Experimental Investigation of a Mach 6 Fixed-geometry Inlet Featuring a Swept External-internal Compression Flow Field PDF Download
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Author: Marvin G. Torrence
Publisher:
ISBN:
Category : Aerodynamics, Hypersonic
Languages : en
Pages : 66
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Book Description
Author: Marvin G. Torrence
Publisher:
ISBN:
Category : Aerodynamics, Hypersonic
Languages : en
Pages : 66
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Book Description
Author:
Publisher:
ISBN:
Category :
Languages : en
Pages : 730
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Author:
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ISBN:
Category : Aeronautics
Languages : en
Pages : 952
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Author:
Publisher:
ISBN:
Category : Government publications
Languages : en
Pages : 2834
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Author:
Publisher:
ISBN:
Category : Aeronautics
Languages : en
Pages : 460
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Author:
Publisher:
ISBN:
Category : Aeronautics
Languages : en
Pages : 644
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Author: United States. Superintendent of Documents
Publisher:
ISBN:
Category : United States
Languages : en
Pages : 1408
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Author: Earl C. Watson
Publisher:
ISBN:
Category : Boundary layer control
Languages : en
Pages : 32
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Author: National Aeronautics and Space Administration (NASA)
Publisher: Createspace Independent Publishing Platform
ISBN: 9781722219024
Category :
Languages : en
Pages : 48
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Book Description
Three-dimensional sidewall-compression scramjet inlets with leading-edge sweeps of 30 deg and 70 deg were tested in the Langley Hypersonic CF4 Tunnel at Mach 6 and with a ratio of specific heats of 1.2. The parametric effects of leading-edge sweep, cowl position, contraction ratio, and Reynolds number were investigated. The models were instrumented with 42 static pressure orifices that were distributed on the sidewalls, base plate, and cowl. Schlieren movies were made of each test for flow visualization of the effects of the internal flow spillage on the external flow field. 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 (i.e., sidewall impingement locations) and the observed increase in spillage with increasing leading-edge sweep. The primary effect of moving the cowl forward was capturing the flow that would have otherwise spilled out ahead of the cowl. Increasing the contraction ratio 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. Holland, Scott D. Langley Research Center RTOP 506-40-41-02...
Author: Scott D. Holland
Publisher:
ISBN:
Category : Aerodynamics, Hypersonic
Languages : en
Pages : 52
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