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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.

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At least 541 records · Page 30
All sourcesNASA NTRSDOE OSTI
NASA NTRS1976

Upper wing surface boundary layer measurements and static aerodynamic data obtained on a 0.015-scale model (42-0) or the SSV orbiter configuration 140A/B in the LTV HSWT at a Mach number of 4.6 (LA58)

The purpose of the test was to investigate the nature of the Orbiter boundary layer characteristics at angles of attack from -4 to 32 degrees at a Mach number of 4.6. The effect of large grit, employed as transition strips, on both the nature of the boundary layer and the force and moment characteristics were investigated along with the effects of large negative elevon deflection on lee side separation. In addition, laminar and turbulent boundary layer separation phenomena which could cause asymmetric flow separation were investigated.

Ball, J. W.↗
DOE OSTI2020

Lift Equivalence and Cancellation for Airfoil Surge–Pitch–Plunge Oscillations

A NACA 0018 airfoil in freestream velocity is oscillated in longitudinal, transverse, and angle-of-attack directions with respect to the freestream velocity, known as surge, plunge, and pitch. The lift-based equivalence method introduces phase shifts between these three motions to construct in-phase sinusoidal components for maximum lift, waveform construction. Lift cancellation is also determined with the exact negative pitch and plunge motion amplitudes found from the equivalence method to achieve out-of-phase wave destruction. Lift cancellation occurs when a combination of these motions is sought to obtain a constant lift magnitude throughout the oscillation cycle. To achieve both equivalence and cancellation of lift, a prescribed pure pitch amplitude through the Theodorsen theory equates the corresponding equivalent plunge amplitude and pitch–plunge phase shift. These Theodorsen, linear superposition findings of pitch–plunge are leveraged toward the Greenberg theory to determine a closed-form, surge–pitch–plunge solution through the addition of a surge–plunge phase shift and optimal surge amplitude for lift cancellation. The lift cancellation surge–pitch–plunge amplitudes define the equivalence amplitude investigated here and theoretically limit the experiment to combinations of the first lift harmonic of the Greenberg theory. The analytical results are then compared with experimental lift force measurements and dye visualization. Finally, the normalized lift differences due to unsteady wake and boundary-layer behavior are examined to explore the extents of the Greenberg theory for these cases of lift-based equivalence and cancellation.

42 ENGINEERING↗
NASA NTRS1974

Analysis of dynamic stall using unsteady boundary-layer theory

The unsteady turbulent boundary layer and potential flow about a pitching airfoil are analyzed using numerical methods to determine the effect of pitch rate on the delay in forward movement of the rear flow reversal point. An explicit finite difference scheme is used to integrate the unsteady boundary layer equations, which are coupled at each instant of time to a fully unsteady and nonlinear potential flow analysis. A substantial delay in forward movement of the reversal point is demonstrated with increasing pitch rate, and it is shown that the delay results partly from the alleviation of the gradients in the potential flow, and partly from the effects of unsteadiness in the boundary layer itself. The predicted delay in flow-reversal onset, and its variation with pitch rate, are shown to be in reasonable agreement with experimental data relating to the delay in dynamic stall. From the comparisons it can be concluded (a) that the effects of time-dependence are sufficient to explain the failure of the boundary layer to separate during the dynamic overshoot, and (b) that there may be some link between forward movement of the reversal point and dynamic stall.

Scruggs, R. M.↗
NASA NTRS1976

Laser velocimetry applied to transonic and supersonic aerodynamics

Measurements obtained with laser velocimetry in a Mach 2.9 separated turbulent boundary layer and in the transonic flow past a two-dimensional airfoil section are presented and compared to data realized by conventional techniques. Agreement in mean velocities was realized where the pressure measurements could be considered reliable; however, in regions of instantaneous reverse velocities, the laser results were found to be consistent with the physics of the flow whereas the pressure data were not. Streamwise turbulence intensities are also presented. In the transonic airfoil study, velocity measurements obtained immediately outside the upper surface boundary layer of a 6-inch chord NACA 64A010 airfoil are compared to edge velocities inferred from surface pressure measurements. For free-stream Mach numbers of 0.6 and 0.8, the agreement in results was very good. "Dual scatter" optical arrangements in conjunction with a single particle, counter-type signal processor were employed in these investigations.

