Document
IAC-05-D2.6.01
X-43A HYPERSONIC VEHICLE TECHNOLOGY DEVELOPMENT
Randall T. Voland Lawrence D. Huebner Charles R. McClinton (retired) NASA Langley Research Center, Hampton, VA, USA ABSTRACT NASA recently completed two major programs in Hypersonics: Hyper-X, with the record-breaking flights of the X-43A, and the Next Generation Launch Technology (NGLT) Program. The X-43A flights, the culmination of the Hyper-X Program, were the first-ever examples of a scramjet engine propelling a hypersonic vehicle and provided unique, convincing, detailed flight data required to validate the design tools needed for design and development of future operational hypersonic airbreathing vehicles. Concurrent with Hyper-X, NASA’s NGLT Program focused on technologies needed for future revolutionary launch vehicles. The NGLT was “competed” by NASA in response to the President’s redirection of the agency to space exploration, after making significant progress towards maturing technologies required to enable airbreathing hypersonic launch vehicles. NGLT quantified the benefits, identified technology needs, developed airframe and propulsion technology, chartered a broad University base, and developed detailed plans to mature and validate hypersonic airbreathing technology for space access. NASA is currently in the process of defining plans for a new Hypersonic Technology Program. Details of that plan are not currently available. This paper highlights results from the successful Mach 7 and 10 flights of the X-43A, and the current state of hypersonic technology.
INTRODUCTION Benefits of airbreathing launch systems are safety, mission flexibility, robustness, and operating costs as highlighted in figure 2. Safety benefits result from H. Julian Allen made an important observation in the characteristics such as abort capability and power 1958 25th Wright Brothers Lecture: “Progress in density. Horizontal takeoff and powered landing allows aeronautics has been brought about more by revolu- tionary than evolutionary changes in methods of ability to abort over most of the flight, both ascent and decent. High lift-to-drag ratio (L/D) allows longer-range propulsion.” Steam engines replaced sails and glide for a large landing footprint. Power density, or the introduced mass transportation on the sea and lead to quantity of propellant pumped, is 1/10 that of a vertical the railroads. The internal combustion engine replaced take off rocket due to lower thrust loading (T/W), the horse for private transportation and lead to the smaller vehicle weight and higher specific impulse airplane. The jet engine replaced the piston engine, (Isp), defined as thrust per pound of propellant—a and revolutionized the airliner, taking it routinely above measure of efficiency. Power density is a large factor in the weather and beyond the seas. Modern rockets catastrophic failures. Recent analysis indicates that opened the space age for the bold and wealthy. In the safety increases by several orders of magnitude are 21st century, revolutionary applications of airbreathing possible using airbreathing systems. Mission flexibility propulsion will make space travel routine and interconti- results from horizontal takeoff and landing, the large nental travel as easy as intercity travel today. One of landing (unpowered) footprint and high L/D. Utilization NASA’s focuses today remains the application of revo- of aerodynamic forces rather than thrust allows efficient lutionary propulsion systems for space transportation.
orbital plane changes during ascent, and a wider launch window. Robustness and reliability can be built Figure 1 represents the first application of this into airbreathing systems because of large margins and revolutionary propulsion technology to space access reduced weight-growth sensitivity, and the low thrust missions with an airline-sized space vehicle that serves required for smaller, horizontal takeoff systems. Cost as a high-speed, high-altitude, first-stage launch models under development for airbreathing launch platform for an expendable or reusable rocket-powered systems indicate about one order-of-magnitude second stage . This modest approach significantly reduction in operating cost is possible.
improves launch safety and reliability. In addition, it places the world on a new path to reduced cost for The development of reusable launch vehicles holds space access which cannot be achieved with great promise as the key to unlocking the vast potential evolutionary improvements of current systems.
of space for business exploitation. Only when access to space is assured in a system that provides routine operation with airline-like safety and affordable cost will businesses be willing to take the risks and make the investments necessary to realize this great potential.
Rocket-powered vehicles are approaching their limits in terms of these parameters ; switching to a new approach is the only way to achieve significant improvements. Airbreathing vehicles that are capable of hypersonic speeds can transform access to space, just like turbojets transformed the airline business.
