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NASA NTRS1987

Pressure field due to drag reducing outer layer devices in turbulent boundary layers

The wall static pressure in the vicinity of drag reducing outer layer devices in flat wall turbulent boundary layers has been measured and compared with an inviscid theory. Symmetric and cambered airfoil devices have been examined at small angles of attack and very low chord Reynolds numbers. Airfoil devices impose a sequence of strong favorable and adverse pressure gradients on the boundary layer whose drag is to be reduced. At very small angles of attack (+ or - 2 deg), this pressure field extends up to about three chord lengths downstream of the trailing edge of an airfoil device. Also examined are the pressures on the upper and lower surfaces of a symmetric airfoil device in the freestream and near the wall. The freestream pressure distribution around an airfoil section is altered by the wall proximity. The relevance of lift enhancement caused by wall proximity to drag reduction has been discussed. The pressure distributions on the flat wall beneath the symmetric airfoil devices are predicted well by the inviscid theory. However, the remaining pressure distributions are predicted only qualitatively, presumably because of strong viscous effects.

Bandyopadhyay, P. R.↗

NASA NTRS1988

A transonic wind tunnel wall interference prediction code

A small disturbance transonic wall interference prediction code has been developed that is capable of modeling solid, open, perforated, and slotted walls as well as slotted and solid walls with viscous effects. This code was developed by modifying the outer boundary conditions of an existing aerodynamic wing-body-pod-pylon-winglet analysis code. The boundary conditions are presented in the form of equations which simulate the flow at the wall, as well as finite difference approximations to the equations. Comparisons are presented at transonic flow conditions between computational results and experimental data for a wing alone in a solid wall wind tunnel and wing-body configurations in both slotted and solid wind tunnels.

Phillips, Pamela S.↗

NASA NTRS1988

Turbofan forced mixer lobe flow modeling. 1: Experimental and analytical assessment

A joint analytical and experimental investigation of three-dimensional flowfield development within the lobe region of turbofan forced mixer nozzles is described. The objective was to develop a method for predicting the lobe exit flowfield. In the analytical approach, a linearized inviscid aerodynamical theory was used for representing the axial and secondary flows within the three-dimensional convoluted mixer lobes and three-dimensional boundary layer analysis was applied thereafter to account for viscous effects. The experimental phase of the program employed three planar mixer lobe models having different waveform shapes and lobe heights for which detailed measurements were made of the three-dimensional velocity field and total pressure field at the lobe exit plane. Velocity data was obtained using Laser Doppler Velocimetry (LDV) and total pressure probing and hot wire anemometry were employed to define exit plane total pressure and boundary layer development. Comparison of data and analysis was performed to assess analytical model prediction accuracy. As a result of this study a planar mixed geometry analysis was developed. A principal conclusion is that the global mixer lobe flowfield is inviscid and can be predicted from an inviscid analysis and Kutta condition.

Barber, T.↗

NASA NTRS1988

Aerodynamic applications of an efficient incompressible Navier-Stokes solver

A robust and accurate upwind scheme based on flux-difference splitting for three-dimensional, incompressible flows in general coordinates is described. Viscous effects are included through a thin-layer approximation to the Navier-Stokes equations. Turbulent flows are simulated with extensions to the Baldwin-Lomax (1978) turbulence model. The effects of accuracy and gridding on the computed flow field results are assessed. The results obtained from the code for vortical flows over a delta wing, a double delta wing, and a tangent-ogive forebody compare well with experimental data.

Hartwich, Peter M.↗

NASA NTRS1989

Viscous-inviscid interaction and local grid refinement via truncation error injection

A methodology is presented which makes it possible to decouple a complex problem having multiple disparate length scales into problems of single length scale so that they can be solved more efficiently on a computer. The method is applied to a viscous transonic flow over an airfoil. It is found that accurate prediction of the flow over an airfoil can be obtained by solving the Euler equations on a relatively coarse global grid with viscous effects computed separately on a boundary-layer type grid and injected into the global grid solution as a combination of vorticity and trucation error.

