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  • 1
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    Aerodynamic heat transfer to a hypersonic research aircraft model /X-24C/ at Mach 6 (1976)
    In:  Other Sources
    Details
    Publication Date: 2011-08-17
    Description: The paper reports on results of heat-transfer tests conducted on a 1/29-scale model of the X-24C-12I hypersonic research aircraft configuration in a Mach 6 tunnel at a Reynolds number of thirteen million using the phase-change heat transfer technique. Sequences of phase-change heat transfer pattern photographs are presented showing windward side and leeward side heating processes. Theoretical predictions of dimensionless heat transfer coefficients along a data line on lower fuselage and on fuselage side bracket the experimental values. A turbulent heating theory gives good agreement with data when shifted to a new virtual origin.
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft; 13; Dec. 197
    Format: text
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    NASA TECHNICAL REPORTS
  • 2
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    Mach 6 flowfield survey at the engine inlet of a research airplane (1977)
    In:  Other Sources
    Details
    Publication Date: 2011-08-17
    Description: A flowfield survey was conducted to better define the nature of vehicle forebody flowfield at the inlet location of an airframe-integrated scramjet engine mounted on the lower surface of a high-speed research airplane to be air launched from a B-52 and rocket boosted to Mach 6. The tests were conducted on a 1/30-scale brass model in a Mach-6 20-in. wind tunnel at Reynolds number of 11,200,000 based on distance to engine inlet. Boundary layer profiles at five spanwise locations indicate that the boundary layer in the area of the forebody centerline is more than twice as thick as the boundary layer at three outboard stations. It is shown that the cold streak found in heating contours on the centerline of the forebody is caused by a thickening of the boundary layer on the centerline, and that this thickening decreases with angle of attack.
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft; 14; Apr. 197
    Format: text
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    NASA TECHNICAL REPORTS
  • 3
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    Inlet boundary-layer shapes on four aircraft forebodies at Mach 6 (1978)
    In:  Other Sources
    Details
    Publication Date: 2011-08-17
    Description: Boundary-layer shape measurements at the engine inlet on four different hypersonic aircraft forebody designs (with no engine on the forebody) are reported. The measurements provide a qualitative assessment of the effectiveness of various forebody geometries as engine inlet precompression surfaces. The designs, tested in a hypersonic tunnel at Mach 6 and a nominal freestream Reynolds number of 30,500,000, included a semiconical forebody, a configuration similar to a slab delta wing, a conical nose blended into a flat surface, and a conical, complex forebody shape. Boundary layer height as a function of forebody compression is shown for each design.
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft; 15; Jan. 197
    Format: text
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    NASA TECHNICAL REPORTS
  • 4
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    Hypersonic Airbreathing Missile (1978)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A hypersonic airbreathing missile using dual mode scramjet engines for propulsion is described. The fuselage is constructed of a material with a high heat sink capacity and is covered with a thermal protective shield and lined with an internal insulating blanket. The engine airframe integration uses the flat lower portion of the lower fuselage to precompress the air entering the scramjet engines. The precompression of air entering the scramjet inlets increases as the angles of attack. This feature results in a highly maneuverable missile which can accelerate as it banks into a turn.
