Stress, buckling, and vibration analysis of shells of revolution
Stress, buckling and vibration analysis of shells of revolution by numerical integration and finite difference methods, summarizing computer programs
Anderson, M. S.↗
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Computer program calculates transonic velocities in turbomachines
Computer program, TSONIC, combines velocity gradient and finite difference methods to obtain numerical solution for ideal, transonic, compressible flow for axial, radial, or mixed flow cascade of turbomachinery blades.
Katsanis, T.↗
The compressible turbulent boundary layer on a blunt swept slab with and without leading- edge blowing
Finite difference method for solving equations for compressible turbulent boundary layers on swept infinite cylinders
Beckwith, I. E.↗
A numerical procedure to calculate the inviscid flow field about a space shuttle orbiter traveling at a supersonic/hypersonic velocity
Finite difference method for calculating inviscid flow field around supersonic/hypersonic space shuttle
Grossman, B.↗
Finite difference calculation of the behavior of a discontinuous signal near a caustic
Finite difference method for predicting sonic boom overpressure signature near caustic
Krupp, J. A.↗
Modified thermal analyzer digital computer program (TAP)
Computer program for solving n-dimensional transient or steady state heat flow problems by creating electrical analogy of problem and solving by finite difference method
Knittle, D. E.↗
Nonsimilar solutions for laminar film condensation of liquid metals
Liquid metals vapor drag and electromagnetic fields effects on laminar film condensation, using finite difference method
Denny, V. E.↗
Calculation of relaxing turbulent boundary layers downstream of tangential slot injection
Turbulent boundary layers calculation downstream of compressible relaxing slot injection flows, using finite difference method based on eddy viscosity model
Bushnell, D. M.↗
Dynamic buckling of shallow arches
Symmetric and asymmetric dynamic buckling of shallow elastic arches under uniform loads, using nonlinear finite difference method
Barton, F. W.↗
A Numerical Study of Displacement Body and Curvature Effects on Incompressible and Compressible Laminar Boundary Layers
This technique has been applied to study such effects on incompressible flow around cylinders at moderate to low Reynolds numbers and for compression ramps at hypersonic Mach numbers by employing a finite difference method to obtain numerical solutions. The results indicate that the technique can be applied successfully in both regimes and does predict the correct trend in regions of large curvature and displacement body effects. It was concluded that curvature corrections should only be attempted in cases where all displacement effects can be fully accounted for.
Wornom, S. F.↗
Flow field predictions for a slab delta wing at incidence
Theoretical results are presented for the structure of the hypersonic flow field of a blunt slab delta wing at moderately high angle of attack. Special attention is devoted to the interaction between the boundary layer and the inviscid entropy layer. The results are compared with experimental data. The three-dimensional inviscid flow is computed numerically by a marching finite difference method. Attention is concentrated on the windward side of the delta wing, where detailed comparisons are made with the data for shock shape and surface pressure distributions. Surface streamlines are generated, and used in the boundary layer analysis. The three-dimensional laminar boundary layer is computed numerically using a specially-developed technique based on small cross-flow in streamline coordinates. In the rear sections of the wing the boundary layer decreases drastically in the spanwise direction, so that it is still submerged in the entropy layer at the centerline, but surpasses it near the leading edge. Predicted heat transfer distributions are compared with experimental data.
Conti, R. J.↗
Numerical solutions of the Navier-Stokes equations for the supersonic laminar flow over a two-dimensional compression corner
Numerical solutions have been obtained for the supersonic, laminar flow over a two-dimensional compression corner. These solutions were obtained as steady-state solutions to the unsteady Navier-Stokes equations using the finite difference method of Brailovskaya, which has second-order accuracy in the spatial coordinates. Good agreement was obtained between the computed results and wall pressure distributions measured experimentally for Mach numbers of 4 and 6.06, and respective Reynolds numbers, based on free-stream conditions and the distance from the leading edge to the corner. In those calculations, as well as in others, sufficient resolution was obtained to show the streamline pattern in the separation bubble. Upstream boundary conditions to the compression corner flow were provided by numerically solving the unsteady Navier-Stokes equations for the flat plate flow field, beginning at the leading edge. The compression corner flow field was enclosed by a computational boundary with the unknown boundary conditions supplied by extrapolation from internally computed points.