Goble, Brian D.↗

NASA NTRS1988

Robust vortex methods for three-dimensional incompressible flows

Vortex methods for the numerical simulation of incompressible three-dimensional flows at high Reynolds number are discussed. Particular emphasis is placed on schemes that prevent the excessive generation of mesh elements, include viscous effects, and are relatively inexpensive even as the number of elements becomes large. Several simulations are presented involving the interaction of vortex rings.

Chua, K.↗

NASA NTRS1990

An analysis of the viscous flow through a compact radial turbine by the average passage approach

A steady, three-dimensional viscous average passage computer code is used to analyze the flow through a compact radial turbine rotor. The code models the flow as spatially periodic from blade passage to blade passage. Results from the code using varying computational models are compared with each other and with experimental data. These results include blade surface velocities and pressures, exit vorticity and entropy contour plots, shroud pressures, and spanwise exit total temperature, total pressure, and swirl distributions. The three computational models used are inviscid, viscous with no blade clearance, and viscous with blade clearance. It is found that modeling viscous effects improves correlation with experimental data, while modeling hub and tip clearances further improves some comparisons. Experimental results such as a local maximum of exit swirl, reduced exit total pressures at the walls, and exit total temperature magnitudes are explained by interpretation of the flow physics and computed secondary flows. Trends in the computed blade loading diagrams are similarly explained.

Heidmann, James D.↗

NASA NTRS1989

Dynamics of a class of vortex rings

The contour dynamics method is extended to vortex rings with vorticity varying linearly from the symmetry axis. An elliptic core model is also developed to explain some of the basic physics. Passage and collisions of two identical rings are studied focusing on core deformation, sound generation and stirring of fluid elements. With respect to core deformation, not only the strain rate but how rapidly it varies is important and accounts for greater susceptibility to vortex tearing than in two dimensions. For slow strain, as a passage interaction is completed and the strain relaxes, the cores return to their original shape while permanent deformations remain for rapidly varying strain. For collisions, if the strain changes slowly the core shapes migrate through a known family of two-dimensional steady vortex pairs up to the limiting member of the family. Thereafter energy conservation does not allow the cores to maintain a constant shape. For rapidly varying strain, core deformation is severe and a head-tail structure in good agreement with experiments is formed. With respect to sound generation, good agreement with the measured acoustic signal for colliding rings is obtained and a feature previously thought to be due to viscous effects is shown to be an effect of inviscid core deformation alone. For passage interactions, a component of high frequency is present. Evidence for the importance of this noise source in jet noise spectra is provided. Finally, processes of fluid engulfment and rejection for an unsteady vortex ring are studied using the stable and unstable manifolds. The unstable manifold shows excellent agreement with flow visualization experiments for leapfrogging rings suggesting that it may be a good tool for numerical flow visualization in other time periodic flows.

Shariff, Karim↗

NASA NTRS1989

Improved vortex methods for three-dimensional flows

Robust numerical methods are developed for three-dimensional incompressible vortical flows, using Lagrangian vortex elements. A successful scheme must be able to handle regions of intense vortex stretching and vortex reconnection with reasonable accuracy (without diverging). Here, consideration is given to vortex particles, also commonly called vortons or vortex sticks. The following issues are discussed: (1) use of delta-function elements and weak solutions of the vorticity equation; (2) use of smoothed elements and the choice of the smoothing function; (3) representation of viscous effects and the redistribution of element strength; and (4) conservation laws (are they satisfied?). The various proposed schemes have been tested on flows involving a strong interaction between two vortex rings.

