AAWSAP DIRD, Space Access: Where We’ve Been and Where We Could Go, March 2010

UNCLASSIFIED//FOR OFFICIAL USE Oilti Defense Intelligence Reference Document Acquisition Threat Support 8 March 2010 !COD: 1 December 2009 DIA-08-1001-006 Space Access: Where We've Been ... and Where We Could Go UNCLASSIFIED/ /FOA OFFI@IAL USE er•tv UNCLASSIFIED//FQA QFFl61alil l!ISI!! 91ft I Space Access: Where We've Been . . . and Where We Could Go Prepared by: Acquisition Support Division (DW0-3) Defense Warning Office Directorate for Analysis Defense Intelligence Agency Author: AAP Person 67 Administrative Note COPYRIGHT WARNING: Further dissemination of the photographs in this publication is not authorized. This product is one in a series of advanced technology reports produced in FY 2009 under the Defense Intelligence Agency, Defense Warning Office's Advanced Aerospace Weapon System Applications (AAWSA) Program. Comments or questions pertaining to this document should be addressed tolAAP Person 1 I, AAWSA Program Manager, Defense Intelligence Agency, ATTN: CLAR/DW0-3, Bldg 6000, Washington, DC 20340-5100. UNCLASSIFIED//FQA QFFICl1'1k W&li 8HLY ii UNCLASSIFIED/;'FOR OFFI@IAL l:ISE OHL¥ Contents Introduction ............................................................................................................v Propulsion Perspective........................................................................................... 1 Hypersonic Configuration Concepts........................................................................ 2 Thermodynamics and Materials ............................................................................ 16 The Qu Tube ......................................................................................................... 23 Rocket Propulsion ................................................................................................ 25 Up-and-Down Operations ..................................................................................... 30 Launch Options .................................................................................................... 33 Atmospheric Variations ........................................................................................ 35 Conclusion............................................................................................................ 37 Appendix A: Historical Perspective ....................................................................... 43 Appendix B: Aeropropulsion Integrated Vehicle ................................................... 46 Appendix C: TAV Operational Costs ...................................................................... 47 Appendix D: Landing Ellipses for Hypersonic Gliders ............................................ 48 Figures Figure 1. Hardware Flow ........................................................................................vi Figure 7. Detailed Design Analyses Show the Weight Trends are as Much a Figure 9. NASA Langley Wing-Body Configuration WB-004 with Critical Areas for Figure 15. Both Delta Planform Lifting Body (Dynasoar) and Model 176 Offer Figure 2. Impact of Air-Breathing Rocket ................................................................1 Figure 3. HSVS Hypersonic Cruise Aircraft Showing True Skin Temperature ...........3 Figure 4. Hypersonic Rocket-Powered Glider Hypersonic Air- Breathing Cruiser ......4 Figure 5. Delta-Lifting Body Designs .......................................................................5 Figure 6. Martin Marietta X-24 A & B Research Gliders............................................7 Function of Configuration Family as Lift-to-Drag Ratio ............................9 Figure 8. High-Performance Hypersonic Glide Aircraft ..........................................10 Wing Bodies Identified ...........................................................................11 Figure 10. FDL-7C/D and FDL-7MC Lifting-Body Configuration..............................12 Figure 11. FDL-7C/D with a DuPont Retractable Inward-Turning Inlet .................12 Figure 12. Comparison of FDL-7C/D and Model 176 ..............................................13 Figure 13. Sufficient Cross Range (L/D) Means There is No Waiting to Return ..... 14 Figure 14. Hypersonic Glider Characteristics .........................................................15 Superior Landing Performance .............................................................16 Figure 16. McDonnell Aircraft Company Roll-Bonded Titanium Structure ..............17 UNCLASSIFIED//FOR OFFICIO! 1155 ON! X iii UNCLASSIFIED/ /POI': OPPIEIJct t:191! BHL'f Figure 17. Model 176 in the McDonnell Douglas Hypervelocity Impulse Tunnel ....18 Figure 18. FDL-7C/D, Model 176 Entry Temperature Distribution .........................19 Figure 19. Even At Mach 12, Embedded Vortices in the Boundary Layer Alter the Local Heat Transfer ..............................................................................19 Figure 20. Thermographic Phosphor Image of Model 176 at Near-Maximum Angle of Attack .....................................................................................20 Figure 21. From L/D Maximum to Maximum Angle of Attack, There Is Always a Cool Sub-layer Adjacent to the Wall .....................................................20 Figure 22. This 1988 SEP Bordeaux SiC/SiC Panel Could Sustain Temperatures of up to 3,000°F........................................................................................21 Figure 23. UBE Corporation's Tyranno Cloth ..........................................................21 Figure 24. A Porous Nickel Tip Oozing Water ........................................................21 Figure 25. FDL ASSET Flight-Tested From Orbital Speeds to Evaluate 1960s Materials ..............................................................................................22 Figure 26. Heat Pipe Shuttle Leading Edge Designed and Built by McDonnell Figure 27. Boost-Glide Strategic Vehicle with Pratt & Whitney XLR-129 Rocket Douglas Astronautics ...........................................................................22 Engine Installed ...................................................................................25 Figure 28. XLR-129 ...............................................................................................25 Figure 29. Two Rocket Air-Breathing Rocket Cycles to Mach 5.5 ...........................27 Figure 30. HOTOL Evolution: From Aerodynamic Optimum Configuration to Practical Launcher Configuration .........................................................27 Figure 31. LACE Air-Breathing Rocket ...................................................................29 Figure 32. The FDL-7 Class of Vehicles ..................................................................29 Figure 33. Takeoff and Landing Speeds of Minimum-Sized Launchers ...................30 Figure 34. Horizontal launch Not Practical Unless Weight Ratio Less Than Four ...31 Figure 35. Propellant Tanks That Are Not Reentry Vehicles Greatly Reduce System Weight .....................................................................................33 Figure 36. Simple Horizontal Integration and Vertical Launch Provides Rapid Launch Capability .................................................................................34 Figure 37. A Vertical Launch Complex Provides Vertical Toss Back Booster Recovery and