SEARCH · Search NASA

Results for “vortex generators”

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.

SourceResearch connectionFrom yearThrough year

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

At least 325 records · Page 18

All sources NASA NTRS DOE OSTI

NASA NTRS1982

Supercritical maneuvering fighter configuration. Wind-tunnel investigation at Mach numbers of 0.60 to 0.95

A wind tunnel investigation was conducted to study the application of supercritical technology to highly maneuverable combat aircraft. The configuration studied has a leading-edge sweep of 45 deg and an aspect ratio of 3.28. Two supercritical-wing shapes were tested at Mach numbers from 0.60 to 0.95 with angles of attack from -2 deg to 17 deg. On supercritical wing was designed to achieve a high level of transonic maneuver performance at a Mach number of 0.90; however, excessive flow separation developed on this wing at a Mach number of 0.85. A second supercritical wing was tested which had significantly reduced flow separation and improved drag characteristics at a Mach number of 0.85 and maintained the performance of the original wing at the higher Mach numbers. Leading-edge vortex generators did not improve the performance of the second wing; however, a sharp leading-edge flap produced sizable drag reductions at Mach numbers from 0.60 to 0.90.

Mann, M. J.↗

NASA NTRS1974

Dynamic stability characteristics in pitch, yaw, and roll of a supercritical-wing research airplane model

The aerodynamic damping in pitch, yaw, and roll and the oscillatory stability in pitch and yaw of a supercritical-wing research airplane model were determined for Mach numbers of 0.25 to 1.20 by using the small-amplitude forced-oscillation technique. The angle-of-attack range was from -2 deg to 20 deg. The effects of the underwing leading-edge vortex generators and the contributions of the wing, vertical tail, and horizontal tail to the appropriate damping and stability were measured.

Boyden, R. P.↗

NASA NTRS1983

Dynamic Distortion in a Short S-Shaped Subsonic Diffuser with Flow Separation

An experimental investigation of the time varying distortion at the diffuser exit of a subscale HiMAT forebody and inlet was conducted at Mach 0.9 in the Lewis 8 by 6 foot Supersonic Wind Tunnel. A transitory separation was detected within the subsonic diffuser. Vortex generators were installed to eliminate the flow separation. Results from a study of the instantaneous pressure variations at the diffuser exit are presented. The time unsteady total pressures at the diffuser exit are computer interpolated and presented in the form of a movie showing the transitory separation. Limited data showing the instantaneous distortion levels is also presented.

Stumpf, R.↗

NASA NTRS1983

Waterspout, Gust Fronts and Associated Cloud Systems

Nine waterspouts observed on five experimental days during the GATE period of observations are discussed. Primary data used are from 2 aircraft flying in different patterns, one above the other between 30 and 300 m. There is strong evidence associating whirl initiation with cumulus outflow. Computations prepared from estimates of convergence with the region suggest the possibility of vortex generation within 4 minutes. This analysis supports (1) the importance cumulus outflows may have in waterspout initiation and (2) the possibility that sea surface temperature gradients may be important in enabling waterspout development from modest size cumuli.

Simpson, J.↗

NASA NTRS1983

Dynamic distortion in a short s-shaped subsonic diffuser with flow separation

An experimental investigation of the time varying distortion at the diffuser exit of a subsonic HiMAT forebody and inlet was conducted at Mach 0.9 in the Lewis 8 by 6 foot Supersonic Wind Tunnel. A transitory separation was detected within the subsonic diffuser. Vortex generators were installed to eliminate the flow separation. Results from a study of the instantaneous pressure variations at the diffuser exit are presented. The time unsteady total pressures at the diffuser exit are computer interpolated and presented in the form of a movie showing the transitory separation. Limited data showing the instantaneous distortion levels are also presented. Previously announced in STAR as N83-26838

Stumpf, R.↗

NASA NTRS1984

Space Shuttle AFRSI OMS pod environment test using model 81-0 test fixture in the Ames Research Center 9x7-foot supersonic wind tunnel (OS-314A/B/C)

A test was conducted in the NASA/Ames Research Center 9x7-foot Supersonic Wind Tunnel to help resolve an anomaly that developed during the STS-6 orbiter flight wherein sections of the Advanced Flexible Reusable Surface Insulation (AFRSI) covering the OMS pods suffered some damage. A one-third scale two-dimensional shell structure model of an OMS pod cross-section was employed to support the test articles. These consisted of 15 AFRSI blanket panels form-fitted over the shell structures for exposure to simulated flight conditions. Of six baseline blankets, two were treated with special surface coatings. Two other panels were configured with AFRSI sections removed from the OV099 orbiter vehicle after the STS-6 flight. Seven additional specimens incorporated alternative designs and repairs. Following a series of surface pressure calibration runs, the specimens were exposed to simulated ascent and entry dynamic pressure profiles. Entry conditions included the use of a vortex generator to evaluate the effect of shed vortices on the AFRSI located in the area of concern.

