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At least 397 records · Page 22

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NASA NTRS1974

An eddy-viscosity treatment of the unsteady turbulent boundary layer on a flat plate in an expansion tube

An analysis is presented for the relaxation of a turbulent boundary layer on a semiinfinite flat plate after passage of a shock wave and a trailing driver gas-driven gas interface. The problem has special application to expansion tube flows. The flow-governing equations have been transformed into the Lamcrocco variables. The numerical results indicate that a fully turbulent boundary layer relaxes faster to the final steady-state values of heat transfer and skin-friction than a fully laminar boundary layer.

Gupta, R. N.↗

NASA NTRS1979

Longitudinal vortices in a transitioning boundary layer

Naturally occurring spanwise variations of the streamwise velocity component, characteristic of longitudinal vortices embedded in a transitioning boundary layer were explored using hot-wire anemometers. A vibrating ribbon introduced stable or unstable Tollmien-Schlichting waves into the laminar boundary layer. These damped or growing disturbances always developed a strong three dimensional pattern even though no spanwise perturbations were artificially induced. Changing the radius of the leading edge and other modifications to the flat plate, wind tunnel and boundary layer did not alter the spanwise wavelength of the vortices.

Anders, J. B.↗

NASA NTRS1989

Multigrid solution of the Navier-Stokes equations on triangular meshes

A Navier-Stokes algorithm for use on unstructured triangular meshes is presented. Spatial discretization of the governing equations is achieved using a finite element Galerkin approximation, which can be shown to be equivalent to a finite volume approximation for regular equilateral triangular meshes. Integration steady-state is performed using a multistage time-stepping scheme, and convergence is accelerated by means of implicit residual smoothing and an unstructured multigrid algorithm. Directional scaling of the artificial dissipation and the implicit residual smoothing operator is achieved for unstructured meshes by considering local mesh stretching vectors at each point. The accuracy of the scheme for highly stretched triangular meshes is validated by comparing computed flat-plate laminar boundary layer results with the well known similarity solution, and by comparing laminar airfoil results with those obtained from various well-established structured quadrilateral-mesh codes. The convergence efficiency of the present method is also shown to be competitive with those demonstrated by structured quadrilateral-mesh algorithms.

Mavriplis, Dimitri J.↗

NASA NTRS1990

Multigrid solution of the Navier-Stokes equations on triangular meshes

Mavriplis, D. J.↗

NASA NTRS2006

Shuttle Return To Flight Experimental Results: Protuberance Effects on Boundary Layer Transition

The effect of isolated roughness elements on the windward boundary layer of the Shuttle Orbiter has been experimentally examined in the Langley Aerothermodynamic Laboratory in support of an agency-wide effort to prepare the Shuttle Orbiter for return to flight. This experimental effort was initiated to provide a roughness effects database for developing transition criteria to support on-orbit decisions to repair damage to the thermal protection system. Boundary layer transition results were obtained using trips of varying heights and locations along the centerline and attachment lines of 0.0075-scale models. Global heat transfer images using phosphor thermography of the Orbiter windward surface and the corresponding heating distributions were used to infer the state of the boundary layer (laminar, transitional, or turbulent). The database contained within this report will be used to formulate protuberance-induced transition correlations using predicted boundary layer edge parameters.

