Lawrence D. Huebner

Hyper-X flight engine ground testing for X-43 flight risk reduction

Airframe-integrated scramjet engine testing has been completed at Mach 7 flight conditions in the NASA Langley 8-Foot High Temperature Tunnel as part of the NASA Hyper-X program. This test provided engine performance and operability data, as well as design and database verification, for the Mach 7 flight tests of the Hyper-X research vehicle (X-43), which will provide the first-ever airframe-integrated scramjet data in flight. The Hyper-X Flight Engine, a duplicate Mach 7 X-43 scramjet engine, was mounted on an airframe structure that duplicated the entire three-dimensional propulsion flowpath from the vehicle leading edge to the vehicle trailing edge. This model was also tested to verify and validate the complete flight-like engine system. This paper describes the subsystems that were subjected to flight-like conditions and presents supporting data. The results from this test help to reduce risk for the Mach 7 flights of the X-43. Nomenclature 8-Ft. HTT NASA Langley 8-Foot High Temperature Tunnel AETB Alumina-Enhanced Thermal Barrier AOA, α angle of attack (degrees) BLA Boeing Lightweight Ablator C D drag coefficient C L lift coefficient C M pitching moment coefficient ∆C p,cowl pressure coefficient increment due to cowl opening ∆C p,fuel pressure coefficient increment due to p pressure (psia) Q heat flux (BTU/ft 2 sec) q dynamic pressure SCFM standard cubic feet per minute SiH 4 gaseous silane SiO 2 silicon dioxide T temperature (o R) TPS Thermal Protection System VFS Vehicle Flowpath Simulator (used with the HXFE) VIL Vehicle in the Loop X spindle streamwise location with origin at the rotation point of the horizontal wing X-43 flight vehicle designation for the Hyper-X research vehicles β angle of sideslip (degrees) φ fuel equivalence ratio Subscripts comb facility combustor condition t total condition freestream condition isolator condition downstream of the internal inlet and upstream of the combustor ∞ Introduction NASAs Hyper-X program will advance technologies for vehicles utilizing hypersonic air-breathing propulsion from the laboratory to the flight environment by obtaining data on a hydrogen-fueled, airframe-integrated, dual-mode supersonic combustion ramjet (scramjet) propulsion system in flight. 1 These data will provide the first flight validation of analytical and computational techniques and wind-tunnel test techniques used to design and analyze this class of vehicle. The Hyper-X program is jointly conducted by The flight-test project phase of the program involves the fabrication and flight testing of three unpiloted, autonomous Hyper-X research vehicles, designated X-43. These vehicles are fabricated by a contractor team led by …

Calibration of the Langley 8-Foot High Temperature Tunnel for Hypersonic Airbreathing Propulsion Testing

The NASA Langley 8-Foot High Temperature Tunnel has recently been modified to produce a unique testing capability for hypersonic airbreathing propulsion systems. Prior to these modifications, the facility was used primarily for aerothermal loads and structural verification testing at true flight total enthalpy conditions for Mach numbers between 6 and 7. One of the recent modifications was an oxygen replenishment system which allows operating airbreathing propulsion systems to be tested at true flight total enthalpies. Following the modifications to the facility, calibration runs were performed at total enthalpies corresponding to flight Mach numbers of 6.3 and 6.8 to establish the flow characteristics of the facility with its new capabilities. The results of this calibration, as well as modifications to tunnel combustor hardware prior to calibration to improve tunnel flow quality, are described in this paper.

Propulsion Airframe Integration Test Techniques for Hypersonic Airbreathing Configurations at NASA Langley Research Center (Invited)

ABSTRACT The scope and significance of propulsion airframe integration (PAI) for hypersonic airbreathing vehicles is presented through a discussion of the PAI test techniques utilized at NASA Langley Research Center. Four primary types of PAI model tests utilized at NASA Langley for hypersonic airbreathing vehicles are discussed. The four types of PAI test models examined are the forebody/inlet test model, the partial-width/truncated propulsion flowpath test model, the powered exhaust simulation test model, and the full-length/width propulsion flowpath test model. The test technique for each of these four types of PAI test models is described, and the relevant PAI issues addressed by each test technique are illustrated through the presentation of recent PAI test data. Nomenclature AF axial force ATE aftbody trailing edge C A axial force coefficient C D drag coefficient C L lift coefficient C M pitch moment coefficient C N normal force coefficient C P pressure coefficient NF normal force M Mach number P static pressure PM pitch moment q dynamic pressure SNPR static nozzle pressure ratio Xaft axial distance from CTE α, ΑΟΑ angle of attack γ ratio of specific heats

EXHAUST SIMULATION TESTING OF A HYPERSONIC AIRBREATHING MODEL AT TRANSONIC SPEEDS

An experimental study was performed to examine jet -effects for an airframe -integrated, scramjet -rocket combined -cycle vehicle configuration at transonic test conditions. This investigation was performed by testing an exi sting exhaust simulation wind tunnel model, known as Model 5B, in the NASA Langley 16 -Ft. Transonic Tunnel. Tests were conducted at freestream Mach numbers from 0.7 to 1.2, at angles of attack from -2 to +14 degrees, and at up to seven nozzle static press ure ratio values for a set of horizontal -tail and body -flap deflections. The model aftbody, horizontal tails, and body flaps were extensively pressure instrumented to provide an understanding of jet -effects and control -surface/plume interactions, as well as for the development of analytical methodologies and calibration of computational fluid dynamic codes to predict this type of flow phenomenon. At all transonic test conditions examined, the exhaust flow at the exit of the internal nozzle was over -expand ed, generating an exhaust plume that turned toward the aftbody. Pressure contour plots for the aftbody of Model 5B are presented for freestream transonic Mach numbers of 0.70, 0.95, and 1.20. These pressure data, along with shadowgraph images, indicated the impingement of an internal plume shock and at least one reflected shock onto the aftbody for all transonic conditions tested. These results also provided evidence of the highly three -dimensional nature of the aftbody exhaust flowfield. Parametric tes ting showed that angle -of -attack, static nozzle pressure ratio, and freestream Mach number all affected the exhaust -plume size, exhaust -flowfield shock structure, and the aftbody -pressure distribution, with Mach number having the largest effect. Integrati on of the aftbody pressure data showed large variations in the pitching moment throughout the transonic regime.