Horizontal Landing Facilities for the Hypersonic Gliders 34 Figure 38. A 1964 MDC Astronautics, St. Louis, Briefing........................................35 Figure 39. Earth's Atmosphere ..............................................................................36 Figure 40. FDL-5 Scale Model of A Stage and One-Half..........................................37 Figure 41. The FDL-7 and Model 176 Class of Hypersonic Gliders..........................37 Figure 42. Hypersonic Decelerating.......................................................................38 Figure 43. Where We Are Today ............................................................................39 Figure 44. Where We Could Be If We Can Recapture the Engineering Confidence and Expertise of the Apollo/Saturn V Era .............................................40 Figure 45. TAV Operational Costs ..........................................................................47 Figure 46. Landing Ellipse .....................................................................................48 Tables Table 1. Characteristics of Selected Flight Dynamics Laboratory Hypersonic Glider Configurations During the 1958-68 Timeframe ...............................6 Table 2. Elements of the Space Infrastructure Shown in Figure 44 .......................41 UNCLASSIFIED/;CfOR. OFFI&IAI: W&li &Ptklf iv UNCLASSIFIED/fFOR OFFIE!IAL l::l!H! 8HL'I Space Access: Where We've Been . . . and Where We Could Go Introduction Development of commercial access to space by our budding space­ faring civilization is a straightforward effort dominated by propulsion and reliability. The initial focus should be on schedulable, dependable access to and from low Earth orbit (LEO). For years we have known the means to accomplish such a task but have lacked a dedicated organized effort. The key requirement is to develop a robust and not necessarily a low-cost infrastructure, without which commercial exploitation of LEO and the moon will not be possible. This is a matter of skill; operational hardware based on durable, reliable, and demonstrated components; and operational systems. It is not necessarily a matter of technology. However, technology discovery and development are necessary for future space travel beyond Earth's environs. This paper addresses these issues by providing a running account of the historical details associated with the development of the myriad systems proposed and tested to provide access to space. Among the many advances in space access that will be possible in the future,1 the key technology developments will be in the area of propulsion, because without these we are confined to our solar system by flight times limited to a project team's functional life. The Pioneer spacecraft were fortunate to be monitored for 20 years. However, the issue facing our spaceflight organizations is the lack of a durable, consistent, schedulable, and frequent hardware system to and from space assets such as the International Space Station. In October 1958, the author's job in the vertical wind tunnel at Wright-Patterson Air Force Base abruptly changed; hypersonic and high-temperature flows became a new focus. What was then the Aircraft Laboratory was to become the Air Force Flight Dynamics Laboratory (AFFDL), with a focus on space flight. Al Draper of the AFFDL began working with a select group of aerospace firms on hypersonic gliders. The initial requirement from the Air Force was to quickly find operational access to space. Technology application, hardware design and fabrication with an innovative application, and extending the industrial capabilities of the time were very much the issue, as exemplified by the Lockheed A-12/SR-71. When asked about space access at the time, a group of Aerospace Corporation UNCLASSIFIED/ /FOR. OFFICI CIs. flili Qllls.¥ V UNCLASSIFIED/,«POil Offl@IAL YSli QPU,¥ veterans responded, "It was not a technology issue; it was a hardware issue." In a keynote address to the Aeronautical Revolutionary Concepts Workshop sponsored by the Vehicle Applications Panel of the National Research Council and held at NASA Ames in July 1984, then-Assistant Secretary of Commerce for Productivity, Technology, and Innovation Dr. D. Bruce Merrifield identified the problem of translating ideas into products as preparing technology for product manufacture. Dr. Merrifield drew an analogy between this step and Major League Baseball's farm system, which prepares skilled but untrained players for the major leagues. The United States assigns projects to accomplish technology tasks so the flow of production­ ready hardware is always improving and is not fixed (see Figure 1). Innovation Focus: Preparation of Technology for Application ---r-----7to10years-----~1 ------~--""••"_, ,__...;;.;.:, ----,....-=----_-_-_,---...,.._ ".........., •"- _ .....:.., .. --"'"'--... ............. ·····•.. ,••· ' •, .. Prototype .... Idea Generation .______., .______, PrOCluct and Pilot lntenm + Commercialization ·. Develooment Plant I Manufacture lnve~n Translation ··•.... ··•. ••······· ...•••· •.................. ................•1:i:t· fA•• ·····••..................... ROLP Gap in effective No Gap in 100/o Cost 900/o Cost preparation for transformation application to production in United States Or.O. Bruce Merrifield Oeptment of Commerce 1983 Innovation In Aeronautics Assistant Secretary for Productivity, Wortcshop, NASA AMES Technology and Innovation Figure 1. Hardware Flow Saturn I and Saturn V could be readied for a moon flight in such a short time because most of their hardware was based on a frozen design, proven production processes, and adaptation of existing hardware. Using a similar approach, current industrial capabilities can create the next practical system for accessing space. In the late 1950s and early 1960s, the U.S. Air Force was working toward an operational capability analogous to its B-52 fleets: flight operations when required or "on demand." After NASA was assigned UNCLASSIFIED//iOA: OiilCIOls. n&li Qfslls.¥ vi UNCLASSIFIED//Pelt err1e1J11t l:191!!! 9HL'I responsibility for space access, that Air Force's focus switched to surveillance, communication, and Global Positioning System satellites. In the late 1950s, there existed a predisposition-forced by the military competition between the United States and the former Soviet Union-to use rockets derived from military ballistic missiles. That decision curtailed efforts to develop alternatives to chemical rockets together with practical commercial developments. With the orbiting of Sputnik, the aircraft path to space, as represented by the X series of planes, ended with the X-15. With the X-15's demise, all efforts to fly aircraft to space ended, replaced by the more familiar (but less practical) strategy of loudly blasting to space with expendable rockets derived from undertested ballistic missile hardware, as documented in early failures. Like their ballistic missile progenitors, current expendable rockets can be launched only once. With the exception of the experimental Delta Clipper developed and operated by William Gaubatz and the late Pete Conrad, no operational launcher has ever successfully aborted. In this context, a reusable launcher is simply an expendable with some parts reused a few times. Thus, neither the United States nor the Soviet Union/Russia has ever realized a truly commercial approach to space travel, although