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  • 1
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    Effect of upstream sidewall boundary layer removal on an airfoil test (1984)
    In:  CASI
    Details
    Publication Date: 2006-02-14
    Description: Sidewall boundary layer effects were investigated by applying partial upstream sidewall boundary layer removal in the Langley 0.3-m transonic cryogenic tunnel. Over the range of sidewall boundary layer displacement thickness of these tests the influence on pressure distribution was found to be small for subcritical conditions; however, for supercritical conditions the shock position was affected by the sidewall boundary layer. For these tests (with and without boundary layer remove) comparisons with predictions of the GRUMFOIL computer code indicated that Mach number corrections due to the sidewall boundary layer improve the agreement for both subcritical and supercritical conditions. The results also show that sidewall boundary layer removal reduces the magnitude of the sidewall correction; however, a suitable correction must still be made.
    Keywords: AERODYNAMICS
    Type: Wind Tunnel Wall Interference Assessment and Correction, 1983; p 143-163
    Format: text
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    NASA TECHNICAL REPORTS
  • 2
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    Model experience in the Langley 0.3-m transonic cryogenic tunnel (1981)
    In:  Other Sources
    Details
    Publication Date: 2011-08-18
    Description: The model building, development, and testing experience gained during 8 years of operation of the 0.3-m Transonic Cryogenic Tunnel (TCT) is summarized. The summary is divided into four portions: (1) models tested in the 0.3-m TCT's original octagonal test section; (2) models tested in the present two dimensional test section; (3) models tested as a part of tunnel calibration and the development of advanced technology airfoils; and (4) development of a new way to construct two dimensional airfoil models. Design requirements imposed on the models by high Reynolds number testing at cryogenic temperatures are reviewed.
    Keywords: AERODYNAMICS
    Type: High Reynolds Number Res. - 1980; p 53-73
    Format: text
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    NASA TECHNICAL REPORTS
  • 3
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    An external insulation system for a cryogenic wind tunnel (1984)
    In:  Other Sources
    Details
    Publication Date: 2018-12-01
    Description: The thermal insulation system of the 0.3-Meter Transonic Cryogenic Tunnel (0.3-m TCT) at the NASA Langley Research Center is described in text, photographs, and drawings. The system is designed to operate from room temperature down to about 77.4 K, the temperature of liquid nitrogen at 1 atmosphere. A detailed description is given of the primary insulation system which consists of glass fiber mats, a 3-part vapor barrier, and a dry nitrogen positive-pressure purge system. Also described are several secondary insulation systems required for the test section, actuators, and tunnel supports. An appendix briefly describes the original insulation system which is considered inferior to the one presently in place. Time required for opening and closing portions of the insulation system for modification or repair to the tunnel has been reduced, typically, from a few days for the original thermal insulating system to a few hours for the present system.
    Keywords: RESEARCH AND SUPPORT FACILITIES (AIR)
    Format: text
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    NASA TECHNICAL REPORTS
  • 4
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    Support interference of wind tunnel models: A selective annotated bibliography (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: This bibliography, with abstracts, consists of 143 citations arranged in chronological order by dates of publication. Selection of the citations was made for their relevance to the problems involved in understanding or avoiding support interference in wind tunnel testing throughout the Mach number range. An author index is included.
    Keywords: RESEARCH AND SUPPORT FACILITIES (AIR)
    Type: NASA-TM-81909-SUPPL , NAS 1.15:81909-SUPPL
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 5
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    A fan pressure ratio correlation in terms of Mach number and Reynolds number for the Langley 0.3 meter transonic cryogenic tunnel (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Calibration data for the two dimensional test section of the Langley 0.3 Meter Transonic Cryogenic Tunnel were used to develop a Mach number-Reynolds number correlation for the fan pressure ratio in terms of test section conditions. Well established engineering relationships combined to form an equation which is functionally analogous to the correlation. A geometric loss coefficient which is independent of Reynolds number or Mach number was determined. Present and anticipated uses of this concept include improvement of tunnel control schemes, comparison of efficiencies for operationally similar wind tunnels, prediction of tunnel test conditions and associated energy usage, and determination of Reynolds number scaling laws for similar fluid flow systems.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1752 , L-13713
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 6
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    Increasing Metal Fracture Toughness (1982)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: In technique developed at Langley Research Center several thin sheets of metal are diffusion-brazed together in vacuum furnace to create thick piece of metal that retains much of fracture toughness of its thin components. Technique is expected to make many of high-strength stainless steels, not currently suitable, usable at cryogenic temperatures.
