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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 595 records · Page 33
All sourcesNASA NTRSDOE OSTI
NASA NTRS1990

In-flight flow field analysis on the NASA F-18 high alpha research vehicle with comparisons to ground facility data

In-flight flow visualization results of the vortical flow on the forebody and leading-edge extensions (LEX) of an F-18 research aircraft have been presented for angles of attack from 15.8 to 42.5 deg and for sideslip angles up to 7.5 deg. Water tunnel results using a 3-percent scale F-18 model and a variety of wind tunnel results are used for comparison and interpretation of the flight results. The LEX vortex core breakdown point moved forward with increasing angle of attack. For a constant angle of attack, the windward LEX vortex core breakdown moves forward and inboard with sideslip and the leeward vortex core breakdown moves aft and outboard. For a constant angle of attack, the windward location of interaction moved aft with increasing sideslip and the leeward interaction moved forward.

Del Frate, John H.↗
NASA NTRS1991

Seals research at the University of Akron

The flow visualization aims at determining the flow regimes and the sealing mechanisms that are at work in the brush seal, regardless of whether the seal is linear or circular. A vision system coupled with an image processing system is used to analyze the nature of the flow. Pressure measurements were made to study the effect of the number of brushes as well as the brush structure on the seal effectiveness. Details of a numerical approach to flow simulation in one hydrostatic pocket is presented. The effects of the jet, shaft velocity, and clearance on the flow patterns and pressure generation are described.

Braun, M. Jack↗
NASA NTRS1992

Advanced nozzle and engine components test facility

A test facility for conducting scaled advanced nozzle and engine component research is described. The CE-22 test facility, located in the Engine Research Building of the NASA Lewis Research Center, contains many systems for the economical testing of advanced scale-model nozzles and engine components. The combustion air and altitude exhaust systems are described. Combustion air can be supplied to a model up to 40 psig for primary air flow, and 40, 125, and 450 psig for secondary air flow. Altitude exhaust can be simulated up to 48,000 ft, or the exhaust can be atmospheric. Descriptions of the multiaxis thrust stand, a color schlieren flow visualization system used for qualitative flow analysis, a labyrinth flow measurement system, a data acquisition system, and auxiliary systems are discussed. Model recommended design information and temperature and pressure instrumentation recommendations are included.

Beltran, Luis R.↗
NASA NTRS1992

LDV measurements on a rectangular wing with a simulated glaze ice accretion

LDV measurement results are presented for the upper surface of a rectangular semispan wing with and without simulated glaze ice accretion. Inspection of the model centerline flow field indicates that a large region of reverse flow exists aft of the ice horn on the iced model. At alpha = 0 deg, this region extends to 7 percent chord, while at alpha = 4 deg the bubble grows to more than 12 percent chord. At alpha = 8 deg, the time-averaged separation bubble is measured well beyond 50 percent chord. Experimental and computational flow visualization support these findings. The flow in the vicinity of the ice shape contains many of the features of flow over a backward-facing step.

Khodadoust, A.↗
NASA NTRS1992

Advanced nozzle and engine components test facility

A test facility for conducting scaled advanced nozzle and engine component research is described. The CE-22 test facility, located in the Engine Research Building of the NASA Lewis Research Center, contains many systems for the economical testing of advanced scale-model nozzles and engine components. The combustion air and altitude exhaust systems are described. Combustion air can be supplied to a model up to 40 psig for primary air flow, and 40, 125, and 450 psig for secondary air flow. Altitude exhaust can be simulated up to 48,000 ft, or the exhaust can be atmospheric. Descriptions of the multiaxis thrust stand, a color schlieren flow visualization system used for qualitative flow analysis, a labyrinth flow measurement system, a data acquisition system, and auxiliary systems are discussed. Model recommeded design information and temperature and pressure instrumentation recommendations are included.

