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At least 217 records · Page 12

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NASA NTRS1977

Application of a multi-level grid method to transonic flow calculations

A multi-level grid method has been studied as a possible means of accelerating convergence in relaxation calculations for transonic flows. The method employs a hierarchy of grids, ranging from very coarse (e.g., 8 x 2 mesh cells) to fine (e.g., 128 x 32); the coarser grids are used to diminish the magnitude of the smooth part of the residuals, hopefully with far less total work than would be required with, say, optimal SLOR iterations on the finest grid. The method was applied to the solution of the transonic small-disturbance equation for the velocity potential in the conservation form. Nonlifting transonic flow past a parabolic-arc airfoil is the example studied, with meshes of both constant and variable step size.

South, J. C., Jr.↗

NASA NTRS1978

Transonic wing analysis using advanced computational methods

This paper discusses the application of three-dimensional computational transonic flow methods to several different types of transport wing designs. The purpose of these applications is to evaluate the basic accuracy and limitations associated with such numerical methods. The use of such computational methods for practical engineering problems can only be justified after favorable evaluations are completed. The paper summarizes a study of both the small-disturbance and the full potential technique for computing three-dimensional transonic flows. Computed three-dimensional results are compared to both experimental measurements and theoretical results. Comparisons are made not only of pressure distributions but also of lift and drag forces. Transonic drag rise characteristics are compared. Three-dimensional pressure distributions and aerodynamic forces, computed from the full potential solution, compare reasonably well with experimental results for a wide range of configurations and flow conditions.

Henne, P. A.↗

NASA NTRS1978

The shock tube as a device for testing transonic airfoils at high Reynolds numbers

A performance analysis of gas-driven shock tubes shows that transonic airfoil flows with chord Reynolds numbers in the range of 100 million can be generated behind the primary shock in a large shock tube. A study of flow over simple airfoils has been carried out at low and intermediate Reynolds numbers to assess the testing technique. Results obtained from schlieren photos and airfoil pressure measurements show that steady transonic flows similar to those observed for the airfoils in wind tunnels can be generated within the available testing time in a shock tube with either properly-contoured test section walls or a properly-designed slotted-wall test section. The study indicates that the shock tube is a useful facility for studying two-dimensional high Reynolds number transonic airfoil flows.

Cook, W. J.↗

NASA NTRS1978

Design and physical characteristics of the Transonic Aircraft Technology (TACT) research aircraft

The Transonic Aircraft Technology (TACT) research program provided data necessary to verify aerodynamic concepts, such as the supercritical wing, and to gain the confidence required for the application of such technology to advanced high performance aircraft. An F-111A aircraft was employed as the flight test bed to provide full scale data. The data were correlated extensively with predictions based on data obtained from wind tunnel tests. An assessment of the improvement afforded at transonic speeds in drag divergence, maneuvering performance, and airplane handling qualities by the use of the supercritical wing was included in the program. Transonic flight and wind tunnel testing techniques were investigated, and specific research technologies evaluated were also summarized.

Painter, W. D.↗

NASA NTRS1979

The design of supercritical wings by the use of three-dimensional transonic theory

A procedure was developed for the design of transonic wings by the iterative use of three dimensional, inviscid, transonic analysis methods. The procedure was based on simple principles of supersonic flow and provided the designer with a set of guidelines for the systematic alteration of wing profile shapes to achieve some desired pressure distribution. The method was generally applicable to wing design at conditions involving a large region of supercriterical flow. To illustrate the method, it was applied to the design of a wing for a supercritical maneuvering fighter that operates at high lift and transonic Mach number. The wing profiles were altered to produce a large region of supercritical flow which was terminated by a weak shock wave. The spanwise variation of drag of this wing and some principles for selecting the streamwise pressure distribution are also discussed.

Mann, M. J.↗

NASA NTRS1976

Some recent progress in transonic flow computation

Although the development of a finite difference relaxation procedure to solve the steady form of equations of motion gave birth to the study of computational transonic aerodynamics and considerable progress has been made using the small disturbance theory, no general analytical solution method yet exists for transonic flows that include three dimensional unsteady, and viscous effects. Two techniques are described which are useful in computational transonic aerodynamics applications. The finite volume method simplifies the application of boundary conditions without introducing the constriction associated with small disturbance theory. Governing equations are solved in a Cartesian coordinate system using a body-oriented and shock-oriented mesh network. Only the volume and surface normal directions of the volume elements must be known. The other method, configuration design by numerical optimization, can be used by aircraft designers to develop configurations that satisfy specific geometric performance constraints. Two examples of airfoil design by numerical optimization are presented.

