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  • 1
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    An improved viscid/inviscid interaction procedure for transonic flow over airfoils (1985)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A new interacting boundary layer approach for computing the viscous transonic flow over airfoils is described. The theory includes a complete treatment of viscous interaction effects induced by the wake and accounts for normal pressure gradient effects across the boundary layer near trailing edges. The method is based on systematic expansions of the full Reynolds equation of turbulent flow in the limit of Reynolds numbers, Reynolds infinity. Procedures are developed for incorporating the local trailing edge solution into the numerical solution of the coupled full potential and integral boundary layer equations. Although the theory is strictly applicable to airfoils with cusped or nearly cusped trailing edges and to turbulent boundary layers that remain fully attached to the airfoil surface, the method was successfully applied to more general airfoils and to flows with small separation zones. Comparisons of theoretical solutions with wind tunnel data indicate the present method can accurately predict the section characteristics of airfoils including the absolute levels of drag.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3805 , NAS 1.26:3805 , RE-682
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 2
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    In-flight measurements of wing ice shapes and wing section drag increases caused by natural icing conditions (1986)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Aircraft icing flight research was performed in natural icing conditions with a twin engine computer type STOL aircraft. In-flight measurements were made of the icing cloud environment, the shape of the ice accretion on the wing, and the corresponding increase in the wing section drag. Results are presented for three icing encounters. On one flight, the wing section drag coefficient increased 35 percent over the uniced baseline for cruise conditions while a 43 percent increase was observed at an aircraft angle of attack of 6.2 degrees.
    Keywords: AERODYNAMICS
    Type: NASA-TM-87301 , E-3013 , NAS 1.15:87301
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 3
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    Viscous wing theory development. Volume 1: Analysis, method and results (1986)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Viscous transonic flows at large Reynolds numbers over 3-D wings were analyzed using a zonal viscid-inviscid interaction approach. A new numerical AFZ scheme was developed in conjunction with the finite volume formulation for the solution of the inviscid full-potential equation. A special far-field asymptotic boundary condition was developed and a second-order artificial viscosity included for an improved inviscid solution methodology. The integral method was used for the laminar/turbulent boundary layer and 3-D viscous wake calculation. The interaction calculation included the coupling conditions of the source flux due to the wing surface boundary layer, the flux jump due to the viscous wake, and the wake curvature effect. A method was also devised incorporating the 2-D trailing edge strong interaction solution for the normal pressure correction near the trailing edge region. A fully automated computer program was developed to perform the proposed method with one scalar version to be used on an IBM-3081 and two vectorized versions on Cray-1 and Cyber-205 computers.
    Keywords: AERODYNAMICS
    Type: NASA-CR-178156 , NAS 1.26:178156 , RE-725-VOL-1
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 4
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    Double-branched vortex generator (1985)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: In order to assess the suitability of using a double branched vortex generator in parametric studies involving vortex interactions, an experimental study of the main vortex and secondary flows produced by a double branched vortex generator was conducted in a 20-by-40 cm indraft wind tunnel. Measurements of the cross flow velocities were made with a five hole pressure probe from which vorticity contours and vortex parameters were derived. The results showed that the optimum configuration consisted of chord extensions with the absence of a centerbody.
