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  • 1
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    Comparison of Performance of Two Aerodynamically Similar 14-inch-diameter Single-stage Compressor Rotors of Different Chord Length (1958)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Type: NACA-RM-E57I03
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 2
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    Documentation of Two- and Three-Dimensional Hypersonic Shock Wave/Turbulent Boundary Layer Interaction Flows (1989)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Experimental data for a series of two- and three-dimensional shock wave/turbulent boundary layer interaction flows at Mach 7 are presented. Test bodies, composed of simple geometric shapes, were designed to generate flows with varying degrees of pressure gradient, boundary-layer separation, and turning angle. The data include surface-pressure and heat-transfer distributions as well as limited mean-flow-field surveys in both the undisturbed and the interaction regimes. The data are presented in a convenient form for use in validating existing or future computational models of these generic hypersonic flows.
    Keywords: AERODYNAMICS
    Type: NASA-TM-101075 , A-89048 , NAS 1.15:101075
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 3
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    Hypersonic flows as related to the National Aerospace Plane (1995)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The object of Cooperative Agreement NCC2-452 was to identify, develop, and document reliable turbulence models for incorporation into CFD codes, which would then subsequently be incorporated into numerical design procedures for the NASP and any other hypersonic vehicles. In a two-pronged effort, consisting of an experimental and a theoretical approach, several key features of flows over complex vehicles were identified, and test bodies were designed which were composed of simple geometric shapes over which these flow features were measured. The experiments were conducted in the 3.5' Hypersonic Wind Tunnel at NASA Ames Research Center, at nominal Mach numbers from 7 to 8.3 and Re/m from 4.9 x 10(exp 6) to 5.8 x 10(exp 6). Boundary layers approaching the interaction region were 2.5 to 3.7 cm thick. Surface and flow field measurements were conducted, and the initial boundary conditions were experimentally documented.
    Keywords: AERODYNAMICS
    Type: NASA-CR-199365 , NAS 1.26:199365
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 4
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    Hypersonic flows as related to the national aerospace plane (1992)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The study in the last 6 months has observed a clear evidence that the current two-equation models tend to under-predict flow separation and over-predict heat transfer rate near flow re-attachment regions. In hypersonic flow calculations, these model deficiencies appear to be even more pronounced. This is particularly true in the incapability of the model to predict the extent of the flow separation. Two major deficiencies of the current two-equation models in predicting complex hypersonic flows have been reported, i.e., under-prediction of flow separation and over-prediction of peak heat transfer rate. Two modifications to the k - epsilon model were reported and tested over a range of flows. Based on our limited study, the modified models have been found to give better agreements in both surface pressure and heat transfer predictions for several complex shock-wave boundary-layer interaction flows. However, in order to confirm our observation, more calculations will be performed in the future study covering a wider range of flows and conditions than reported here.
    Keywords: AERODYNAMICS
    Type: NASA-CR-191980 , NAS 1.26:191980
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 5
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    Hypersonic flows as related to the National Aerospace plane (1991)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Experimental data for a series of 2-D and 3-D shock wave/boundary layer interaction flows at Mach 8.2 are presented. The test bodies, composed of simple geometric shapes fastened to a flat plate test bed, were designed to generate flows with varying degrees of pressure gradient, boundary layer separation, and turning angle. The data include surface pressure and heat transfer distributions as well as limited mean flowfield surveys both in the undisturbed and interaction regimes. The data are presented in a convenient form to be used to validate existing or future computational models of these hypersonic flows.
    Keywords: AERODYNAMICS
    Type: NASA-CR-187985 , NAS 1.26:187985
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 6
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    Performance of Typical Rear-Stage Axial-Flow Compressor Rotor Blade Row at Three Different Blade Setting Angles (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: A comparison of the performance of a single-stage rotor run at three different blade setting angles is presented. The rotor was of a design typical for a last stage of a multistage compressor. At each setting angle, the rotor blade row was operated from 53 to 100 percent of equivalent maximum speed (850 ft/sec tip speed) at constant inlet pressure. Hot-wire anemometry was used to observe rotating-stall and surge patterns in time unsteady flow. Results indicated that an increase in peak pressure ratio and an increase in maximum equivalent weight flow were obtained at each speed investigated when the blade setting angle was decreased. An increase in peak efficiency was achieved with decrease in blade setting angle for part of the range of speeds investigated. However, the peak efficiencies for the three blade setting angles were approximately the same at the maximum speed investigated. The flow ranges for all three configurations were about the same at minimum speed and decreased at almost the same rate when the rotative speed was increased through part of the range of speeds investigated. At maximum speed, the flow range for the smallest setting angle was considerably less than the flow range for the other two configurations. A decrease in efficiency and flow range for the smallest blade setting angle at maximum speed can be attributed primarily to a Mach number effect. In addition, because of the difference in projected axial chord lengths at the casing wall, some effect on performance could be expected from the change in three-dimensional flow occurring at the tip. Rotating-stall characteristics for the two smaller blade setting angles were essentially the same. Only surge could be detected for the largest blade setting angle in the unstable-flow region of operation.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-MEMO-11-27-58E , E-117
    Format: application/pdf
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    NASA TECHNICAL REPORTS
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