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Rubbert, P. E.

Publications and source records attributed to Rubbert, P. E..

17 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1986

A new approach to the solution of boundary value problems involving complex configurations

A new approach for solving certain types of boundary value problems about complex configurations is presented. Numerical algorithms from such diverse fields as finite elements, preconditioned Krylov subspace methods, discrete Fourier analysis, and integral equations are combined to take advantage of the memory, speed and architecture of current and emerging supercomputers. Although the approach has application to many branches of computational physics, the present effort is concentrated in areas of Computational Fluid Dynamics (CFD) such as steady nonlinear aerodynamics, time harmonic unsteady aerodynamics, and aeroacoustics. The most significant attribute of the approach is that it can handle truly arbitrary boundary geometries and eliminates the difficult task of generating surface fitted grids.

Rubbert, P. E.↗
NASA NTRS1984

Flow prediction for propfan configurations using Euler equations

An Euler code has been developed for the analysis of a wing-mounted propfan configuration. Surface-fitted grids are used to represent the wing, fuselage, and nacelle geometry. The propeller is simulated by an actuator disk along a computational plane. A grid embedding technique is employed to capture detailed flow field resolution in the vicinity of the engine exhaust plume. Results of a NASA turboprop configuration are compared with test data. Specific issues on grid embedding and methods of resolving them are discussed.

Yu, N. J.↗
NASA NTRS1983

Transonic flow simulation of prop-fan configurations

This paper presents the results of a study that used a three-dimensional transonic analysis method to analyze flow around a wing-mounted prop-fan configuration. The transonic analysis method employs a cell-oriented finite volume approach to solve the full potential equation in conservative form over a numerically generated surface-fitted grid. The propeller slipstream effects are simulated using a linearized transpiration boundary condition. The results of analyses of flow over clean wing/body, wing/body/nacelle, and wing/body nacelle with slipstream are compared to test data.

Samant, S. S.↗
NASA NTRS1982

Some ideas and opportunities concerning three-dimensional wind-tunnel wall corrections

Opportunities for improving the accuracy and reliability of wall corrections in conventional ventilated test sections are presented. The approach encompasses state-of-the-art technology in transonic computational methods combined with the measurement of tunnel-wall pressures. The objective is to arrive at correction procedures of known, verifiable accuracy that are practical within a production testing environment. It is concluded that: accurate and reliable correction procedures can be developed for cruise-type aerodynamic testing for any wall configuration; passive walls can be optimized for minimal interference for cruise-type aerodynamic testing (tailored slots, variable open area ratio, etc.); monitoring and assessment of noncorrectable interference (buoyancy and curvature in a transonic stream) can be an integral part of a correction procedure; and reasonably good correction procedures can probably be developd for complex flows involving extensive separation and other unpredictable phenomena.

Rubbert, P. E.↗
NASA NTRS1980

Grid generation for general three-dimensional configurations

A suitable grid system for complex three dimensional configurations such as wing/body/nacelle shape for the solution of nonlinear transonic flow problems was constructed. Two approaches were explored based on Thompson's body fitted coordinate concept. The most general approach was to divide the computational domain into multiple rectangular blocks where the configuration itself was also represented by a set of blocks whose structure follows the natural lines of the configuration. The block structured grid system was adaptable to complex configurations and gives good grid quality near physical corners. However, it introduced algorithm issues for the flow solution concerning the treatment of nonanalytic grid block boundaries and nonstandard grid cells. These issues were explored in relation to the grid generation. A more limited approach treats a wing/body configuration with only a single rectangular block in computational space. The issues involving nonstandard cells were avoided, but other limitations on grid resolution appear. Both a linear and a nonlinear system of grid generation equations were developed including methods of grid control.