Figure 3. Future Hypersonic Air Vehicles Figure 1. Revolutionary Launch Vehicle Figure 4. Historical Perspective NASA funded scramjet technology development, focused mostly on the propulsion cycle efficiency with numerous ground tests in wind tunnels (fig. 4) over the past 40 years . Starting in the mid 1960’s, NASA built and tested a hydrogen-fueled and -cooled scramjet engine that verified scramjet cycle efficiency, structural integrity, first-generation design tools, and engine system integration. Starting in the early 1970’s, NASA designed and demonstrated a fixed-geometry, airframe-integrated scramjet “flowpath” capable of Figure 2. Benefits of Hypersonic Airbreathing Vehicles propelling a hypersonic vehicle from Mach 4 to 7 in wind tunnel tests. Starting in the mid 1980’s, NASA NASA has funded hypersonic vehicle and propulsion teamed with the Department of Defense in the National technology for over 40 years, aiming at futuristic AeroSpace Plane (NASP) Program to demonstrate space launch capabilities . This work represents the hypersonic technologies required for a hypersonic, next frontier in air vehicle design. Recent U.S. scramjet-based, combined-cycle powered, single- industry focus has been on avionics, stealth, and stage-to-orbit launch vehicle. Under the NASP program methods of making the last generation of aircraft more NASA focused on engine definition and testing for both effective. Evolutionary vehicle changes will give way the aerodynamic lines/cycle efficiency and hydrogen- to revolutionary changes when a new propulsion cooled engine structure.
system is available. Studies by the Next Generation Launch Technology (NGLT) Program reconfirmed that NASA developed the concept for the Hyper-X (X-43A) hypersonic systems using turbine-based “low-speed” Program in 1995-1996 as a result of several blue- engines combined with scramjets for higher speed ribbon panel recommendations that flight experiments operation, up to at least Mach 7 and eventually to of airframe-integrated scramjet-propulsion systems be Mach 13-15, are the preferred long-term approach for the next major step in hypersonic research. The airbreathing space access vehicles. Clearly, these experts agreed that, at a minimum and as a first step, turbine-based systems are also needed for super- a vehicle must fly with an airframe-integrated sonic cruise/hypersonic dash or pure hypersonic supersonic-combustion ramjet (scramjet) propulsion cruise aircraft of the future (fig. 3). system. Consequently NASA initiated the Hyper-X program to provide the flight data to validate the cruise thrust at Mach 9.68. Many technical results design tools and methods to be used in the from these flights will be discussed herein. Details development of future hypersonic vehicles. The on the preparation and execution of these flights are program goals were established to provide perform- addressed in references 14-17.
ance data to reduce development risks for subse- quent, operational vehicles; to advance performance- The second flight trajectory is illustrated in figure 6.
prediction capabilities for airbreathing hypersonic The launch vehicle was dropped from the B-52 flying vehicles; to flight-validate airframe-integrated scramjet at Mach 0.8 and 40,000 feet. The booster ignited after performance and design methods; and to flight- 5-seconds free fall to about 39,500 feet. The launch validate other selected, key technologies. vehicle executed a 1.9g pull-up, followed by a 0.7g pushover to achieve nearly level flight at 95,000 ft.
The Hyper-X team developed the X-43A as a small- Following burnout and stage separation, the engine scale research vehicle to provide flight data for a opened for about 30 seconds: 5 seconds of fuel-off hydrogen-fueled, airframe-integrated scramjet tare, 11 seconds of powered flight (at about Mach engine . In addition, aerodynamic, thermal, structural, 6.83 and dynamic pressure of 980 psf), another 5 guidance, flush-air-data-system, and other data were seconds of unpowered steady tare, followed by 10 to be obtained. Test plans called for boosting each of seconds of parameter identification (PID) maneu- three X-43A research vehicles to the required test vers . The engine cowl then closed, and the vehicle condition by a drop-away booster. The resulting 12' flew a controlled descent over 300 NM to “splash- long vehicle is illustrated in figure 5. The development down” in the Pacific Ocean. The powered portion of of the X-43A and its systems are detailed in this flight included 1.5 seconds for ignition, followed references 5-13. by hydrogen fuel ramp up and ramp down to generate a large database.
