Engineering Papers⌕ Search

SEARCH · Engineering Papers

Results for “BOUNDARY LAYER SEPARATION”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 343 records · Page 19
All sourcesNASA NTRSDOE OSTI
NASA NTRS1985

Three-dimensional inviscid analysis of radial turbine flow and a limited comparison with experimental data

The three-dimensional inviscid DENTON code is used to analyze flow through a radial-inflow turbine rotor. Experimental data from the rotor are compared with analytical results obtained by using the code. The experimental data available for comparison are the radial distributions of circumferentially averaged values of absolute flow angle and total pressure downstream of the rotor exit. The computed rotor-exit flow angles are generally underturned relative to the experimental values, which reflect the boundary-layer separation at the trailing edge and the development of wakes downstream of the rotor. The experimental rotor is designed for a higher-than-optimum work factor of 1.126 resulting in a nonoptimum positive incidence and causing a region of rapid flow adjustment and large velocity gradients. For this experimental rotor, the computed radial distribution of rotor-exit to turbine-inlet total pressure ratios are underpredicted due to the errors in the finite-difference approximations in the regions of rapid flow adjustment, and due to using the relatively coarser grids in the middle of the blade region where the flow passage is highly three-dimensional. Additional results obtained from the three-dimensional inviscid computation are also presented, but without comparison due to the lack of experimental data. These include quasi-secondary velocity vectors on cross-channel surfaces, velocity components on the meridional and blade-to-blade surfaces, and blade surface loading diagrams. Computed results show the evolution of a passage vortex and large streamline deviations from the computational streamwise grid lines. Experience gained from applying the code to a radial turbine geometry is also discussed.

Choo, Y. K.↗
NASA NTRS1986

Numerical simulation of vortical flow over an elliptical-body missile at high angles of attack

Numerical solutions to the Reynolds-averaged Navier-Stokes equations are given for the flow about an elliptical body missile (3:1 ellipse) at a Mach number of 2.5 and a unit Reynolds number of 6.56 x 10 to the 6th/m. At high angles of attack, the flow is dominated by large-scale free vortices which occur in the lee-side flow field due to crossflow boundary-layer separation. Emphasis is focused on the accurate prediction of the lee-side vortical flow. Solutions are presented for both symmetric and asymmetric (body rolled 45 deg) configurations at 10 deg and 20 deg angle of attack. The computed results are compared with experimental surface pressure coefficients and vapor-screen photographs. Excellent agreement is obtained in all cases.

Newsome, R. N.↗
NASA NTRS1985

Three-dimensional inviscid analysis of radial-turbine flow and a limited comparison with experimental data

The three-dimensional inviscid DENTON code is used to analyze flow through a radial-inflow turbine rotor. Experimental data from the rotor are compared with analytical results obtained by using the code. The experimental data available for comparison are the radial distributions of circumferentially averaged values of absolute flow angle and total pressure downstream of the rotor exit. The computed rotor-exit flow angles are generally underturned relative to the experimental values, which reflect the boundary-layer separation at the trailing edge and the development of wakes downstream of the rotor. The experimental rotor is designed for a higher-than-optimum work factor of 1.126 resulting in a nonoptimum positive incidence and causing a region of rapid flow adjustment and large velocity gradients. For this experimental rotor, the computed radial distribution of rotor-exit to turbine-inlet total pressure ratios are underpredicted due to the errors in the finite-difference approximations in the regions of rapid flow adjustment, and due to using the relatively coarser grids in the middle of the blade region where the flow passage is highly three-dimensional. Additional results obtained from the three-dimensional inviscid computation are also presented, but without comparison due to the lack of experimental data. These include quasi-secondary velocity vectors on cross-channel surfaces, velocity components on the meridional and blade-to-blade surfaces, and blade surface loading diagrams. Computed results show the evolution of a passage vortex and large streamline deviations from the computational streamwise grid lines. Experience gained from applying the code to a radial turbine geometry is also discussed.

