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  • AERODYNAMICS  (83)
  • AIRCRAFT PROPULSION AND POWER  (50)
  • FLUID MECHANICS AND HEAT TRANSFER  (35)
  • GENERAL
  • 1975-1979  (169)
  • 1977  (169)
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  • 1975-1979  (169)
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  • 1
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    Vortex lattice prediction of subsonic aerodynamics of hypersonic vehicle concepts (1977)
    In:  Other Sources
    Details
    Publication Date: 2011-08-17
    Description: The vortex lattice method introduced by Lamar and Gloss (1975) was applied to the prediction of subsonic aerodynamic characteristics of hypersonic body-wing configurations. The reliability of the method was assessed through comparison of the calculated and observed aerodynamic performances of two National Hypersonic Flight Research Facility craft at Mach 0.2. The investigation indicated that a vortex lattice model involving 120 or more panel elements can give good results for the lift and induced drag coefficients of the craft, as well as for the pitching moment at angles of attack below 10 to 15 deg. Automated processes for calculating the local slopes of mean-camber surfaces may also render the method suitable for use in preliminary design phases.
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft; 14; Oct. 197
    Format: text
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    NASA TECHNICAL REPORTS
  • 2
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    Effect of endwall cooling on secondary flows in turbine stator vanes (1977)
    In:  CASI
    Details
    Publication Date: 2012-05-19
    Description: An experimental investigation was performed to determine the effect of endwall cooling on the secondary flow behavior and the aerodynamic performance of a coreturbine stator vane. The investigation was conducted in a cold-air, full-annular cascade, where three-dimensional effects could be obtained. Two endwall cooling configurations were tested. In the first configuration, the cooling holes were oriented so that the coolant was injected in line with the inviscid streamline direction. In the second configuration, the coolant was injected at an angle of 15 deg to the inviscid streamline direction and oriented toward the vane pressure surface. In both cases the stator vanes were solid and uncooled so that the effect of endwall cooling could be obtained directly. Total-pressure surveys were taken downstream of the stator vanes over a range of cooling flows at the design, mean-radius, critical velocity ratio of 0.778. Changes in the total-pressure contours downstream of the vanes were used to obtain the effect of endwall cooling on the secondary flows in the stator. Comparisons were made between the two cooled-endwall configurations and with the results obtained previously for solid endwalls.
    Keywords: AIRCRAFT PROPULSION AND POWER
    Type: AGARD Secondary Flows in Turbomachines; 29 p
    Format: text
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    NASA TECHNICAL REPORTS
  • 3
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    Application of laser velocimetry to aircraft wake-vortex measurements (1977)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: The theory and use of a laser velocimeter that makes simultaneous measurements of vertical and longitudinal velocities while rapidly scanning a flow field laterally are described, and its direct application to trailing wake-vortex research is discussed. Pertinent measurements of aircraft wake-vortex velocity distributions obtained in a wind tunnel and water towing tank are presented. The utility of the velocimeter to quantitatively assess differences in wake velocity distributions due to wake dissipating devices and span loading changes on the wake-generating model is also demonstrated.
    Keywords: AERODYNAMICS
    Type: Wake Vortex Minimization; p 157-192
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 4
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    Aerodynamic characteristics at Mach numbers from 0.33 to 1.20 of a wing-body design concept for a hypersonic research airplane (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An experimental investigation of the static aerodynamic characteristics of a model of one design concept for the proposed National Hypersonic Flight Research Facility was conducted in the Langley 8 foot transonic pressure tunnel. The experiment consisted of configuration buildup from the basic body by adding a wing, center vertical tail, and a three module or six module scramjet engine. The freestream test Mach numbers were 0.33, 0.80, 0.90, 0.95, 0.98, 1.10, and 1.20 at Reynolds numbers per meter ranging from 4.8 x 1 million to 10.4 x 1 million. The test angle of attack range was approximately -4 deg to 22 deg at constant angles of sideslip of 0 deg and 4 deg; the angle of sideslip ranged from about -6 deg to 6 deg at constant angles of attack of 0 deg and 17 deg. The elevons were deflected 0 deg, -10 deg, and -20 deg with rudder deflections of 0 deg and 15.6 deg.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1044 , L-11723
    Format: application/pdf
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  • 5
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    Low-speed wind tunnel investigation of an advanced supersonic cruise arrow-wing configuration (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A preliminary assessment of possible means for improving the low speed aerodynamic characteristics of advanced supersonic cruise arrow wing configurations and to extend the existing data base of such configurations has been made. Principle configuration variables included wing-leading and trailing-edge flap deflection, fuselage nose strakes, and engine exhaust nozzle deflection. Results showed that deflecting the wing leading edge apex flaps downward provided improved longitudinal stability but resulted in reduced directional stability. The model exhibited relatively low values of directional stability over the operational angle of attack range and experienced large asymmetric yawing moments at high angles of attack. The use of nose strakes was found to be effective in increasing the directional stability and eliminating the asymmetric yawing moment.
