in excess of any anticipated demands. Melting at temperatures at about 1300° C, the nickel base superalloys have operated at temperatures up to 1000° C, uncooled, and higher with cooling, or in excess of 80% of the absolute melting temperature. For increased high temperature performance, these alloys have also been prepared as directionally solidified (DS). coarse Brained, precision cast blades, and as single crystal (SC) alloys with preferred crystalline orien tations. The literature and prior

32 experience are abundant for temper- ature-stress-life time applications of these materials for a broad range of operating conditions. Further, a number of coatings have also been developed of the MCrAlY type where M is generally Ni, Co, Fe, or combinations of these elements and the Cr. A1, and Y are as indicated. These coatings have excellent life performance and are routinely applied as needed. The major negatives of the super- alloys are low melting temperatures and a relatively high specific gravity. 2.5.2.3 Dispersion Strengthened (Stabilized) Alloys Where oxides are used as the disper- soids, the resultant alloys are ODS (oxide dispersion stabilized) alloys; where carbides are the dispersoids, the alloys are carbide dispersion strengthened (CDS). Only refractory oxides are inert in most metallic matrixes, and the oxide does not contribute to the strength through interaction with the matrix; carbides and other intermetallics are generally wetted by the metallic matrix and therefore can contribute directly to strengthening through resultant bonding forces. The ODS alloys have reached a state of development where significant increases in applications are being achieved. (See Figure 2-4.) Oxide volume contents from as little as 0.5% to as much as 10% have been used in both pure metal bases and complex alloys such as the commercial nickel base alloy IN-100 (Ni-Co-Cr-Mo- Al-Ti- V-B- Zr-C). Strength generally increases with increasing oxide volume content, as do the elastic modulus and hardness, whereas formability, ductility, and toughness decrease. The critical structural parameter is the interparticle spacing among dispersoids, making it HYPERSONIC TECHNOLOGY FOR MILITARY APPLICATION desirable to use exceedingly fine disper- soids (0.01 to 0.2 microns) to achieve high strength levels, long time stability at very high fractional melting temper- atures, combined with demonstrated practical formability, useful ductility, and toughness. With these very fine dispersoids one needs only 1 to 2 volume percent of oxide to achieve the desired high temperature stability and strength- ening. For still greater stiffness, higher volume contents of oxides can be added (10-30%), with further benefits in strength but loss of ductibility. Dispersion strengthened (as well as composite alloys) are strengthened through the stored energy of cold work in processing. The very fine inter- particle spacing pins high levels of dislocations (cold work); and the avoidance of significant alloy recovery or recrystallization preserves the strengthening benefits effectively almost to the melting temperature of the alloy. Recent studies have now established preferred alloy processing methods to incorporate the fine oxide particulates reproducibility; in fact, several process- ing technologies are being used success- fully. Highly successful ODS materials have been reported based on Mg, A1, Ag, Cu. Fe, Ni, Co, Pt. stainless and heat resist- ant alloys, the superalloys, titanium, and others. A lloyed matrixes are used to achieve selected properties, such as corrosion or oxidation resistance, and high strength, but pure metals are also used effectively for high strength and high conductivity as with Cu. Ag and A1. Refractory carbides are also of interest for selected CDS alloys. Because the carbides are rarely inert in n~ost metallic systems, strengthening is often enhanced by bonding reactions between the very fine carbide disper- soids and the metallic matrix. For maximum alloy stability at the h ighest

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STATUS OF HYPERSONIC TECHNOLOGIES possible temperatures, only the refrac- tory carbides such as HfC, ZrC, and SiC are of interest. As a result of such bonding, strengthening is achieved and retained at low or high temperatures even after recrystallization, unlike the case for the ODS alloys. At very high temperature, say, above about 0.6 of the absolute melting temperature, strength properties fall off more rapidly than for the ODS alloys. Refractory carbides, with much lower values of heat of formation than refractory oxides, tend to be coarser than the oxides. Further, the ultrafine carbides, if prepared separately for subsequent blending interactions, may form oxide films that will modify the expected interactions. Thus, carbides tend to be coarser than oxides, and are always used in larger quantities, for example, from 10-25 volume percent for maximum strengthening. Combinations of fine oxides at low volume content with fine carbides at high volume content recently have been shown to provide interesting and useful combinations of properties, each accord- ing to its interactions with the matrix. For specific applications intermetallics other than the carbides can be used effectively to achieve other combinations of properties; these include the refrac- tory borides, nitrides, silicides, etc. Relatively little developmental work has been done with higher melting alloys than those of Fe, Ni, and Co. Platinum has been dispersion stabilized with ThO2 and showed very large increases in strength in long time tests at 1300° C. Rupture strength values for lives of 100 tO 1 000 hours were 10 to 20 times greater than for pure platinum and four times greater than for the best known alloy, namely Pt-40% Rh. The upper temperature limit for ODS alloys has not vet been determined. In terms of oxidation and certain corrosion phenom- ena, these precious metal ODS alloys 33 offer some interesting properties. Minor efforts have been undertaken with Nb, Mo, Ta, and W as high temper- ature ODS alloys. Molybdenum has been dramatically strengthened to 1600° C by fine dispersions of HfC, produced by conventional melting technology. Powder metallurgy blending should result in higher temperature performance with ODS alloys, with carbide dispersion strengthening of combinations of both. 2.5.2.4 Carbon-Carbon Composites and Ceramic Composite Materials For temperatures to 2200° C, carbon-carbon composite materials have shown excellent properties and perfor- mance on the shuttle leading edges and nose. New, improved manufacturing methods continue to be developed. Sec- tion sizes of interest for airframes are being fabricated. Carbon-carbon com- posites show excellent combinations of properties: very low specific gravity, high modulus, low thermal conductivity, improving strength at high specific levels with increasing temperature, excellent thermal and mechanical shock resistance and toughness Oxidation is a problem but rates are not catastrophic, as with some of the refractory metals in an oxidizing atmosphere; however, coat- ing will undoubtedly be required for long life performance. At somewhat louver temperatures, from about 1 4C0° to ~ 9G0° C, ceramic composite materials show promise. Manufacturing technologies are being developed, and reinforcement with refractory and carbon fibers look promising. C)xidat~oll r esistance is excellent; strength levels are attractive; but thermal stresses and thermal and mechanical shock are more severe than with the carbon-carbon composites. Specific gravity, specific strength, and specific modulus Values are all excellent.