Choo, Y. K.↗
NASA NTRS1987

Experimental evaluation of honeycomb/screen configurations and short contraction section for NASA Lewis Research Center's altitude wind tunnel

An experimental investigation was conducted in the high speed leg of the 0.1 scale model of the proposed Altitude Wind Tunnel to evaluate flow conditioner configurations in the settling chamber and their effect on the flow through the short contraction section. The lowest longitudinal turbulence intensity measured at the contraction-section entrance, 1.2%, was achieved with a honeycomb plus three fine-mesh screens. Turbulence intensity in the test section was estimated to be between 0.1 and 0.2% with the honeycomb plus three fine mesh screens in the settling chamber. Adding screens, however, adversely affected the total pressure profile, causing a small defect near the centerline at the contraction section entrance. No significant boundary layer separation was evident in the short contraction section.

Burley, Richard R.↗
NASA NTRS1986

A summary of Reynolds number effects on some recent tests in the Langley 0.3-meter transonic cryogenic tunnel

Reynolds number effects noted from selected test programs conducted in the Langley 0.3-Meter Transonic Cryogenic Tunnel (0.3-m TCT) are discussed. The tests, which cover a unit Reynolds number range from about 2.0 to 80.0 million per foot, summarize effects of Reynolds number on: (1) aerodynamic data from a supercritical airfoil, (2) results from several wall interference correction techniques, and (3) results obtained from advanced, cryogenic tests techniques. The test techniques include: (1) use of a cryogenic sidewall boundary layer removal system, (2) detailed pressure and hot wire measurements to determine test section flow quality, and (3) use of a new hot film system suitable for transition detection in a cryogenic wind tunnel. The results indicate that Reynolds number effects appear most significant when boundary layer transition effects are present and at high lift conditions when boundary layer separation exists on both the model and the tunnel sidewall.

Johnson, C. B.↗
NASA NTRS1987

Numerical simulations of unsteady airfoil-vortex interactions

Numerical methods based on the thin layer Navier-Stokes, Euler and transonic small disturbance equations, all three of which use the perturbation form of an implicit numerical algorithm, are used to formulate the unsteady interactions of a moving vortex with a stationary airfoil. Results are presented for both transonic and subsonic flows, as well as for weak and strong interactions. In general, the results exhibit an overwhelming influence of the vortex on the flow field around the airfoil, especially when the convecting vortex is within one chord of the airfoil. It is noted that even the ostensibly weak interactions produced vortex/shock-induced boundary layer separation.

Srinivasan, G. R.↗
NASA NTRS1987

Optimal heading change with minimum energy loss for a hypersonic gliding vehicle

A three state model is presented for analyzing the problem of optimal changes in heading with minimum energy loss for a hypersonic gliding vehicle. A further model order reduction to a single state model is examined using singular perturbation theory. The optimal solution for the reduced problem defines an optimal altitude profile dependent on the current energy of the vehicle, and the corresponding optimal lift and bank angle. A separate boundary layer analysis, based on an expansion of the necessary conditions about the reduced solution, is used to account for altitude and flight path angle dynamics and to derive a guidance law in feedback form. The guidance law is evaluated for a hypothetical vehicle.

Calise, Anthony J.↗
NASA NTRS1987

On the effect of leading edge blowing on circulation control airfoil aerodynamics

In the present context the term circulation control is used to denote a method of lift generation that utilizes tangential jet blowing over the upper surface of a rounded trailing edge airfoil to determine the location of the boundary layer separation points, thus setting an effective Kutta condition. At present little information exists on the flow structure generated by circulation control airfoils under leading edge blowing. Consequently, no theoretical methods exist to predict airfoil performance under such conditions. An experimental study of the flow field generated by a two dimensional circulation control airfoil under steady leading and trailing edge blowing was undertaken. The objective was to fundamentally understand the overall flow structure generated and its relation to airfoil performance. Flow visualization was performed to define the overall flow field structure. Measurements of the airfoil forces were also made to provide a correlation of the observed flow field structure to airfoil performance. Preliminary results are presented, specifically on the effect on the flow field structure of leading edge blowing, alone and in conjunction with trailing edge blowing.