Collette, J. G. R.↗

NASA NTRS1984

Two-dimensional wake characteristics of inlet vanes for open-circuit wind tunnels

This paper summarizes the near-field flow characteristics measured downstream of a half-scale two-dimensional wind-tunnel model of the inlet vanes designed for the National Full-Scale Facilities Complex at NASA Ames Research Center. Variations on this baseline were tested to determine how the downstream flow field is affected. Evaluations of the various configurations were made by the use of hot-wire surveys across the flow-field. These traverses yielded the properties of total pressure, boundary layers, and turbulence in the wake of the vanes. It was found that large variations in the flow field characteristics downstream of the vanes are achieved by the use of various arrangements of splitter vanes, vortex generators, screens, tail extensions, and honeycomb. Separation on the boat-tails of the vanes can be controlled and turbulence reduced by suitable combinations of screens or honeycomb or both. The penalties associated with each modification in terms of increased pressure loss are also presented.

Dudley, M. R.↗

NASA NTRS1985

Reflection plane tests of a wind turbine blade tip section with ailerons

Tests were conducted in the NASA Langley 30 by 60 foot Wind Tunnel on a full scale 7.31 m (24 ft) long tip section of a wind turbine rotor blade. The blade tip section was built with ailerons on the trailing edge. The ailerons, which spanned a length of 6.1 m (20 ft), were designed so that two types could be evaluated: the plain and the balanced. The ailerons were hinged on the suction surface at the 0.62 X chord station behind the leading edge. The purpose of the tests was to measure the aerodynamic characteristics of the blade section for: an angle of attack range from 0 deg to 90 deg aileron deflections from 0 deg to -90 deg, and Reynolds numbers of 0.79 and 1.5 x 10 to the 6th power. These data were then used to determine which aileron configuration had the most desirable rotor control and aerodynamic braking characteristics. Tests were also run to determine the effects of vortex generators, leading edge roughness, and the gaps between the aileron sections on the lift, drag, and chordwise force coefficients of the blade tip section.

Savino, J. M.↗

NASA NTRS1985

Some recent advances in computational aerodynamics for helicopter applications

The growing application of computational aerodynamics to nonlinear helicopter problems is outlined, with particular emphasis on several recent quasi-two-dimensional examples that used the thin-layer Navier-Stokes equations and an eddy-viscosity model to approximate turbulence. Rotor blade section characteristics can now be calculated accurately over a wide range of transonic flow conditions. However, a finite-difference simulation of the complete flow field about a helicopter in forward flight is not currently feasible, despite the impressive progress that is being made in both two and three dimensions. The principal limitations are today's computer speeds and memories, algorithm and solution methods, grid generation, vortex modeling, structural and aerodynamic coupling, and a shortage of engineers who are skilled in both computational fluid dynamics and helicopter aerodynamics and dynamics.

Mccroskey, W. J.↗

NASA NTRS1985

Ground effects on V/STOL and STOL aircraft: A survey

The flow fields encountered by jet- and fan-powered Vertical/Short Takeoff and Landing (V/STOL) aircraft operating in ground effect are reviewed and their general effects on the aerodynamic characteristics are discussed. The ground effects considered include: (1) the suckdown experienced by a single jet configuration in hover; (2) the fountain flow and additional suckdown experienced by multiple jet configurations in hover; (3) the ground vortex generated by jet and jet flap configurations in short takeoff and landing (STOL) operation and the associated aerodynamic and hot-gas-ingestion effects; and (4) the change in the downwash at the tail due to ground proximity. After over 30 years of research on V/STOL aircraft, the general flow phenomena are well-known and, in most areas, the effects of ground proximity can be established or can be determined experimentally. However, there are some anomalies in the current data base which are discussed.

Kuhn, R. E.↗

NASA NTRS1986

Statistical prediction of dynamic distortion of inlet flow using minimum dynamic measurement. An application to the Melick statistical method and inlet flow dynamic distortion prediction without RMS measurements

The Melick method of inlet flow dynamic distortion prediction by statistical means is outlined. A hypothetic vortex model is used as the basis for the mathematical formulations. The main variables are identified by matching the theoretical total pressure rms ratio with the measured total pressure rms ratio. Data comparisons, using the HiMAT inlet test data set, indicate satisfactory prediction of the dynamic peak distortion for cases with boundary layer control device vortex generators. A method for the dynamic probe selection was developed. Validity of the probe selection criteria is demonstrated by comparing the reduced-probe predictions with the 40-probe predictions. It is indicated that the the number of dynamic probes can be reduced to as few as two and still retain good accuracy.