~4 Because of their high temperature capability, favorable properties' and low density, carbor~-carbon composites repre- sent enabling technology for advanced propulsion systems and hypersonic vehicle structures. These systems have a spectrum of temperature requirements in uncooled nose sections, leading edges of infers and wings in hypersonic air- craft design, from 1630° to 2200° C. The current generation of superalloys is limited to approximately 1000° C. With improvements now being developed, such as directional solidification, oxide dispersion strengthening and RST, use temperatures as high as 1100° C are probable. However, to operate at 1400° C and higher, carbon-carbon and ceramic composites, ceramics, and coated refractory-based alloys are the remaining viable materials. To date much of the mechanical property data for design of carbon- carbon has been obtained on test cou- pons at elevated temperatures, although other properties of the material may be more critical because of the specific structure] concept. The properties of carbon-carbon composites depend strongly on the processing history of the material. The thermal expansion, tensile strength, bimodularity, shear defor- mation, and notch sensitivity are all interrelated and depend strongly on the processing of the composite. During manufacture, crease cracks and waviness sometimes appear. These are caused by thermal stresses arising from the heating of the preformed structure to the graphitization temperature and occur because of the anisotropy of both the thermal expansion and the elastic modulus. Small changes in the con- stituents, geometry, or processing might lead t,-, ~ rash of failures unless the mechanisms of failures are understood and processing variables are very ca~-et~uliy controlled. Defects in tl~e composites can affect the performance of tl~e end product; the causes and the signif icance of Avoids and delaminat~ons~ HYPERSONIC TECHNOLOGY FOR MILITARY APPLICATION in particular, must be identified. Such defects can originate during one or more of the processing steps, e.g. during pre- impregnation lay up, impregnation, and high temperature pyrolysis and graphiti- zation steps. The most fundamental cause of defects in the high temperature processing phases is the thermal expan- sion anisotropy of crystalline graphite parallel and perpendicular to the basal plane, with variation of thermal expan- sion in different directions in the composite, producing high stresses at elevated temperatures, subsequently giving rise to structural flaws. Another aspect of the thermomechanical behavior at high temperatures that should be examined is the potential for stress relaxation by creep, particularly above 1650° C. All carbon-carbon composites in hypersonic vehicles will encounter oxidizing environments, and the single most critical factor that could limit the use of such materials for structural applications is reliable oxidation protection. They will need coatings to protect them from oxidation. Long term oxidation protection of carbon-carbon composites at temperatures over 1650° C will require coatings and material combinations different from those used successfully on the space shuttle. Outer coatings of SiC and Si3N~ cannot be used at temperatures over 1900° C because reaction products disrupt the SiO2 layer necessary for oxidation protection. Oxide coatings may have to be used. Criteria for selection of oxide coatings include melting point, vapor pressure, thermal expansion, and reactivity. Fundamental issues ot thermodynamics and kinetics limit the material choices to a small number of oxides. Joining of carbon-carbon composite panels presents major problems. Fa~- tener openings must be protecte~i from oxidation. Fastener materials nest have the same coefficients of expansior1 a::

STATUS OF HYPERSONIC TECHNOLOGIES the composite, since any protrusion above the surface could produce a shockwave causing very high heating at the shock interface. Gap fillers will be needed (as do the shuttle tiles). Tolerances on thin carbon-carbon skins will be very close since the fasteners cannot be used to form the skins to the substructure as is done with metals, and shims cannot be used. The carbon- carbon material may be limited to a relatively low operating strain. Carbon- carbon composites are weak in shear, exacerbating the problem of load transfer through the fastener assembly. Also, if the processing of the material is not rigorously controlled and the graphite fibers are not uniformly bonded to the graphitized matrix, hysteresis in thermal expansion and contraction can occur, varying from pane} to panel thus complicating the thermal-structural analysis. Major use of carbon-carbon com- posites can be specified with confidence, only after several technical issues are resolved and understood, including the best fibers to use and the relation of the matrix-fiber interface to properties, the best precursors, the processing steps that will minimize cracks and defects, and a detailed data base on the proper- ties and behavior of carbon-carbon composites in the intended applications. 2.5.2.5 Processing Developments for Advanced Metallic Materials Since about 1950 a significant number of processing technologies have beers developed for incorporating fine refractory particulates into metallic matrixes. Recently, Benjamin and co- workers at the International Nickel Company, developed a mechanical alloy- ing process for blending ultrafine Y2O3 into nickel base superalloys. Following hot extrusion of the metallic master alloy powder m ix with the Y2O3, the a] toys are zone anneal -processed (ZAP), 35 undergoing directional recrystallization and extreme grain gr(3wt;h of the highly cold worked extruded product just below the melting temlie' bout c;. This process produces an oriented, elongated very coarse grain structure that has extended the operating temperature of the nickel base alloy from about 900 to 11 00° C without loss of creep resistance, a phenomenal improvement in high tem- perature operating pert ;~rrnance. A simpler technique by Smith and co- workers at MIT simplified the overall processing developed by Benjamin and achieved similar properties at 1 100° C. At 1100° C, at high engineering stresses and long life, the resultant ODS nickel- base superalloy recrystallized to an oriented, very coarse "rained structure and showed operating capability at almost 90 percent of the melting tem- perature (absolute) of the basic nickel base alloy. The Pt-ThO2 alloy discussed above showed outstanding 1300° C prop- erties at greater than 80 percent of the (absolute) melting point of pure Pt. The upper temperature limit for high strength applications has not been reached as of now. The processing technologies required for the production of ODS and CDS alloys are important. The results recounted above demonstrate the tre- mendous strides made in increasing the temperature potential for alloys embody- ing such structures. In particular, the processing tech- nologies currently available make it imperative to apply the same processes to the refractory metals and their alloys: Nb, Mo, Ta, and W' with emphasis on Nb and Mo and their selected alloys. Because of their relatively low specific gravities, very high useful temperatures, and outstanding elastic modulus in the case of molybdenum, the potential exists for high strength operation at 1400° to 1750° C. This would mean transcrystal- line deformation and fracture modes (relatively much more ductile titan in the

36 HYPERSONIC TECHNOLOGY FOR MILITARY APPLICATION intercrystalline failure mode) which would result in increased ductility. The more simple niobium (molybdenum) ODS alloys might indeed also be ductile near room temperature (Iow ductile to brittle transition temperature compared to molybdenum.) 2.5.2.6 Other Developments The development of the Osprey and Liquid Dynamic Compaction (LDC) "spray atomization and collection" processes has opened new potential for producing bulk, complex, reactive metal products and shapes. By combining rapid solidification processing (1000°C/s) to minimize segre- gation during solidification, and minimiz- ing exposure to the atmosphere, the alu- minides, in particular, should benefit on both counts. This also would be true for the refractory metals and alloys; however, the problem of extremely high melting temperatures vis-a-vis atomiza- tion has not yet been adequately ad- dressed and deserves attention. The still more recent development of fine solid particle (oxides, carbides, etc.) injection into an atomized liquid metallic spray offers unusual potential for inject- ing fine particulates with uniform dis- tribution during production of sheet, strip, plate or preform bodies while achieving high as-deposited densities, typically 96+ percent, in conventional metallic alloy deposition. Very fine structures, relatively free of segregation, and at low contamination levels are indicated to be quite practical. Sup- plementing the usual Osprey and LDC processes, a linear atomizer has been patented and used to sharply increase rates of deposition and to make possible the production of large, flat, uniform cross-sectional sheets, strip, and plate. The LDC production of dispersion strengthened alloys with very fine dispersoids, oxides, carbides, borides, etc., at low volume content ( 1 to 3 volume percent) to produce ductile, tough strong alloys, or metal matrix composites (oxides, carbides, borides, etc.) using coarser dispersoids and high volume content (10 to 30 volume per- cent) appears feasible and should be studied, in particular because bulk shapes can be produced at high density, convertible to full density by hot working or hot isostatic pressing (HIP). 2.5.3 Coatings Coatings will play a large role in insuring stability of the major structural materials at the high temperatures that hypersonic vehicles will be subjected to in flight at the high Mach numbers. In the case of compressor and turbine blades of modern jet engines, coatings are mandatory to insure adequate life. In previous applications of nickel base alloys to thin gauge honeycomb face sheets of high temperature structures, thermal cycling for short periods of time to 1100° C causes internal oxidation to occur. Data on several nickel base alloys thermally cycled 10 times for one hour each time to temperatures of 1200° C indicate that internal oxidation will occur from both sides of a minimum gauge sheet using up a good portion of the load carrying capability of the material. (See Figure 2-5.) In the case of the aluminides an external coating may -be needed as well as a barrier coating in the fibers for the metal matrix composites to prevent inter- diffusion. Simple binary diffusion couples can be used to determine optimum coating materials for these applications. Extended use of carbon-carbon com- posites at elevated temperature is limited by the extent to which they can be pro- tected from oxidation. Oxidation pro- tection has concentrated on the use of exterior coatings such as silicon nitride and silicon carbide. However, their use alone is not sufficient to protect the carbon-carbon substructure against