Carter, J. E.↗
Digital-computer normal shock position and restart control of a Mach 2.5 axisymmetric mixed-compression inlet
Digital computer control of a mixed-compression inlet is discussed. The inlet was terminated with a choked orifice at the compressor face station to dynamically simulate a turbojet engine. Inlet diffuser exit airflow disturbances were used. A digital version of a previously tested analog control system was used for both normal shock and restart control. Digital computer algorithms were derived using z-transform and finite difference methods. Using a sample rate of 1000 samples per second, the digital normal shock and restart controls essentially duplicated the inlet analog computer control results. At a sample rate of 100 samples per second, the control system performed adequately but was less stable.
Neiner, G. H.↗
Computer program: Jet 3 to calculate the large elastic plastic dynamically induced deformations of free and restrained, partial and/or complete structural rings
A user-oriented FORTRAN 4 computer program, called JET 3, is presented. The JET 3 program, which employs the spatial finite-element and timewise finite-difference method, can be used to predict the large two-dimensional elastic-plastic transient Kirchhoff-type deformations of a complete or partial structural ring, with various support conditions and restraints, subjected to a variety of initial velocity distributions and externally-applied transient forcing functions. The geometric shapes of the structural ring can be circular or arbitrarily curved and with variable thickness. Strain-hardening and strain-rate effects of the material are taken into account.
Wu, R. W.↗
Displacement interaction and surface curvature effects on hypersonic boundary layers.
The title problem was studied employing implicit finite-difference methods to obtain numerical solutions to a composite set of boundary-layer equations valid to second order. Results are given for flow up a two-dimensional cubic compression ramp for free-stream Mach numbers of 6, 8, and 12.25 and for free-stream Re/inch equal to 85,800 and 25,800 at a wall-to-stagnation temperature ratio of 0.223. Comparisons with independent theories and experimental results are given. Nonsingular separation was produced at a free-stream Mach number of 12.25. For all cases considered, displacement and curvature effects canceled one another when a consistent treatment of inviscid and viscous curvature corrections was employed - the second-order theory virtually reproducing the first-order results.
Wornom, S. F.↗
Pressure-gradient effects on hypersonic turbulent skin friction and boundary-layer profiles.
Local skin friction, total-temperature profiles, and pitot-pressure profiles were measured on the wall of a Mach-7.4 wind tunnel. The wall to adiabatic wall temperature ratio was varied from 0.3 to 0.5. Boundary-layer characteristics were compared with those predicted by a finite-difference method. Local skin friction was predicted to within 15%. Pressure-gradient effects on the temperature and Mach number distributions and the shape factor (displacement thickness/momentum thickness) were underpredicted, but the velocity distributions were closely predicted.
Hopkins, E. J.↗
Simulation of a steady-state integrated human thermal system.
The mathematical model of an integrated human thermal system is formulated. The system consists of an external thermal regulation device on the human body. The purpose of the device (a network of cooling tubes held in contact with the surface of the skin) is to maintain the human body in a state of thermoneutrality. The device is controlled by varying the inlet coolant temperature and coolant mass flow rate. The differential equations of the model are approximated by a set of algebraic equations which result from the application of the explicit forward finite difference method to the differential equations. The integrated human thermal system is simulated for a variety of combinations of the inlet coolant temperature, coolant mass flow rate, and metabolic rates. Two specific cases are considered: (1) the external thermal regulation device is placed only on the head and (2) the devices are placed on the head and the torso. The results of the simulation indicate that when the human body is exposed to hot environment, thermoneutrality can be attained by localized cooling if the operating variables of the external regulation device(s) are properly controlled.
Hsu, F. T.↗
Numerical computation of multishocked, three-dimensional supersonic flow fields with real gas effects.
A computational procedure is presented which is capable of determining the supersonic flow field surrounding three-dimensional wing-body configurations such as a delta-wing space shuttle. The governing equations in conservation-law form are solved by a finite difference method using a second-order noncentered algorithm between the body and the outermost shock wave, which is treated as a sharp discontinuity. Secondary shocks which form between these boundaries are captured automatically, and the intersection of these shocks with the bow shock posed no difficulty. Resulting flow fields about typical blunt nose shuttle-like configurations at angle of attack are presented. The differences between perfect and real gas effects for high Mach number flows are shown.
Kutler, P.↗