Lee, K. D.↗
NASA NTRS1979

A higher order panel method for linearized supersonic flow

The basic integral equations of linearized supersonic theory for an advanced supersonic panel method are derived. Methods using only linear varying source strength over each panel or only quadratic doublet strength over each panel gave good agreement with analytic solutions over cones and zero thickness cambered wings. For three dimensional bodies and wings of general shape, combined source and doublet panels with interior boundary conditions to eliminate the internal perturbations lead to a stable method providing good agreement experiment. A panel system with all edges contiguous resulted from dividing the basic four point non-planar panel into eight triangular subpanels, and the doublet strength was made continuous at all edges by a quadratic distribution over each subpanel. Superinclined panels were developed and tested on s simple nacelle and on an airplane model having engine inlets, with excellent results.

Ehlers, F. E.↗
NASA NTRS1978

A Mach line panel method for computing the linearized supersonic flow over planar wings

A method is described for solving the linearized supersonic flow over planar wings using panels bounded by two families of Mach lines. Polynomial distributions of source and doublet strength lead to simple, closed form solutions for the aerodynamic influence coefficients, and a nearly triangular matrix yields rapid solutions for the singularity parameters. The source method was found to be accurate and stable both for analysis and design boundary conditions. Similar results were obtained with the doublet method for analysis boundary conditions on the portion of the wing downstream of the supersonic leading edge, but instabilities in the solution occurred for the region containing a portion of the subsonic leading edge. Research on the method was discontinued before this difficulty was resolved.

Ehlers, F. E.↗
NASA NTRS1978

Computational aerodynamics requirements: The future role of the computer and the needs of the aerospace industry

The commercial airplane builder's viewpoint on the important issues involved in the development of improved computational aerodynamics tools such as powerful computers optimized for fluid flow problems is presented. The primary user of computational aerodynamics in a commercial aircraft company is the design engineer who is concerned with solving practical engineering problems. From his viewpoint, the development of program interfaces and pre-and post-processing capability for new computational methods is just as important as the algorithms and machine architecture. As more and more details of the entire flow field are computed, the visibility of the output data becomes a major problem which is then doubled when a design capability is added. The user must be able to see, understand, and interpret the results calculated. Enormous costs are expanded because of the need to work with programs having only primitive user interfaces.

Rubbert, P. E.↗
NASA NTRS1978

An improved higher order panel method for linearized supersonic flow

An improved higher order panel method for linearized supersonic flow is described. Each panel, defined by four points on the surface, is divided into eight subpanels in such a way that all subpanel and panel edges are contiguous. By prescribing a quadratic distribution of the doublet on each subpanel, the doublet strength is made strictly continuous on the paneled surface. A linear source distribution is also used. Numerical results are smoother and in better agreement with experiment than the previous method with less strict continuity. A brief discussion of superinclined panels used to eliminate interior interference in nacelles is included.

Ehlers, F. E.↗
NASA NTRS1977

Boundary-value problem of configurations with compressible free vortex flow

A self-consistent version of the compressible boundary-value problem for configurations with leading-edge vortex separation is formulated, based on the assumption that the compressible flow field is controlled by the linearized potential equation. The stream surface boundary condition and the zero pressure jump condition of the compressible free vortex flows are analyzed; application of the Goethert rule permits the compressible nonlinear boundary-value problem for the subsonic flow domain to be transformed into an equivalent nonlinear incompressible problem. The compressibility corrections developed are used in numerical calculations of subsonic leading-edge vortex flows about planar wing geometries. The sample calculations, employing an inviscid flow model in which the wing and vortex sheets are represented by piecewise continuous quadratic doublet sheet distributions, are applicable to high subsonic Mach numbers.

Brune, G. W.↗
NASA NTRS1976

A higher order panel method for linearized supersonic flow

A higher order panel method using linearly varying source and quadratically varying doublet distributions for computing linearized supersonic flow over arbitrary wings and bodies is described. With tangential mass flux boundary conditions on circular cones at angles of attack, the source panel method gave pressure distributions which are in excellent agreement with exact nonlinear theory. The numerical source panel method was demonstrated to be very stable by randomly panelling an axially symmetric body. For wings with thickness and camber, both the source and doublet panel methods yield excellent agreement with exact linearized solutions. Computational experiments indicate that the use of exact surface paneling in lieu of linearized boundary conditions for wings provides better agreement with higher order theory and with experiment.