STATUS OF HYPERSONIC TECHNOLOGIES oxidation at some of the temperatures to be encountered. One reason for this is the thermal expansion mismatch between the substrate and the coating. Carbon- carbon 2-D composites have low in-plane and high through-thickness thermal expansion. During cooling from the processing temperature, thermal stresses are generated that craze the coating. Defects in the coating, such as pinholes and thin spots, lead to premature degradation of the substrate, and the steady state and dynamic fatigue loads of flight can cause additional cracking. To overcome this problem the use of glass-forming sealants based on silicon dioxide that will flow at the temper- atures encountered to seal cracks and pinholes has proven very advantageous. This sealant remains intact both chemically and physically over a wide range of environments. It has proven very effective on the nose cap and leading edges of the shuttle. However, for the times at the contemplated highest temperatures of hypersonic flight most coatings will either melt, evap- orate, or decompose. For example, some of the carbon-carbon composite structure will encounter temperatures of at least 1900° C, at pressures of 0.1 atmosphere or less. Under these conditions, a silicon dioxide sealant could react with the silicon carbide coating, forming silicon monoxide, which at this pressure and temperature can boil off as a gas. Hence, barrier coatings or different sealants are needed. 2.6 The Structural Challenge The challenge confronting the struc- tural designer of a hypersonic vehicle is reflected in the structural weight frac- tion, which is the weight of the whole structure divided by the take-off gross weight (TOGW). System requirements and propulsion systems presently envisioned require fuel fractions in the neighborhood of .75 for single stage to orbit. Some design studies are predicat 37 ing structural weight fractions in the vicinity of . 15. This leaves .10 of the TOGW for the payload and all the sub- systems which make the airplane a use- ful system. As will be demonstrated with historical data, a structural weight fraction of .15 requires a giant step function increase in the capabilities of materials and structural concepts. If we assume the payload fraction is .05, then a 10,000 pound payload requires an airplane with a TOGW of 200,000 pounds. If the structural weight frac- tion increases by .01 at the expense of the payload fraction, i.e., the payload fraction decreases to .04, then the TOGW increases to 250,000 pounds to carry the same 10,000 pound payload. Thus a 7 percent increase in the struc- tural weight fraction (.01 /. 15) in the preliminary design stage with fixed payload results in a 25 percent increase in TOGW. Contrarily, a 7 percent increase in the structural weight fraction for an airplane with a fixed TOGW of 200,000 pounds means a 20 percent decrease in payload. To return to the question of feasible structural fractions, summary weight statements are shown in Table 2-B for the shuttle orbiter and for the C- 141 military logistics transport. Both use mostly aluminum alloys for their struc- tures. The orbiter can use aluminum alloys because a passive thermal pro- tection system maintains the temperature of the aluminum structure at acceptable levels. The first four items, wing, tail, body and alighting gear, comprise the struc- ture of the vehicle. Note that these add up to a structural weight fraction of .297 for the orbiter and of .273 for the C-141. However, the thermal protection system of the orbiter should be con- sidered as a part of the structural weight fraction; when this is added to the first four items, the structural weight fraction becomes .409. The fuel

38 HYPERSONIC TECHNOLOGY FOR MILITARY APPLICATION fraction for the C-141 is less than one- half of that required for a single stage to orbit hypersonic vehicle. It should also be observed that airplanes must have important sub-systems such as propulsion, electronic, air conditioning, etc.: the designers of these sub-systems will demand their fair share (in their view) of the total weight fraction. Structural weight fractions of other transport category aircraft such as the L- 101 ] and the 747 are approximately .30. Transport category aircraft are designed to maneuver limit load factors of 2.5, gusts in the atmosphere may induce slightly larger load factors. Structural weight fractions of fighter type aircraft such as the F-15 and F-18 are approximately .32; these aircraft are designed to maneuver limit factors in excess of 7. The smallest structural weight fraction of a vehicle is .1 1 for a supertanker (to this committee's knowl- edge) and that is of approximately 1200 million pounds! 2.6.1 Determination of Structural Mass Fraction for High Performance Vehicles Several different criteria are used to evaluate the optimum design concepts for various portions of the vehicle based on its configuration, the trajectory flown, the aerothermal environment en- countered, and a number of other fac- tors. Portions of the vehicle may be stiffness critical, others buckling or crippling critical; some areas may be designed by fracture toughness, strength, thermal fatigue, creep, compression, oxidation, etc. To determine the structural mass of the component to be evaluated we must note the probable failure modes that the component could encounter. The mass of the component is distributed according to the primary failure modes that size the structure for the vehicle under consideration. In this procedure, the structural mass of the component is distributed to prevent fail- ure in the critical (primary) failure mode. These primary failure modes determine the local thickness and stiff- ness of the structure, and thus "size" the structure for that particular type of vehicle. The durability and damage tolerance allowable is determined by analysis of fatigue, crack growth, fracture toughness, and stress corrosion or oxidation for each material. When these factors are taken into account together with gauge tolerance and other considerations of fastening and joining, a more realistic structural weight fraction is then obtained than that which is predominantly a function of strength and stiffness. 2.6.2 Structural Concepts Hypersonic vehicle performance depends intimately on such structural concepts as material choice, active/- passive cooling, insulation, thermal protection, thermal/structural behavior, configuration, fabrication, and manufac- turability. Vehicle structural concepts depend on the mission profile, e.g., a high-Mach cruiser with frequent use and short turn-around times will have dif- ferent requirements than an orbital vehicle used frequently or a hypersonic reentry vehicle used only once. The technology base for structures running for sustained periods at elevateat temperatures needs to be developed. The mechanics of hot structures is not well understood in such areas as vehicle leading edges, vehicle nose or lip regions, airframe surfaces including control surfaces, internal load-carrying structures, or local regions of severe aerothermal loads. Leading edges, for instance, are a critical design area because of extremely high localized heating, high thermal gradients and stresses, coupled with aerodynamic loads,

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ORBITER ORBITER C-141 C-141 . WEIGHT FRACTION WEIGHT FRACTION COMPONENT Pounds %of TOGW Pounds %of TOGW Wing 14,718 0.06134,350 0.109 Tail 2,891 0.0125,907 0.019 Body 41,540 0.17234,523 0.110 Alighting Gear 12,569 0.05210,935 0.035 Surface Control 2,255 0.009 3,652 0.012 Propulsion 34,027 0.141 25,212 0.080 Aux. Power Plant 3,941 0.016 523 0.002 Hydraulics 2,293 0.009 2,689 0.009 Electrical 4,041 0.017 2,683 0.009 Electronics 5,063 0.021 2,871 0.009 Furnishings 2,130 0.009 5,017 0.016 Air Conditioning 3,649 0.015 2,537 0.008 Thermal Protecti on 27,154 0.112 0.000 Crew 1,252 0.005 2,756 0.009 Fuel 23,856 0.099 111,480 0.354 Payload 60,000 0.249 70,000 0.222 J TOGW 241r379 1.000 315,135 1.000 TABLE 2-B Summary Weight Statements for Shuttle Orbiter and C-141