the Soviets came close to taking the first step with the since-terminated Energia/Buran system. Both the United States and the Soviet Union/Russia historically have generated a large number of concepts that could fly directly to space and return on a sustained, frequent, scheduled basis. An all-up air breather such as the NASP was to solve that problem and fly directly to space and return. Developing an operational mach 12 to 14 aircraft with air-breathing propulsion presents a serious design, engineering, and fabrication challenge analogous to the SR-71 Blackbird. UNCLASSIFIED/,,roA 8FFI&IAk WSE 8HLY vii UNCLASSIFIED/ /FOR: 9FFl@IAL WIiii QPIL¥ Propulsion Perspective In exiting Earth's atmosphere, the propulsion system and configuration are inexorably linked. A hypersonic glider exits the atmosphere on either a rocket booster or a first stage of a two-stage-to-orbit aircraft. As such, it usually exits the atmosphere quickly, and the key exit design considerations are the high transonic aerodynamic and the mechanical loads encountered in the exit trajectory. Whether for a new rocket launcher or the U.S. space shuttle, the phenomenon is the same: the peak mechanical loads occur during exit. In this case, the exit aerodynamics are important but not vital. The vital aerodynamics and thermodynamics (aerothermodynamics) are in the entry glide, where thermal loads are maximal and must be controlled. The vehicle must always be controlled in flight so its attitude and direction are within limits set by the aerothermodynamics. The angle-of-attack limits are very close for high-performance hypersonic gliders, as their glide angle of attack is 11 to 15 degrees, not the 45 degrees of the space shuttle. Even the Russian Buran had a lower glide angle of attack than the shuttle; a TsAGI report given to the author by Vladimir Neyland shows it to have been about 30 to 35 degrees. 2 Like the Buran, the high-performance glider is best controlled by an automatic integrated flight control system that monitors the thermodynamic state of the vehicle, as well as its aerodynamic and trajectory states. The sensor array provides real-time information to the control system that can maintain the correct attitude in a manner a human controller could not accomplish. So it is this phase of the flight that designs the hypersonic glider. The exception is when powered by an All Rocket air-breathing rocket (HOTOL, Skylon, and LACE), which must remain lower in the atmosphere until reaching the air­ breathing rocket transition to conventional rocket. The configuration for the air-breathing rocket is different, as it must have a retractable air inlet in the mach O to 5 range but does not determine the vehicle configuration. The impact is significant, as the carried oxidizer is reduced in the heaviest initial portion of the flight, as shown in Figure 5. 35 ton payload 2 for a Delta Clipper-type design with 334 ton 101 ton TOGW an aerospike nozzle tested by 28.6 ton 15.9 ton DEW 678 m2 428 m2 Konstantin Feotkiskov. The example is from a Senior Capstone Design Study Figure 2. Impact of Air-Breathing Rocket Team from Parks College, Saint Louis University, circa 1992, and is based on the engineering reports the author was permitted to read from the library of Konstantin Feotkiskov, an aerospace designer and cosmonaut. The question, as always, is, why bother with air-breathing systems at all if they are that much of a challenge? The answer is to consider a partial air-breathing system based on available hydrogen/oxygen rockets that operate to about mach S.S. It operates in a flight region where the carried oxidizer quantities are the greatest. An operational system is sought that is capable of a large number of flights per year. The fewer resources required for launch, the greater ease with which the system can operate and the greater potential to operate from more bases. LACE Rooket Aero Spike Konstanbn FeotkIskov ..._Under expanded ·, Over expand'e<I UNCLASSIFIED/,erQA QFFI&il.t.k WIiii &Ptllf 1 UNCLASSIFIED//FOR OFFl&IAk W&li 0PU,¥ The Russian design bureaus are to thank for arriving at a concept that eliminated the noisy and hazardous air-breather takeoff and for increasing the operational flexibility of the British HOTOL concept. Glebe Lozino-Lozinski had a concept for a spacecraft with a 7-metric-ton payload carried atop an Antonov An-225, with a second An-225 carrying the liquid hydrogen and launch facilities and staff.3 The An-225 was in fact a mobile launch facility; it could literally launch a satellite for any facility that could accommodate a B-747 or an MDC-11. With Rolls Royce or General Electric engines, the An-225 becomes a more easily maintained vehicle with better altitude performance. The An-225's empennage is modified from the An-124's single vertical and horizontal empennage to an 'H' configuration. This permits the powered hypersonic glider to easily lift off the top of the vehicle, as the MBB Sanger wind tunnel test demonstrated. Most commercial transport aircraft larger than ERJ 170 are potential mobile launch platforms for space tourism, point-to-point cargo, or orbital facilities support. Most of the commercial passenger equipment can be removed, with just enough equipment remaining for a launch crew. The fuselage is strengthened and fitted with external mountings for the hypersonic glider. The landing gear need not be modified, as the same maximum weight as the commercial transport will be maintained. The flight control system would be adapted to automatically maintain the correct launch trajectory until separation. A second modified transport would be modified to carry the liquid hydrogen and liquid air to fuel the hypersonic vehicle, along with maintenance and support crew. The intent is to use the automatic launch checkout the author witnessed at Baikanour in 1988, wherein a Soyuz that arrived on its train carrier at 0500 hours launched carrying a Progress capsule at 1715 hours the same day. That should make a local launch possible within hours of arriving at the specified airport launch departure site. These two elements can provide a commercial space launch facility that requires no special or dedicated operational base. Hypersonic Configuration Concepts The configuration and the propulsion system are linked through aerothermopropulsion integration. This approach is not new, as a wide spectrum of configurations and concepts existed in the 1960s. One such McDonnell Aircraft Company concept is shown in Figure 3. This potential operational mach 12 cruise vehicle was developed for the U.S. government as a strike reconnaissance vehicle taking off from a U.S. Air Force base. The concept was to provide on-demand reconnaissance in force operations. However, as was the case with all such efforts in the 1960s, none of the aircraft derived from the "flight-to-space" efforts reached a hardware stage. Individuals working on these projects were convinced that the industrial capability existed to design and fabricate these vehicles, and that such vehicles were technically feasible. The concepts varied widely among different nations, but all had as their goal a transportation system to space that had commercial potential. This discussion is provided to discriminate between rocket-powered hypersonic gliders and hypersonic cruisers with an air­ breathing propulsion system . UNCLASSIFIED/;CFQR 0FFI&l.t.k Wliliii 8Plk¥ 2 -- UNCLASSIFIED//FOR OFFl@IAL l::l!H! 8HL'I McDonnell Aircraft Advanced Design Dept. 1958 to 1967 Mr. H. D.Altis, Director Figure 3. HSVS Hypersonic Cruise Aircraft Showing True Skin Temperature A wide variety of configurations for recoverable spacecraft are possible. But if the requirements for a transportation system capable of traveling to and returning from space are to be met, the configurations spectrum is significantly narrowed. Two basic configuration types emerge. One configuration is for a hypersonic glider powered by either rocket or air-breathing rocket cycle propulsion that can operate as air-breathing propulsion to mach 5.5 or less. A versatile variable-capture, inward-turning inlet4 can be integrated with the vehicle configuration derived from the FDL series of hypersonic gliders developed by the U.S. Air Force Flight Dynamics Laboratory (AFFDL) 5 and the work of the McDonnell Douglas Astronautics Company. Because of the mass ratio to orbit, these configurations are vertical takeoff and horizontal landing vehicles, exemplified by the upper-left vehicle in Figure 4. This vehicle is usually an upper stage in a two-stage-to-orbit rather than a single-stage-to-orbit vehicle. UNCLASSIFIED//EAR OFFICIO I. Uiliii 0PUi!lf 3 UNCLASSIFIED/JFOR OFFICIJltL YSE 8PUslf Rocket Powered Hypersonic Glider Combined Cycle Airbreather Powered Hypersonic Cruiser Figure 4. Hypersonic Rocket-Powered Glider and Hypersonic Air-Breathing Cruiser The second configuration is for air-breathing propulsion systems operating at between mach 6 and mach 14 that require a propulsion-configured vehicle, where the underside of the vehicle is an integral part of the propulsion system (forming most of the air­ capturing inlet). This is typified by the lower-right vehicle in Figure 4. The thermally integrated, air-breathing, combined-cycle configuration concept is derived from the McDonnell Douglas (St. Louis) Advanced Design organization. The vehicle concept initially conceived in the late 1950s and early 1960s was an air-breathing propulsion­ configured vehicle accelerated by a main rocket in the aft end of the body, as shown in Figure 3. The vehicle's underside is the propulsion system; the engine is in the engine module. Both basic shapes are functions of tau-that is, for a given planform area, the cross­ sectional distribution is determined by the volume required. Tau was reported in D. Kuchemann's book on supersonic aerodynamics6 as: V 'C = total (1) s1.5 plan The only configuration discussed in the book in any detail is the rocket-powered hypersonic glider. The hypersonic glider has greater near-term potential to become an operational system, considering the failure of the National Aerospace Plane (NASP) to reach a functional hardware stage. UNCLASSIFIED/ /FOR OFFICIO! 1 !SE Ql!L¥ 4 UNCLASSIFIED/fFOR OFFl&IAL YSIE er•tv Whatever goes into orbit must enter the atmosphere many times if it is to be a sustained-use vehicle. If it is to be a commercial vehicle, then the flexibility to land wherever the commercial customers are is essential. Consider how successful FedEx, UPS, or DHL would be if there were only two pickup and delivery sites in the United States and a few more elsewhere in the world. A ballistic capsule has even fewer landing options, and a saltwater landing and recovery is too costly to be commercially feasible. What is needed is a hypersonic glider with the flexibility to enter when necessary, without waiting, and to land at different operational bases, just as a transport might. There were three serious competitors in the United States with respect to hypersonic glider configurations: the AFFDL at Wright-Patterson Air Force Base, the McDonnell Douglas Corporation (MDC), and the Lockheed Corporation. NASA Ames and NASA Langley were also generating hypersonic configurations, but NASA's views on hypersonic gliders (fundamentally research and development projects) and their glide range requirements differed from those of the three organizations listed above. That difference is clearly exemplified by the difference between the operational requirements of an experimental aircraft (such as X-1, X-2, X-10, X-15, or X-20) that flies infrequently and at the convenience of the research organization and those of an operational Air Force or Navy aircraft that must be able to fly on any day in almost any weather when needed (also a Russian spacecraft operational rule). From the middle of the 1960s to the early 1970s, the U.S. Air Force and NASA had disagreements over the operational capability of these aircraft and their requirements. As a result, each went its own development direction, and much of the originality and practicality of the AFFDL concepts has not been reflected in the space access configurations developed by NASA. There was a final attempt to apply the AFFDL's philosophy of a high lift-to-drag (L/D) ratio delta planform configuration to the NASA space shuttle, as detailed in the article "A Delta Shuttle Orbiter" in the January 1971 issue of Astronautics and Aeronautics. 7 Figure 5 shows the array of delta planform configurations the AFFDL considered during the 1958-68 timeframe. Figures. Delta-Lifting Body Designs. Array of delta-lifting body designs shows configuration is not limited to high hypersonic lift-to-drag ratios, high cross range, and large size. 8 UNCLASSIFIED/,<FOR OFFI&IAL YSIE 8HLY 5 UNCLASSIFIED/,CFOA OFFl€1Ak Uliliii ONkY The AFFDL's approach was to design a hypersonic performance configuration that would minimize the waiting time in orbit to return to the continental United States (CONUS). This resulted in configurations with sharper leading edges and smaller nose radii than found in NASA and Russian configurations. All of the material, structural, and thermodynamic details related to the sharper configurations were tested and verified in ground test facilities and flight tests (BGRV and ASSET). Characteristics of selected AFFDL hypersonic glider configurations are identified in Table 1. Table 1. Characteristics of Selected Flight Dynamics Laboratory Hypersonic Glider Configurations During the 1958-68 Timeframe # Model Observation l FDL-24B Flat bottom, sharp leading edges, conventional tails, as designed 2 All Body Glider, similar to Russian BOR vehicles Upnirned spatular nose, conventional tail s 3 ASSET Test vehicle to evaluate aerodynamics, thermodynamics and materi als, based on nose of DynaSoar 4 FDL-7MC Flat bottom, sharp leading edges, variable geometry wing, ex perimentally developed tail X configuration 5 Blunt nose, wing-body DynaSoar type configuration 6 Spaturlar Nose Version of DynaSoar type First integration of 2-dirnensional nose (less drag) on a hypersonic glider (R.D. Newmann) 7 FDL-8 Flat bottom, sharp leading edges, outboard tails 8 HL-10 ASA Ames flat up-swept with bottom, round upper body, high dihedral angle tai ls 9 X-24A NASA Langley round body, high dihedral angle tails LO Star Body based on Russian Star Body type configuration Configuration 2 was a higher wing-loading, relatively blunt all-body with an upswept spatular nose that is not unlike Russia's Bor series of Lozino-Lozinski hypersonic gliders. When the author was at Wright-Patterson, interest in this waned quickly because of the limited cross range available. Because of the longitudinal extent of the former Soviet Union compared with the United States, the minimum L/D ratio to ensure a landing on the continental land mass was less for the former Soviet Union than it was for the United States-1.7 for the Soviet Union versus 2. 