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  • 1
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    Dual-Mode Combustion of Hydrogen in a Mach 5, Continuous-Flow Facility (2001)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Results of an experimental and numerical study of a dual-mode scramjet combustor are reported. The experiment consisted of a direct-connect test of a Mach 2 hydrogen-air combustor with a single unswept-ramp fuel injector. The flow stagnation enthalpy simulated a flight Mach number of 5. Measurements were obtained using conventional wall instrumentation and a particle-imaging laser diagnostic technique. The particle imaging was enabled through the development of a new apparatus for seeding fine silicon dioxide particles into the combustor fuel stream. Numerical simulations of the combustor were performed using the GASP code. The modeling, and much of the experimental work, focused on the supersonic combustion mode. Reasonable agreement was observed between experimental and numerical wall pressure distributions. However, the numerical model was unable to predict accurately the effects of combustion on the fuel plume size, penetration, shape, and axial growth.
    Keywords: Inorganic, Organic and Physical Chemistry
    Type: Journal of Propulsion and Power; 17; 6; 1313-1318
    Format: application/pdf
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  • 2
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    Velocity Measurement in a Dual-Mode Supersonic Combustor using Particle Image Velocimetry (2001)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Temporally and spatially-resolved, two-component measurements of velocity in a supersonic hydrogen-air combustor are reported. The combustor had a single unswept ramp fuel injector and operated with an inlet Mach number of 2 and a flow total temperature approaching 1200 K. The experiment simulated the mixing and combustion processes of a dual-mode scramjet operating at a flight Mach number near 5. The velocity measurements were obtained by seeding the fuel with alumina particles and performing Particle Image Velocimetry on the mixing and combustion wake of the ramp injector. To assess the effects of combustion on the fuel air-mixing process, the distribution of time-averaged velocity and relative turbulence intensity was determined for the cases of fuel-air mixing and fuel-air reacting. Relative to the mixing case, the near field core velocity of the reacting fuel jet had a slower streamwise decay. In the far field, downstream of 4 to 6 ramp heights from the ramp base, the heat release of combustion resulted in decreased flow velocity and increased turbulence levels. The reacting measurements were also compared with a computational fluid dynamics solution of the flow field. Numerically predicted velocity magnitudes were higher than that measured and the jet penetration was lower.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: AIAA Paper 2001-1761 , AIAA/NAL-NASDA-ISAS 10th International Space Planes and Hypersonic Systems and Technologies Conference; Apr 24, 2001 - Apr 27, 2001; Kyoto; Japan
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  • 3
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    Flowfield investigation of a supercruise fighter model (1985)
    In:  Other Sources
    Details
    Publication Date: 2011-08-19
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft (ISSN 0021-8669); 22; 289-295
    Format: text
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  • 4
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    Particle generation experience in Langley's 16-foot transonic tunnel (1985)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: The Langley 16-Foot Transonic Tunnel is a single-return, continuous-flow, atmospheric tunnel which uses air exchange for cooling. The tunnel speed is continuously variable from Mach 0 to 1.30. The test section is a regular octagon in cross-section with slots at the corners of the octagon. The laser velocimeter optics, laser, photomultipliers, etc. are mounted in the test section plenum chamber which surrounds the test section. The test volume is approximately a 1-meter cube about the tunnel center line centered on tunnel station 40.84 meters. The seeding system particle generators are mounted on the upstream side of the fourth set of turning vanes approximately 49 meters upstream of the test volume. At this point the tunnel has a diameter of 17.68 meters which meant that installation of the generators was a difficult procedure. It might be noted that particle generation had to be continuous as when the generators were turned off the data rate rapidly deteriorated to zero.
    Keywords: INSTRUMENTATION AND PHOTOGRAPHY
    Type: Wind Tunnel Seeding Systems for Laser Velocimeters; p 149-167
    Format: application/pdf
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  • 5
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    Modification to the Langley 8-foot high temperature tunnel for hypersonic propulsion testing (1987)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Described are the modifications currently under way to the Langley 8-Foot High Temperature Tunnel to produce a new, unique national resource for testing 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 (4 and 5) 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.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NASA-TM-100486 , NAS 1.15:100486 , AIAA PAPER 87-1887
    Format: application/pdf
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  • 6
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    Investigation of a supersonic cruise fighter model flow field (1985)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An investigation was conducted in the Langley 16-Foot Transonic Tunnel to survey the flow field around a model of a supersonic cruise fighter configuration. Local values of angle of attack, side flow, Mach number, and total pressure ratio were measured with a single multi-holed probe in three survey areas on a model previously used for nacelle/nozzle integration investigations. The investigation was conducted at Mach numbers of 0.6, 0.9, and 1.2, and at angles of attack from 0 deg to 10 deg. The purpose of the investigation was to provide a base of experimental data with which theoretically determined data can be compared. To that end the data are presented in tables as well as graphically, and a complete description of the model geometry is included as fuselage cross sections and wing span stations. Measured local angles of attack were generally greater than free stream angle of attack above the wing and generally smaller below. There were large spanwise local angle-of-attack and side flow gradients above the wing at the higher free stream angles of attack.
