Figure 3.10. Wetted area parameter from Figure 3.9 correlates with Kuchemann's tau yielding a geometric relationship to describe the delta planform configurations of different cross-sectional shape. VT = Vtotal.

tau = 0.10 vicinity. Airbreathing space launchers are in the range of 0.18 to 0.20. Rocket-powered hypersonic gliders are in the range of 0.22 to 0.26 tau. A correlating equation provides a means of translating Kuchemann tau into the S parameter, Swet/V°o66i7. As implied in Figure 3.10, as tau increases, the value of S decreases, meaning that the volume is increasing faster than the wetted area— crucial for a hypersonic aircraft, as skin friction is a significant part of the total drag. Later in the chapter this parameter will be related to the size and weight of a converged design as a function of the industrial capability to manufacture the spacecraft.

There are a wide variety of configurations possible. But if the requirements for a transportation system to space and return are to be met, the configurations spectrum is significantly narrowed [Thompson and Peebles, 1999]. Two basic configurations types are selected. One is for all-rocket and airbreathing rocket cycle propulsion systems that can operate as airbreathing systems to about Mach 6. For the rocket propulsion and airbreathing rocket propulsion concepts that are limited to Mach 6 or less, a versatile variable capture, inward-turning inlet [DuPont, 1999] can be integrated into the vehicle configuration derived from the FDL series of hypersonic gliders (see Figure 3.14) developed by the Flight Dynamics Laboratory [Buck et al., 1975] and the work of the McDonnell Douglas Astronautics Company. Because of

Figure 3.11. Hypersonic rocket powered glider for airbreathing Mach <6 and hypersonic combined cycle powered aircraft for airbreathing Mach >6.

the mass ratio to orbit, these are generally vertical takeoff and horizontal landing vehicles (VTOHL). This is the upper left vehicle in Figure 3.11. The second is for airbreathing propulsion systems that require a propulsion-configured vehicle where the underside of the vehicle is the propulsion system. The thermally integrated air-breathing combined cycle configuration concept is derived from the McDonnell Douglas, St. Louis, Advanced Design organization. This is a family of rocket hypersonic airbreathing accelerators and cruise vehicles [HyFac, 1970]. Depending on the mass ratio of vehicle these can take off horizontally (HTOL) or be launched vertically (VTOHL) and always land horizontally. The initial 1960s vehicle concept was propulsion configuration accelerated by a main rocket in the aft end of the body. Today it can retain this concept or use a rocket-based combined cycle propulsion concept. In any case, individual rockets are usually mounted in the aft body for space propulsion. This is the lower right vehicle in Figure 3.11. Both are functions of tau, that is, for a given planform area, the cross-sectional distribution is determined by the required volume.

Both this hypersonic glider based on the FDL-7C and the hypersonic airbreathing aircraft in Figure 3.11 have hypersonic lift-to-drag ratios in excess of 2.7. That means un-powered cross-ranges in excess of 4500 nautical miles and down-ranges on the order of the circumference of the Earth. So these two craft can depart from any low-altitude orbit in any location and land in the continental United States (CONUS) or in continental Europe (CONEU). Both are stable over the entire glide regime. The zero-lift drag can be reduced in both by adding a constant width section to create a spatular configuration. The maximum width of this section is generally the pointed body half-span. The pointed configurations are shown in Figure 3.11. No hypersonic winged-cylindrical body configurations were considered, as these have poor total heat load characteristics and limited down-range capability. As a strap-on booster the configuration is acceptable. The key to achieving the NASA goals of reduced payload to orbit continues to be flight rate and, as in the case of the transcontinental railroad, the scheduled services were supplied when as little at 300 statute miles of track (out of 2,000) had been laid

Figure 3.12. Wind-tunnel model configurations for tail effectiveness determination over hypersonic to subsonic speed regime (Mach 22 to 0.3).

Figure 3.12. Wind-tunnel model configurations for tail effectiveness determination over hypersonic to subsonic speed regime (Mach 22 to 0.3).

[Ambrose, 2000]. So our flights to Earth orbit need to be as frequent as they can be scheduled.

Vertical fin configuration has presented low-speed stability problems for many hypersonic glider configurations such as X-24A, M2/F2, HL-19 and others. The high dihedral angle verticals for three of the four configurations in Figure 3.12 are representative of the vertical fin orientation. The "X" fin configuration was the result of an extensive wind-tunnel investigation by McDonnell Douglas and the AFFDL that covered Mach 22 to Mach 0.3. A total of four tail configurations were investigated over the total Mach number range and evaluated in terms of stability and control; they are shown in Figure 3.12. All of the configurations, except the first "X" tail configuration had serious subsonic roll-yaw instabilities at lower speeds. The "X" tail configuration has movable trailing edge flaps on the lower anhedral fins, and upper surfaces are all movable pivoting control surfaces at approximately 45 degrees dihedral angle. This combination provided inherent stability over the entire Mach number range from Mach 22 to landing.