Winckelmans, G.↗

NASA NTRS1990

Supersonic flameholding by attached oblique shock waves

The object of this investigation was to explore various regimes of shock-induced combustion that may occur as a result of supersonic flow of a reactive gas mixture past a simple compression ramp or wedge. Attention was confined to combinations of wedge angles, fuel-air mixture ratios and freestream velocities that permit the resulting oblique shock or oblique detonation wave to remain attached. Computational fluid dynamic (CFD) simulations, both inviscid and viscous, were performed and analyzed in order to investigate two modes of compression ramp flameholding; namely, complete oblique detonation wave (ODW) and shock-induced combustion. In addition, the possibility of flame persistence or hysteresis was explored. It was preliminarily concluded that inclusion of viscous effects may be essential to accurately predict the occurrence of the various flameholding modes.

Fort, James A.↗

NASA NTRS1989

Aerodynamic stability and heating analyses for the aeroassist flight experiment vehicle

Since ground based flow simulations are presently unable to model flight conditions expected for AOTVs (Aeroassist Orbital Transfer Vehicle) and other hypersonic space vehicles, computer codes are being developed to provide design parameters necessary for structure, guidance, and control aspects. Over the past four years, VRFLO (Viscous Reactive Flow) has been written to model finite-rate chemistry and viscous effects for a variety of aerobrake bodies. VRFLO includes a number of unique features that are summarized as follows: (1) Grid generation is an integral part of the code for several aerobrake configurations which includes the wake flow region; (2) The formulation is valid for three air chemical models; (3) An ADI central difference technique is used to solve the Navier-Stokes and species continuity equations in split groups; and (4) Grid density and numerical damping are minimized by shock-fitting and conformal mapping of body points.

Mcgary, J.↗

NASA NTRS1990

Grid generation and flow computation about a Martian entry vehicle

A number of vehicles are currently being proposed for a manned mission to Mars. One of these vehicles has a modified blunt-nosed cone configuration. Experimental results were obtained for this vehicle in 1968. They show lift-over-drag ratios comparable to those needed for Mars entry. Computations are performed to verify the earlier results and to further describe the flight characteristics of this vehicle. An analytical method is used to define the surface of this vehicle. A single-block volume grid is generated around the vehicle using the algebraic Two-Boundary Grid Generation algorithm (TBGG) and transfinite interpolation. Euler solutions are then obtained from a Langley Aerodynamic Upward Relaxation Algorithm (LAURA) at Mach 6.0 and angles of attack of 0, 6, and 12 deg. The lift coefficient determined from the LAURA code agree very well with the experimental results. The drag and pitching moment coefficients, however, are underestimated by the code since viscous effects are not considered. Contour plots of the flowfield show no evidence of separation for angles of attack up to 12 deg.

Stewart, J. E.↗

NASA NTRS1990

A method for the design of transonic flexible wings

Methodology was developed for designing airfoils and wings at transonic speeds which includes a technique that can account for static aeroelastic deflections. This procedure is capable of designing either supercritical or more conventional airfoil sections. Methods for including viscous effects are also illustrated and are shown to give accurate results. The methodology developed is an interactive system containing three major parts. A design module was developed which modifies airfoil sections to achieve a desired pressure distribution. This design module works in conjunction with an aerodynamic analysis module, which for this study is a small perturbation transonic flow code. Additionally, an aeroelastic module is included which determines the wing deformation due to the calculated aerodynamic loads. Because of the modular nature of the method, it can be easily coupled with any aerodynamic analysis code.

Smith, Leigh Ann↗

NASA NTRS1991

Navier-Stokes simulation of the supersonic combustion flowfield in a ram accelerator