STATUS OF HYPERSONIC TECHNOLOGIES _ which together prevent a safe, reusable, 4) reliable, light leading edge. Active cooling is also required, with attendant technology development required for hypersonic vehicles. An indication of the thermal environment that must be sustained in hypersonic flight is shown in Figure 2-6. Multiple materials are necessary for local conditions on the vehicle. Active cooling is essential at the higher thermal fluxes that are characteristic of the ~ · ~ Hypersonic regime. Specific structural technology base requirements for hypersonic flight are: 1) Vehicle leading edge - Analysis tools are needed for active cooled edges (convective, transpiration), heat-pipe cooled edges, radiatively cooled edges, or for other alternate means of leading-edge cooling. Verification of these analytical procedures is necessary through testing at suf- ficient scale before the technology is ready for design or development. 2) Nose or lip regions - Again, analyt- ical tools are necessary for active, passive, or reactive cooled noses or lips. A durable light structure is desired. Verification through testing beyond proof-of-concept experiments is important before the technology is ready for development purposes. Airframe surfaces and control surfaces - Hypersonic vehicles will run hot. Local temperatures will often exceed 1650° C. Active cooling everywhere is not practical. Structural element (spar, rib, stiffener, panel, etc.) technology at elevated temperatures needs develop- ment. Testing is necessary to validate structural concepts with and without thermal protection and for various cooling concepts and joining concepts. 39 Internal structure - Practical means need to be developed to design internal structures which can carry significant loads while running hot, or for cold internal structures that attach to hot structures. There is a technology requirement to develop procedures to conduct simultaneous thermal, structural, and dynamic analysis. Again, verification through component testing where aero- dynamic and thermal loads are imposed is necessary before the technology is ready for design or development purposes. 5) Local aerothermal loads - Intense local loads often drive structural concepts. A comprehensive data base and analytical tools need to be developed to allow durable and safe structural concepts at minimum weight. 2.6.3 Development Process Figure 2-7 illustrates the develop- ment process for producing structural concepts for a hypersonic vehicle with specified mission/performance goals given an adequate technology base. Testing of structural concepts at the component level and assembled level in a realistic thermal/structural environment would be a large part of the development process. The result of this development procedure would be critical structural hardware experience. Presumably the structural concepts developed are designed to real- istic criteria, ultralight, temperature resistant, and durable. The next iog,ical step to advance a technology toward use as structural hardware is often discounted This step is fabrication or demonstration of manufacturability Very often, high technology en~leas~ors (such as the development of hypersonic vehicles) stretch-out or fa ~1 because of lack of attention lo the industry which must

40 produce the parts. Successful structural concepts for hypersonic vehicles will require careful development of com- ponent fabrication processes, manufac- turing process methods and procedures, fastening technology, joining methods, assembly techniques, inspection methods, sealing technology, methods of repair or rework, and maintainability consider- ations. A lack of technological maturity in any of these areas leads to incom- plete structural hardware experience. Successful activity in these areas would in principle lead to structural concepts that are not only ultralight, temperature resistant, and durable but also pro- ducible, maintainable, and repairable. The final step of the development process is flight testing to assess safety-of-flight, operations, support- ability, and vehicle performance. How rugged, how reliable, how accurate, and how operationally useful the hypersonic vehicle is would follow from this flight test program. 2.6.4 Status of Structural Concepts for Hypersonic Vehicles The NASP program has an aggressive Technology Maturation Plan to develop the technology base to design an exper- imental hypersonic X-plane. Prior to this NASP program, hypersonic technol- ogy base development had been spotty at best for 20 years with the exception of technology for reentry vehicles and the shuttle orbiter. After careful study of those elements of the Technology Maturation Plan directed at structural concepts, we conclude that the projected technology base has generic as well as design specific validity for hypersonic vehicles. The Technology Maturation Plan should result in a structural data base valid for hot structures a few feet by a few feet. There probably are adequate test facil- ities to provide a credible thermal HYPERSONIC TECHNOLOGY FOR MILITARY APPLICATION structural loading of such test articles applicable to some NASP components. We are, however, concerned about the lack of large scale test facilities required to validate components at full- scale. Of concern is the ability to adequately test a wing-fuselage carry through section, a tank-fuselage section, or an actively-coolec! full airframe section. It might be impossible to adequately ground test a full-size representative section in a combined aerothermal environment. Such testing can be done only in flight. The attend- ant technical risk in proceeding to flight with only moderate confidence in the flight vehicle design is accordingly high. We are also concerned about the eventual feasibility of producing hypersonic vehicle structures that are durable and easy to maintain, and economical in small or large quantities. The design and fabrication of actively- cooled structures must consider factors such as regenerative cooling channels' piping, valves, etc. Such plumbing must be considered as part of the structural weight fraction. The high-temperature materials for hypersonic aircraft may require great care to process, fashion or maintain. It is worth considering mul- tiple production sources of key vehicle sections. A good deal more emphasis must be placed on production issues before judgments of the operational utility of hypersonic vehicles are approached. 2.6.5 Structural Design Considerations The attainment of routine hypersonic flight places before the structural designer a host of new design conditions which accompany the large amounts of thermal energy generated in the bound- ary layer. An air-breathing aircraft passes through flight regimes from subsonic through supersonic to reach hypersonic, all in two to three n~inu-tes.

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STATUS OF HYPERSONIC TECHNOLOGIES Stagnation temperatures quickly reach 500°+ C at low altitudes and 2200° C at high altitudes. The common aircraft aluminum alloys, which melt at about 600° C, become worthless as structural materials at about 400° C. Thermal stresses are induced by the transient nature of the flight profile. For example, a steel alloy wing structure may undergo thermal stresses of 75,000 to 150,000 psi during the transient temperature gradients induced as the aircraft attains hypersonics cruise if specific design details cannot minimize the adverse effects of thermal gradients. If all the structural parts reach 550° C, a ten-foot-span wing of stainless steel would expand over an inch in length. Internal push-pull rods and hydraulic lines might differ in temperature by several hundred degrees, and, in length, by intolerable amounts. These are some of the simpler thermal phenomena that confront the structural designer. 2.6.5.1 The Structural Designer's Approach, Circa 1988 It is informative and instructive to consider courses of action that the structural designer can take. Several courses of action are apparent. The designer would naturally explore first the possibilities of using new materials. Obviously, if a material were available that had a coefficient of thermal expansion comparable to pyrex glass, and had elevated temperature strength properties comparable to those of 7075-T6 at normal temperatures, then the thermal problem would be trivial. Strength properties are important but such factors as corrosion resistance, susceptibility to thermal stresses, maturity of the manufacturing infra- structure to translate materials into structural forms and availability are of equal significance. Historically, the record shows that many years are required to gather a data base sufficient 41 for design and to bring the manufactur- ing and material producing infrastructure to an acceptable level of maturity. Good properties at high temperatures necessarily require high melting temper- atures. Melting point data may, however, be misleading. For example, the available titanium alloys do not fulfill the promise that the high melting temperature of pure titanium would indicate. Thus, at 200° C, which is approximately one-third of its melting point, 2024-T4 aluminum still retains about 80 percent of its normal temper- ature tensile yield strength. At a comparable percentage of its melting temperature, i.e., at about 550° C, titanium has only about 25 percent of its normal temperature strength. The high temperatures have other effects that must be anticipated. Thin · ~ s On monocoque or semi-monocoque con- struction which is not free to expand can fail by thermal buckling. Thick skin construction leads to larger thermal stresses because of the larger temper- ature gradients. The structural designer can also devise preventive measures. For example, the load-carrying structural members can be covered with a passive thermal protection system (TPS) such as that used in the orbiter. A passive thermal protection system can a) reduce the rate of heat flow into the structure and, hence, b) Increase the time neces- sary for the load carrying structure to reach a given temperature. Since the insulation material has a certain heat capacity, it can also serve as a sink for a portion of the thermal energy. Problems accompany the use of ~ passive thermal protection system. If the coating is bonded tightly to the structure, then the strains of the coating will be transferred to the structure, but, more seriously, the coating will be unable to expand freely