7 for the United States. Configuration 3 was a subscale research vehicle to evaluate the thermodynamic and materials for hypersonic gliders. The nose and leading edge radii were full-scale size. ASSET was successfully flown on a Thor intermediate-range ballistic missile (IRBM) booster. One that was recovered after an ocean landing is on display in the U.S. Air Force Museum in Dayton, Ohio. Configuration 6 was the first two-dimensional nose applied to a conventional winged-body (configuration 5) in the United States. Configuration 4 was a product of cooperation between the AFFDL (Alfred Draper) and McDonnell Douglas Astronautics Company (Robert Masek) to develop a vehicle to support the Manned Orbiting Laboratory (MOL). This concept was briefed to the U.S. UNCLASSIFIED//FOR OFFICIO! 1ISliii AIIL¥ 6 UNCLASSIFIED/; FOR OFFICIAL U.!I! er~t'f Air Force in 1964, and elements of that configuration will be shown later. The intent was a 9- to 12-person vehicle for crew rotation that could alternatively carry supplies to the orbital station on a regular, frequent schedule (about one flight per week per vehicle) . The variable geometry switchblade wing permitted landing with heavy loads returning from space and eventually horizontal takeoff. The experimentally determined configuration feature was the tail configuration. This configuration was wind tunnel tested and demonstrated inherent stability and control at speeds ranging from mach 22 to landing speed. Configuration 6 was a product of cooperation between the AFFDL (Richard D. Neumann) and McDonnell Douglas Astronautics Company (Robert Krieger) to reduce the drag of hypersonic gliders. Based on the physics that a two-dimensional wedge has less drag than a right circular cone of the same volume, these engineers devised the "spatular leading edge." The wind-body configuration formed the basis of the X-20 and DynaSoar configurations that had a limited hypersonic L/D ratio, primarily because of drag. With the spatular nose, the nose wave drag could be reduced by 35 to 40 percent, thus increasing the hypersonic L/D ratio. Configuration 6 was derived from the conventional wing body, configuration 5. Configuration 10 is an adaptation of the Russian "Star Body" concept that can enter in one of three orientations and need not always have one side facing the flow (compression side). The theory was that in a damaged situation, one of the three sides would be available for a safe entry. The limitation of this configuration concept is a small internal volume and a high ratio of wetted (surface) area per planform area that reduces the hypersonic L/D ratio. The X-24B was based on the FDL-8 Martin Mar•tt ■ X-23A configuration. The different approaches to hypersonic glider configuration are best exemplified by Figure 6. The X­ 24A, built by Martin Marietta at its Denver, Colorado, facilities, is a round fuselage configuration with outboard high-dihedral-angle vertical tails. All the configurations of this type have serious lateral-directional stability problems at low speeds and tend to roll about the horizontal axis through the fuselage . One designer, the Russian Glebe Lozino-Lozinski, solved the problem by employing variable dihedral Figure 6. Martin Marietta X-24 A& BResearch tails. The AFFDL solved the problem by Gliders. X-24A based on USAF PRIME configuration. using nonround configurations; that is, the quest for high hypersonic L/D ratios led to the solution of the low speed problem. Under an AFFDL program, Martin Marietta modified the X-24A into a flat-bottomed configuration with trailing edge elevens called the X-24B, shown in Figure 6. Comments by Bill Dana, the NASA pilot who flew the X-15 and the X-24A/B, about the change in the slow speed performance of the X-24B confirmed the advantage of the AFFDL approach.9 l' 0[ USA f Mod1lied Th• X·24 I UNCLASSIFIED//EAR AEEJCJOP 1155 All! X 7 UNCLASSIFIED//FOR OFFl&IAl l:ISlii 8,.lY The design parameters that largely determine a spacecraft's weight are its configuration and the amount of wetted or surface area relative to the planform area. The hypersonic gliders shown in Figure 5 have differing values of wetted area to planform area. Another important factor is the presence of wings, such as for configurations 5 and 6, which are wing bodies with a relatively thin wing or no wing, such as the lifting­ body FDL-class hypersonic glider (configurations 2, 4, or 7). In this case the lifting bodies have a shape advantage that reduces the amount of surface area that is thin or subject to high heating. In the 1960s, when the U.S. Air Force's high-performance lifting body was competing with NASA's modest-performance wing body, there was much debate regarding the weight of these lifting concepts compared with that of a ballistic capsule (see Appendix A). At that time, with the large sea-recovery fleets, ballistic capsules were the only entry vehicles in either the United States or the former Soviet Union. A number of studies in the early-to-mid-1960s attempted to rectify and quantify the weight of a lifting entry vehicle compared with a ballistic capsule. In all the discussion in the Mercury, Gemini, and Apollo programs, the cost of the sea recovery was almost taken for granted, so the focus was on the cost of the vehicle itself, not the entire vehicle system. The government assembled a chart representing the relative weight of hypersonic entry systems-from ballistic to high-performance (high L/D ratio) gliders-collected from contractor and government reports. The relative weight was the system weight compared with that of a ballistic capsule with the same payload capacity. The result was a correlation curve that showed the high­ performance wing-body gliders could weigh as much as twice what a comparable payload ballistic capsule weighed. This correlation was based on the L/D ratio of the vehicle. Apollo has an L/D ratio of about 0.5, but the system was still a ballistic vehicle with a very limited cross range. One correlation of the data is: W/W0 =l + 0.1259aL/D)-0.1029aL/D)2 + 0.0621aL/D)3 (2) W = the weight of a ballistjc capsule with the same payload 0 This correlation yields a high hypersonic L/D ratio glider with a weight almost twice that of the ballistic capsule. In this correlation, different configuration concepts were mixed and correlated as a single data set. A report cited in Appendix A (Stephens, 1965) concluded, "Weight factor for lifting spacecraft results primarily from larger surface area and only secondarily from the associated spacecraft environment and may be as large as a factor of two greater than ballistic spacecraft." Engineers at the McDonnell Douglas Astronautics Company examined the database and concluded the large weight impact for a lifting spacecraft was as much a function of the configuration as the L/D ratio. The engineers set out to separate the database into families of like configurations. Where gaps existed, they established a configuration that provided the L/D ratio sought that was based on the configuration rules for that family. Three fam ilies were identified. The SV family configurations were based on circular/elliptical cross-section configurations that were characteristic of the HL-10 and X-24A NASA configuration concepts. The FDL family configurations were based on the trapezoidal delta planform configuration. And the MRS family configurations were based on a McDonnell Douglas modified version of the FDL family, with an emphasis on creating metal-radiative thermal-protection shingles that were flat, thereby reducing the cost of the shingle and perhaps introducing an element of hardware interchangeability. Altogether, 10 configurations from among the 3 families were 8 UNCLASSIFIED//FOR OFFICilAk Wliliii OPtllf UNCLASSIFIED/;'FOR: OFFIEiIAk W&lii &Pn.Y designed, weighed, and performed using the same industrial fabrication capability. The result was a curve representing each configuration family, as shown in Figure 7. The three families are represented by the three parallel straight lines. SV FAMILY 1.5 10 • w 0.5 W - W •(1 L/DJ - 0 +11 .1 - - - ~----- MRS FAMILY 0 1 2 3 Hypersonic LID Figure 7. Detailed Design Analyses Show the Weight Trends are as Much a Function of Configuration Family as Lift-to-Drag Ratio W/W0 = K{l+0.090l!