The FDL-7 derived hypersonic gliders have a higher lift-to-drag ratio configuration than those similarly developed by Mikoyan and Lozino-Lozinski in Russia as the "BOR" family of configurations because of operational requirements. Some of the first studies performed for NASA by McDonnell Aircraft Company and Lockheed [Anon, McDonnell, 1970; Anon, Lockheed, 1967] identified as a need, the ability to evacuate a disabled or damaged space station immediately, returning to Earth without waiting for the orbital plane to rotate into the proper longitude (see Chapter 2). Unfortunately, many of these studies were not published in the open

Figure 3.13. BOR V after return from hypersonic test flight at Mach 22. The one-piece carbon-carbon nose section is outlined for clarity. The vertical tails are equipped with a root hinge, so at landing the tails are in the position shown by the dashed line. Thus BOR V is stable in low-speed flight. If the variable dihedral were not present, BOR V would be laterally and directionally unstable at low speeds.

Figure 3.13. BOR V after return from hypersonic test flight at Mach 22. The one-piece carbon-carbon nose section is outlined for clarity. The vertical tails are equipped with a root hinge, so at landing the tails are in the position shown by the dashed line. Thus BOR V is stable in low-speed flight. If the variable dihedral were not present, BOR V would be laterally and directionally unstable at low speeds.

technical literature and were subsequently destroyed. For a Shuttle or CRV configuration that waiting might last seven to eleven orbits, depending on inclination, or, in terms of time, from 10.5 to 16.5 hours for another opportunity for entry: that might be too long in a major emergency. In order to accomplish a "no waiting'' descent with the longitudinal extent of the United States, that requires a hypersonic lift-to-drag ratio of 2.7 to 2.9. The hypersonic gliders based on the FDL-7 series of hypersonic gliders have demonstrated that capability. Given the longitudinal extent of the former USSR, that requirement translates into a more modest hypersonic lift-to-drag ratio of 1.7 to 1.9. So Lozino-Lozinski BOR hypersonic gliders meet that requirement to land in continental Russia without waiting. This hypersonic lift-to-drag ratio means that, if the deorbit rocket retrofiring is ground-controlled, Russian spacecraft could be precluded from reaching the United States. The BOR class of vehicles is now being realized not in Russia but in the United States, as the CRV is in fact an adaptation of the BOR V vehicle. Such a BOR vehicle is shown in Figure 3.13 after recovery from a hypersonic flight beginning at about Mach 22 [Lozinski, 1986]. The BOR V picture was given to the author by Glebe Lozino-Lozinski at the IAF Congress held in Malaga, Spain. Lozinski was very familiar with the subsonic lateral-directional instability for this high dihedral angle fin configuration, and in the 1960s constructed a turbojet powered analog that investigated this problem. The solution was to make the aft fins capable of variable dihedral (a power hinge was mounted in the root of each fin) so that at high Mach numbers the fins were at about

Figure 3.14. FDL-7 C/D (top) compared with Model 176 (bottom).

plus 45 degrees, as shown in Figure 3.13. However, when slowing down to transonic and subsonic Mach numbers, the dihedral angle was decreased, so that at landing the fins were at a minus 10 degrees, as shown by the dashed outline in Figure 3.13. So the BOR class of vehicle was a variable geometry configuration that could land in continental Russia; its stability could be maintained over the entire flight regime, from Mach 22 to zero.

The Model 176 began with the collaboration between Robert Masek of McDonnell Douglas and Alfred Draper of AFFDL in the late 1950s on hypersonic control issues. After a series of experimental and flight tests with different configurations the "X" tail configuration and the FDL-7C/D glider configurations emerged (Figure 3.12) as the configuration that was inherently stable over the Mach range and had Earth circumferential glide range. The result was the FDL-7MC and then the McDonnell Douglas Model 176. Figure 3.14 compares the two configurations. In the early 1960s both configurations had windshields for the pilots to see outside (see Figure 3.19). However, with today's automatic flight capability visual requirements can be met with remote viewing systems. The modified FDL-7 C/D configuration was reshaped to have flat panel surfaces, and the windshield provisions were deleted, but it retains all of the essential FDL-7 characteristics. To assure the lift-to-drag ratio for the circumferential range glide, the Model 176 planform was reshaped for a parabolic nose to increase the lift and decrease the nose drag. A spatular nose would have also provided the necessary aerodynamic margin; however, the original configuration was retained, with just the windshield provisions (Figure 3.16) deleted. The Model 176 was proposed for the Manned Orbiting Laboratory (MOL) described in Chapter 2. It was a thoroughly designed and tested configuration with a complete all-metal thermal protection system that