The NASA Hyper-X program employed a low-cost approach to design, build, and flight test three small, airframe-integrated scramjet-powered research vehicles at Mach 7 and 10. The research vehicles were dropped from the NASA Dryden B-52, rocket-boosted to test point by a modified Pegasus first stage, separated from the booster, and then operated in autonomous flight.
Tests were conducted at approximately 100,000 ft. at a dynamic pressure of about 1000 psf.
Figure 6. Flight 2 Trajectory The flight 3 trajectory was somewhat different than flight 2. The B-52 flight conditions were the same; however, the launch vehicle executed a 2.5g pull-up to a flight path angle of over 30 degrees, followed by 0.5g push over to achieve nearly level flight at 110,000 ft. Following burnout and stage separation, the engine opened for about 20 seconds: 3 seconds of fuel-off tare, 11 seconds of powered flight (at about Mach 9.68 and dynamic pressure of 930 psf), and another 6 seconds of unpowered steady tare.
Figure 5. X-43A Vehicle Geometry (No cowl-open PID maneuvers were performed due to cowl survival concerns.) The engine cowl closed, FLIGHT RESULTS and the vehicle flew a controlled descent over 800NM to “splash-down” in the Pacific Ocean.
The first Mach 7 flight was attempted June 2, 2001. This During descent, PID maneuvers were successfully flight failed when the launch vehicle went out of control performed at each successively decreasing Mach early in the flight. The second flight occurred March number . The powered portion of flight included 5 27, 2004. The flight was completely successful, and seconds of silane-piloted operation at two fuel demonstrated acceleration of the X-43A vehicle equivalence ratio settings, followed by hydrogen- during climbing flight at Mach 6.83. Preparations for only fueled operation at two settings, then ramp the third flight started before the second flight was down . This conservative approach assured good completed. The third flight occurred November 16, data before trying to operate the small engine 2004. This final flight successfully demonstrated without piloted fuel.
The following discussion highlights flight 2 and flight 3 The data obtained from these flights meets the primary data and results from design tool validation studies. Hyper-X Program objective: validation of design The flight results contain much commonality, so the methods including experimental, analytical and results are presented in terms of flight sequence and computational methods. The data was released to the technology areas, including results from both US hypersonic community in the form of complete 20,21 successful flights. These are designated as F2 for the classified data packages to more than 20 US successful 2nd flight that targeted Mach 7, and F3 for Government and industrial entities and well over 50 the 3rd flight that targeted Mach 10. classified and unclassified papers and presentations.
EXPERIMENTAL MEASUREMENTS X-43A Free Flight The X-43A vehicles were well instrumented, with Following stage separation from the rocket booster over 200 measurements of surface pressure; over (boost and stage separation discussed in references 100 thermocouples to measure surface, structure 23-27), the X-43A F2 research vehicle stabilized to and environmental temperatures; and discrete local 2.5° AOA, and the engine cowl was opened. After 5 strain measurements on the hot wings and tail seconds of tare force measurements, the engine was structure. The flight management unit included a ignited with a pyrophoric silane-hydrogen fuel mixture highly accurate 3-axis measurement of acceleration (left side of fig. 8), then switched to pure hydrogen and rates. In addition, over 500 data words were fuel for about 10 seconds of powered flight. The extracted from the flight computer. Instrumentation contrail from the pure-hydrogen fuel was not visible in density is illustrated in figure 7 by external and the raw image from the High-Altitude Observatory internal wall pressure and temperature on the lower (HALO) aircraft (right side of fig. 8).
body surface. Internal engine instrumentation on the body side is more dense in order to capture internal flow details of the engine.