Beltran, Luis R.↗
NASA NTRS1992

An experimental investigation of the flow in a diffusing S-duct

Compressible, subsonic flow through a diffusing S-duct has been experimentally investigated. Benchmark aerodynamic data are presented for flow through a representative S-duct configuration. The collected data would be beneficial to aircraft inlet designers and is suitable for the validation of computational codes. Measurements of the 3D velocity field and total and static pressures were obtained at five cross-sectional planes. Surface static pressures and flow visualization also helped to reveal flow field characteristics. All reported tests were conducted with an inlet centerline Mach number of 0.6 and a Reynolds number, based on the inlet centerline velocity and duct inlet diameter, of 2.6 x 10(exp 6). The results show that a larger region of streamwise flow separation occurred within the duct. Details about the separated flow region, including mechanisms which drive this complicated flow phenomenon, are discussed. Transverse velocity components indicate that the duct curvature induces strong pressure driven secondary flows, which evolve into a large pair of counter-rotating vortices. These vortices convect the low momentum fluid of the boundary layer towards the center of the duct, degrading both the uniformity and magnitude of the total pressure profile.

Wellborn, Steven R.↗
NASA NTRS1993

Effects of flow-path variations on internal reversing flow in a tailpipe offtake configuration for ASTOVL aircraft

A one-third-scale model of a generic tailpipe offtake system for an advanced short takeoff, vertical landing (ASTOVL) aircraft was tested at the NASA Lewis Research Center Powered Lift Facility. The basic model consisted of a tailpipe with a center body to form an annulus simulating turbine outflow with no swirl; twin offtake ducts with elbows at the ends to turn the flow to a downward direction; flow control nozzles at the ends of the elbows; and a blind flange at the end of the tailpipe to simulate a closed cruise nozzle. The offtake duct-to-tailpipe diameter ratio was 0.74. Modifications of a generic nature were then made to this basic configuration to measure the effects of flow-path changes on the flow and pressure-loss characteristics. The modifications included adding rounded entrances at the forward edges of the offtake openings, blocking the tailpipe just aft the openings instead of at the cruise nozzle, changing the location of the openings along the tailpipe, removing the center body, and varying the Mach number (flow rate) over a wide range in the tailpipe ahead of the openings by changing the size of the flow control nozzles. The tests were made with unheated air at tailpipe-to-ambient pressure ratios from 1.4 to 5. Results are presented and compared with performance graphs, total-pressure contour plots, paint streak flow visualization photographs, and a flow-angle probe traverse at the offtake entrance.

Mcardle, Jack G.↗
NASA NTRS1993

Unsteady, three-dimensional, Navier-Stokes simulations of multistage turbomachinery flows

Hub corner stall regions, endwall boundary layers, tip leakage flows and airfoil wakes combine to form the 3D flowfields found in axial turbomachines. Turbomachinery flow fields are also inherently unsteady because of the relative motion between rotor and stator airfoils. This relative motion causes complex time-varying aerodynamic interactions to occur between the different aerodynamic structures and the rotor and stator airfoils. It is necessary to understand the 3D unsteady aerodynamics associated with these interactions in order to design turbomachines that are both lightweight and compact as well as reliable and efficient. The current study uses a time-accurate 3D thin-layer Navier-Stokes zonal approach to investigate the unsteady aerodynamics of multistage turbines and compressors. Relative motion between rotor and stator airfoils is accounted for by the use of systems of patched and overlaid grids. Time-averaged surface pressures, surface flow visualizations, and time-averaged flow field contours have been computed for a 1 1/2-stage turbine and are in good agreement with experimental data. This favorable comparison represents an initial validation of the current method for unsteady computations of multistage turbomachinery flows.

Gundy-Burlet, Karen L.↗
NASA NTRS1993

Effects of flow-path variations on internal reversing flow in a tailpipe offtake configuration for ASTOVL aircraft

A one-third-scale model of a generic tailpipe offtake system for an advanced short takeoff, vertical landing (ASTOVL) aircraft was tested at the NASA Lewis Research Center Powered Lift Facility. The basic model consisted of a tailpipe with a center body to form an annulus simulating turbine outflow with no swirl; twin offtake ducts with elbows at the ends to turn the flow to a downward direction; flow control nozzles at the ends of the elbows; and a blind flange at the end of the tailpipe to simulate a closed cruise nozzle. The offtake duct-to-tailpipe diameter ratio was 0.74. Modifications of a generic nature were then made to this basic configuration to measure the effects of flow-path changes on the flow and pressure-loss characteristics. The modifications included adding rounded entrances at the forward edges of the offtake openings, blocking the tailpipe just aft the openings instead of at the cruise nozzle, changing the location of the openings along the tailpipe, removing the center body, and varying the Mach number (flow rate) over a wide range in the tailpipe ahead of the openings by changing the size of the flow control nozzles. The tests were made with unheated air at tailpipe-to-ambient pressure ratios from 1.4 to 5. Results are presented and compared with performance graphs, total-pressure contour plots, paint streak flow visualization photographs, and a flow-angle probe traverse at the offtake entrance.