(L/D)] K =1.000 for the SV Family (3) K =0.9297 for the FDL Family K =0.7622 for the MRS Family This illustrates that the configuration and its individual wetted area to planform area can vary as much in their spacecraft weight as they can in their L/D ratio. The high L/D-ratio configurations had weights comparable to same-payload ballistic capsules of: sv = 1.9 FOL= 1.7 MRS =1.4 So the penalty for having a lifting-body configuration is less than expected if the configuration characteristics are taken into consideration in the design and weighing of the spacecraft. In addition, the reason the lifting spacecraft with high performance was considered was to eliminate the need for sea recovery and therefore the cost of a recovery fleet and the damage incurred by the spacecraft in a saltwater landing. The 9 UNCLASSIFIED//iOA: OiilCI0ls. !lili QPlls.V UNCLASSIFIED//FOR OFFI@IAL l:ISE OHL¥ goal was to be able to recover the spacecraft at any airport in CONUS, to eliminate the need for an overseas recovery site, and to eliminate the waiting required until a lower L/D ratio could land in CONUS (up to 14 orbits for the Apollo capsule, or 21 hours). In an emergency, that may be too long. The AFFDL's goal for the spacecraft to support the Manned Orbiting Laboratory was no waiting but to be able to reach CONUS from any arbitrary MOL position in its orbit. This was considered possible in the 1964-65 briefs to the government with respect to MOL, specifically the Model 176 configuration the MDC proposed for the MOL support in 1964. The hypersonic glider based on the FDL-7C and the hypersonic air-breathing aircraft in Figure 8 both have hypersonic L/D ratios in excess of 2.7. In very practical terms, that means unpowered cross ranges in excess of 4,500 nautical miles and down ranges on the order of the Earth's circumference. So these two craft can depart from any location of a low-altitude orbit and land in CONUS or in continental Europe. Both are dynamically stable over the entire glide regime. FDL-7 CID Hypersonic Glider Blended Body Hypersonic Cruiser Figure 8. High-Performance Hypersonic Glide Aircraft. Rocket boost-glide and air-breather cruiser. The wing-body, cylindrical fuselage advocates have strongly criticized the lifting bodies, contending that they are poorer configurations and much more complicated than the conventional-wisdom wing-body configurations (see Figure 9). However, that is far from the truth . The structural specialist sees this configuration as a lightweight propellant tank and assumes it is this consideration that drives the design. Rather, that observation introduces problems for all other technical disciplines that are far more difficult to rectify than a noncylindrical tank or a cylindrical tank in a nonsymmetrical cross section. The lone lifting surface with trailing edge controls introduces control issues just as it did for the space shuttle. UNCLASSIFIED/;'POlt err1e1J1tt U.!I!! er~t I 10 UNCLASSIFIED/fFOA OFFI&iIAk Ulilii ONkY system A ALang l Wing- yl inder nfigu rati n n ep Inadequate lateral-directional stability and control WB-004 Wing training edge control s Unsymmetrical vortices can produce super­ sonic control problems Cylinder heating extends up to 55° to 60° Poor hypersonic lift-to-drag ratio Glide angle of attack Complex very high ~ 45° compared curved TPS to 15° for high UO design provide limited limits glide range Thin wings control capability encounter for a military severe heating maneuver vehicle from both sides i.e. "hot structure" Figure 9. NASA Langley Wing-Body Configuration WB-004 With Generally Critical Areas for Wing Bodies Identified With a high entry angle of attack, the cross flow over the cylinder produces high heating rates beyond the mid-cylinder line. With a lower L/D ratio, the down and cross ranges are limited as to what might be achieved but more in line with NASA one­ missed-orbit criterion. The thin wings are heated on both sides to create added thermal problems, as well as added surface area to increase drag. Al Draper and his team, together with Bob Masek's team at McDonnell Douglas Astronautics, worked long and diligently to arrive at the FDL-7 /Model 176 configurations shown in Figure 10. The insert photo is Dale Reed's model of the FDL­ 7MC radio-controlled model at NASA Dryden. The AFFDL and MDC configurations were inherently stable at all operational angles of attack from at least mach 22 to landing speed. The remainder of this report will focus on the characteristics of t his class of lifting body . The statements in Figu re 10 were all based on wind tunnel data. A real advantage of the trapezoidal shape was not only flat metallic shingles but heating on the sides and upper surface that was at least three-fifths that of the conventional shapes. The glide range was such that this configuration could land in CONUS from any location on any inclination orbit from its current orbit with no waiting. UNCLASSIFIED//fOR OFFI&il.lJ.k 11ilii CNP X 11 UNCLASSIFIED/ /POR: OPPIEIICL t:191!!! 9HL'I Lifting body configuration AFFDL-7C/D Lifting Body Flight tested nose eliminates thin wings with tip design assures Configuration Concept wing-like lift. High lift-to-drag 100 °C temperature ratio and global glide range All-flying control surfaces provide stability and control from Mach 02 to 22 Heat pipe leading edges assures high LID leading edge radii. Flat pane TPS system Trapezoidal cross section increases lift-to-drag ratio and reduces side Anhedral fixed fins and upper surface with trailing edge heating by 1/3 flaps assure subsonic lateral-directional stability and control Figure 10. FDL-7C/D and FDL-7MC Lifting-Body Configuration. Offers inherent stability and control with sufficient volume and high lift-to-drag ratio. The switchblade wing version of the FDL-7MC was the preferred version for 1983 studies that were part of the McDonnell Douglas TAV (transatmospheric vehicle) effort; that vehicle was powered by either an Aerojet Sacramento air turboramjet or an air­ breathing rocket propulsion system. The inward-turning, variable-capture area inlet10 provides the correct engine airflow from landing speeds to mach 5.5, as illustrated in Figure 11. Figure 11. FDL-7C/D with a DuPont Retractable Inward-Turning Inlet The propellant tanks were cylindrical-segment, multilobe structures with bulkheads and stringers to support the flat, metal-radiative thermal-protection shingles (very