    Keywords: AERODYNAMICS
    Type: NASA-TM-74043
    Format: application/pdf
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  • 6
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    Effect of endwall cooling on secondary flows in turbine stator vanes (1977)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The effect of endwall cooling on the secondary flow behavior and the aerodynamic performance of a core turbine stator vane was determined. The investigation was conducted in a cold-air, full-annular cascade, where three-dimensional effects were obtained. Two endwall cooling configurations were tested. In the first configuration, the cooling holes were oriented so that the coolant was injected in line with the inviscid streamline direction. In the second configuration, the coolant was injected at an angle of 15 deg to the inviscid streamline direction and oriented towards the vane pressure stator. In both cases the stator vanes were solid and uncooled so that the effect of endwall cooling was obtained directly. Total-pressure surveys were taken downstream of the stator vanes over a range of cooling flows at the design, mean-radius, critical velocity ratio of 0.778. Changes in the total-pressure contours downstream of the vanes were used to obtain the effect of endwall cooling on the secondary flows in the stator.
    Keywords: AIRCRAFT PROPULSION AND POWER
    Type: NASA-TM-X-73568 , AGARD 49th Propulsion and Energetics Panel Conf.; Mar 28, 1977 - Apr 04, 1977; The Hague
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 7
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    Computation of viscous transonic flow about a lifting airfoil (1977)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: The Reynolds averaged Navier-Stokes equations are solved numerically for the viscous transonic flow about a stationary NACA 64A010 airfoil in free air. This paper presents descriptions of the numerical method, turbulence models employed, and boundary conditions appropriate to simulation of free-air flight. Computed results are presented for the airfoil at a free-stream Mach number of 0.8, angles of attack of 0 and 2 deg, and a Reynolds number based on a chord of 4 x 10 to the 6th. For the lifting case, unsteady periodic motion was calculated along the aft portion of the airfoil and in its wake. Recent experimental results obtained by Johnson indicate periodicity aft of the shock closely approximates the computed frequency, but the amplitude of the disturbances was significantly less than the calculated amplitude.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 77-679 , Conference on Fluid and Plasmadynamics; Jun 27, 1977 - Jun 29, 1977; Albuquerque, NM
    Format: text
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  • 8
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    Compliant wall surface motion and its effect on the structure of a turbulent boundary layer (1977)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: The status of the continuing compliant wall drag reduction research at NASA-Langley Research Center is discussed. Preliminary surface motion calculations are reported along with compliant surface design concepts and their numerical models. A compliant drag reduction theory based on stabilizing the turbulent substructure is proposed and previous experiments have been examined relative to that theory. Results of recent low speed compliant surface experiments have been reported which include initial attempts to measure local compliant surface motion.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: Biennial Symposium on Turbulence in liquids; Sep 22, 1975 - Sep 24, 1975; Rolla, MO
    Format: text
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    NASA TECHNICAL REPORTS
  • 9
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    An experimental and computational investigation of the steady and unsteady transonic flow fields about an airfoil in a solid-wall test channel (1977)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: An experimental and computational investigation of the steady and unsteady transonic flow field about a thick airfoil is described. An operational computer code for solving the two-dimensional, compressible Navier-Stokes equations for flow over airfoils was modified to include solid-wall, slipflow boundary conditions to properly assess the code and help guide the development of improved turbulence models. Steady and unsteady flow fields about an 18% thick circular arc airfoil at Mach numbers of 0.720, 0.754, and 0.783 and a chord Reynolds number of 11 x 10 to the 6th are predicted and compared with experiment. For the first time, computed results for unsteady turbulent flows with separation caused by a shock wave were obtained which qualitatively reproduce the time-dependent aspects of experiments. Features such as the intensity and reduced frequency of airfoil surface-pressure fluctuations, oscillatory regions of trailing-edge and shock-induced separation, and the Mach number range for unsteady flows were all qualitatively reproduced.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 77-678 , Conference on Fluid and Plasmadynamics; Jun 27, 1977 - Jun 29, 1977; Albuquerque, NM
    Format: text
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  • 10
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    Identification of Spey engine dynamics in the augmentor wing jet STOL research aircraft from flight data (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The development and validation of a spey engine model is described. An analysis of the dynamical interactions involved in the propulsion unit is presented. The model was reduced to contain only significant effects, and was used, in conjunction with flight data obtained from an augmentor wing jet STOL research aircraft, to develop initial estimates of parameters in the system. The theoretical background employed in estimating the parameters is outlined. The software package developed for processing the flight data is described. Results are summarized.
    Keywords: AIRCRAFT PROPULSION AND POWER
    Type: NASA-CR-152054
    Format: application/pdf
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