Ehlers, F. E.↗
NASA NTRS1976

A higher order panel method for general analysis and design applications in subsonic flow

A higher-order panel method is described for numerical solution of boundary-value problems relating to steady inviscid irrotational incompressible subsonic fluid flow in a domain. Both Neumann and Dirichlet boundary conditions are treated; two types of auxiliary conditions are used to remove the degrees of freedom that arise from specifying only the derivative of the perturbation velocity potential. Four general network types and two expansions of the induced potential kernel are employed in the numerical solution. Some results are presented which illustrate the modeling options and numerical characteristics of the method.

Johnson, F. T.↗
NASA NTRS1975

A Three-Dimensional Solution of Flows over Wings with Leading-Edge Vortex Separation. Part 1: Engineering Document

A method of predicting forces, moments, and detailed surface pressures on thin, sharp-edged wings with leading-edge vortex separation in incompressible flow is presented. The method employs an inviscid flow model in which the wing and the rolled-up vortex sheets are represented by piecewise, continuous quadratic doublet sheet distributions. The Kutta condition is imposed on all wing edges. Computed results are compared with experimental data and with the predictions of the leading-edge suction analogy for a selected number of wing planforms over a wide range of angle of attack. These comparisons show the method to be very promising, capable of producing not only force predictions, but also accurate predictions of detailed surface pressure distributions, loads, and moments.

Brune, G. W.↗
NASA NTRS1975

A three-dimensional solution of flows over wings with leading-edge vortex separation

The application of a new, general, potential flow computational technique to the solution of the subsonic, three-dimensional flow over wings with leading-edge vortex separation is presented. The present method is capable of predicting forces, moments, and detailed surface pressures on thin, sharp-edged wings of rather arbitrary planform. The wing geometry is arbitrary in the sense that leading and trailing edges may be curved or kinked and the wing may have arbitrary camber and twist. The method employs an inviscid flow model in which the wing, the rolled-up vortex sheets, and the wake are represented by piecewise continuous quadratic doublet sheet distributions. The Kutta condition is imposed along all wing edges. Strengths of the doublet distributions as well as shape and position of the free fortex sheet spirals are computed in iterative fashion starting with an assumed initial sheet geometry. The method is verified by numerous computed results.

Weber, J. A.↗
NASA NTRS1975

Advanced panel-type influence coefficient methods applied to subsonic flows

An advanced technique for solving the linear integral equations of three-dimensional subsonic potential flows (steady, inviscid, irrotational and incompressible) about arbitrary configurations is presented. It involves assembling select, logically consistent networks whose construction comprises four tasks, which are described in detail: surface geometry definition; singularity strength definition; control point and boundary condition specification; and calculation of induced potential or velocity. The technique is applied to seven wing examples approached by four network types: source/analysis, doublet/analysis, source/design, and doublet/design. The results demonstrate the forgiveness of the model to irregular paneling and the practicality of combined analysis/design boundary conditions. The appearance of doublet strength mismatch is a valuable indicator of locally inadequate paneling.

Johnson, F. T.↗
NASA NTRS1972

Sideslip of wing-body combinations

A small-disturbance theory is developed for predicting the aerodynamics of an airplane in sideslip. Second-order terms involving the interaction between sideslip angle and angle of attack, sideslip angle and wing camber, etc., are retained. It is found that the second-order terms can produce the dominant sideslip effects when the dihedral of the lifting surfaces is small. Numerical implementation of the theory requires a solution procedure capable of producing accurate velocity gradients in the first-order solution.

Rubbert, P. E.↗