Figure 3.15. Model 176 in the McDonnell Douglas Hypervelocity Impulse Tunnel (circa 1964).

had the same weight of ceramic tile and carbon-carbon concepts used for the US Shuttle, but was sturdier. A wind-tunnel model of the McDonnell Douglas Astronautics Company Model 176 installed in the McDonnell Aircraft Company Hypervelocity Impulse Tunnel for a heat transfer mapping test is shown in Figure 3.15. Note that conforming to the piloting concepts of the 1960s it has a clearly distinct windshield that is absent from the configuration concept in Figure 3.14. The model is coated with a thermographic phosphor surface temperature mapping system [Dixon and Czysz, 1964]. This system integrated with semiconductor surface temperature heat transfer gauges [Dixon, 1966] permits the mapping of the heat transfer to the model and full-scale vehicle. The model permitted accurate thermal mapping to the heat transfer distribution on the body and upper fins. From this data the full-scale surface temperatures for a radiation shingle thermal protection system could be determined and the material and thermal protection system appropriate for each part of the vehicle determined.

The important determinations that resulted from these heat transfer tests are that the sharp leading-edge, flat-bottomed, trapezoidal cross-section reduced the heating to the sides and upper surfaces. The surface temperatures of the thermal protection shingles are shown in Figure 3.16. In the range of angles of attack corresponding to maximum hypersonic lift-to-drag ratio the sharp leading-edge corner separates and reduces the upper surface heating. Because of the separation, the isotherms are parallel to the lower surface and to 2,100 to 2,400°F (1,149 to 1,316°C) cooler than on the compression surface. The upper control fins are hot, but there are approaches and materials applicable to control surfaces. The temperatures shown are radiation equilibrium temperatures. The temperatures with asterisks are the radiation equilibrium temperatures if not employing thermal management.

Figure 3.16. FDL-7 C/D, Model 176 entry temperature distribution. Upper surface heating minimized by cross-section geometry.

Thermally managed with nose water transpiration cooling (demonstrated in flight test in 1966) and heat pipe leading edges (demonstrated at NASA Langley in 1967-68) these temperatures of the nose and leading edges are 212°F and 1,300°F (100°C and 704°C) respectively.

Except for the tail control surfaces, the vehicle is a cold aluminum/titanium structure protected by metal thermal protection shingles. Based on the local heat transfer and surface temperature, the material and design of the thermal protection system was determined, as shown in Figure 3.17. It employs a porous nose tip with about a one-half inch (12.3 mm) radius, such the Aerojet Corporation's diffusion bonded platelet concept. In arc-tunnel tests in the 1960s, a one-half-inch radius sintered nickel nose tip maintained a 100°C wall temperature in a 7200 R (4,000 K) stagnation flow for over 4300 seconds utilizing less than a kilogram of water. The one-half-inch (12.3 mm) radius leading edges and the initial portion of the adjacent sidewall forms a sodium-filled, Hastelloy X heat pipe system that maintains the structure at approximately constant temperature. Above the heat pipe, sidewall are insulated Inconel honeycomb shingles, and above those and over the top, are diffusion-bonded multicell titanium. The compression side (underside) is coated columbium (niobium) insulated panels or shingles similar to those on the compression side of the X-33, that protect the primary structure shown in Figure 3.18. The upper all-flying surfaces and the lower trailing flap control surfaces provide a significant challenge. Instead of very high temperature materials that can still have sufficient differential heating to warp the surfaces significantly, the approach was to adapt the heat pipe concept to heat pipes contained within honeycomb cells

Figure 3.17. FDL-7 C/D, Model 176 materials, thermal protection systems distribution based on temperature profile in Figure 3.16.
Figure 3.18. McDonnell Aircraft Company Roll-Bonded Titanium Structure (circa 1963), from Advanced Engine Development at Pratt & Whitney SAE [Mulready, 2001]. Today this structure would be super-plastically formed and diffusion-bonded from RSR titanium sheets.

perpendicular to the surface. In that way the control surfaces are more isothermal reducing thermal bending and reducing the overall material temperature.