Ballhaus, W. F.↗

NASA NTRS1978

Finite elements and finite differences for transonic flow calculations

The paper reviews the chief finite difference and finite element techniques used for numerical solution of nonlinear mixed elliptic-hyperbolic equations governing transonic flow. The forms of the governing equations for unsteady two-dimensional transonic flow considered are the Euler equation, the full potential equation in both conservative and nonconservative form, the transonic small-disturbance equation in both conservative and nonconservative form, and the hodograph equations for the small-disturbance case and the full-potential case. Finite difference methods considered include time-dependent methods, relaxation methods, semidirect methods, and hybrid methods. Finite element methods include finite element Lax-Wendroff schemes, implicit Galerkin method, mixed variational principles, dual iterative procedures, optimal control methods and least squares.

Hafez, M. M.↗

NASA NTRS1979

Computational optimization and wind tunnel test of transonic wing designs

A practical procedure for the optimum design of transonic wings is demonstrated. The procedure uses an optimization program based on the method of feasible directions coupled with an aerodynamic analysis program which solves the three-dimensional potential equation for subsonic through transonic flow. Two new wings for the A-7 aircraft were designed by using the optimization procedure to achieve specified surface pressure distributions. The new wings, along with the existing A-7 wing, were tested in the Ames 11-foot transonic wind tunnel. The experimental data show that all of the performance goals were met.

Haney, H. P.↗

NASA NTRS1979

Numerical computation of aeroelastically corrected transonic loads

A numerical scheme is presented for the computation of transonic aerodynamic loads on flexible wings. The method consists of iteratively applying the loads computed by a 3D transonic aerodynamics code to a structural model to obtain elastic twist, and then recomputing the loads. Because this iteration is performed concurrently with the iterations performed in computing the aerodynamics, flexible loads are obtained in roughly the same amount of computing time as required to obtain rigid loads. Applications of this method to a flexible supercritical transonic transport wing are presented and compared with model test data.

Chipman, R.↗

NASA NTRS1979

Experiments in unsteady transonic flow

The static and dynamic response of a 0.5-m-chord airfoil were measured in the NASA-Ames 11- by 11-Foot Transonic Wind Tunnel. The effects of mean angle of attack, Reynolds number, oscillation mode, and frequency were investigated over a range of subsonic and transonic Mach numbers. Unsteady pressure distributions and loads on an oscillating NACA 64A010 airfoil are discussed. The unsteady pressure distributions are compared with classical subsonic theory and with newer unsteady aerodynamic codes. The experimental data are also used to assess the validity of linearity and modal superposition in the transonic-flow regime.

Davis, S. S.↗

NASA NTRS1979

Artificial compressibility methods for numerical solutions of transonic full potential equation

New methods for transonic flow computations based on the full potential equation in conservation form are presented. The idea is to modify slightly the density (due to the artificial viscosity in the supersonic region), and solve the resulting elliptic-like problem iteratively. It is shown that standard discretization techniques (central differencing) as well as some standard iterative procedures (SOR, ADI, and explicit methods) are applicable to the modified transonic mixed-type equation. Calculations of transonic flows around cylinders and airfoils are discussed with special emphasis on the explicit methods that are suitable for vector processing on the STAR 100 computer.

Hafez, M.↗

NASA NTRS1980

Transonic wing DFVLR-F4 as European test model

A transonic wing, the DFVLR-F4 was designed and tested as a model in European transonic wind tunnels and was found to give performance improvements over conventional wings. One reason for the improvement was the reduction of compression shocks in the transonic region as the result of improved wing design.

Redeker, G.↗

NASA NTRS1976

Application of a multi-level grid method to transonic flow calculations

A multi-level grid method was studied as a possible means of accelerating convergence in relaxation calculations for transonic flows. The method employs a hierarchy of grids, ranging from very coarse to fine. The coarser grids are used to diminish the magnitude of the smooth part of the residuals. The method was applied to the solution of the transonic small disturbance equation for the velocity potential in conservation form. Nonlifting transonic flow past a parabolic arc airfoil is studied with meshes of both constant and variable step size.

