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  • 1
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    Wind tunnel blockage study of a generic three-dimensional sidewall compression scramjet inlet at Mach 10 (1991)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: A large scale model of a generic three-dimensional sidewall compression scramjet inlet has been designed based on the results of a computational parametric study for testing in the 31-inch Mach 10 Hypersonic Wind Tunnel at the NASA Langley Research Center. In order to increase the instrumentation density in interaction regions for a highly instrumented model, it is desirable to make the model as large as possible. When the cross-sectional area of a model becomes large relative to the inviscid core size of the tunnel, the effects of blockage must be considered. In order to assess these effects, a blockage model (an inexpensive, much less densely instrumented version of the configuration) was fabricated for preliminary testing. Since it was desired to determine both the effect of the model on the performance of the wind tunnel and also to determine if the inlet would start, the model possessed a total of 32 static pressure orifices distributed on the forebody plane and sidewalls; seventeen static pressure orifices on the tunnel wall and 3 pitot probes on the model monitored the tunnel performance. This paper presents the design considerations in the development of the wind tunnel model and the blockage aspects of the effects of contraction ratio, cowl location, Reynolds number, and angle of attack.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-0294
    Format: text
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  • 2
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    A computational parametric study of three-dimensional sidewall compression scramjet inlets at Mach 10 (1990)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A computational parametric study of three-dimensional sidewall compression scramjet inlets has been performed. The parameters considered for the study include the leading edge sweep angle, varied between 30 and 60 degrees, and the cowl position, located at the throat and at two forward positions. Additionally, the effects of laminar and turbulent boundary layers as well as adiabatic and cold wall boundary conditions are assessed. The parametric study is performed for a Mach number of 10 and a unit freestream Reynolds number of 2 x 10 to the 6th/ft at a geometric contraction ratio of 5. Comparisons among the various configurations are made in terms of the gross parameters of mass capture, throat Mach number, total pressure recovery, and area weighted internal pressure ratios, as well as more detailed flow field phenomena comparisons. The inclusion of one 0-deg sweep computation indicates that the leading edge may be swept by up to 30 deg before significant changes in the gross flow field parameters occur. It is found that a 45-deg sweep configuration provides a good compromise of mass capture, total pressure recovery, and internal compression.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 90-2131
    Format: text
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  • 3
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    Contraction ratio effects in a generic sidewall compression scramjet inlet - A computational and experimental investigation (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Computed and experimental data on a generic three-dimensional sidewall compression scramjet inlet with a leading edge sweep of 45 degrees at Mach 20 are presented. A three-dimensional Navier-Stokes code was adapted to perform preliminary parametric studies leading to the present configuration. Following the design phase, the code was then employed as an analysis tool to provide a better understanding of the flow field and the experimental static and pitot pressure data. The model possessed 240 static pressure orifices distributed on the forebody plane, sidewalls, and cowl and was tested in the 31 Inch Mach 10 facility at the NASA Langley Research Center. Pitot rakes were employed to map the entrance and exit planes. The contraction ratio was observed to have a dominant effect on the inlet shock structure and performance; these effects are the emphasis of the present report. In addition to pressure measurements, oil flows were used for further visualization and comparison with computation.
    Keywords: AIRCRAFT PROPULSION AND POWER
    Type: AIAA PAPER 91-1708
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  • 4
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    Mach 6 testing of two generic three-dimensional sidewall compression scramjet inlets in tetrafluoromethane (1990)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Three-dimensional sidewall compression scramjet inlets with leading edge sweeps of 30 and 70 degrees have been tested in the Langley Hypersonic CF4 Tunnel at Mach 6 and a ratio of specific heats of 1.2. The effects of cowl position, contraction ratio, and Reynolds number were investigated. The models were instrumented with 42 static pressure orifices distributed on the sidewalls, baseplate, and cowl. Schlieren movies were made of each test for flow visualization of the entrance plane and cowl region. In order to obtain an approximate characterization of the flow field, a modification to two-dimensional inviscid oblique shock theory was derived to accommodate the three-dimensional effects of leading edge sweep. This theory qualitatively predicted the reflected shock structure/sidewall impingement locations and the observed increase in spillage (flow upturning) with increasing leading edge sweep. The primary effect of moving the cowl forward is capturing the flow which would have otherwise spilled out ahead of the cowl. Increasing the contraction ratio (moving the sidewalls closer together) increases the number of internal shock reflections and hence incrementally increases the sidewall pressure distribution. Significant Reynolds number effects were noted over a small range of Reynolds number.
