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Reubush, David E.

Publications and source records attributed to Reubush, David E..

15 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS2004

Numerical Investigation of Dual-Mode Scramjet Combustor with Large Upstream Interaction

Dual-mode scramjet combustor configuration with significant upstream interaction is investigated numerically, The possibility of scaling the domain to accelerate the convergence and reduce the computational time is explored. The supersonic combustor configuration was selected to provide an understanding of key features of upstream interaction and to identify physical and numerical issues relating to modeling of dual-mode configurations. The numerical analysis was performed with vitiated air at freestream Math number of 2.5 using hydrogen as the sonic injectant. Results are presented for two-dimensional models and a three-dimensional jet-to-jet symmetric geometry. Comparisons are made with experimental results. Two-dimensional and three-dimensional results show substantial oblique shock train reaching upstream of the fuel injectors. Flow characteristics slow numerical convergence, while the upstream interaction slowly increases with further iterations. As the flow field develops, the symmetric assumption breaks down. A large separation zone develops and extends further upstream of the step. This asymmetric flow structure is not seen in the experimental data. Results obtained using a sub-scale domain (both two-dimensional and three-dimensional) qualitatively recover the flow physics obtained from full-scale simulations. All results show that numerical modeling using a scaled geometry provides good agreement with full-scale numerical results and experimental results for this configuration. This study supports the argument that numerical scaling is useful in simulating dual-mode scramjet combustor flowfields and could provide an excellent convergence acceleration technique for dual-mode simulations.

Mohieldin, T. O.↗
NASA NTRS2004

Review of X-43A Return to Flight Activities and Current Status

This paper provides an overview and status of the return to flight activities for the X-43A scramjet flight demonstrator after the first flight mishap. The first flight was attempted on June 2, 2001 and resulted in vehicle destruction by range safety when the booster went out of control early in the flight. In the time since the mishap much work has been done to examine the causes of the failure and make modifications to the booster to insure that the boost for the second flight will be successful. In addition, all other aspects of the flight have been examined to maximize the probability of a successful flight.

Reubush, David E.↗
NASA NTRS2001

Hyper-X Program Status

This paper provides an overview of the objectives and status of the Hyper-X program, which is tailored to move hypersonic, airbreathing vehicle technology from the laboratory environment to the flight environment. The first Hyper-X research vehicle (HXRV), designated X-43, is being prepared at the Dryden Flight Research Center for flight at Mach 7. Extensive risk reduction activities for the first flight are completed, and non-recurring design activities for the Mach 10 X-43 (third flight) are nearing completion. The Mach 7 flight of the X-43, in the spring of 2001, will be the first flight of an airframe-integrated scramjet-powered vehicle. The Hyper-X program is continuing to plan follow-on activities to focus an orderly continuation of hypersonic technology development through flight research.

McClinton, Charles R.↗
NASA NTRS2001

Hyper-X Stage Separation: Simulation Development and Results

This paper provides an overview of stage separation simulation development and results for NASA's Hyper-X program; a focused hypersonic technology effort designed to move hypersonic, airbreathing vehicle technology from the laboratory environment to the flight environment. This paper presents an account of the development of the current 14 degree of freedom stage separation simulation tool (SepSim) and results from use of the tool in a Monte Carlo analysis to evaluate the risk of failure for the separation event. Results from use of the tool show that there is only a very small risk of failure in the separation event.

Reubush, David E.↗
NASA NTRS1999

Hyper-X Stage Separation: Background and Status

This paper provides an overview of stage separation activities for NASA's Hyper-X program; a focused hypersonic technology effort designed to move hypersonic, airbreathing vehicle technology from the laboratory environment to the flight environment. This paper presents an account of the development of the current stage separation concept, highlights of wind tunnel experiments and computational fluid dynamics investigations being conducted to define the separation event, results from ground tests of separation hardware, schedule and status. Substantial work has been completed toward reducing the risk associated with stage separation.

Reubush, David E.↗
NASA NTRS1996

High Reynolds number analysis of an axisymmetric afterbody with flow separation

The ability of a three-dimensional Navier-Stokes method, PAB3D, to predict nozzle afterbody flow at high Reynolds number was assessed. Predicted surface pressure coefficient distributions and integrated afterbody drag are compared with experimental data obtained from the NASA-Langley 0.3 m Transonic Cryogenic Tunnel. Predicted afterbody surface pressures matched experimental data fairly closely. The change in the pressure coefficient distribution with Reynolds number was slightly over-predicted. Integrated afterbody drag was typically high compared to the experimental data. The change in afterbody pressure drag with Reynolds number was fairly small. The predicted point of flow separation on the nozzle was slightly downstream of that observed from oilflow data at low Reynolds numbers and had a very slight Reynolds number dependence, moving slightly further downstream as Reynolds number increased.

Carlson, John R.↗
NASA NTRS1992

A historical perspective on hypersonic research at the NACA/NASA Langley Research Center (1944-1984)

A survey of some of the highlights on hypersonic research and technology development conducted at NASA Langley Research Center are presented. Attention is given to the range of disciplines that are being and have been explored at Langley. This review is also intended to supplement previous surveys written around one particular discipline while this survey covers a range of disciplines.

Reubush, David E.↗
NASA NTRS1992

Flow field over the wing of a delta-wing fighter model with vortex control devices at Mach 0.6 to 1.2

As part of a cooperative research program between NASA, McDonnell Douglas Corporation, and Wright Research and Development Center, a flow field investigation was conducted on a 7.52 percent scale windtunnel model of an advanced fighter aircraft design. The investigation was conducted in the Langley 16 ft Transonic Tunnel at Mach numbers of 0.6, 0.9, and 1.2. Angle of attack was varied from -4 degrees to 30 degrees and the model was tested at angles of sideslip of 0, 5, and -5 degrees. Data for the over the wing flow field were obtained at four axial survey stations by the use of six 5 hole conical probes mounted on a survey mechanism. The wing leading edge primary vortex exerted the greatest influence in terms of total pressure loss on the over the wing flow field in the area surveyed. A number of vortex control devices were also investigated. They included two different apex flaps, wing leading edge vortex flaps, and small large wing fences. The vortex flap and both apex flaps were beneficial in controlling the wing leading edge primary vortex.