42 and may crack. This suggests the pos- sibility of applying the insulation in a "fish-scale" or"roof-shingle" pattern. In this manner, the insulation is free to expand. Another possibility is to "float" the structure in an insulated shell. The transfer of air loads from the shell to the main load carrying members could be effected by a network of posts, and if these posts could be arranged so that the loads are predominantly compression, then glass or ceramics can be used. The structural designer can also devise ways to cool the structure and in this manner use the best properties of the materials. Cooling the structure involves many more complications. The cooling of the structure requires free access of the cooling agent; at this moment, invention is required. Trans- piration cooling offers another possibility if a suitable porous material can be created. 2.6.5.2 The Aerodynamic Heating Phenomena A proper assessment of the effects of aerodynamic heating requires an understanding of the source of the thermal energy. During flight a thin boundary layer, in which the air velocity drops from flight speed to zero, envelops the aircraft. We understand the con- sequences of the boundary layer phen- omena for low speed flight, but at higher speeds the dissipation of energy within the boundary layer represents an important additional factor. In fact, the degradation of mechanical energy into heat is approximately proportional to the square of the velocity gradient. As heat is so produced near the surfaces, some conduction into or out of the surfaces is to be expected. Thus, there is also a thermal as well as a momentum boundary layer within which convection, con- duction, and dissipation are balanced. HYPERSONIC TECHNOLOGY FOR MILITARY APPLICATION Two parameters of chief interest to the structural designer are the recovery factor and the heat transfer coefficient. The recovery factor relates stagnation temperature arid free-stream temperature to the adiabatic wall temperature. Once this number is known, the adiabatic wall temperature can be found. This is the temperature at the surface of the object if the object is insulated against heat transfer. Having found the adiabatic wall temperature, the problem is one of convective heat transfer from the boundary layer to the structure. Here we find the biggest hiatus in present knowledge. Some information exists on the heat-transfer coefficient for steady heat flow but much remains to be done before the transient problem can be solved. The magnitude of the recovery fac- tor ranges between 0.80 and 0.95. It depends chiefly upon the nature of the flow, i.e., laminar or turbulent, on the Prandtl number (a dimensionless combin- ation of specific heat at constant pressure, viscosity, and thermal conduc- tivity) and upon the shape of the solid boundary to the flow. The heat-transfer coefficient is the constant of proportionality which relates the heat flow into the structure and the temperature difference between the structure and its recovery temperature. It is conveniently combined with a char- acteristic length and thermal conductiv- ity into a dimensionless number called the Nusselt number. The Nusselt number varies with Reynolds number, Prandtl number, shape of the solid boundary and the nature of the flow, i.e., laminar or turbulent. All of these dimensionless parameters depend on the properties of the fluid, and these, in turn, are all functions of the temperature. The exterior surface of the structure is a boundary of the flow across which there will be interaction between condi- tions in the structure and conditions in

STATUS OF HYPERSONIC TECHNOLOGIES the boundary layer. Heat transfer in the boundary layer depends on the tem- perature at the surface of the structure. As the surface temperature rises, the gradient will change, and the rate of heat flow into the structure will change. If the flow conditions are not changing with time, the necessary trial procedure in determining the correct balance between heat transfer in the boundary layer and heat flow into the structure may not be too difficult. If the flow conditions are transient, then the trial procedure becomes quite cumbersome. A necessary objective of research would be to define the degree of interdependence, but it is likely that there will be important flight configurations in which the interaction cannot be neglected. Again, the structural designer here needs certain detailed information about the flow, though this is perhaps of minor interest to the aerodynamicist. In the past, it was the structural designer who demanded knowledge of the spanwise and chordwise pressure distributions and who felt the greatest need for detailed information on transient aerodynamic forces. So in this new field, it is again the structural designer who needs inti- mate knowledge of the origin and trans- fer of thermal energy in the boundary layer. 2.6.5.3 The Temperature Distribution in the Structure All three basic heat-transfer pro- cesses: convection, conduction, and radiation, are present in the aircraft temperature distribution problem. Heat enters the structure from the boundary layer, is conducted to distant portions of the structure, and escapes at the surface by radiation. Heat transfer out of the structure by convection and by radiation can be considered as additional boundary conditions to the main process of heat conduction into remote portions of the structure. Determination of the temper- ature distribution in the structure, 43 therefore, depends chiefly on the solu- tion of the heat-conduction equation in solids. This equation discloses that the mass density, specific heat, and thermal conductivity of the structure are the material properties present. All three are functions of the local temperature. Unfortunately, data about the variation of these material properties with tem- perature are meager. But even if such data existed, the solution of the heat conduction equation can be obtained only by numerical methods. The flight history of the vehicle will determine the type of boundary condi- tions that the temperature distribution must satisfy. Two types are possible: those that depend on time, and those that do not. If the vehicle is accelerating, diving, or climbing, then transient conditions exist on the surface of the structure, and time becomes a very important fac- tor in the formulation of the boundary conditions. On the other hand, if the vehicle is in level, constant speed flight, then the boundary conditions do not depend on time. In both cases, the temperature distribution in the structure varies continuously with time. It should also be noted that the entire flight history must be specified. Thus it is insufficient to specify a flight at constant velocity because the temper- ature distribution at the initiation of constant velocity will determine the character of the subsequent heat flow, and this initial temperature distribution depends on how the vehicle attained this velocity. Experimental data on temperature distributions are also quite meager because of an absence of a suitable laboratory source of thermal energy, an absence of experimental techniques, and inadequate knowledge of what is impor- tant.

44 2.6.5.4 Structural Effects of High Temperatures The objectives of the structural designer are to design a structure that fulfills the following requirements. a. Strength, i.e., the structure must be strong enough to withstand all conditions imposed on it which are within the established limitations of that particular mode! without permanent deformations or failures of any member. Rigidity, i.e., the structure must possess sufficient rigidity to prevent such aeroelastic phenomena as flutter or static divergence. c. Longevity, i.e., the structure must meet all requirements for a specified number of missions. d. Damage tolerance, i.e., the structure must be able to resist failure due to the presence of flaws, cracks or other damage for a specified period of unrepaired usage. e. Efficiency, i.e., the vehicle must be able to meet its performance speci- fications as to speed, climb, etc. which means obtaining the required strength and rigidity for the least possible weight. f. Economy, i.e., the above objectives must be obtainable at a cost estab- lished by the importance of the mission or by the expense of alter- native methods of accomplishing the same mission. A gross measure of the efficiency of the structure is given by the structural weight fraction. Measures of the struc- tural integrity of the structure are given by the margins of safety. These result from extensive calculations under a number of critical conditions to which all other conditions are compared. HYPERSONIC TECHNOLOGY FOR MILITARY APPLICATION These critical conditions are either obviously worse than others or have been shown from experience to be the most severe. In the case of the added complication of high temperatures, the relative severity of different conditions is far from obvious, and there is insufficient accrued experience to delineate the most severe temperature distributions. Consequently, much analytical and experimental work must be done before design criteria can be formulated. High temperatures in the structure have three primary adverse effects: · Stresses are induced by the differ- ential expansion of different por- tions of the structure. · The properties of the material are adversely affected. · The shape of the structure is dis- torted, which may lead to nonlinear interactions with the flight loads. If it can be assumed that the struc- ture behaves elastically and if the tem- perature distribution is known, then in theory the stresses can be calculated. But, if the materials behave inelastically, as they will at the high temperatures, then there are no proven methods avail- able. Certain strength properties of current materials at high temperatures are known, but there are wide gaps in this picture. Time is again a very important parameter because failure at elevated temperatures is not instan- taneous but may take time to develop. Also, after exposure to high tempera- tures for a certain amount of time, the elastic properties of the material at normal temperatures can also be affected. It is envisioned that the structural designer may be able or may be forced to specify the amount of time which the vehicle can sustain a certain flight path or flight maneuver. The factor of time also introduces new modes of structural behavior. Thus creep and creep-buckling under constant load can

