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  • 1
    Publication Date: 2018-12-01
    Description: An efficient grid-interfacing zonal algorithm has been developed for computing the three-dimensional transonic flow field about wing/nacelle multicomponent configurations. The algorithm uses the full-potential formulation and the AF2 fully-implicit approximate factorization scheme. The flow field position is computed using a component-adaptive grid approach in which separate grids are employed for the individual components in the multicomponent configuration, where each component grid is optimized for a particular geometry such as the wing or nacelle. The wing and nacelle component grids are allowed to overlap, and flow field information is transmitted from one grid to another through the overlap region using trivariate interpolation. This paper presents a discussion of the computational methods used to generate both the wing and nacelle component grids, the technique used to interface the component grids, and the method used to obtain the inviscid multicomponent flow field solution. Computed results and correlations with experiment are presented to illustrate application of the analysis.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-2430
    Format: text
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  • 2
    Publication Date: 2013-08-31
    Description: A highly efficient computer analysis was developed for predicting transonic nacelle/inlet flowfields. This algorithm can compute the three dimensional transonic flowfield about axisymmetric (or asymmetric) nacelle/inlet configurations at zero or nonzero incidence. The flowfield is determined by solving the full-potential equation in conservative form on a body-fitted curvilinear computational mesh. The difference equations are solved using the AF2 approximate factorization scheme. This report presents a discussion of the computational methods used to both generate the body-fitted curvilinear mesh and to obtain the inviscid flow solution. Computed results and correlations with existing methods and experiment are presented. Also presented are discussions on the organization of the grid generation (NGRIDA) computer program and the flow solution (NACELLE) computer program, descriptions of the respective subroutines, definitions of the required input parameters for both algorithms, a brief discussion on interpretation of the output, and sample cases to illustrate application of the analysis.
    Keywords: AERODYNAMICS
    Type: NASA-CR-166528 , LG83ER0163 , NAS 1.26:166528
    Format: application/pdf
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  • 3
    Publication Date: 2013-08-31
    Description: The flow field in supersonic mixed compression aircraft inlets at angle of attack is calculated. A zonal modeling technique is employed to obtain the solution which divides the flow field into different computational regions. The computational regions consist of a supersonic core flow, boundary layer flows adjacent to both the forebody/centerbody and cowl contours, and flow in the shock wave boundary layer interaction regions. The zonal modeling analysis is described and some computational results are presented. The governing equations for the supersonic core flow form a hyperbolic system of partial differential equations. The equations for the characteristic surfaces and the compatibility equations applicable along these surfaces are derived. The characteristic surfaces are the stream surfaces, which are surfaces composed of streamlines, and the wave surfaces, which are surfaces tangent to a Mach conoid. The compatibility equations are expressed as directional derivatives along streamlines and bicharacteristics, which are the lines of tangency between a wave surface and a Mach conoid.
    Keywords: AERODYNAMICS
    Type: NAS 1.26:167941 , NASA-CR-167941
    Format: application/pdf
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  • 4
    Publication Date: 2013-08-31
    Description: An efficient grid-interfacing zonal algorithm was developed for computing the three-dimensional transonic flow field about wing/nacelle configurations. the algorithm uses the full-potential formulation and the AF2 approximate factorization scheme. The flow field solution is computed using a component-adaptive grid approach in which separate grids are employed for the individual components in the multi-component configuration, where each component grid is optimized for a particular geometry such as the wing or nacelle. The wing and nacelle component grids are allowed to overlap, and flow field information is transmitted from one grid to another through the overlap region using trivariate interpolation. This report represents a discussion of the computational methods used to generate both the wing and nacelle component grids, the technique used to interface the component grids, and the method used to obtain the inviscid flow solution. Computed results and correlations with experiment are presented. also presented are discussions on the organization of the wing grid generation (GRGEN3) and nacelle grid generation (NGRIDA) computer programs, the grid interface (LK) computer program, and the wing/nacelle flow solution (TWN) computer program. Descriptions of the respective subroutines, definitions of the required input parameters, a discussion on interpretation of the output, and the sample cases illustrating application of the analysis are provided for each of the four computer programs.
    Keywords: AERODYNAMICS
    Type: NAS 1.26:166529 , LG83ER0164 , NASA-CR-166529
    Format: application/pdf
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  • 5
    Publication Date: 2013-08-31
    Description: A computer program was developed which is capable of calculating the flow field in the supersonic portion of a mixed compression aircraft inlet operating at angle of attack. The supersonic core flow is computed using a second-order three dimensional method-of-characteristics algorithm. The bow shock and the internal shock train are treated discretely using a three dimensional shock fitting procedure. The boundary layer flows are computed using a second-order implicit finite difference method. The shock wave-boundary layer interaction is computed using an integral formulation. The general structure of the computer program is discussed, and a brief description of each subroutine is given. All program input parameters are defined, and a brief discussion on interpretation of the output is provided. A number of sample cases, complete with data listings, are provided.
    Keywords: AERODYNAMICS
    Type: NASA-CR-168002 , NAS 1.26:168002
    Format: application/pdf
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  • 6
    Publication Date: 2018-12-01
    Description: A highly efficient computer analysis has been developed for predicting transonic nacelle/inlet flowfields. This algorithm can compute the three-dimensional transonic flowfield about axisymmetric or asymmetric nacelle/inlet configurations at zero or nonzero incidence. The flowfield is determined by solving the full-potential equation in conservative form on a body-fitted curvilinear computational mesh. The difference equations are solved using the AF2 approximate factorization scheme. The effects of boundary layer viscous entrainment are approximated in the inviscid algorithm by applying a surface transpiration velocity which is determined from the calculated boundary layer growth. Computed results and correlations with existing methods and experiment are presented to illustrate application of the analysis.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1417
    Format: text
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  • 7
    Publication Date: 2011-08-18
    Description: Previously cited in issue 07, p. 965, Accession no. A82-19778
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 22; 873-881
    Format: text
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  • 8
    Publication Date: 2011-08-18
    Description: Previously cited in issue 18, p. 2841, Accession no. A82-37477
    Keywords: AERODYNAMICS
    Type: (ISSN 0001-1452)
    Format: text
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  • 9
    Publication Date: 2011-08-17
    Description: Program uses method of characteristics for steady three-dimensional flow to calculate flow field in supersonic portion of mixed-compression aircraft inlet at non-zero angle of attack. Results agree well with experimental data except in regions of high viscous interaction. Flow field for variety of mixed-compression inlets can be calculated. Input includes geometry and attack of inlet. Output consists of list of parameters, solution planes, and description of shock waves. Program is written in FORTRAN IV for batch execution on CDC 6000-series.
    Keywords: MECHANICS
    Type: LEW-13279 , NASA Tech Briefs (ISSN 0145-319X); 5; 1; P. 74
    Format: text
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  • 10
    Publication Date: 2013-08-31
    Description: The calculation procedure is based on the method of characteristics for steady three-dimensional flow. The bow shock wave and the internal shock wave system were computed using a discrete shock wave fitting procedure. The general structure of the computer program is discussed, and a brief description of each subroutine is given. All program input parameters are defined, and a brief discussion on interpretation of the output is provided. A number of sample cases, complete with data deck listings, are presented.
    Keywords: AERODYNAMICS
    Type: NASA-TM-78947 , E-9694
    Format: application/pdf
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