Johnson, D. A.↗
NASA NTRS1978

Flying-hot-wire study of two-dimensional mean flow past an NACA 4412 airfoil at maximum lift

Hot-wire measurements have been made in the boundary layer, the separated region, and the near wake for flow past an NACA 4412 airfoil at maximum lift. The Reynolds number based on chord was about 1,500,000. The main instrumentation was a hot-wire probe mounted on the end of a rotating arm. A digital computer was used to control synchronized sampling of hot-wire data at closely spaced points along the probe arc. Ensembles of data were obtained at several thousand locations in the flow field. The data include intermittency, two components of mean velocity, and twelve mean values for double, triple, and quadruple products of two velocity fluctuations. The data are available on punched cards in raw form and also after use of smoothing and interpolation routines to obtain values on a fine rectangular grid aligned with the airfoil chord. The data are displayed in the paper as contour plots.

Coles, D.↗
NASA NTRS1979

Effects of turbulence model selection on the prediction of complex aerodynamic flows

Numerical simulations of viscous transonic flow over a circular-arc airfoil and in a diffuser are described. The simulations are made with a new computer program designed to serve as a tool in the development of improved turbulence models for complex flows. The program incorporates zero-, one-, and two-equation eddy viscosity models and includes a variety of subsonic and supersonic boundary conditions. The airfoil flow contains a shock-separated boundary-layer interaction that has resisted previous attempts at simulation. The diffuser flow also contains a shock-boundary-layer interaction, which has not been simulated previously. Calculations using standard turbulence models, developed originally for incompressible unseparated flows, are described. Results indicate that although there are interesting differences in predictions between the various models, none of them predict the flows accurately. Suggestions for improved turbulence models are discussed.

Coakley, T. J.↗
NASA NTRS1981

Measurements of the boundary layer and near wake of a supercritical airfoil at cruise conditions

The flow behavior within the upper-surface boundary layer and near wake of a supercritical airfoil operating at cruise conditions is discussed. Experimental results obtained from wind-tunnel tests are presented which provide a more detailed description of the flow in these regions than was previously available. Mean streamwise velocity profiles measured by pitot-pressure-probe and laser-velocimeter techniques were found to be in excellent agreement. Other mean-flow properties obtained by the laser-velocimeter technique were the local flow angles in the viscous layers and the static pressures at the edges of the boundary layer and wake. The data set also includes measurements of the turbulence intensity and turbulent Reynolds-stress distributions as obtained by the laser-velocimeter technique. To assess the effects of the shock wave, a less extensive set of measurements was realized at a subcritical test condition. The two test conditions (Mach number at free-stream conditions = 0.72, airfoil section lift coefficient = 0.76 and Mach number of free-stream conditions = 0.5, airfoil section lift coefficient = 0.75) provide a good test for state-of-the-art prediction methods because the upper-surface-boundary layer is separated just upstream of the trailing edge in both cases.

Johnson, D. A.↗
NASA NTRS1982

A new surface-streamline flow-visualization technique

This paper describes a new surface-streamline flow-visualization technique that is suitable for use in low-speed wind tunnels or other low-speed gas flows. The technique provides a permanent record of ink traces that show surface-streamline direction and shape. The visualization of the endwall surface flow of the horseshoe-vortex system formed around the base of a cylinder by a separating turbulent boundary layer is used as an example. The results of the new technique are compared with those obtained from a conventional flow-visualization technique. Good agreement was found between the two, with the new technique appearing to give more-accurate surface streamlines.