    Keywords: LAUNCH VEHICLES AND SPACE VEHICLES
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 5
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    Analysis of various descent trajectories for a hypersonic-cruise, cold-wall research airplane (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The probable descent operating conditions for a hypersonic air-breathing research airplane were examined. Descents selected were cruise angle of attack, high dynamic pressure, high lift coefficient, turns, and descents with drag brakes. The descents were parametrically exercised and compared from the standpoint of cold-wall (367 K) aircraft heat load. The descent parameters compared were total heat load, peak heating rate, time to landing, time to end of heat pulse, and range. Trends in total heat load as a function of cruise Mach number, cruise dynamic pressure, angle-of-attack limitation, pull-up g-load, heading angle, and drag-brake size are presented.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-7860 , L-9778
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 6
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    Configuration heating for a hypersonic research airplane concept having a 70 deg swept double-delta wing (1978)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The heating on a candidate hypersonic research airplane configuration has been examined experimentally at Mach 6 by the phase-change-paint technique. The configuration has a double-delta wing with tip fins. Phase-change-paint diagrams give heating data for the model top, side, and bottom, with and without deflected elevons for an angle-of-attack range of 0 deg to 24 deg. Nominal Reynolds numbers are on the order of 15,000,000 with supplementary data at length Reynolds number of 4,000,000, which moves the model from the predominantly turbulent into the predominantly laminar regime. Also, intermediate Reynolds numbers were investigated on the lee side for one angle of attack.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NASA-TP-1143 , L-11841
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 7
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    Review of design and operational characteristics of the 0.3-meter transonic cryogenic tunnel (1979)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: The fundamentals of cryogenic testing are validated both analytically and experimentally employing the 0.3-m transonic cryogenic tunnel. The tunnel with its unique Reynolds number capability has been used for a wide variety of aerodynamic tests. Techniques regarding real-gas effects have been developed and cryogenic tunnel conditions are set and maintained accurately. It is shown that cryogenic cooling, by injecting nitrogen directly into the tunnel circuit, imposes no problems with temperature distribution or dynamic response characteristics.
    Keywords: RESEARCH AND SUPPORT FACILITIES (AIR)
    Type: International Symposium on Cryogenic Wind Tunnels; Apr 03, 1979 - Apr 05, 1979; Southampton
    Format: text
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    NASA TECHNICAL REPORTS
  • 8
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    Conceptual study of hypersonic airbreathing missiles (1978)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: The purpose of this paper is to report recent results of an in-house conceptual study to evaluate the performance potential and research needs of airbreathing hypersonic missiles. An alkylated-borane (noncryogenic) fueled, dual-mode, ramjet/scramjet propulsion system structured with a Rene 41 inlet and a carbon-carbon combustor was assumed along with a Lockalloy heat sink fuselage structure and beryllium wings and control surfaces. Performance for an air-launched baseline missile with a 961 pound staging weight containing a 100 pound payload indicated excellent long range cruise, moderate acceleration and high maneuverability potential. A sizing study indicates that Mach 6 cruise ranges of the order of 2500 nautical miles for payloads of 300 pounds can be achieved with moderate size missile carry weights (9000 lbs.). Aerodynamic heating analyses indicate that unprotected heat-sink structures with internal insulation are feasible for ranges of several hundred miles. For ranges of several thousands of miles a multiwall radiation shield (Inconel/titanium) was selected for protection of the internally insulated heat sink structure.
    Keywords: AIRCRAFT DESIGN, TESTING AND PERFORMANCE
    Type: AIAA PAPER 78-6 , Aerospace Sciences Meeting; Jan 16, 1978 - Jan 18, 1978; Huntsville, AL
    Format: text
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    NASA TECHNICAL REPORTS
  • 9
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    Review of design and operational characteristics of the 0.3-meter transonic cryogenic tunnel (1979)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The past 6 years of operation with the NASA Langley 0.3 m transonic cryogenic tunnel (TCT) show that there are no insurmountable problems associated with cryogenic testing with gaseous nitrogen at transonic Mach numbers. The fundamentals of the concept were validated both analytically and experimentally and the 0.3 m TCT, with its unique Reynolds number capability, was used for a wide variety of aerodynamic tests. Techniques regarding real-gas effects were developed and cryogenic tunnel conditions can be set and maintained accurately. Cryogenic cooling by injecting liquid nitrogen directly into the tunnel circuit imposes no problems with temperature distribution or dynamic response characteristics. Experience with the 0.3 m TCT, indicates that there is a significant learning process associated with cryogenic, high Reynolds number testing. Many of the questions have already been answered; however, factors such as tunnel control, run logic, economics, instrumentation, and model technology present many new and challenging problems.
    Keywords: AERODYNAMICS
    Type: NASA-TM-80123 , Intern. Symp. on Cryogenic Wind Tunnels; Apr 03, 1979 - Apr 05, 1979; Southampton
    Format: application/pdf
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    NASA TECHNICAL REPORTS
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