A computational study of the ram accelerator, a ramjet-in-tube device for accelerating projectiles to ultrahigh velocities, is presented. The analysis is performed using a fully implicit TVD scheme that efficiently solves the Reynolds-averaged Navier-Stokes equations and the species continuity equations associated with a finite rate combustion model. Previous analyses of this concept were based on inviscid assumptions. The present results indicate that viscous effects are of primary importance; in all the cases studied, shock-induced combustion always started in the boundary layer. The effects of Mach number, mixture composition, pressure, and turbulence are investigated for various configurations. Two types of combustion processes, one stable and the other unstable, were observed depending on the inflow conditions. In the unstable case, a detonation wave is formed, which propagates upstream and unstarts the ram accelerator. In the stable case, a solution that converges to steady-state is obtained, in which the combustion wave remains stationary with respect to the ram accelerator projectile. The possibility of stabilizing the detonation wave by means of a backward facing step is also investigated. In addition to these studies, two numerical techniques were tested. These two techniques are vector extrapolation to accelerate convergence, and a diagonal formulation that eliminates the expense of inverting large block matrices that arise in chemically reacting flows.

Yungster, Shaye↗

NASA NTRS1990

Prediction of inviscid stagnation pressure losses in supersonic inlet flows

An effort is made to quantify the stagnation pressure losses associated with shock-wave systems that may be present in such high Mach number flows as those of scramjet hypersonic diffusers. If the shock-related contribution turns out to be much larger than that attributable to viscous effects, a designer could introduce methods for the minimization of the shock system's scale; in particular, the size of the normal shock should be reduced. The angles presently treated may be approached during vehicle maneuvering or other transients.

Azevedo, David J.↗

NASA NTRS1991

Analysis of flow on cones and cylinders using discrete vortex methods

Discrete vortex methods have been developed to investigate the vortex flows on cones and two-dimensional cylinders. The cone problem was solved by assuming that, to meet conical flow and zero-force conditions, the vortices move radially away from the body at a given cross-section. The two-dimensional cylinder problem was solved by limiting the velocity along the zero streamline surrounding the vortex field to two times freestream velocity and by limiting the lateral movement of the vortices. Variations in vortex position and strength with time were determined by taking into account the rate at which circulation is generated at separation points on the body. The calculated vortex positions and strengths were in good agreement with available experimental data. Viscous effects could be accounted for by adding empirically determined damping terms to the velocity equations. The models indicate that different types of asymmetry occur for the cone and two-dimensional cylinder. Asymmetry onset boundaries determined by the discrete vortex method show the same trend as experiment.

Gainer, Thomas G.↗

NASA NTRS1990

Analysis of cone-derived waveriders by hypersonic small-disturbance theory

Hypersonic small-disturbance theory is used to analyze slender waveriders derived from axisymmetric flows past circular cones. Viscous effects are accounted for by means of laminar boundary-layer theory. The shape of the waverider is determined by specifying the upper freestream surface, from which the lower compression surface can then be obtained. When this is done, the lift, drag, and pitching moment are determined in terms of quadratures over the shock layer in the base plane. They are functions of freestream Mach number, freestream Reynolds number based on the length of the waverider, some measure of slenderness, and other parameters relating to the shape of the waverider. The functional relationships can be cast in terms of similarity laws. The lift-to-drag ratio is determined for a wide range of shapes and parameters.

Rasmussen, Maurice L.↗

NASA NTRS1991

The design/analysis of flows through turbomachinery - A viscous/inviscid approach

A new design/analysis system for the flows through turbomachinery is currently being developed for studying turbomachinery problems with an axisymmetric viscous/inviscid 'average-passage' throughflow code. The advantage of this approach, compared to streamline curvature codes, is that the solutions obtained simulate some of the unsteadiness, compressibility and viscous effects of a multistage turbomachine. The design/analysis system consists of three elemental parts, the axisymmetric block grid generator, the blade surface element code, and the axisymmetric flow code. Each element of the system will be discussed and the flow solutions for three axisymmetric geometries will be shown compared to experimental data where available. The computations are shown to be in very good agreement with test data for SR7 spinner body and transonic boattail geometry obtained in the wind tunnels at NASA Lewis Research Center. The VIADAC Rotor 67 Fan results were compared to PARC2D calculated results and shown to be in very good agreement.

Miller, D. P.↗