    Keywords: AERODYNAMICS
    Type: NASA-TM-88201 , REPT-86064 , NAS 1.15:88201
    Format: application/pdf
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  • 5
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    Application of finite element and remeshing technique to shock interference on a cylindrical leading edge (1988)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: The problem of planar oblique shock impingement on a cylindrical leading edge in hypersonic flow is modeled using a Galerkin-Runge Kutta finite element method. The method utilizes a four stage Runge-Kutta time stepping scheme to solve the compressible Euler equations. Freestream Mach numbers of 6.5, 8.0 and 16.0 are studied. The computed surface pressure distributions consistently agree well with available experimental data. The peak pressure amplification ranges from 5.45 at M = 6.5 to approximately 17.0 at M = 16.0. Stagnation point heat transfer rate amplifications are calculated from the inviscid solution using the method of Fay and Riddell. The value and wall location of the peak pressure and heat transfer rate amplifications are extremely sensitive to the location of the impinging shock/bow shock intersection point.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-0368
    Format: text
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  • 6
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    An LDA investigation of three-dimensional normal shock-boundary layer interactions in a corner (1987)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Nonintrusive, three-dimensional, measurements have been made of a normal shock wave-turbulent boundary layer interaction. The measurements were made in the corner of the test section of a continuous supersonic wind tunnel in which a normal shock wave had been stabilized. LDA, surface pressure measurement and flow visualization techniques were employed for two freestream Mach number test cases: 1.6 and 1.3. The former contained separated flow regions and a system of shock waves. The latter was found to be far less complicated. The reported results are believed to accurately define the flow physics of each case and may be used as benchmark data to verify three-dimensional computer codes.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 87-1369
    Format: text
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  • 7
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    An adaptive finite element scheme for the Euler and Navier-Stokes equations (1987)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: The implementation of two explicit finite element schemes for the solution of the compressible Euler and Navier-Stokes equations is presented. The schemes can be employed with general unstructured triangular meshes in two dimensions. Either scheme can therefore be used as the basic solver in a solution adaptive mesh procedure in a direct manner. The particular adaptive approach which is advocated here is intended for the solution of steady state problems only and involves an adaptive regeneration of the grid at prescribed stages during the false transient. The grid regeneration is accomplished by a mesh generator which has the capability of generating triangular grids over computational domains of arbitrary shape. The procedure is illustrated by solving transonic flows over multi-airfoil configurations and high speed flows, involving shock interactions, past circular cylinders.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 87-1172
    Format: text
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  • 8
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    The role of freestream turbulence scale in subsonic flow separation (1985)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The clarification of the role of freestream turbulence scale in determining the location of boundary layer separation is discussed. Modifications to the test facility were completed. Wind tunnel flow characteristics, including turbulence parameters, were determined with two turbulence generating grids, as well as no grid. These results are summarized. Initial results on the role of scale on turbulent boundary layer separation on the upper surface of an airfoil model are also discussed.
    Keywords: AERODYNAMICS
    Type: NASA-CR-176237 , NAS 1.26:176237 , IPR-2
    Format: application/pdf
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  • 9
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    Comparison of analytical and experimental performance of a wind-tunnel diffuser section (1986)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Wind tunnel diffuser performance is evaluated by comparing experimental data with analytical results predicted by an one-dimensional integration procedure with skin friction coefficient, a two-dimensional interactive boundary layer procedure for analyzing conical diffusers, and a two-dimensional, integral, compressible laminar and turbulent boundary layer code. Pressure, temperature, and velocity data for a 3.25 deg equivalent cone half-angle diffuser (37.3 in., 94.742 cm outlet diameter) was obtained from the one-tenth scale Altitude Wind Tunnel modeling program at the NASA Lewis Research Center. The comparison is performed at Mach numbers of 0.162 (Re = 3.097x19(6)), 0.326 (Re = 6.2737x19(6)), and 0.363 (Re = 7.0129x10(6)). The Reynolds numbers are all based on an inlet diffuser diameter of 32.4 in., 82.296 cm, and reasonable quantitative agreement was obtained between the experimental data and computational codes.
    Keywords: AERODYNAMICS
    Type: NASA-TM-88795 , E-3130 , NAS 1.15:88795 , Annual Conference of the National Technical Association; Jun 23, 1986 - Jun 28, 1986; Washington, DC; United States
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  • 10
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    Algebraic grid generation about wing-fuselage bodies (1987)
    In:  Other Sources
    Details
    Publication Date: 2011-08-19
    Description: An algebraic procedure for generating boundary-fitted grids about wing-fuselage configurations is presented. A wing-fuselage configuration consists of two aircraft components specified by cross sections and mathematically represented by Coons' patches. Several grid blocks are constructed to cover the entire region surrounding the configuration, and each grid block maps into a computational cube. Grid points are first determined on the six boundary surfaces of a block and then in the interior. Grid points on the surface of the configuration are derived from the intersection of planes with the Coons' patch definition. Approximate arc length distributions along the resulting grid curves concentrate and disperse grid points. The two-boundary technique and transfinite interpolation are used to determine grid points on the remaining boundary surfaces and block interiors.
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft (ISSN 0021-8669); 24; 868-872
    Format: text
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