Mclachlan, B. G.↗
NASA NTRS1988

A low-drag nose-body

In the present investigation of the application of convex-curvature viscous drag reduction to an axisymmetric body, the qualification of a modified mixing-length formulation in several complex two-dimensional and axisymmetric turbulent boundary layers that are primarily curved is followed by the design of a low-drag nose-body for low drag at high subsonic speeds. The longitudinal distribution of the cross-sectional area ratio is found to be critical to boundary layer separation, requiring implementation not over a single long span of curvature but over three short spans. The possibility of a 'net' drag reduction is examined in relation to 'equivalent' nose-bodies.

Bandyopadhyay, P. R.↗
NASA NTRS1988

Singular perturbation analysis of the atmospheric orbital plane change problem

A three-state model is presented for the aeroassisted orbital plane change problem. A further model order reduction to a single state model is examined using singular perturbation theory. The optimal solution for this single state model compares favorably with the exact numerical solution using a four-state model; however, a separate boundary layer solution is required to satisfy the terminal constraint on altitude. This, in general, involves the solution of a two-point boundary value problem, but for a two-state model. An approximation is introduced to obtain an analytical control solution for lift and bank angle. Included are numerical simulation results of a guidance law derived from this analysis, along with comparison to earlier work by other researchers.

Calise, A. J.↗
NASA NTRS1988

Parametric study of grid size, time step and turbulence modeling on Navier-Stokes computations over airfoils

An upwind-biased implicit approximate factorization algorithm is applied to several steady and unsteady turbulent flows. The thin layer form of the compressible Navier-Stokes equation is used. Both the flux vector splitting and flux difference splitting methods are used to determine fluxes, and the results are compared. Flux difference splitting predicts results more accurately than flux vector splitting on a given mesh size, but, in its present implementation, is more severely limited by the maximum CFL number for unsteady time accurate flows. Physical aspects of the computations are also examined. An equilibrium turbulent boundary layer model computes generally better steady and unsteady results than a nonequilibrium model when there is little to no boundary layer separation. Conversely, when a significant region of separation exists, the nonequilibrium model performs in better agreement with experiment.

Rumsey, Christopher L.↗
NASA NTRS1989

Documentation of Two- and Three-Dimensional Hypersonic Shock Wave/Turbulent Boundary Layer Interaction Flows

Experimental data for a series of two- and three-dimensional shock wave/turbulent boundary layer interaction flows at Mach 7 are presented. Test bodies, composed of simple geometric shapes, were designed to generate flows with varying degrees of pressure gradient, boundary-layer separation, and turning angle. The data include surface-pressure and heat-transfer distributions as well as limited mean-flow-field surveys in both the undisturbed and the interaction regimes. The data are presented in a convenient form for use in validating existing or future computational models of these generic hypersonic flows.

Kussoy, Marvin I.↗
NASA NTRS1989

Transonic airfoil design for helicopter rotor applications

Despite the fact that the flow over a rotor blade is strongly influenced by locally three-dimensional and unsteady effects, practical experience has always demonstrated that substantial improvements in the aerodynamic performance can be gained by improving the steady two-dimensional charateristics of the airfoil(s) employed. The two phenomena known to have great impact on the overall rotor performance are: (1) retreating blade stall with the associated large pressure drag, and (2) compressibility effects on the advancing blade leading to shock formation and the associated wave drag and boundary-layer separation losses. It was concluded that: optimization routines are a powerful tool for finding solutions to multiple design point problems; the optimization process must be guided by the judicious choice of geometric and aerodynamic constraints; optimization routines should be appropriately coupled to viscous, not inviscid, transonic flow solvers; hybrid design procedures in conjunction with optimization routines represent the most efficient approach for rotor airfroil design; unsteady effects resulting in the delay of lift and moment stall should be modeled using simple empirical relations; and inflight optimization of aerodynamic loads (e.g., use of variable rate blowing, flaps, etc.) can satisfy any number of requirements at design and off-design conditions.