    Keywords: FABRICATION TECHNOLOGY
    Type: LAR-12805 , NASA Tech Briefs (ISSN 0145-319X); 6; 3; P. 341
    Format: text
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    NASA TECHNICAL REPORTS
  • 7
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    Studies of sidewall boundary layer in the Langley 0.3 meter transonic cryogenic tunnel with and without suction (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Boundary layer measurements on the sidewalls of the Langley 0.3 Meter Transonic Cryogenic Tunnel were made to determine the effectiveness of the passive boundary layer bleed system over a Reynolds number range from 20 to 200 x 10 to the sixth power per meter at Mach numbers from 0.30 to 0.76. The tunnel sidewall boundary layer displacement thickness was about 2 percent of the width of the test section without the boundary layer bleed. Measured velocity profiles correlated well with the defect law of Hama. With the boundary layer bleed equivalent to about 2 percent of the test section mass flow, the boundary layer displacement thickness reduced to about 1 percent of the test section width, which is generally considered acceptable for testing airfoils. It was also noticed that effectiveness of the bleed was nearly independent of the Mach number and Reynolds number over the range of conditions tested. A comparison of the measured suction effectiveness of the bleed with the finite difference and integral methods of boundary layer calculation showed good agreement.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NASA-TP-2096 , L-15437 , NAS 1.60:2096
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 8
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    Cooling system for high speed aircraft (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The system eliminates the necessity of shielding an aircraft airframe constructed of material such as aluminum. Cooling is accomplished by passing a coolant through the aircraft airframe, the coolant acting as a carrier to remove heat from the airframe. The coolant is circulated through a heat pump and a heat exchanger which together extract essentially all of the added heat from the coolant. The heat is transferred to the aircraft fuel system via the heat exchanger and the heat pump. The heat extracted from the coolant is utilized to power the heat pump. The heat pump has associated therewith power turbine mechanism which is also driven by the extracted heat. The power turbines are utilized to drive various aircraft subsystems, the compressor of the heat pump, and provide engine cooling.
    Keywords: AIRCRAFT DESIGN, TESTING AND PERFORMANCE
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 9
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    Orbiter thermal pressure drop characteristics for shuttle orbiter thermal protection system components: High density tile, low density tile, densified low density tile, and strain isolation pad (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Pressure drop tests were conducted on available samples of low and high density tile, densified low density tile, and strain isolation pads. The results are presented in terms of pressure drop, material thickness and volume flow rate. Although the test apparatus was only capable of a small part of the range of conditions to be encountered in a Shuttle Orbiter flight, the data serve to determine the type of flow characteristics to be expected for each material type tested; the measured quantities also should serve as input for initial venting and flow through analysis.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA-TM-81891
    Format: application/pdf
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  • 10
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    Flow rate/pressure drop data gathered from testing a sample of the Space Shuttle Strain Isolation Pad (SIP): Effects of ambient pressure combined with tension and compression conditions (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Tests were conducted on a sample of strain isolation pad (SIP) typical of that used in the shuttle orbiter thermal protection system to determine the characteristics of SIP internal flow. Data obtained were pressure drop as a function of flow rate for a range of ambient pressures representing various points along the Shuttle trajectory and for stretched and compressed conditions of the SIP. Flow was in the direction of the weave parallel to most of the fibers. The data are plotted in several standard engineering formats in order to be of maximum utility to the user. In addition to providing support to the Space Shuttle Program, these data are a source of experimental information on flow through fiberous (rather than the more usual sand bed type) porous media.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA-TM-84591 , NAS 1.15:84591
    Format: application/pdf
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