All of the data from the X-43A flights were successfully telemetered and captured by multiple air and ground stations. The instrumentation health and performance were excellent: very few lost instruments/parameters, extremely low noise content, no significant calibration issues, no significant delay or time-lag issues, and extremely limited TM drop outs. Accuracy of these measurements benefited from day-of-flight atmospheric Figure 8. X-43A Powered Flight Video From Army measurements by weather balloons. These measure- HALO Aircraft ments provide a small change in flight Mach number and dynamic pressure vis-à-vis atmospheric conditions Following powered flight, with the engine cowl open, and winds from historical atmospheric tables. Flight 2 the vehicle performed pre-programmed PID maneuvers Best Estimated Trajectory (BET) resulted in higher to quantify the X-43A aerodynamic stability and control dynamic pressure and Mach number, but only a trivial derivatives . These maneuvers lasted 15 seconds, change in angle of attack (AOA). Flight 3 BET resulted after which the engine cowl was closed, the vehicle in lower dynamic pressure and higher Mach number glided another 300 miles, and splashed into the Pacific and AOA. The flight trajectory reconstruction is Ocean. During the descent of both flights, the vehicle discussed in reference 22.
Figure 7. Lower Surface Instrumentation continued to perform a series of PID maneuvers at The green band in figure 9 illustrates pretest Monte each decreasing Mach number to Mach 2. This Carlo predictions of acceleration (using an unclassified segment of the flight provided a large, unique aerody- representative propulsion database). The heavy blue namic database for sharp leading edge lifting body line depicts flight data trends. The vehicle deceleration configurations that will be used to validate wind tunnel is greater than predicted, both with cowl closed and data and computational tools. open because of two factors: 1) actual flight conditions vs predicted (2/3 of the error), and 2) vehicle Scramjet Powered Vehicle Performance drag is higher than predicted (1/3 of the error). How- ever, the drag was within the uncertainty associated 30, pg. 831 For Mach 7, flight 2 the X-43A was commanded to fly at with the wind tunnel database . The uncertainty 2.5 degrees angle of attack during the cowl-open was not resolved/reduced before flight because it did portion of the flight. However, as the fuel is turned-on/off not threaten the outcome of the engine tests. Flight and the throttle adjusted, the pitching moment changes data will be used to help resolve the wind tunnel significantly. Figure 9 illustrates the measured angle of uncertainty for future flights/missions.
attack—from cowl open to fuel off and the start of the Mach 7 PID maneuvers. During the scramjet-powered Under scramjet power the F2 vehicle acceleration segment, the AOA was maintained to 2.5° ± 0.2°, was positive, and varies with throttle position. The except during flameout. Some efforts were made in the increment in acceleration is about as predicted, which flight control system to provide feed-forward control. For confirms the predicted engine thrust to within less Mach 10, flight 3, the vehicle was commanded to fly at than 2% . It should be noted that the engine throttle 1.0 degree angle of attack during the cowl open was varied over a wide range without engine unstart segment. The vehicle maintained pitch control about the or flame-out. Under scramjet power the F3 vehicle same as during the powered segment of F2. cruised (thrust = drag) at the reference fuel equivalence ratio with 2% silane pilot, and engine For both F2 and F3 the fuel sequencing for powered thrust was in agreement with predictions .
flight starts with a silane/hydrogen mixture to assure ignition, then transitions to pure hydrogen fuel. The The predicted scramjet performance is also confirmed by ignition sequence for F2 requires about 1.5 seconds. the excellent comparison of pre-test predicted and flight With transition to pure hydrogen fuel, the throttle is scramjet flowpath wall pressure (fig. 10). Data is ramped up to either a predetermined or controlled max- presented from vehicle nose to tail for F2 (fig. 10(a)), and imum value, and then decreased as the fuel is depleted from cowl leading edge to cowl trailing edge for F3 (fig.
from the tanks. The resulting vehicle performance is cha- 10(b)). The Mach 7 data is clearly operating in “dual racterized by vehicle acceleration, as shown in figure 9. mode,” with sonic flow in the isolator dissipating the inlet The ignition sequence for F3 was different—the silane shocks. The Mach 10 data exhibits classical pure remains on for the first two fueled conditions, requiring 5 supersonic combustion mode, and the combustor seconds of piloted data. Then the same equivalence pressure is shock dominated. The pretest prediction for ratio test conditions were run with only hydrogen. Mach 7 was made using the SRGULL code, with a) Flight 2, Mach 6.83 b) Flight 3, Mach 9.68 Figure 9. X-43A Axial Acceleration and Angle of Attack During Powered Flight a) F2, Mach 6.83. b) F3, Mach 9.68.