    Keywords: AIRCRAFT PROPULSION AND POWER
    Type: AIAA PAPER 90-0530
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  • 5
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    Internal shock interactions in propulsion/airframe integrated three-dimensional sidewall compression scramjet inlets (1992)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The advantages and design requirements of propulsion/airframe integration for high Mach number flight are studied in terms of the 3D sidewall compression scramjet inlet. The present work addresses in a parametric fashion the inviscid effects of leading edge sweep, sidewall compression, and inflow Mach number on the internal shock structure in terms of inlet compression and mass capture. The source of the Mach number invariance with leading edge sweep for a constant sidewall compression class of inlet is identified, and a previously undocumented spillage phenomenon in a constant effective wedge angle class of inlets is discussed.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 92-3099
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  • 6
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    30 x 60 foot wind tunnel test highlights for an over-the-tail advanced turboprop configuration (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: This paper presents results from a recent investigation of the static aerodynamic and stability characteristics of a two-surface advanced turboprop aircraft. The conceptual design places Hamilton Standard SR-7 turboprop blades close to the horizontal and vertical tail for potential acoustic shielding. Evaluation of the data shows generally favorable effects of power on aircraft stability and control, and that lateral directional trim can be achieved with one engine inoperative. The tests did show a marked effect of the direction of propeller rotation on thrust minus drag performance.
    Keywords: AIRCRAFT DESIGN, TESTING AND PERFORMANCE
    Type: AIAA PAPER 91-0681
    Format: text
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  • 7
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    Limitations of the method of characteristics when applied to axisymmetric hypersonic nozzle design (1990)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A design study of axisymmetric hypersonic wind tunnel nozzles was initiated by NASA Langley Research Center with the objective of improving the flow quality of their ground test facilities. Nozzles for Mach 6 air, Mach 13.5 nitrogen, and Mach 17 nitrogen were designed using the Method of Characteristics/Boundary Layer (MOC/BL) approach and were analyzed with a Navier-Stokes solver. Results of the analysis agreed well with design for the Mach 6 case, but revealed oblique shock waves of increasing strength originating from near the inflection point of the Mach 13.5 and Mach 17 nozzles. The findings indicate that the MOC/BL design method has a fundamental limitation that occurs at some Mach number between 6 an 13.5. In order to define the limitation more exactly and attempt to discover the cause, a parametric study of hypersonic ideal air nozzles designed with the current MOC/BL method was done. Results of this study indicate that, while stagnations conditions have a moderate affect on the upper limit of the method, the method fails at Mach numbers above 8.0.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 90-0192
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  • 8
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    Effects of vibrational nonequilibrium on axisymmetric hypersonic nozzle design (1991)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A vibrational nonequilibrium Navier-Stokes computational algorithm is used to determine the flow conditions in several existing nozzles. Vibration freezes near the throat for typical stagnation conditions in these nozzles. The freezing causes the gas to behave as though the ratio of specific heats is constant. It is shown that the thick boundary layers in hypersonic nozzles create problems in their design using classic techniques. As a result, existing nozzles such as the NASA Langley Mach 17 Nitrogen Tunnel may be underexpanded and operate with poor test section conditions. A design technique based on the Navier-Stokes equations including the effects of vibrational nonequilibrium are required for high quality flow.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-0297
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  • 9
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    Exploratory wind tunnel investigation of the stability and control characteristics of a three-surface, forward-swept wing advanced turboprop model (1990)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: The purpose of the present investigation was to parametrically study the stability and control characteristics of a forward-swept wing three-surface turboprop model through an extended angle of attack range, including the deep-stall region. As part of a joint research program between North Carolina State University and NASA Langley Research Center, a low-speed wind tunnel investigation was conducted with a three-surface, forward-swept wing, aft-mounted, twin-pusher propeller, model, representative of an advanced turboprop configuration. The tests were conducted in the NASA Langley 12-Foot Low-Speed Wind Tunnel. The model parameters varied in the test were horizontal tail location, canard size, sweep and location, and wing position. The model was equipped with air turbines, housed within the nacelles and driven by compressed air, to model turboprop power effects. A three-surface, forward-swept wing configuration that provided satisfactory static longitudinal and lateral/directional stability was identified. The three-surface configuration was found to have greater longitudinal control and increased center of gravity range relative to a conventional (two-surface) design. The test showed that power had a large favorable effect on stability and control about all three axis in the post-stall regime.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 90-3074 , AIAA Applied Aerodynamics Conference; Aug 20, 1990 - Aug 22, 1990; Portland, OR; United States
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  • 10
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    Comparison of Experimental and Computational Aerothermodynamics of a 70-deg Sphere-Cone (1996)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: Numerical solutions for hypersonic flows of carbon-dioxide and air around a 70-deg sphere-cone have been computed using an axisymmetric non-equilibrium Navier-Stokes solver. Freestream flow conditions for these computations were equivalent to those obtained in an experimental blunt-body heat-transfer study conducted in a high-enthalpy, hypervelocity expansion tube. Comparisons have been made between the computed and measured surface heat-transfer rates on the forebody and afterbody of the sphere-cone and on the sting which supported the test model. Computed forebody heating rates were within the estimated experimental uncertainties of 10% on the forebody and 15% in the wake except for within the recirculating flow region of the wake.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: AIAA Paper 96-1867 , 31st AIAA Thermophysics Conference; Jun 18, 1996 - Jun 20, 1996; New Orleans, LA; United States
    Format: application/pdf
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