STATUS OF HYPERSONIC TECHNOLOGIES occur. Superimposed upon the stresses caused by thermal expansion are the ordinary flight loads caused by gusts, pullouts, etc. In some cases, these can be added linearly, but in other cases there will be interactions. Here again the interactions which bring the struc- ture closer to failure must be considered in the design. 2.6.5.5 Summary It is apparent from this brief exam- ination that the invention of new mater- ials is the panacea for hypersonic flight. The structural designer must rely on the material scientist to invent the material and can only offer words of encourage- ment and point out what properties should be optimized. It is perhaps less apparent what other structural research goals should be set. Unfortunately, personal opinions cannot be entirely divorced from evalu- ations that must precede the establish- ment of research goals. The following steps should be taken at this time. 1. Create experimental facilities that can enable the structural designer to arrive at an optimum structure by cut-and-try methods. 2. Study flow patterns to evaluate the effect of shape in reducing the boundary layer temperature. 3. Study the possibility of minimizing thermal gradients and, therefore, stresses by design alterations; vie., with the same flight program comparison of stress distributions in structures with systematic differ- ences in the following charac- teristics. 4. Study the feasibility of controlling the heat flow by the following 45 means: a. lightweight passive thermal protection systems b. internal impedances such as joints c. active cooling d. other ingenious means In certain part, the above items assume data that are not yet available. Some of these deficiencies concern data of a very basic nature, including: a. High-temperature properties of aircraft materials b. Effects of heating rate on material properties Heat-transfer coefficients for the entire flight regime, from supersonic to hypersonic d. Recovery factors These data can be obtained only experimentally. Some of the other deficiencies concern techniques, i.e., the state of the art needs to be advanced. The following are included in this category: a. Efficient methods of calculating temperature distribution b. Efficient methods of calculating temperature stresses and strains The ramifications of the aerodynamic heating problem are so wide that, undoubtedly, there are good arguments for drastically altering or adding to the above. 2.7 Role of Computational Fluid Dynamics Computational fluid dynamics (CFD) has become the principal tool for aerodynamic and propulsion-flow design of hypersonic vehicles. Three cir- cumstances have contributed to this. First, existing ground-based experimental facilities are unable to realistically

46 simulate the combined thermal, dynamic, and chemical flow conditions of hyper- sonic flight. Second, the power of current supercomputers now enables 3-D- flow simulations of realistic flight conditions using the full Reynolds- averaged Navier-Stokes ~ equations. Third, the extent of flow detail obtained in computer simulations is inherently much greater than experiments can pro- vide. The inadequacy of simulation in ground-based experimental facilities is especially acute for the high-velocity, high altitude portions of hypersonic trajectories involving non-equilibrium air chemistry. Under these conditions, chemical reaction rates are important, and realistic air flow simulation requires that the flight parameters of velocity, free stream density, free stream ther- mochemical state, and physical scale all be simulated together. Though realistic air flow simulations are impossible in present experimental facilities, they are, at least in principle, possible in CFD simulations. Ballistic ranges can simulate velocity, free stream density, and free stream chemical state, but not the physical scale of flight vehicles. Even if money were no object, and a full-scale experimental hypersonic flow facility were envisioned, known technology does not provide a guide to produce a clean air flow around a sta- tionary test object with the requisite velocity, density, temperature, and chemical state. Computers, however, can simulate the proper free-stream and physical scale conditions, while providing detailed information on temperature, density, pressure, and chemical species con- centration at every point in the flow field provided the physical behavior can be adequately modeled, within present limits of computational power. Such detail, of course, also yields information, on surface skin friction, heating, lift, drag, thrust, and pitching moments. In consequence, the potential overall ability HYPERSONIC TECHNOLOGY FOR MILITARY APPLICATION of computers to simulate the proper thermodynamic and chemical aspects of hypersonic flight, while providing extraordinarily detailed information, has made it apparent that CFD continues to be essential for the design of hypersonic vehicles. We must note, however, that some level of physical approximation will be required to bring the numerical problem within practical computational capabilities. The modeling of turbulence is one example. It follows that experi- mental verification of the physical validity of the numerical models is crucial to the central role of CFD in hypersonics. The anticipated near-future feasibil- ity of scramjet propulsion has added to the role of CFD some major elements missing from the design of past hyper- sonic vehicles. Hypersonic flow problems associated with ballistic missiles, Apollo, and shuttle, for example, were those of unpowered atmosphere entry and descent. But hypersonic vehicles that must ascend through the atmosphere with air- breathing engines and descend during unpowered reentry involve much more complex hypersonic flows, as noted earlier. Air-breathing ramjet and scramjet flows can interact strongly with the external flow over the vehicle body, as discussed more fully in section 2.2. Experimental simulations of the integra- tion of external aerodynamic flows with hypersonic propulsive flows are not feasible for flight conditions, but computer simulations are. Thus CFD is an especially critical technology in the design of hypersonic vehicles that involve major problems of airframe/- engine integration. 2.7.1 Status of Hypersonic CFD The current state of CFD codes for hypersonic flow depends on the type of computation being made, whether for

STATUS OF HYPERSONIC TECHNOLOGIES external aerodynamics, inlet, combustion, or nozzle flow. Some hypersonic vehi- cles, such as boost-glide, involve only external aerodynamic flows, whereas NASP involves all types. 2.7.1.1 External Aerodynamics Computational fluid dynamics codes for this type of flow are the most advanced. With present supercomputers and codes, we can make 3-D computa- tions of aerodynamic and heating para- meters using the Reynolds-averaged Navier-Stokes equations for hypersonic flow over fuselage-wing-tail configur- ations, provided we know the location of boundary layer transition. The aero- space industry already has such codes for perfect gas and equilibrium-air chemistry. Three-dimensional codes for nonequilibrium (reaction rate dependent) chemistry are available now in a few laboratories, while others, currently being developed by aerospace companies, are expected soon. The main limitation to the realization of these codes is that the location of, or some criterion for, transition must be known. A second limitation is that present semi-empirical models for the turbulence stresses and heat flux while acceptable, are not as accurate as would be desirable for hypersonic wall-bounded flows having strong transverse or streamwise pressure gradients. With continued calibration of hypersonic codes, and exploration of new or modified turbulence models, the degree of CFD realism should improve with time. 2.7.1.2 Inlet Flows In hypersonic inlets, pronounced 3-D aspects of flow can be produced by interaction of shock waves with a boundary layer, and by the presence of internal corners. The relatively simple parabolized form of the Reynolds- averaged Navier-Stokes equations (PNS) 47 is useful for such flows if the inlet passage is slender and there is no streamwise flow separation. PNS codes have been used for some time through- out the aerospace industry. Hypersonic codes for 3-D flow with equilibrium air- chemistry using the full Reynolds- averaged Navier-Stokes equations also exist widely. Analogous inlet codes, however, do not yet exist widely, al- though some with reaction rate rather than equilibrium air chemistry are under development in various places and should be available for application in the near future. As in the case of external aero- dynamic flows, the main limitations to the physical reality of inlet codes are in the uncertain empirical criterion used for transition from laminar to turbulent flow when it occurs on portions of inlet surfaces, and the uncertain semi- empirical models used for hypersonic turbulence stresses and heat transfer. There is an additional special limit- ation of present inlet codes. They are unable to treat realistically the high altitude flight conditions of flow over cowl lips with small radii of curvature. Maximum cowl-lip heating on a vehicle like NASP occurs at high altitudes and high Mach numbers. The physical flow conditions in flight are of a type not yet investigated realistically by either CFD or experiment. In high altitude flight, the shock wave thickness on a cowl-lip is not negligible compared to shock detachment distance. Instead, the flow can be of the "merged" type where- in the lip bow wave is merged together with the lip boundary layer. The inte- raction of a relatively thick fuselage bow shock impinging on a merged lip shock layer has not yet been explored. Present CFD simulations and experi- mental investigations have corresponded to low-altitude flow conditions wherein the shock wave thickness is negligible compared to the detachment distance. Under these conditions, unlike high altitude conditions, it is unnecessary to compute through the shock wave struc