Bare, E. Ann↗
NASA NTRS1989

Pressure and heat transfer investigation of a modified NASP baseline configuration at M = 6

A cooperative NASA Langley-Boeing investigation was conducted in the Langley eight-Foot High Temperature Tunnel to obtain hypersonic pressure and heat transfer data. In this investigation a large scale (1/20), modified version of the National Aero-Space Plane configuration known as the 'Government Baseline' was tested at a nominal Mach number of 6; at two Reynolds numbers (0.6 and 1.6 million per foot); and at angles of attack from about 0 to 15 deg. There were several purposes for the investigation: to provide a windward and leeward pressure and heat transfer data base for a realistic configuration for verification of computational methods, to provide these data for a large-scale model, and to provide these data for true temperature conditions because of concern about data from low temperature tunnels.

Reubush, David E.↗
NASA NTRS1987

Effects of the installation and operation of jet-exhaust yaw vanes on the longitudinal and lateral-directional characteristics of the F-14 airplane

A wind tunnel investigation was conducted in the Langley 16-Foot Transonic Tunnel to determine the effects of the installation and use of jet exhaust yaw vanes on the longitudinal and lateral-directional characteristics of the F-14 aircraft. The model was tested at Mach numbers from 0.70 to 1.25 at angles of attack from 0 deg to 4.3 deg. Compressed air was used to simulate nozzle exhaust flow from jet off up to a nozzle pressure ratio of 8. The results of the investigation show that the yaw vanes can augment the rudders to provide directional control, but further investigation will be necessary to optimize the deflection schedule associated with the various nozzle power settings.

Reubush, David E.↗
NASA NTRS1987

Static internal performance of a two-dimensional convergent-divergent nozzle with thrust vectoring

A parametric investigation of the static internal performance of multifunction two-dimensional convergent-divergent nozzles has been made in the static test facility of the Langley 16-Foot Transonic Tunnel. All nozzles had a constant throat area and aspect ratio. The effects of upper and lower flap angles, divergent flap length, throat approach angle, sidewall containment, and throat geometry were determined. All nozzles were tested at a thrust vector angle that varied from 5.60 tp 23.00 deg. The nozzle pressure ratio was varied up to 10 for all configurations.

Bare, E. Ann↗
NASA NTRS1987

Effects of afterbody boattail design and empennage arrangement on aeropropulsive characteristics of a twin-engine fighter model at transonic speeds

The effects of empennage arrangement and afterbody boattail design of nonaxisymmetric nozzles on the aeropropulsive characteristics of a twin-engine fighter-type model have been determined in an investigation conducted in the Langley 16-Foot Transonic Tunnel. Three nonaxisymmetric and one twin axisymmetric convergent-divergent nozzle configurations were tested with three different tail arrangements: a two-tail V-shaped arrangement; a staggered, conventional three-tail arrangement; and a four-tail arrangement similar to that on the F-18. Two of the nonaxisymmetric nozzles were also vectorable. Tests were conducted at Mach numbers from 0.60 to 1.20 over an angle-of-attack range from -3 deg to 9 deg. Nozzle pressure ratio was varied from 1 (jet off) to approximately 12, depending on Mach number. Results indicate that at design nozzle pressure ratio, the medium aspect ratio nozzle (with equal boattail angles on the nozzle sidewalls and upper and lower flaps) had the lowest zero angle of attack drag of the nonaxisymmetric nozzles for all tail configurations at subsonic Mach numbers. The drag levels of the twin axisymmetric nozzles were competitive with those of the medium-aspect-ratio nozzle at subsonic Mach number.

Bangert, Linda S.↗
NASA NTRS1987

Investigation of a delta-wing fighter model flow field at transonic speeds

The paper reports a flow-field investigation on a 7.52-percent scale model of an advanced fighter aircraft design conducted in the NASA-Langley 16-ft Transonic Tunnel. The effects of free-stream Mach number, angle-of-attack, angle of sideslip, and various vortex control devices on the local flow values were studied. The model was tested at Mach numbers of 0.6, 0.9, and 1.2 and the angles of sideslip of 0 and +/- 5 deg; the model angle-of-attack was varied from -4 to 30 deg. Results are presented in terms of contour plots of local total pressure recovery. The dominant influence on the over-wing flow field was found to be the wing leading-edge vortex which first appears in the survey region at an angle-of-attack of 8 deg and increases in strength and influence with increasing angle-of-attack, finally dominating the entire survey region at very high angles-of-attack.

Bare, E. Ann↗
NASA NTRS1987

Modification to the Langley 8-Foot High Temperature Tunnel for hypersonic propulsion testing

This paper describes the modifications currently underway to the Langley 8-Foot High Temperature Tunnel to produce a new, unique national resource for testing of hypersonic air-breathing propulsion systems. The current tunnel, which has been used for aerothermal loads and structures research since its inception, is being modified with the addition of a LOX system to bring the oxygen content of the test medium up to that of air, the addition of alternate Mach number capability to augment the current M = 7 capability, improvements to the tunnel hardware to reduce maintenance downtime, the addition of a hydrogen system to allow the testing of hydrogen powered engines, and a new data system to increase both the quantity and quality of the data obtained. The paper discusses both the modifications and the development thereof.

Reubush, David E.↗