Langston, L. S.↗
NASA NTRS1982

Optical studies of shock generated transient supersonic base flows

A shock tube employing interferometric and schlieren techniques is used to study transient base flow phenomena following shock wave passage over two plane bluff bodies: a hemicircular cylinder and a cylinder with the Galileo Jovian probe profile. An attempt is made to understand the physics of transition from transient to steady state flow, and to provide code verification for a study employing the Illiac IV computer. Transient base flow interactions include a series of shock diffraction, regular, and Mach reflections, coupled with boundary layer development, separation, and recompression. Vorticity generation and transport underlie these features. The quantitative verification of the computer code includes comparisons of transient pressure and density fields, near wake geometries, and bow shock standoff distances.

Liang, P.-Y.↗
NASA NTRS1983

SC3 - A wing concept for supersonic maneuvering

A supersonic wing design concept that overcomes the problems associated with nonlinear inviscid pressure distribution is presented together with a summary of the supporting theoretical and experimental development program. The key design feature is a conical panel supporting a controlled supercritical conical camber (SC3) crossflow expansion and recompression which permits lift on the upper surface to be obtained without producing an adverse pressure gradient or crossflow shock wave strong enough to separate the boundary layer. The role of aerodynamics in the design implementation is discussed, as are the concept development, the initial design validation on a fabricated and tested SC3 design, and the body and canard interaction effect. The design and testing of an isolated wing with a planform and thickness distribution representative of a practical application are described. Wind-tunnel results show that significant performance gains can be obtained for fighter aircraft using this concept.

Mason, W. H.↗
NASA NTRS1984

Acoustic excitation: A promising new means of controlling shear layers

Techniques have long been sought for the controlled modification of turbulent shear layers, such as in jets, wakes, boundary layers, and separated flows. Relatively recently published results of laboratory experiments have established that coherent structures exist within turbulent flows. These results indicate that even apparently chaotic flow fields can contain deterministic, nonrandom elements. Even more recently published results show that deliberate acoustic excitation of these coherent structures has a significant effect on the mixing characteristics of shear layers. Therefore, we have initiated a research effort to develop both an understanding of the interaction mechanisms and the ability to use it to favorably modify various shear layers. Acoustic excitation circumvents the need for pumping significant flow rates, as required by suction or blowing. Control of flows by intentional excitation of natural flow instabilities involves new and largely unexplored phenomena and offers considerable potential for improving component performance. Nonintrusive techniques for flow field control may permit much more efficient, flexible propulsion systems and aircraft designs, including means of stall avoidance and recovery. The techniques developed may also find application in many other areas where mixing is important, such as reactors, continuous lasers, rocket engines, and fluidic devices. It is the objective of this paper to examine some potential applications of the acoustic excitation technique to various shear layer flows of practical aerospace systems.

Stone, J. R.↗
NASA NTRS1984

Loss reduction in axial-flow compressors through low-speed model testing

A systematic procedure for reducing losses in axial-flow compressors is presented. In this procedure, a large, low-speed, aerodynamic model of a high-speed core compressor is designed and fabricated based on aerodynamic similarity principles. This model is then tested at low speed where high-loss regions associated with three-dimensional endwall boundary layers flow separation, leakage, and secondary flows can be located, detailed measurements made, and loss mechanisms determined with much greater accuracy and much lower cost and risk than is possible in small, high-speed compressors. Design modifications are made by using custom-tailored airfoils and vector diagrams, airfoil endbends, and modified wall geometries in the high-loss regions. The design improvements resulting in reduced loss or increased stall margin are then scaled to high speed. This paper describes the procedure and presents experimental results to show that in some cases endwall loss has been reduced by as much as 10 percent, flow separation has been reduced or eliminated, and stall margin has been substantially improved by using these techniques.

Wisler, D. C.↗
NASA NTRS1984

Some structural features of unsteady separating turbulent shear flows

Some physical features of unsteady separating turbulent boundary layers are presented for practical Reynolds numbers and reduced frequencies for helicopter and turbomachinery flows. Upstream of detachment in moderate amplitude flows, the flow is quasi steady, i.e., the phase averaged flow is described by the steady-free stream flow structure. Results show that oscillation waveform and amplitude strongly influence the detached flow behavior.

Simpson, R. L.↗