similar to those fabricated by Goodrich Aerospace for the now-defunct X-33). The nose was UNCLASSIFIED/lFOR. OFFI&I.t.k Wliliii 8Plk¥ 12 UNCLASSIFIED/ ;5P81l Offl@IAL YSli QPU,¥ transpiration cooled with a low-rate water-porous spherical nose. The sharp leading edges (the same leading edge radius was used for the nose tip) were liquid-metal heat pipes. This approach was tested successfully during the 1964-68 timeframe and was found to be equal in weight and far more durable than a comparable ceramic tile/carbon-carbon system. The AFFDL's experience with carbon-carbon leading edges on the ASSET test vehicle convinced the Air Force it needed a more durable solution. Figure 12 compares the FDL-7C/D and the McDonnell Douglas Model 176. The Model 176 had a power law nose that was essentially a curved spatular nose and a higher sweep angle, resulting in the same usable volume but with a higher hypersonic L/D. Sacrificed were the flat-panel thermal-protection shingles over part of the fore body. Both configurations retained the X-tail configuration developed by Gil Gaumer of McDonnell Douglas. At the time, the launch vehicle would have been a Martin Titan IIIC. Had an engine with the performance of the Pratt & Whitney XLR-129 been available, there would have been lateral recoverable, fuel/oxidizer tanks on either side of the vehicle, with all of the engines installed in the hypersonic glider. This was similar to the Lockheed Star Clipper (see Figure 38). FDL-7 CID &a e a e&a circa 1962 Modified 5 meters ..,., Gffl Wd1 circa 1964 5 meters Model 176 Modified Figure 12. Comparison of FDL-7C/D (top) and Model 176 {bottom) UNCLASSIFIED/ ,«FOil OPFISl.t.L Y&liii 0PIL¥ 13 UNCLASSIFIED/}FOR OFFI@IAL 1481: 8HLY The author was aware of three people-James S. McDonnell, the AFFDL's Albert Draper, and Russia's Glebe Lozino-Lozinski-who clearly understood the need for a long cross­ range and down-range capability, not just for one missed orbit. Critics will observe that Lozino-Lozinski had limited his BOR vehicles to an L/D ratio of 1.7 to 1.8 and not the 2.7 to 3.0 required for Earth circumferential glide range. First, the longitudinal extent of the former Soviet Union was twice that of CONUS, and an Earth circumferential glide range was not necessary to ensure recovery within the continental Soviet Union; therefore, a lesser L/D ratio was acceptable. Second, in personal conversations with the author, Lozino-Lozinski indicated a Russian government agency forced him to limit the glide range to ensure recovery in continental Russia and prevent escape to the United States. In a further step to prevent escape, when the vehicle was in range of CONUS, ground control disabled its deorbit system. The need for a long cross-range and down-range capability so there is no waiting in orbit in the case of an emergency or military need is presented graphically in Figure 13. Interestingly, the greatest lateral-range (cross-range) requirement for no waiting is for 55° orbital inclination, the usual Russian orbital inclination. The nominal U.S. orbital inclination is 28.5 degrees, with a waiting time of 8 orbits (approximately 12 hours) for a space shuttle-class glider. At the International Space Station orbital inclination, the orbital waiting time for a shuttle-class glider is 6 orbits. In comparison, Apollo's orbital waiting time was about 14 orbits, provided the return trajectory included an Earth­ parking orbit before entry into the Earth's atmosphere. The FDL-7 and Model 176 class of gliders could immediately enter a return glide from their orbits. This provides a significant advantage for the International Space Station operators and vehicle crew, who need only enter hypersonic gliders attached to an orbital station and initiate deorbit procedures to be on the ground in less than 90 minutes in an emergency. Time (orbits) Required to Land in California or Florida Orbital Altitude is 200 nautical miles, 230.1 statute miles, 370.4 km 100 -----------------------~ ,..... .,"' ., 6b ~ '--' Minim um later I Range For Once A Da Landing 60 1-+----->d------'ri------+-r----- -~------i <l 0 '.d "' .5 u ..£1 40 l-l------',.,__---l-nao.~ ~---+------+---+--i'-------.4--------l 20 1---""<,-"""'d------+--+---+---,<---lo"'--- Eanlope ----1 ] '.S... 0 Mo•el 176 Operadq Apollo 0'---------'-----=::..L------L------'-----' 0 2000 3000 4000 5000 Spacecraft Lateral Range (nautical miles) Figure 13. Sufficient Cross Range {L/D) Means There is No Waiting to Return UNCLASSIFIED/fFOlil &FFI@IAL U.!I!! t>IIL I 14 UNCLASSIFIED//POil OFFIEiIAk Ulilii QPIL¥ Figure 14 compares the glide ranges and L/D ratios of selected hypersonic gliders. The shuttle would be in the area of the nominal wing body on the righthand chart. From Figure 13, the no-waiting cross range (lateral range) is 3,600 to 4,400 nautical miles. That means the hypersonic L/D ratio needs to be in the 2.7 to 3.2 range. Appendix D has down and lateral ranges shown as landing ellipses. The key to a successful landing is a subsonic L/D ratio that is in the 4.5-or-greater range. The NASA round-bottom configurations were not capable of that subsonic L/D ratio. The X-24B was in that category and was therefore easily landed compared with the X-24A or Prime vehicles. These high-performance gliders were unique to the AFFDL. The intent in case of a fire would be to immediately evacuate to the hypersonic gliders and then depressurize the station to control any fire (remember that Mercury and Gemini could be depressurized). A crew could then be launched to recover the operation and repair the station. j ·e <O a ::i .S ~ -~ § (I,) ~ ij ~ >­ o::: I C ei a 30000 •· 20,0CJ0 1----+---+---,..,,...--i1----i1---1 Nominal Lifting Body 10,000 1----1■ --------+-----I 0 .....,___..___.___...___..____..____ 4.0 T MDC data 2.0 t---t----t--:::~ -t-'---+---1 1.0 ~ =--....-. 1----1--- Apollo + Par 0 0 2,000 4,000 6,000 2.0 3.0 4.0 5.0 6.0 7.0 Lateral Ran e (nautical miles) Su sonic UD Figure 14. Hypersonic Glider Characteristics There is always the question about landing these blended bodies. Figure 15 shows the blended body handling qualities to be very good, and therefore a pilot's fear factor is less than with the X-15. Bill Dana said the X-24A was difficult to handle when landing but that he could land the X-24B almost with no hands as it flared automatically. To the author's knowledge, all of the wind tunnel tests showed inherent static and dynamic stability over the entire speed range. UNCLASSIFIED//FOR AEEJCJOL flilii QPIL¥ 15 UNCLASSIFIED//P91t 9FFl@IAL YSE 8HLY Minim um Acceptable Initial Speed Based on Accelerated Stall M argin Psychological LimitDue to Excessive Initial Speed Limit Due to Margin Fl are Capability 0 100 120 140 160 180 200 220 240 260 Fl are Initiation Speed (percent of minimum spee d) %Vmin Figure 15. Both Delta Planform Lifting Body {Dynasoar) and Model 176 Offer Superior Landing Performance Thermodynamics and Materials The structure of Model 176 was based on diffusion-bonding and super-plastic forming of flat titanium sheets. Forty years ago, the method was called "roll bonding" and executed with the titanium sealed within an evacuated steel envelope and 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 super-plastic forming industrial capabilities. The photo in Figure 16 is from a Society of Automotive Engineers book titled Advanced Engine Development at Pratt & Whitney: The Inside Story of Eight Special Projects 146­ 1971, by Dick Mulready. Chapter 6, "Boost Glide and the XLR-129 - mach 20 at 200,000 Feet," shows a surviving remnant from the 1960s program. The super-plastic-forming, diffusion bonding that was so difficult in early 1960 is now an accepted fabrication procedure. One of the F-1 S's major bulkheads was fabricated from titanium sheet elements using this procedure instead of machining away more than 90 percent of a titanium forging. Had the procedure been adopted as a product­ manufacturing method, it would have eliminated the almost 2-year manufacturing cycle in acquiring titanium forging of the wing spars and major bulkheads. Note that only 2.5 percent of the thermal heating enters the primary structure. 