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  • 1
    Publication Date: 2018-12-01
    Description: A quasi-three-dimensional boundary layer method for conical inviscid flows is presented. The model was developed to characterize the flow fields over slender delta wings, particularly the shed vortex sheet. A steady, incompressible, laminar boundary layer is assumed and a solution is obtained with the Smith (1966) finite difference code. Predictions are compared with oil flow patterns and surface pressure measurements for 74 and 80 deg delta wings at 0-20 deg angles of attack. The model is capable of accurately predicting the secondary separation lines. Inviscid velocities derived from panel methods lead, however, to inaccurate pressure distributions.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-2175
    Format: text
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  • 2
    Publication Date: 2016-06-07
    Description: A new inverse boundary layer method is developed and applied to incompressible flows with laminar separation and reattachment. Test cases for two dimensional flows are computed and the results are compared with those of other inverse methods. One advantage of the present method is that the calculation of the inviscid velocities may be determined at each marching step without having to iterate. The inverse method was incorporated with the direct method to calculate the incompressible, conical flow over a slender delta wing at incidence. The location of the secondary separation line on the leeward surface of the wing is determined and compared with experiment for a unit aspect ratio wing at 20.5 deg incidence. The viscous flow in the separated region was calculated using prescribed skin friction coefficients.
    Keywords: AERODYNAMICS
    Type: NASA. Langley Research Center Vortex Flow Aerodynamics, Vol. 1; p 115-133
    Format: application/pdf
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  • 3
    Publication Date: 2018-12-01
    Description: A three-dimensional, laminar boundary-layer method is applied to the incompressible flow over a slender delta wing at incidence. The predictor-corrector finite-difference scheme of Matsuno is used to difference the governing equations. The method has the advantages that no iterations are required to advance the solution and the cross-flow derivatives are formed independent of the cross-field direction. The difference scheme is demonstrated to yield accurate numerical results when compared to the exact solution of the three-dimensional boundary-layer equations for parabolic flow over a moving flat plate. The method is applied to delta wings of various sweep angles at angles of attack up to 20 deg., with the inviscid solution determined using a higher-order, three-dimensional panel method.
    Keywords: AERODYNAMICS
    Type: SAE PAPER 851818
    Format: text
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  • 4
    Publication Date: 2011-08-19
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft (ISSN 0021-8669); 22; 602-608
    Format: text
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