    Keywords: AERODYNAMICS
    Type: NASA-TM-86361 , L-15884 , NAS 1.15:86361
    Format: application/pdf
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  • 7
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    Performance Test of Laser Velocimeter System for the Langley 16-foot Transonic Tunnel (1985)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An investigation in the Langley 16-Foot Transonic Tunnel has been conducted in which a laser velocimeter was used to measure free-stream velocities from Mach 0.1 to 1.0 and the flow velocities along the stagnating streamline of a hemisphere-cylinder model at Mach 0.8 and 1.0. The flow velocity was also measured at Mach 1.0 along the line 0.533 model diameters below the model. These tests determined the performance characteristics of the dedicated two-component laser velocimeter at flow velocities up to Mach 1.0 and the effects of the wind tunnel environment on the particle-generating system and on the resulting size of the generated particles. To determine these characteristics, the measured particle velocities along the stagnating streamline at the two Mach numbers were compared with the theoretically predicted gas and particle velocities calculated using a transonic potential flow method. Through this comparison the mean detectable particle size (2.1 micron) along with the standard deviation of the detectable particles (0.76 micron) was determined; thus the performance characteristics of the laser velocimeter were established.
    Keywords: LASERS AND MASERS
    Type: NASA-TP-2502 , L-15940 , NAS 1.60:2502 , AVSCOM-TR-85-B-4
    Format: application/pdf
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  • 8
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    Effect of thrust vectoring and wing maneuver devices on transonic aeropropulsive characteristics of a supersonic fighter (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The aeropropulsive characteristics of an advanced fighter designed for supersonic cruise were determined in the Langley 16-Foot Transonic Tunnel. The objectives of this investigation were to evaluate the interactive effects of thrust vectoring and wing maneuver devices on lift and drag and to determine trim characteristics. The wing maneuver devices consisted of a drooped leading edge and a trailing-edge flap. Thrust vectoring was accomplished with two dimensional (nonaxisymmetric) convergent-divergent nozzles located below the wing in two single-engine podded nacelles. A canard was utilized for trim. Thrust vector angles of 0 deg, 15 deg, and 30 deg were tested in combination with a drooped wing leading edge and with wing trailing-edge flap deflections up to 30 deg. This investigation was conducted at Mach numbers from 0.60 to 1.20, at angles of attack from 0 deg to 20 deg, and at nozzle pressure ratios from about 1 (jet off) to 10. Reynolds number based on mean aerodynamic chord varied from 9.24 x 10 to the 6th to 10.56 x 10 to the 6th.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2119 , L-15526 , NAS 1.60:2119
    Format: application/pdf
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  • 9
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    Flow-field investigation of a supercruise fighter model (1984)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: A NASA Langley investigation was conducted in the 16-foot Transonic Tunnel to survey the flow field around a model of a Supersonic cruise fighter configuration. In this investigation, a model of a supersonic cruise fighter configuration formerly utilized in afterbody-nozzle performance investigations was surveyed with a single, multiholed probe to determine local values of angle of attack, side flow, and Mach number. The investigation was conducted at Mach numbers of 0.6, 0.9, and 1.2 at angles of attack from 0 to 10 deg. The purpose of the investigation was to provide a data base of experimental data for use in verification of theoretical methods, and to compare the experimental data with predictions from currently available theoretical techniques. Results from this investigation show that local angles of attack were generally greater than free stream above the wing and generally less than free stream below the wing. Also there were large spanwise gradients above the wing at the higher angles of attack. The comparisons of experimental data with theoretical predictions show that the theoretical techniques give a qualitative estimate of the flow-field but will require much work to give good quantitative results.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-1331
    Format: text
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  • 10
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    Effect of Reynolds number on the subsonic boattail drag of several wing-body configurations (1976)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An investigation was conducted in a transonic cryogenic tunnel to determine the effect of varying Reynolds number on the boattail drag of several wing-body configurations. This study was made at 0 deg angle of attack at Mach numbers from 0.6 to 0.9 for Reynolds numbers up to 67 x 1 million (based on distance from the nose to the start of the boattail). Results indicate that as the Reynolds number was increased the boattail static pressure coefficients in the expansion region of the boattail became more negative while those in the recompression region became more positive. Results show that there was only a small effect of Reynolds number of boattail pressure drag.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-8238 , L-10853
    Format: application/pdf
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