Liechty, Derek S.↗

NASA NTRS1988

Flight test results for the Daedalus and Light Eagle human powered aircraft

The results of the flight test program of the Daedalus and Light Eagle human powered aircraft in the winter of 1987/88 are given. The results from experiments exploring the Light Eagle's rigid body and structural dynamics are presented. The interactions of these dynamics with the autopilot design are investigated. Estimates of the power required to fly the Daedalus aircraft are detailed. The system of sensors, signal conditioning boards, and data acquisition equipment used to record the flight data is also described. In order to investigate the dynamics of the aircraft, flight test maneuvers were developed to yield maximum data quality from the point of view of estimating lateral and longitudinal stability derivatives. From this data, structural flexibility and unsteady aerodynamics have been modeled in an ad hoc manner and are used to augment the equations of motion with flexibility effects. Results of maneuvers that were flown are compared with the predictions from the flexibility model. To extend the ad hoc flexibility model, a fully flexible aeroelastic model has been developed. The model is unusual in the approximate equality of many structural natural frequencies and the importance of unsteady aerodynamic effects. the Gossamer Albatross. It is hypothesized that this inverse ground effect is caused by turbulence in the Earth's boundary layer. The diameters of the largest boundary layer eddies (which represent most of the turbulent kinetic energy) are proportional to altitude; thus, closer to the ground, the energy in the boundary layer becomes concentrated in eddies of smaller and smaller diameter. Eventually the eddies become sufficiently small (approximately 0.5 cm) that they trip the laminar boundary layer on the wing. As a result, a greater percentage of the wing area is covered with turbulent flow. Consequently the aircraft's drag and the pow er required both increase as the aircraft flies closer to the ground. The results of the flight test program are examined critically, especially with respect to future applications. Maneuvers that allow the observation of stability derivatives for flexible aircraft are detailed. Considerations for the design of autopilots for future human power aircraft and high-altitude RPV's are discussed. Techniques useful in estimating the power required to fly aircraft of very high lift to drag ratio are described.

Sullivan, R. Bryan↗

NASA NTRS1982

Observations and implications of natural laminar flow on practical airplane surfaces

The results of natural laminar flow (NLF) experiments conducted by NASA to determine if modern aircraft structures can benefit from NLF as do sailplanes are presented. Seven aircraft, ranging from a Cessna 210 to a Learjet 28/29, with relatively stiff skins were flown in production configurations with no modifications. Measurements were made of the boundary-layer laminar to turbulent transition locations on various aerodynamic surfaces, the effect of a total loss of laminar flow, the effect of the propeller slipstream on the wing boundary-layer transition and the boundary-layer profiles, the wing section profile drag, the effect of flight through clouds, and insect debris contamination effects. Favorable pressure gradients for NLF were concluded to be feasible up to a transition Reynolds number of 11 million. Laminar flows were observed in propeller slipstreams, and insects were found to cause transition 1/4 of the time.

Holmes, B. J.↗

NASA NTRS1987

Boundary-layer receptivity and laminar-flow airfoil design

Boundary-layer receptivity examines the way in which external disturbances generate instability waves in boundary layers. Receptivity theory is complementary to stability theory, which studies the evolution of disturbances that are already present in the boundary layer. A transition prediction method which combines receptivity with linear stability theory would directly account for the influence of free-stream disturbances and also consider the characteristics of the boundary layer upstream of the neutral stability point. The current e sup N transition prediction methods require empirical correlations for the influence of environmental disturbances, and totally ignore the boundary layer characteristics upstream of the neutral stability point. The regions where boundary-layer receptivity occurs can be separated into two classes, one near the leading edges and the other at the downstream points where the boundary layer undergoes rapid streamwise adjustments. Analyses were developed for both types of regions, and parametric studies which examine the relative importance of different mechanisms were carried out. The work presented here has focused on the low Mach number case. Extensions to high subsonic and supersonic conditions are presently underway.

Kerschen, Edward J.↗

NASA NTRS1990

The semidiscrete Galerkin finite element modeling of compressible viscous flow past an airfoil