The structure of Model 176 was based on diffusion bonding and superplastic forming of flat titanium sheets. Forty-five years ago the method was called ''roll bonding'' and executed with the titanium sealed within a stainless steel envelope and

Figure 3.19. USAF one-half scale FDL-5 vehicle (reproduced from Astronautics and Aeronautics [Draper et al., 1971]).

processed in a steel rolling plant. With a lot of effort and chemical leaching the titanium part was freed from its steel enclosure. All of that has been completely replaced today by the current titanium diffusion bonding and superplastic forming industrial capability. The picture in Figure 3.18 is from a Society of Automotive Engineers (SAE) publication entitled Advanced Engine Development at Pratt & Whitney by Dick Mulready. The subtitle is "The Inside Story of Eight Special Projects 146-1971.'' In Chapter 6, ''Boost glide and the XLR-129—Mach 20 at 200,000 feet''. The McDonnell Douglas boost-glide strategic vehicle is mentioned, together with the key personnel at McDonnell Aircraft Company. Low thermal conductivity standoffs set the metal thermal protection insulated shingles off from this wall so that there is an air gap between them. The X-33 applied the metal shingle concept but with significant improvement in the standoff design and thermal leakage, in the orientation of the shingles, and in the thickness and weight of the shingles. This is one aspect of the X-33 that can be applied to future spacecraft for a more reliable and repairable TPS than ceramic tiles. The titanium diffusion bonded and superplastically formed wall was both the primary aircraft structure and the propel-lant tank wall. The cryogenic propellants were isolated from the metal wall by a metal foil barrier and sealed insulation on the inside of the propellant tank.

The United States Air Force Flight Dynamics Laboratory fabricated a half-scale mock-up of the stage and a half of the FDL-5 configuration [Draper et al., 1971] shown on the right side of Figure 3.14, and presented in Figure 3.19. The strap-on tanks provided propellants to about Mach 6 or 7 and then the mission continued on internal propellants. Note the windshields installed in this 1960s mock-up. This was a vertical launch, horizontal landing configuration, as shown in Figure 3.19. The intent was to provide the United States Air Force with an on-demand hypersonic aircraft that could reach any part of the Earth in less than a half-hour and return to its launch base or any base within the continental United States (CONUS).

Initial operating capability -j;; ;;

Initial operating capability -j;; ;;

Figure 3.20. Individual Model 176 launch costs for a 100-launch program, as projected in a McDonnell Douglas Astronautics Corporation 1964 brief. RSH, reusable spacecraft hardware; ESH, expendable spacecraft hardware; RSS, reusable spacecraft spares; OOPC, other operational costs; T IIIC, Martin Titan III C cost.

Figure 3.20. Individual Model 176 launch costs for a 100-launch program, as projected in a McDonnell Douglas Astronautics Corporation 1964 brief. RSH, reusable spacecraft hardware; ESH, expendable spacecraft hardware; RSS, reusable spacecraft spares; OOPC, other operational costs; T IIIC, Martin Titan III C cost.

However, in a very short period of time after this mock-up was fabricated, the path the United States took to space detoured and most of this work was abandoned and discarded.

The ultimate intent was to begin operational evaluation flights, with the Model 176 launched on a Martin Titan IIIC, as shown in Figure 3.20. In 1964, the estimated cost was $US 13.2 million per launch for a 100-launch program, or about $US 2,700 per payload pound. As the system was further developed, two strap-on liquid hydrogen-liquid oxygen propellant tanks would be fitted to the Model 176 spaceplane for a fully recoverable system, as shown on the right side of Figure 3.20. The estimated 1964 cost of this version was $US 6.1 million per launch for a 100-launch program, or about $US 1,350 per payload pound. The launch rate for which the launch estimate was made has been lost in history, but to maintain the MOL spacecraft, launch rates on the order of one per week were anticipated for both re-supply and waste return flights. The latter flights could exceed the former in all of the studies the author is familiar with.

One of the most practical operational aspects of the FDL-7 class of hypersonic gliders was that the lifting body configuration forms an inherently stable hypersonic glider. Based on work by McDonnell Douglas Astronautics on control of

hypersonic gliders, the FDL-7 as configured by McDonnell Douglas incorporated an integral escape module. As shown in Figure 3.21, the nose section with fold-out control surfaces was a fully controllable hypersonic glider capable of long glide ranges (though less than the basic vehicle, but greater then the Space Shuttle). So the crew always had an escape system that was workable over the entire speed range. As shown the fold-out control surfaces are representative of a number of different configurations possible.

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