South, J. C., Jr.↗

NASA NTRS1980

A comprehensive comparison between experiment and prediction for a transonic turbulent separated flow

Attempts to predict surface pressure distributions on lifting surfaces have been relatively unsuccessful in the transonic regime when the shock wave is of sufficient strength to produce an extensive region of turbulent separated flow. For these conditions, the viscous flow behavior must be accurately described even to obtain reasonable predictions of surface pressure. The present paper addresses this problem. Detailed comparisons between prediction and experiment are made for a transonic, turbulent boundary-layer separation (freestream Mach number = 0.875) for which the turbulent flow properties (including the turbulent Reynolds stress) had been measured by the laser velocimeter technique from upstream of the separated region through reattachment. The flow was generated on an axisymmetric 'bump' model designed to simulate the flow on an airfoil at transonic conditions. The numerical methods used in the comparisons include the solution of the time-dependent, mass-averaged Navier-Stokes equations, and the solution of the compressible boundary-layer equations by the inverse method. Solutions were obtained for the well established Cebeci-Smith algebraic turbulence model and the more recently developed Wilcox-Rubesin two-equation turbulence model.

Johnson, D. A.↗

NASA NTRS1980

Top inlet system feasibility for transonic-supersonic fighter aircraft applications

Top inlet flow field and inlet performance data are presented which provide preliminary insight into the feasibility of upper-fuselage mounted inlet systems for transonic-supersonic fighter aircraft. Presented data span the Mach 0.2 to 2.0 envelope and enable evaluation of the influence of key aircraft configuration variables - inlet location, wing position, wing leading-edge extension (LEX) planform area, and variable incidence canards - on top inlet performance. The viability of this concept relative to more conventional inlet/airframe integrations is assessed via comparative evaluation of top and conventional inlet flow field parameters at transonic and supersonic speeds. It is shown that the action of the wing LEX vortex system produces a significant improvement in top inlet performance. Currently available transonic-supersonic data indicate that top inlet systems pose a viable configuration option for fighter aircraft requiring moderate angle of attack capability. However, recently acquired data indicate that increased angle of attack capability may be obtained by increasing wing leading-edge sweep angle.

Williams, T. L.↗

NASA NTRS1980

Evaluation of flow quality in two large NASA wind tunnels at transonic speeds

Wind tunnel testing of low drag airfoils and basic transition studies at transonic speeds are designed to provide high quality aerodynamic data at high Reynolds numbers. This requires that the flow quality in facilities used for such research be excellent. To obtain a better understanding of the characteristics of facility disturbances and identification of their sources for possible facility modification, detailed flow quality measurements were made in two prospective NASA wind tunnels. Experimental results are presented of an extensive and systematic flow quality study of the settling chamber, test section, and diffuser in the Langley 8 foot transonic pressure tunnel and the Ames 12 foot pressure wind tunnel. Results indicate that the free stream velocity and pressure fluctuation levels in both facilities are low at subsonic speeds and are so high as to make it difficult to conduct meaningful boundary layer control and transition studies at transonic speeds.

Harvey, W. D.↗

NASA NTRS1981

TAIR: A transonic airfoil analysis computer code

The operation of the TAIR (Transonic AIRfoil) computer code, which uses a fast, fully implicit algorithm to solve the conservative full-potential equation for transonic flow fields about arbitrary airfoils, is described on two levels of sophistication: simplified operation and detailed operation. The program organization and theory are elaborated to simplify modification of TAIR for new applications. Examples with input and output are given for a wide range of cases, including incompressible, subcritical compressible, and transonic calculations.

Dougherty, F. C.↗

NASA NTRS1980

Numerical solutions of transonic flows by parametric differentiation and integral equation techniques

The paper discusses results of an exploratory study of the advantages obtained by combining integral equation methods with the method of parametric differentiation in the treatment of transonic flow problems. In the proposed method, the nonlinear unsteady transonic flow equation for small perturbations is transformed into a linear equation by the use of the method of parametric differentiation. The linear equation is split into a pair of weakly coupled partial differential equations by writing the transformed perturbation potential as the sum of a steady component and an unsteady component. The solution of the steady equation as an integral equation is based on Ogana's treatment (1978). As a test case, the formulation is applied to predict the steady transonic flow over a nonlifting parabolic-arc airfoil.

Sivaneri, N. T.↗