48 sure. Unfortunately, the Navier-Stokes equations have long been known to be very inaccurate for computing the flow structure within shock waves, so these equations cannot be used to get reliable CFD results for the particular conditions existing on cowl lips in high altitude flight. Since the shock on cowl lip heating may be very high, it must be investigated more realistically than it has been, by either direct simulation Monte Cario methods, or by appropriate new experiments, or by continuum equations more appropriate than Navier- Stokes. 2.7.1.3 Combustion Flows Computational fluid dynamics codes for scramjet combustors in hypersonic flight are in a relatively primitive state because of two circumstances. First, models for turbulent mixing in reacting compressible flows have been far less successful to date than have models for boundary layer flows. Second, experi- mental data on hydrogen-air mixing in scramjet combustors for flight faster than Mach number ~ have not been available to provide any code calibration in this important Mach number range. The degree of incomplete fuel-air mix- ing, and the nonuniform distribution of both species concentration and of ther- modynamic state are issues of potentially vital importance. It is unfortunate that the turbulent-mixing and chemically reacting type of flow in combustors, which is yet to be investigated exper- imentally for flight above Mach number 8, is also the same type of flow for which present CFD computations are the most uncertain. 2.7.1.4 Nozzle Flows In the expansion of combustion products through a nozzle, or a partly wall-bounded nozzle, reaction-rate chemistry is essential to the flow HYPERSONIC TECHNOLOGY FOR MILITARY APPLICATION computation. Currently 2-D reaction- rate codes are used to compute nozzle exhaust expanding into ambient air, but 3-D reaction-rate codes for an expansion partly over the surface of a vehicle, and partly into a vehicle-dependent external aerodynamic flow, do not yet exist, although they are under development. Vital to a nozzle flow computation is knowledge of the distribution of species, thermodynamic state, and dynamic quantities exhausting from the combustor. One must also know what chemical reactions are important, and what their rates are, since non-equil- ibrium chemistry is essential in nozzle expansion. In the nozzle flow compu- tation, a new element arises because the exhaust-fuselage boundary layer might relaminarize. Like transition, relaminar- ization is poorly understood for the conditions of hypersonic flight. 2.7.2 Validation of Hypersonic CFD Codes An experimental validation of non- equilibrium hypersonic CFD codes is more difficult than for conventional aircraft codes because of the absence of ground-based test facilities that can stimulate together the total variety of physics represented in hypersonic codes. Different experimental facilities, however, can test different components of the overall hypersonic physics simulated in the codes. Hypersonic wind tunnels, for example, can test a code's ability to simulate perfect-gas flows over complex 3-D geometries, although not always at the desired flight Mach and Reynolds numbers. Shock-tube type facilities, on the other hand, can test a code's ability to simulate high temper- ature thermochemical aspects of a flow, although usually for simplified geomet- ries and not always at the desired Reynolds number. Even though ground facilities cannot test together all interacting components of the physics

STATUS OF HYPERSONIC TECHNOLOGIES represented in a hypersonic CFD code, it is nevertheless essential to test by comparison with experiment as many of a code's physics components as is feasible. Even when this is done, some aspects of a hypersonic code (e.g. transition location, and nonequilibrium radiative heating) cannot be adequately tested by comparison with available ground test experiments. Flight tests may be the only way to thoroughly test the ability of a code to accurately compute such flow field parameters. 2.7.3 Future Role of CFD Overall, hypersonic CFD today appears acceptable for external aero- dynamics and inlet flows provided that the location of transition is known, and for nozzle flows, provided the initial entrance conditions are known. CFD, however, is weak for combustor flows, and is unable to reliably predict the location of transition. We must recog- nize that these current limitations are not inherent to CFD, but are mainly a consequence of the present state of supercomputer development which forces the use of a Reynolds-averaged form of the Navier-Stokes equations. While this · . tlme-averaglng process pro( uces equa- tions that are solvable on current computers in a practical amount of time, it requires that the turbulent stresses and heat flux be modeled, thereby introducing the primary inaccuracies in present day codes. If the full time- dependent Navier-Stokes equations were employed instead, the turbulent eddies would be directly computed rather than modeled, and the degree of CFD realism would be expected to greatly increase. Such calculations for a complex three- dimensional vehicle are outside the reach of today's supercomputers. However, they may be feasible for computing the onset and extent of boundary layer transition on a fuselage using current or next generation supercomputers. This advanced type of computation would require the development of methods for computing through transition in a hypersonic boundary layer on a vehicle flying through a prescribed disturbance field. Knowledge of the disturbance field in the stratosphere would, of course, also be needed. These obstacles to the direct numerical simulation of transition and turbulence (DNST) prob- ably will be overcome in the long range future. This will open up an entirely new level of CFD capability having much greater realism than present capabilities provide. The potential future importance of DNST to the Air Force is great, since it would largely overcome most of the present limitations of CFD. Such a capability would provide a large increase in the effectiveness of CFD applications to the design of aircraft and turbine engines as well as hypersonic vehicles. In view of such a major future potential, the technology of DNST computation is clearly important to the long range interests of the Air Force. Since this potential would serve industry and other agencies as well as the Air Force, it may by itself justify the development of future supercomputers with power suf- ficient to realize this next-generation level of advanced CFD technology. 2.X Experimental Capabilities Ground test requirements for hyper- sonic flight vehicles, even for cases of simulation rather than duplication of flight conditions, impose extreme demands on the equipment in terms of pressure and temperature (see Figure 2- 8~. Further, such facilities are expen- sive, ranging from $1-2 million to in excess of $500 million and take several years to construct or bring on-line. During the 1 960's extensive hypersonic test facilities were constructed in the U.S., and overseas, so that by 1971, 52 major operational aerodynamic test units existed, as shown in Figure 2-9. How