The primary project was the numerical simulation, by a finite element/finite difference method, of the viscous flow about an airfoil. The secondary project involved the numerical simulation of the three-dimensional separated and vortex-dominated flow about a hemispherically capped cylinder in the transonic regime. Preliminary calculations were started for the hemisphere-cylinder at 0 and 5 degree angle of attack. The solution of the flow field about airfoils and wings is required to determine the important parameters of lift, moment, and drag. Viscous effects must be accounted for if the drag is to be accurately calculated. At present there are basically two approaches to the numerical simulation of the flow field, the use of fully viscous models and the inviscid/viscous models. The fully viscous models require the solution of an approximation of the Navier-Stokes equations and therefore should simulate most of the physical mechanisms. A fast, accurate, and computationally efficient inviscid flow solver was recently developed by Hartwich. It is thought that Hartwich's program coupled to a fast, accurate, and computationally efficient boundary layer code, will make an excellent tool for airfoil design. The purpose of the primary project was to develop a compressible boundary layer code using the semidiscrete Galerkin finite element method. The numerical scheme employed used the combination of a Dorodnitsyn formulation of the boundary layer equations, with a finite difference/finite element procedure (semidiscrete Galerkin method), in the solution of the compressible two-dimensional boundary layer equations. A laminar compressible boundary layer code was developed and tested for a NACA 0012 airfoil at a Mach number of 0.5, a Reynolds number of 5000, and zero angle of attack. At present the boundary layer program solves up to, but not beyond, separation.

Meade, Andrew J.↗

NASA NTRS1995

The role of nonlinear critical layers in boundary layer transition

Asymptotic methods are used to describe the nonlinear self-interaction between pairs of oblique instability modes that eventually develops when initially linear spatially growing instability waves evolve downstream in nominally two-dimensional laminar boundary layers. The first nonlinear reaction takes place locally within a so-called 'critical layer', with the flow outside this layer consisting of a locally parallel mean flow plus a pair of oblique instability waves - which may or may not be accompanied by an associated plane wave. The amplitudes of these waves, which are completely determined by nonlinear effects within the critical layer, satisfy either a single integro-differential equation or a pair of integro-differential equations with quadratic to quartic-type nonlinearities. The physical implications of these equations are discussed.

Goldstein, M.E.↗

NASA NTRS1979

Outline of research on oscillating boundary layers

The state of the art in the field of unsteady boundary layers is outlined with emphasis on turbulent boundary layers. The unsteady flows considered are mainly periodic with the external velocity varying around a zero or nonzero mean time value. The principal results obtained on laminar boundary layers are also presented.

Cousteix, J.↗

NASA NTRS1982

Experiments on the stability of the flat-plate boundary layer with suction

Experiments have been conducted in the VPI and SU Stability Wind Tunnel on a flat-plate wind-tunnel model equipped with porous suction panels. Detailed hot-wire measurements were conducted in the laminar boundary layer to investigate the stabilizing effects of suction on growing Tollmien-Schlichting waves, which were introduced into the boundary layer using a vibrating ribbon. Special care was taken to minimize external disturbances and to avoid extraneous experimental bias. The measurements, which included mean-flow and disturbance-amplitude profiles across the boundary layer, showed that suction applied through discrete porous strips can be as effective as suction applied continuously over a much longer streamwise length. The measurements also showed that suction is more effective when placed forward, nearer to the region of neutral stability, than when placed in the region of maximum growth rate. These results also provided meaningful comparison with recent theory.

Reynolds, G. A.↗

NASA NTRS1978

A computer program for calculating laminar and turbulent boundary layers for two-dimensional time-dependent flows

A computer program is described which provides solutions of two dimensional equations appropriate to laminar and turbulent boundary layers for boundary conditions with an external flow which fluctuates in magnitude. The program is based on the numerical solution of the governing boundary layer equations by an efficient two point finite difference method. An eddy viscosity formulation was used to model the Reynolds shear stress term. The main features of the method are briefly described and instructions for the computer program with a listing are provided. Sample calculations to demonstrate its usage and capabilities for laminar and turbulent unsteady boundary layers with an external flow which fluctuated in magnitude are presented.