16 UNCLASSIFIED/ 'EAR AEEJCJOP 1155 All! X • ---------------------- ---------------------- ---------------------- ---------------------- ---------------------- UNCLASSIFIED//POil 8FFIEiIAk Wlilii QJslL¥ Aero Heating 97.5% Radiated to Space metal shingle vacuum multi­ layer insulation - , ---------------------­ --------------------- ,--------------------- diffusion bonded super-plastically formed structure l-1gure 6.i Mel onnc 11 111m111m1 into 1rt' (('m1rrr1y of.!"Im Rohwn1 Figure 16. McDonnell Aircraft Company Roll-Bonded Titanium Structure (circa 1963). Today this structure would be super-plastically formed and diffusion bonded from RSR titaniu m sheets. u Mulready's book mentions the McDonnell Douglas boost-glide strategic vehicle, as well as citing key personnel at McDonnell Aircraft Company. Low thermal conductivity standoffs set off the insulated-metal, thermal-protection-insulated shingles from this wall so that there was an air gap between them. The X-33 applied the metal shingle concept, albeit with significant improvement in the standoff design and thermal leakage, in the orientation, 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 thermal protection system than ceramic tiles. All efforts by the author to obtain information on the shingles manufactured by Goodrich Aerospace have been met with the response, "We lost the contract and are investing in more productive products." The titanium diffusion-bonded and super-plastically formed wall was both the primary aircraft structure and the propellant 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. Significant testing of this structural approach confirmed its superior capabilities as a hypersonic radiation-cooled structure. The Model 176 was proposed for the Manned Orbiting Laboratory. It was a thoroughly designed and tested configuration with a complete, all-metal thermal-protection system that had the same weight as ceramic tile and carbon-carbon concepts used for the U.S. space shuttle but was sturdier and could be repaired in a hanger or in orbit. A wind UNCLASSIFIED//FOR. OFFI&IAk W&IE &Ptklf 17 UNCLASSIFIED/ /POR: err1e1J1tt U.!I!! Oflt I tunnel model of the McDonnell Douglas Astronautics Company Model 176 installed in the McDonnell Aircra~ Company Hypersonic Impulse Tunnel for a heat transfer mapping test is shown in Figure 17. Note that, conforming to the piloting concepts of the 1960s, it has a clearly distinct windshield . The model accomplished thermal mapping to determine the heat transfer distributions on the body and upper fins. Figure 17. Model 176 in the McDonnell Douglas Hypervelocity Impulse Tunnel (circa 1964} for Thermographic Phosphor Heat Transfer Mapping, Including the Upper Fin. Among the important determinations that resulted from these heat transfer tests was that the sharp-leading-edge, flat-bottomed, trapezoidal cross section reduced the heating to the sides and upper surfaces, as shown in Figure 18. In the range of angles of attack corresponding to maximum hypersonic L/D ratio, the sharp leading-edge corner separates and reduces the upper surface heating. Because of this separation, the isotherms are parallel to the lower surface and are 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. With nose water transpiration cooling (demonstrated in a flight test in 1966) and heat pipe leading edges (demonstrated at NASA Langley in 1967-68), the temperatures of the nose and leading edges are 212 °F and 1,300 °F (100 °C and 704 °C ), respectively. The thermal mapping enabled identification of primary flow characteristics in the boundary layer of the vehicle. In UNCLASSIFIED//EAR OFFICIO I. Uliliii 0Ptllf 18 --- 1,200°F 1,500° F 800 °F " 3,800°F RLE=1.6 in. 2,200° F UNCLASSIFIED/ /FOR OFFI@IAL l:ISE OHL¥ Figure 18, radiation equilibrium skin temperature is the skin temperature that results when the radiated thermal energy stemming from the skin temperature equals the input aerodynamic heating minus any conduction into the airframe. 900 °F Upper Surface ! l •4,Joo°F Lower ALE= 0.5 in. Surface • 4,200°F 1,ooo·F Figure 18. FDL-7C/D, Model 176 Entry Temperature Distribution. Upper-surface heating is minimized by cross-section geometry. Figure 19 is a thermographic phosphor image of the model in Figure 17 at a 12­ degree angle of attack (maximum L/D ratio) at mach 12. Even at mach 12, there are vortices embedded in the boundary layer. The three black dots are heat transfer gauges that were used to establish the value for qrefin Figure 19. The stagnation heat transfer was 25 times the reference value. This technique, when calibrated with the ,, - no reference heat transfer gauges, provided a rapid and accurate means to determine heat transfer distributions with a minimum of installed gauges. The technique was adopted by other wind tunnel facilities, including the Arnold Engineering Development Center at Tullahoma, Tennessee. Figure 19. Even at Mach 12, Embedded Vortices in the Boundary Layer Alter the Local Heat Transfer UNCLASSIFIED/ /FOR. OFFICI CIs. flili Qllls.¥ 19 UNCLASSIFIED//EOR OEEICI0L: Pi'li ODIL:¥ On other McDonnell hypersonic configurations with all-movable control surfaces, the interface between the fin and the body became a critical heating issue for the rotating shaft attaching the fin to the body. This was an area of concern on this vehicle, and specially instrumented fins were installed to measure the local heating. Again, the thermographic phosphors were used to map the heating. Figure 20 shows the model in Figure 17 at a maximum 48-degree angle of attack. Fin heating distributions were made at 16-, 24-, 34-, and 48-degree angles of attack 12 and are shown in Figure 21. The brighter the phosphor is, the lower its temperature is (the phosphor darkens as the surface temperature increases). So the area adjacent to the body is at a lower temperature than on the fin. In fact, examining Figure 21 shows that for all angles of attack tested, there was always the cool layer adjacent to the body. So the fin attachment journal/shaft would not be a thermal problem. At angles of attack lower than 16 degrees, the heating became less intense. This tail configuration of a fixed anhedral lower fin with trailing edge controls and an all-movable upper fin provided the control authority over the entire mach range required for stability and control and did not have a thermodynamic issue with fin attachment heating. o =48° Figure 20. Thermographic Phosphor Image of Model 176 at Near-Maximum Angle of Attack FIG. 6. ISODENSITRACER maps for lifting body tail fin for four angles ot attack; contours show qr.ocw,/qun1u,,m, Figure 21. From L/D Maximum to Maximum Angle of Attack, There is Always a C