Figure 10. X-43 Flowpath Pressure Distribution: Design Throttle Position combustion efficiency determined by analysis of comparison of internal wall pressure for the 8-Ft HTT multiple wind tunnel tests, most notably, the 8-Foot test of the HXFE on the VFS and is typical of results High Temperature Tunnel (8-Ft. HTT) test of the from other wind tunnel tests. These data show that the Hyper-X Flight Engine (HXFE) on the Vehicle flowfield in the isolator was separated on the cowl side, Flowpath Simulator (VFS). The Mach 10 pretest and featured a supersonic stream, evidenced by shock prediction was performed using a combination of CFD structure on the body side through most of the combus- tools, with the SHIP code used for the combustor. The tor length. In other words, the isolator was not “pushed” reaction efficiency used in the SHIP code was derived very hard for this design condition. Storch discusses from analysis of data from the HYPULSE Scramjet the implication of this agreement, and the impact on 5 31 Module test. Storch and Ferlemann present a observed combustor performance in reference 5.
detailed discussion of these codes and the pretest Rogers reported a similar trend for the Mach 10, flight 3 predictions for F2 and F3 respectively. data . These results show that ground tests are repre- sentative of flight, if careful attention is paid to modeling Wind Tunnel / Flight Scramjet Comparison the appropriate flow phenomena.
Comparison of the wall pressure measured in four wind Aerothermal/Thermo-Structural Analysis and tunnel tests with F2 data is included in references 6 Boundary Layer Transition and 32. Likewise, results from shock tunnel tests are compared with F3 data in reference 33. Tests with The design of the X-43A research vehicle structure nearly identical fuel equivalence ratio were selected for and thermal protection system depended greatly on comparison. Figure 11 illustrates the resulting accurate estimation of the aerothermal environment, Figure 11. X-43A Flowpath Pressure Distribution Comparison With Wind Tunnel Data: Design Throttle Position which required understanding of the boundary layer trajectory” vs. “design trajectory” is continuing to be state during the entire flight. For design purposes, studied. Results to date generally confirm prediction the lower surface flowpath into the inlet was methods. Accuracy of the predictions appears assumed turbulent due to the inclusion of boundary significantly better than the assumed uncertainty for layer trips near the forebody leading edge. The trips both Mach 7 and 10 flights. For example, figure 13 were required to insure the forebody boundary layer illustrates heating to the leading edge of the vehicle was turbulent for mitigation of any flow separations predicted using the “as flown” trajectory, with along the engine flowpath due to adverse pressure measured temperature within the C-C leading edge gradients. A substantial research and design effort material. Adjacent nodes from the FEA model was executed to ensure proper sizing of the trips bracket the measured temperatures over the boost 30, pg. 853 with minimum induced trip drag . The upper and scramjet-powered flight at Mach 10.
surface, however, was predicted to be laminar during the Mach 7 test point, based on a pre-flight trajectory using a classical correlation methodology (momentum thickness Reynolds number over Mach number of 305).
Figure 13. Flight 3 nose LE temperature confirms thermal model Figure 12. Flight 2 X-43A Upper Surface Temperatures from B-52 Drop to Splash Post Test Analysis Figure 12 provides upper surface temperature time Posttest analysis of the flight data is underway at histories during the entire flight 2 trajectory from the NASA Langley Research Center and elsewhere. This point of release from the B-52. The three upper surface is required to model the “as flown” trajectory to properly thermocouples (T/C) were evenly spaced along the assess thermal loads; to assess inlet mass capture at vehicle centerline starting about midpoint for T/C#19 exactly the flight condition evaluated using analytical and ending near the trailing edge for T/C#21. Note that methods; to evaluate the boundary layer state for BLT by the time the cowl opens and the scramjet is ignited, assessment; and to assess the overall vehicle drag, the entire upper surface appears to be laminar, as engine force, and vehicle acceleration/deceleration at indicated by the dramatic temperature decrease that exact flight conditions and control surface positions.