Schweikhard, W. G.↗

NASA NTRS1985

Ground effects on V/STOL and STOL aircraft - A survey

The flow fields encountered by jet- and fan-powered Vertical/Short Takeoff and Landing (V/STOL) aircraft operating in ground effect are reviewed and their general effects on the aerodynamic characteristics are discussed. The ground effects considered include (1) the suckdown experienced by a single jet configuration in hover, (2) the fountain flow and additional suckdown experienced by multiple jet configurations in hover, (3) the ground vortex generated by jet and jet flap configurations in Short Takeoff and Landing (STOL) operation and the associated aerodynamic and hot-gas-ingestion effects, and (4) the change in the downwash at the tail due to ground proximity. After over 30 years of research on V/STOL aircraft, the general flow phenomena are well known and, in most areas, the effects of ground proximity can be estimated or can be determined experimentally. However, there are some anomalies in the current data base which are discussed.

Eshlemen, J.↗

NASA NTRS1986

Viscous analyses for flow through subsonic and supersonic intakes

A parabolized Navier-Stokes code was used to analyze a number of diffusers typical of a modern inlet design. The effect of curvature of the diffuser centerline and transitioning cross sections was evaluated to determine the primary cause of the flow distortion in the duct. Results are presented for S-shaped intakes with circular and transitioning cross sections. Special emphasis is placed on verification of the analysis to accurately predict distorted flow fields resulting from pressure-driven secondary flows. The effect of vortex generators on reducing the distortion of intakes is presented. Comparisons of the experimental and analytical total pressure contours at the exit of the intake exhibit good agreement. In the case of supersonic inlets, computations of the inlet flow field reveal that large secondary flow regions may be generated just inside of the intake. These strong flows may lead to separated flow regions and cause pronounced distortions upstream of the compressor.

Povinelli, Louis A.↗

NASA NTRS1987

Viscous analyses for flow through subsonic and supersonic intakes

Povinelli, Louis A.↗

NASA NTRS1987

Internal computational fluid mechanics on supercomputers for aerospace propulsion systems

The accurate calculation of three-dimensional internal flowfields for application towards aerospace propulsion systems requires computational resources available only on supercomputers. A survey is presented of three-dimensional calculations of hypersonic, transonic, and subsonic internal flowfields conducted at the Lewis Research Center. A steady state Parabolized Navier-Stokes (PNS) solution of flow in a Mach 5.0, mixed compression inlet, a Navier-Stokes solution of flow in the vicinity of a terminal shock, and a PNS solution of flow in a diffusing S-bend with vortex generators are presented and discussed. All of these calculations were performed on either the NAS Cray-2 or the Lewis Research Center Cray XMP.

Andersen, Bernhard H.↗

NASA NTRS1987

Flow energizer flight tests on a light twin aircraft

Experiments were conducted to explore the use of flow energizers (i.e., horizontally mounted vortex generators), using a special instrumentation subsystem installed on a light twin aircraft. The data, collected for energizer configurations with convergence ratios of 1.2, 1.5, and 1.7, included measurements of pressure on the wing surface, velocity components in the wake of the energizer, and forces on the flow energizer itself. Surface pressure data showed that flow energizer effects are highly localized. The energizer with the smallest convergence ratio tested produced an energizer lift/drag ratio about 75 percent lower that that of the other two configurations. For highly swept planforms, cambered energizers with overlaps of the order of 12-15 percent of the local chord provide the best results.

Ward, D. T.↗

NASA NTRS1986

Low Aspect Ratio Wings at High Angles of Attack

A comprehensive evaluation is made of experimental data compiled to date for the flowfields and aerodynamic forces that occur at high angles of attack for low aspect ratio wings with delta, rectangular, clipped delta, and strake/wing planform geometries. Attention is given to wing leading edge-generated vortex breakdown, aspect ratio and compressibility effects, and strake vortex effects on main wing areas. Although the nonlinear effects created by a wing-body combination significantly alter wing-alone aerodynamics, the wing-alone data presented are vital to the development of prediction methodologies for large angle of attack aerodynamics.

R L Stallings Jr.↗

NASA NTRS1987

Inlets, ducts and nozzles

The internal fluid mechanics research program in inlets, ducts, and nozzles is described. The program consists of a balanced effort between the development of computational tools and the conduct of experimental research. The experiments are designed to better understand the fluid flow physics, to develop new or improved flow models, and to provide benchmark quality data sets for validation of the computational methods. The inlet, duct, and nozzle research program is described according to three major classifications of flow phenomena: highly three-dimensional flow fields; shock-boundary layer interactions; and shear layer control. Specific examples of current and future elements of the research program are described for each of these phenomena. In particular, the highly three-dimensional flow field phenomena is highlighted by describing the computational and experiemental research program in transition ducts having a round-to-rectangular area variation. In the case of shock-boundary layer interactions, the specific details of research for normal shock-boundary layer interactions are described. For shear layer control research in vortex generators and the use of aerodynamic excitation for enhancement of the jet mixing process are described. Future research in inlets, ducts, and nozzles will include more emphasis on three-dimensional full Navier-Stokes methods and corresponding experiments designed to concentrate on the appropriate three-dimensional fluid flow physics.

Abbott, John M.↗