Hassan, Ahmed A.↗
NASA NTRS1988

Passive and active control of jet turbulence

Techniques for controlling the generation, growth, and interactions of large coherent structures in turbulent flows are discussed, reviewing the results of recent experimental investigations. The fundamental principles of active and passive coherent-structure manipulation are outlined; the effects of geometric modification are examined for the case of an elliptic jet; and particular attention is given to passive control via self-excitation, turbulence and noise suppression, control of pairing interaction via two-frequency excitation, and unsteady boundary-layer separation in wall jets. Diagrams and graphs of typical results are provided.

Fazle Hussain, A. K. M.↗
NASA NTRS1983

Aerodynamically-driven condensate layer thickness distributions on isothermal cylindrical surfaces

A simple yet rather general mathematical model is presented for predicting the distribution of condensate layer thickness when aerodynamic shear is the dominant mechanism of liquid flow along the surface. The Newtonian condensate film is treated using well-known thin-layer (lubrication theory) approximations, and condensate supply is taken to be the result of either convective diffusion or inertial impaction. Illustrative calculations for a circular cylinder in a crossflow at Re = 100,000 reveal the consequences of alternate condensate arrival mechanisms and the existence of thicker reverse-flow films behind the position of gas boundary-layer separation. The present formulation is readily generalized to include transient liquid layer flows on noncircular objects of variable surface temperature, as encountered in turbine-blade materials testing or operation.

Rosner, D. E.↗
NASA NTRS1988

A new multipoint thin-film diagnostic technique for fluid dynamic studies

A multielement hot-film sensor has been developed with closely-spaced elements (2.5 mm) in the streamwise direction. In addition to boundary-layer transition detection, the output of these sensors can reveal the locations of boundary-layer separation and reattachment by a 180 deg phase difference in the 0-200-Hz frequency bandwidth between sensor elements on each side of either the separation or reattachment point. This paper describes the features of multielement hot-film sensor arrays and discusses results from their application to a number of low-speed and transonic airfoil tests.

Sewall, W. G.↗
NASA NTRS1989

A natural low-frequency oscillation of the flow over an airfoil near stalling conditions

An experimental and computational study of the low-frequency oscillation observed in the flow over an airfoil at the onset of static stall is presented. Wind-tunnel results obtained with two-dimensional airfoil models show that this phenomena takes place only with a transitional state of the separating boundary layer. It is noted that the flowfield does not involve a Karman vortex street. The experimental results agree well with the results of a two-dimensional Navier-Stokes code. The present study demonstrates that the low-frequency oscillations produce intense flow fluctuations which impart much larger unsteady forces to the airfoil than experienced by bluff-body shedding and which may represent the primary aerodynamics of stall flutter of blades and wings.

Zaman, K. B. M. Q.↗
NASA NTRS1989

Numerical analysis of flow about a total temperature sensor

The unsteady flowfield about an airfoil-shaped inlet temperature sensor has been investigated using the thin-layer and full Navier-Stokes equations. A finite-volume formulation of the governing equations was used in conjunction with a Runge-Kutta time stepping scheme to analyze the flow about the sensor. Flow characteristics for this configuration were established at Mach numbers of 0.5 and 0.8 for different Reynolds numbers. The results were obtained for configurations of increasing complexity; important physical phenomena such as shock formation, boundary-layer separation, and unsteady wake formation were noted. Based on the computational results, recommendations for further study and refinement of the inlet temperature sensor were made.

Von Lavante, Ernst↗