begins at about 70 seconds for the farthest forward T/C Complete nose-to-tail CFD solutions for the actual flight and 85 seconds for the farthest aft T/C. Likewise, at condition are discussed in references 39 and 40. These about 240 seconds the boundary layer transitions from solutions, for cowl closed, cowl open, and powered laminar to turbulent as the vehicle decelerates. These operation show excellent agreement (within a few results are discussed relative to pre-test predictions in percent) with measured acceleration/deceleration in detail in reference 34 along with discussions of the trip flight (fig. 14). They also demonstrate the significant effectiveness on the lower surface.
increase in computational throughput, which permit full 3-D solutions for the entire X-43A vehicle, which were Thermal Loads not possible at earlier stages of the Hyper-X Program.
Preliminary and continuing assessment of thermal Another post-test analysis was to determine the loads for the F2 and F3, compared with design 35-38 engine Isp as tested, and “scale” to a vision vehicle, values, are documented in various sources .
removing effects of small scale, cold fuel, fuel These documents show that the engineering equivalence ratio, operating dynamic pressure, etc.
approach used in the design and development of the This analysis is discussed in reference 41 and flight vehicles was generally conservative. However, unclassified results shown in figure 15. The effective some temperature measurements were higher than impulse developed in this scaling study will certainly anticipated in regions of steep gradients, particularly set the standard for follow-on vehicle configurations.
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33. Rogers, R.C.; Shih, A.T.; and Hass, N.E.: Hyper-X Mach 10 Engine Flowpath Testing: Ground to Flight Comparisons. 28th JANNAF Airbreathing Propulsion Subcommittee Meeting.
Charleston, SC. June 13-17, 2005.
34. Berry, S.A.; Daryabeigi, K.; Auslender, A.H. and Bittner, R.D: Boundary Layer Transition on the X-43A Flight 2. 28th JANNAF Airbreathing Propulsion Subcommittee Meeting. Charleston, SC. June 13-17, 2005.
35. Amundsen, R.M.; Leonard, C.P.; and Bruce, W.E. III: Hyper-X Hot Structures Comparison of Thermal Analysis and Flight Data. 15th Thermal and Fluids Analysis Workshop, Pasadena, CA.
Aug. 30, 2004.
36. Leonard, C.P.; Bruce, W.; and Amundsen, R.M.: Hyper-X Hot Structures Design and Comparison with Flight Data. 28th JANNAF Airbreathing Propulsion Subcommittee Meeting. Charleston, SC. June 13-17, 2005.
37. Cuda, V.; and Woelfel, R.: Comparison of Hyper-X Engine Thermal Analysis and Flight Data. 28th JANNAF Airbreathing Propulsion Subcommittee Meeting. Charleston, SC. June 13-17, 2005.
38. Leonard, C.; Amundsen, R.; and Bruce, W.: Hyper-X Hot Structures Design and Comparison with Flight Data. AIAA 2005-3438. Presented at 13th International Space Planes and Hypersonic Systems and Technology Conference, Capua, Italy. May 2005.
39. Jentink, T.N. and Ferlemann, P.G.: Mach 10 Tip-to-Tail Preflight Prediction, 28th JANNAF Airbreathing Propulsion Subcommittee Meeting.
Charleston, SC. June 13-17, 2005.
40. Meyer, B.M.: Tip to Tail Preflight Prediction.
28th JANNAF Airbreathing Propulsion Subcommittee Meeting. Charleston, SC. June 13-17, 2005.
41. Rock, K.E.; Voland, R.T.; and Witte, D.W.: Hyper-X Mach 7 and 10 Scramjet Operation.
28th JANNAF Airbreathing Propulsion Subcommittee Meeting. Charleston, SC. June 13-17, 2005.