Cebeci, T.↗

NASA NTRS1989

Investigation of CAST-10-2/DOA 2 airfoil in NAE high Reynolds number two-dimensional test facility

A common airfoil model with the CAST 10-2/DOA-2 profile and 228 mm (9 inches) chord length was tested. The tests performed in NAE covered the Mach numbers from 0.3 to 0.8 and Reynolds numbers from 10 to 30 million. The model was tested with transition free and with transition fixed at 5 percent chord for both the upper and the lower surfaces. The data obtained were analyzed for the effects of Reynolds number, transition fixing and Mach number. The role of the boundary layer on the displacement effect, the interaction with the shock wave and the trailing edge separation are examined. The results are summarized as follows: (1) the airfoil performance depends strongly on Reynolds number and transition fixing; (2) with transition fixed, the aerodynamic quantities such as lift, pitching moment and drag show a monotonic variation with Reynolds number; (3) with transition free, the aerodynamic quantities vary less regularly with Reynolds number and a slight parametric dependency is shown. The weak dependency is due to the compensatory effect of the forward shift of the transition position and the thinning of the turbulent boundary layer as Reynolds number increases; (4) the shock Mach number and the shock position are weakly dependent on Reynolds number; and (5) the long extent of the laminar boundary layer at transonic speeds reduces the drag appreciably at low Reynolds numbers. The drag bucket around the design Mach number can be observed below Reynolds number 15 million.

Chan, Y. Y.↗

NASA NTRS1944

Integral methods in the theory of the boundary layer

The application of the well-known basic principle of mechanics, the principle of Jourdain, to problems of the theory of the boundary layerleads tp an equation from which the equations of Von Karman, Leibenson, and Golubev are derived as special cases. The given equation may be employed in other integral methods. The present paper deals with the method of the variation of the thickness of the boundary layer. A number of new approximate formulas valuable in aerodynamic calculations for the fristion distribution are derived from this procedure. The method has been applied only to laminar boundary layers, but it seems probable that it may be generalized to include turbulent layers as well.

Loitsianskii, L G↗

NASA NTRS1995

The Effect of Nonlinear Critical Layers on Boundary Layer Transition

Asymptotic methods are used to describe the nonlinear self-interaction between pairs of oblique instability modes that eventually develops when initially linear and spatially growing instability waves evolve downstream in nominally two-dimensional and spanwise periodic laminar boundary layers. The first nonlinear reaction takes place locally within a so-called 'critical layer' with the flow outside this layer consisting of a locally parallel mean flow plus an appropriate superposition of linear instability waves. The amplitudes of these waves are determined by either a single integro-differential equation or by a pair of integro-differential equations with quadratic to quartic-type nonlinearities.

Goldstein, Marvin E.↗

NASA NTRS2000

Direct Numerical Simulation of Automobile Cavity Tones

The Navier Stokes equation is solved computationally by the Dispersion-Relation-Preserving (DRP) scheme for the flow and acoustic fields associated with a laminar boundary layer flow over an automobile door cavity. In this work, the flow Reynolds number is restricted to R(sub delta*) < 3400; the range of Reynolds number for which laminar flow may be maintained. This investigation focuses on two aspects of the problem, namely, the effect of boundary layer thickness on the cavity tone frequency and intensity and the effect of the size of the computation domain on the accuracy of the numerical simulation. It is found that the tone frequency decreases with an increase in boundary layer thickness. When the boundary layer is thicker than a certain critical value, depending on the flow speed, no tone is emitted by the cavity. Computationally, solutions of aeroacoustics problems are known to be sensitive to the size of the computation domain. Numerical experiments indicate that the use of a small domain could result in normal mode type acoustic oscillations in the entire computation domain leading to an increase in tone frequency and intensity. When the computation domain is expanded so that the boundaries are at least one wavelength away from the noise source, the computed tone frequency and intensity are found to be computation domain size independent.

Kurbatskii, Konstantin↗

NASA NTRS1988

Toward lower drag with laminar flow technology

The most promising method of achieving significantly lower skin-friction drag of an aircraft component is through the stabilization and maintenance of a laminar boundary layer over a large fraction of its surface. In a number of applications, however, laminar flow is unfeasible and/or undesirable, and the most efficient turbulent flow design must be provided. The development of both laminar and turbulent airfoil and wing designs over the past decade is reviewed. High, medium, and low-speed airfoil concepts are discussed, and some overall comparisons are made for different Reynolds number ranges.

Harvey, W. D.↗