Search NASA⌕ Search

Engineering topics

Hoffman, J. D.

Publications and source records attributed to Hoffman, J. D..

At least 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1984

Computation of viscous flow in planar and axisymmetric ducts by an implicit marching procedure

A streamwise marching procedure, approximately 200 times faster than a full Navier-Stokes procedure with comparable accuracy, is presented for solving problems of compressible viscous subsonic flow. Results are presented and compared with experimental data for the cases of developing turbulent flow in a circular pipe; turbulent flow in a two-dimensional S-duct; and turbulent flow in a typical subsonic diffuser. Prior to each main marching step, a preliminary marching step is taken in which the integral continuity equation and an uncoupled form of the streamwise momentum equation are solved simultaneously to obtain the viscous pressure correction. During the main marching step the equations for continuity, streamwise momentum, cross-flow momentum, and energy are solved simultaneously as a coupled system using an implicit finite-difference method, with the viscous pressure correction treated as a source term. The analysis may be used for flows with both favorable and adverse pressure gradients and to predict the location of flow separation.

Towne, C. E.↗
NASA NTRS1982

A computer program for the calculation of the flow field including boundary layer effects for mixed-compression inlets at angle of attack

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.

Vadyak, J.↗
NASA NTRS1982

Implicit marching solution of compressible viscous subsonic flow in planar and axisymmetric ducts

A new streamwise marching procedure was developed and coded for compressible viscous subsonic flow in planar or axisymmetric ducts with or without centerbodies. The continuity, streamwise momentum, cross-flow momentum, and energy equations are written in generalized orthogonal curvilinear coordinates. To allow the use of a marching procedure, second derivatives in the streamwise momentum equation are written as the sum of a known two dimensional imposed pressure field and an unknown one dimensional viscous correction. For turbulent flow, the Reynolds stress and heat flux terms are modeled using two-layer eddy viscosity turbulence models.

Towne, C. E.↗
NASA NTRS1982

Calculation of the flow field including boundary layer effects for supersonic mixed compression inlets at angles of attack

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.

Vadyak, J.↗
NASA NTRS1982

Three-dimensional flow calculations including boundary layer effects for supersonic inlets at angle of attack

An analysis is presented for calculating the steady three-dimensional flow field in supersonic mixed-compression inlets at incidence. A zonal modeling approach is employed to obtain the solution. The supersonic core flow is computed using a second-order pentahedral bicharacteristic algorithm. The bow shock wave and the reflected internal shock train are determined using a three-dimensional discrete shock fitting procedure. The boundary layer flow adjacent to both the centerbody and the cowl is computed using a second-order implicit finite difference method. The flow in a shock wave-boundary layer interaction region is computed using an integral formulation. The culmination of the present research effort is the development of a production-type computer program capable of analyzing flow in a variety of mixed-compression aircraft inlets. Numerical results and experimental correlations are presented to illustrate application of the analysis.

Vadyak, J.↗
NASA NTRS1980

Flow field in supersonic mixed-compression inlets

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.

Bishop, A. R.↗
NASA NTRS1979

Calculation of the three-dimensional flow field in supersonic inlets at angle of attack using a bicharacteristic method with discrete shock wave fitting

An analysis is presented for calculating the flow field in supersonic mixed-compression aircraft inlets operating at angle of attack. The flow field is computed by a steady three-dimensional bicharacteristic method. The bow shock wave and the reflected internal shock wave system are computed by a three-dimensional discrete shock wave fitting procedure. Viscous and thermal diffusion may be included as source terms in the bicharacteristic method. A production type computer program capable of determining the flow field in a variety of axisymmetric mixed-compression supersonic inlets is available. The results of the present analysis agree well with those produced by the two-dimensional method of characteristics when axisymmetric flow fields are computed. For three-dimensional flow fields, the results of the present analysis agree well with experimental data except in regions of high viscous interaction and boundary layer removal. The present analysis does not compute the boundary layer, nor does it account for boundary layer bleed.

Vadyak, J.↗
NASA NTRS1979

Calculation of the flow field in supersonic inlets using a bicharacteristics method with shock wave fitting

The results of weakly viscous flow analysis are presented. The flow field, including molecular transport, is computed with the aid of a bicharacteristic method. The bow shock wave and the internal shock wave are computed with the aid of a three-dimensional shock wave fitting procedure. Characteristic equations are presented, and numerical integration procedure is discussed. Here, an inverse marching scheme is employed in which the solution is obtained on space-like planes of constant x and on space curves defined by the intersections of the internal shock wave with the solid boundaries. The distance between solution planes is arrived at by the Courant-Friedrichs-Lewy stability criterion.

Vadyak, J.↗
NASA NTRS1978

Calculation of the flow field in supersonic mixed-compression inlets at angle of attack using the three-dimensional method of characteristics with discrete shock wave fitting

The influence of molecular transport is included in the computation by treating viscous and thermal diffusion terms in the governing partial differential equations as correction terms in the method of characteristics scheme. The development of a production type computer program is reported which is capable of calculating the flow field in a variety of axisymmetric mixed-compression aircraft inlets. The results agreed well with those produced by the two-dimensional method characteristics when axisymmetric flow fields are computed. For three-dimensional flow fields, the results agree well with experimental data except in regions of high viscous interaction and boundary layer removal.

Vadyak, J.↗
NASA NTRS1978

A computer program for the calculation of the flow field in supersonic mixed-compression inlets at angle of attack using the three-dimensional method of characteristics with discrete shock wave fitting

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.

Vadyak, J.↗
NASA NTRS1973

The analysis of nonequilibrium, chemically reacting, supersonic flow in three dimensions using a bicharacteristics method.

A second-order numerical method of characteristics based on a bicharacteristics scheme was developed for the evaluation of steady, supersonic, nonequilibrium, chemically reacting flows. The absolute accuracy and order of accuracy of the method was shown by comparisons with spherical source flow and axisymmetric, nonequilibrium flows. Numerical results were obtained for elliptical and super-elliptical nozzles. These results illustrate the complex nature of three-dimensional flows and the inadequacy of quasi-three-dimensional methods which neglect cross flows.

Cline, M. C.↗
NASA NTRS1972

Comparison of characteristic schemes for three-dimensional, steady, isentropic flow.

The accuracy and order of three bicharacteristic schemes based on the numerical method of characteristics for three-dimensional, steady, isentropic, supersonic flow were compared based on numerical calculations. Comparisons were made for source flows and for flows in contoured axisymmetric nozzles. The first two methods were tetrahedral line networks, differing only in the manner in which the cross derivatives were evaluated. The third method was the pentahedral line network proposed by Butler. All three methods were numerically stable and produced efficient and accurate solutions. A discussion of the relative merits and disadvantages of each method is presented.

Cline, M. C.↗