so ever, during the 1 970's and l 980's the number of operational facilities was reduced dramatically so that by 1986 only 23 were still in useful condition (see Figure 2-10~. Engine test facilities are similarly restricted (see Figure 2-11~. Indeed, while recent interest has led to some moth-balled facilities being refurbished some are still being consid- ered for destruction. Further, many of these facilities are 20 to 40 years old and do not produce appropriate flow conditions to address the problems presented by hypersonic air-breathing vehicles and their develop- ment. Indeed, recent studies of the upper atmosphere indicate that the com- mon concept of a quiescent and rela- tively ur~iform regime (a goal of some facility developments such as the NASA Quiet Tunnel) may not represent reality. Also, research is needed on facilities and their effect in particular on such phen- omena as boundary layer transition and chemical kinetics. Even where ground testing has been used extensively, such as the space shuttle with 35,000 occupancy hours and other vehicles (see Table 2-C), flight performance is not always well pre- d icted. For the hypersonic regime, extrapolation to portions of the flight envelope will further increase the probability of error. CFD codes will help in some cases; however, code validation in many areas remains to be done and requires the same ground test facilities discussed above and in the following sections. Thus, for the foreseeable future, it appears that it will not be possible to verify advances in many of the hyper- sonic technologies through ground testing. Consequently, flight test programs are needed for components of hypersonic air-breathing vehicles that cannot be adequately tested in ground facilities, for validation of concepts, and proof of CFD codes. Such testing HYPERSONIC TECHNOLOGY FOR MILITARY APPLICATION represents an extension of the ground test concept and is complementary to, not a substitute for complete systems ground testing. Experimental aircraft such as the projected NASP research vehicle will be essential in expanding the data base for large portions of the flight envelope. In the following sections general facility requirements, and then specific requirements for aerodynamics, propulsion, and materials and structures, will be discussed separately. 2.~.1 Test Requirements Test requirements for a hypersonic vehicle capable of flying over a speed range up to orbital (Mach number 25) are extremely severe. Data for aero- dynamic, propulsion, and structural materials are required up to the very high pressures and temperatures at which air becomes a hot plasma com- posed of molecular and atomic particles, ions, and free electrons. In stagnation regions of a hypersonic vehicle at high altitudes, atmospheric oxygen begins to dissociate above about Mach number 7. Both oxygen and nitrogen are almost fully dissociated at Mach number 15, and ionization becomes important at Mach numbers approaching 20. Data are required for pressure distributions, surface friction, temper- ature distributions, heat transfer rates from cooled surfaces, chemical reaction rates at high temperatures, and in locally disturbed flow, fuel/air mixing rates, material properties at high temperatures, and structural response. The key test parameters for aero- dynamic testing are Mach number, Rey- nolds number, Knudsen's number, and facility total enthalpy as well as transients in these variables. For the combustion processes associated with the type of air-breathing cycles being considered, this list is substantially

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STATUS OF HYPERSONIC TECHNOLOGIES longer and allows for less flexibility in terms of simulation through dimension- less variables. Thus, it is necessary to consider Prandtl number, Stanton number, Eckert number, Lewis number, and species chemical reaction rates non- dimensionalized by residence time. For some cases scaling or non-dimension- alization is ineffective and full-scale hardware testing is needed, such as with structural panels and joint sections. (See Appendix C: Glossary and Appendix D: Dimensionless Groups in Fluid Mech- anics for explanations of these dimen- sionless variables.) From an aerodynamic and propulsion standpoint it is necessary to locate the region of boundary layer transition on a slender vehicle forebody to determine flow conditions in the inlet of air- breathing engines and the subsequent combustion process, particularly above Mach number 6 for which scramjet pro- pulsion is envisaged. Thus, full Reynolds number and Mach number should be reproduced in a ground test facility, if full simulation were to be achieved. In contrast to a rocket-powered vehicle such as the space shuttle, which can have a steep ascent trajectory, a hypersonic vehicle using air-breathing propulsion for boost to orbital velocity would require a trajectory over an ex- tended Mach number range in the lower atmosphere to meet the air mass flow requirements of its engine. This implies a high dynamic pressure of about 1500 psf corresponding to which the Reynolds number on a full-scale vehicle would exceed 100 million up to Mach 10 or higher. Although a steeper rate of ascent probably would have to be fol- lowed above Mach number 12 because of temperature limitations on materials, Reynolds numbers of the order of 10 to 20 million can be expected up to above Mach number 20. For free stream Mach numbers above about 10, test requirements become even 51 more severe for slender vehicles because Mach number, free flight Reynolds num- ber, and full enthalpy must be repro- duced in a single test facility as real gas effects become important. In con- trast, a blunt vehicle such a the Apollo capsule or even the space shuttle (which reenters and stays at a high angle of attack - 40 deg - down to about Mach number 10) requires full enthalpy sim- ulation, but not the full Mach number and Reynolds number. Since viscous and high temperature effects are important for virtually all hypersonic testing, facilities using air as a medium are required because other gas media (freon, helium, pure nitrogen) do not provide the right quantitative simulation and there are no satisfactory means of converting all the required data to air. The above requirements hold for both aerodynamic/aerothermodynamic and propulsion testing. For the latter, it is not sufficient to test at free stream conditions corresponding to the lower Mach number at the immediate engine inlet (direct connect testing) because of flow distortion in the actual flight inlet due to ingestion of the thick forebody boundary layer, the impingement of shocks generated upstream and shock wave-boundary layer interactions. Thus full Mach number and forebody geometry are required to test the propulsion system. Another important use of hypersonic ground test facilities will be validation of CFD codes as the dependence on numerical techniques as a design tool is expected to be extensive in the hyper- sonic regime because of physical limit- ations of wind tunnels at the higher Mach numbers. This will require wind tunnels with capabilities not only at the desired test conditions, for example Mach number and Reynolds number for aerodynamic

52 tests, but with well-documented flow quality due to their effect on important phenomena such as boundary layer tran- sition. Most of the existing tunnels were built primarily for lift and drag type measurements or leading edge heat transfer and do not have good enough flow qualities for code validation. Thus, in general, facilities using air as the test medium and operating Rey- nolds numbers of 100 million would be required for full testing of a hypersonic cruise vehicle intended for operation up to Mach numbers 10 or 12, while Reynolds number up to 10 to 20 million and Mach numbers up to 25 would be required for an air-breathing orbital vehicle. Furthermore, full temperature simulation and high Re would be required for a hypersonic lifting vehicle operating above about Mach number 10. 2.~.2 Aerodynamic Test Capabilities Facilities for aerodynamic and pro- pulsion testing in the subsonic, tran- sonic, and supersonic regimes (below Mach number 5) are adequate to meet most future requirements if some facil- ities in need of repair are rehabilitated. Above Mach number 5 there are some 30 major facilities 1 in the United States and eight in Western Europe, the United Kingdom, and Japan. All but one half dozen were built in the 1 950s and 1 960s. All of the newer facilities were built in the early to mid- 1 970s. None are known to have been built after 1 976.2 Of these 30 facilities, seven (vir- tually all in U.S. industry) are on standby status, i.e., they are not pres- ently operational. Only seven hyper- sonic facilities using air are capable of yielding Reynolds numbers based on chord lengths in excess of 20 x 1 o6. (See Table 2-D.) HYPERSONIC TECHNOLOGY FOR MILITARY APPLICATION Of these high Re facilities only the Calspan shock tunnels are able to pro- duce free flight total temperatures above Mach 10, and only the 96-in. shock tun- nel approaches total temperatures for a Mach number well in the teens. Test durations for these facilities are a few milliseconds. From a propulsion standpoint, scramjet research has been undertaken in the 4 ft. diameter Scramjet Test Facility at the NASA Langley Research Center at Mach number 6 and temper- atures up to 4000° R on small models under one sq. ft. in cross-section. Scramjet tests up to Mach number 7 were run in the HRE hypersonic test facility at Plum Brook, Ohio, over a decade ago, but this facility has not · ~ seen in use since. Aside from limited parameter simu- lation, a drawback in most existing facilities is that the flow quality is not good enough for boundary layer transi- tion simulation. Boundary layer transi- tion on wind tunnel models has generally been found to occur at much lower Reynolds numbers than in free flight. This is due mainly to disturbances in the flow emanating from wind tunnel settling chambers and acoustic radiation from nozzle wall boundary layers. A super- sonic facility designed to minimize such disturbances has been under study at the NASA Langley Research Center - the "quiet supersonic tunnel". Present plans are for a Mach number 3.5 capability with possible addition of a Mach number 6 nozzle later. Thus, capabilities for aerodynamic and propulsion testing to meet require- ments in the hypersonic regime are extremely limited below Mach number 10 and virtually non-existent above Mach number 10. Because wind tunnels (even shock tunnels despite their very short running times) are temperature-limited for

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