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  • 11
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    The construction of airfoil pressure models by the plate method: Achievements, current research, technology development and potential applications (1985)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A method of constructing airfoils by inscribing pressure channels on the face of opposing plates, bonding them together to form one plate with integral channels, and contour machining this plate to form an airfoil model is described. The research and development program to develop the bonding technology is described as well as the construction and testing of an airfoil model. Sample aerodynamic data sets are presented and discussed. Also, work currently under way to produce thin airfoils with camber is presented. Samples of the aft section of a 6 percent airfoil with complete pressure instrumentation including the trailing edge are pictured and described. This technique is particularly useful in fabricating models for transonic cryogenic testing, but it should find application in a wide ange of model construction projects, as well as the fabrication of fuel injectors, space hardware, and other applications requiring advanced bonding technology and intricate fluid passages.
    Keywords: RESEARCH AND SUPPORT FACILITIES (AIR)
    Type: NASA-TM-87613 , NAS 1.15:87613
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 12
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    Experimental study of stagnation enthalpy in the Langley hotshot tunnel (1966)
    In:  CASI
    Details
    Publication Date: 2019-05-29
    Description: Enthalpy level determination in hotshot tunnel
    Keywords: THERMODYNAMICS AND COMBUSTION
    Type: NASA-TN-D-3272
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 13
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    Description and initial calibration of the langley hotshot tunnel with some real-gas charts for nitrogen (1963)
    In:  CASI
    Details
    Publication Date: 2019-05-21
    Description: Langley hotshot tunnel and real-gas correction factors for nitrogen flow
    Keywords: ELECTRONICS
    Type: NASA-TN-D-2023
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 14
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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
  • 15
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    Description of the insulation system for the Langley 0.3-Meter Transonic Cryogenic Tunnel (1985)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The thermal insulation system of the Langley 0.3 Meter Transonic Cryogenic Tunnel is described. The insulation 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 sytem consists of glass fiber mats, a three part vapor barrier, and a dry 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. The 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)
    Type: NASA-TM-86274 , L-15723 , NAS 1.15:86274
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 16
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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
  • 17
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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
  • 18
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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
  • 19
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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
  • 20
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    Measurements in the flow field of a cylinder with a laser transit anemometer and a drag rake in the Langley 0.3 m transonic cryogenic tunnel (1985)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An experiment was conducted at the 0.3 m Transonic Cryogenic Tunnel using a Laser Transit Anemometer (LTA) to probe the flow field around a 3.05 centimeter-diameter circular cylinder. Measurements were made along the stagnation line and determination of particle size was evaluated by their ability to follow the flow field. The LTA system was also used to scan into the boundary layer near the 45 degree point on the model. Results of these scans are presented in graphic and tabular form. Flow field particle seeding was accomplished by inbleeding dry service air. The residual moisture (5-10 ppm) condensed and formed ice particles which served as Mie scattering centers for the LTA detection system. Comparison of data taken along the stagnation streamline with theory indicated that these particles tracked the velocity gradient of the flow. Tunnel operating conditions for the tests were a Mach number of 0.3, a pressure of 1.93 x 100000 n/m squared, and a temperature of 225 degrees K. Free stream Mach number and pressure were varied for the particle size determination.
    Keywords: INSTRUMENTATION AND PHOTOGRAPHY
    Type: NASA-TM-86399 , NAS 1.15:86399
    Format: application/pdf
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    NASA TECHNICAL REPORTS
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