Mcardle, Jack G.↗
NASA NTRS2000

Effects of Passive Porosity on Interacting Vortex Flows At Supersonic Speeds

A wind tunnel experiment was conducted in the NASA Langley Research Center (LaRC) Unitary Plan Wind Tunnel (UPWT) to determine the effects of passive surface porosity on vortex flow interaction about a general research fighter configuration at supersonic speeds. Optical flow measurement and flow visualization techniques were used and included pressure-sensitive paint (PSP), schlieren, and laser vapor screen (LVS) These techniques were combined with force and moment and conventional electronically-scanned pressure (ESP) measurements to quantify and to visualize the effects of flow-through porosity applied to a wing leading-edge extension (LEX) mounted to a 65 deg cropped delta wing model.

Erickson, Gary E.↗
NASA NTRS2007

Wind Tunnel Investigation of the Effects of Surface Porosity and Vertical Tail Placement on Slender Wing Vortex Flow Aerodynamics at Supersonic Speeds

A wind tunnel experiment was conducted in the NASA Langley Research Center (LaRC) Unitary Plan Wind Tunnel (UPWT) to determine the effects of passive surface porosity and vertical tail placement on vortex flow development and interactions about a general research fighter configuration at supersonic speeds. Optical flow measurement and flow visualization techniques were used that featured pressure sensitive paint (PSP), laser vapor screen (LVS), and schlieren, These techniques were combined with conventional electronically-scanned pressure (ESP) and six-component force and moment measurements to quantify and to visualize the effects of flow-through porosity applied to a wing leading edge extension (LEX) and the placement of centerline and twin vertical tails on the vortex-dominated flow field of a 65 cropped delta wing model. Test results were obtained at free-stream Mach numbers of 1.6, 1.8, and 2.1 and a Reynolds number per foot of 2.0 million. LEX porosity promoted a wing vortex-dominated flow field as a result of a diffusion and weakening of the LEX vortex. The redistribution of the vortex-induced suction pressures contributed to large nose-down pitching moment increments but did not significantly affect the vortex-induced lift. The trends associated with LEX porosity were unaffected by vertical tail placement. The centerline tail configuration generally provided more stable rolling moments and yawing moments compared to the twin wing-mounted vertical tails. The strength of a complex system of shock waves between the twin tails was reduced by LEX porosity.

Erickson, Gary E.↗
NASA NTRS2011

Large-Scale Low-Boom Inlet Test Overview

This presentation provides a high level overview of the Large-Scale Low-Boom Inlet Test and was presented at the Fundamental Aeronautics 2011 Technical Conference. In October 2010 a low-boom supersonic inlet concept with flow control was tested in the 8'x6' supersonic wind tunnel at NASA Glenn Research Center (GRC). The primary objectives of the test were to evaluate the inlet stability and operability of a large-scale low-boom supersonic inlet concept by acquiring performance and flowfield validation data, as well as evaluate simple, passive, bleedless inlet boundary layer control options. During this effort two models were tested: a dual stream inlet intended to model potential flight hardware and a single stream design to study a zero-degree external cowl angle and to permit surface flow visualization of the vortex generator flow control on the internal centerbody surface. The tests were conducted by a team of researchers from NASA GRC, Gulfstream Aerospace Corporation, University of Illinois at Urbana-Champaign, and the University of Virginia

Hirt, Stefanie↗
NASA NTRS2016

Sweeping Jet Optimization Studies

Progress on experimental efforts to optimize sweeping jet actuators for active flow control (AFC) applications with large adverse pressure gradients is reported. Three sweeping jet actuator configurations, with the same orifice size but di↵erent internal geometries, were installed on the flap shoulder of an unswept, NACA 0015 semi-span wing to investigate how the output produced by a sweeping jet interacts with the separated flow and the mechanisms by which the flow separation is controlled. For this experiment, the flow separation was generated by deflecting the wing's 30% chord trailing edge flap to produce an adverse pressure gradient. Steady and unsteady pressure data, Particle Image Velocimetry data, and force and moment data were acquired to assess the performance of the three actuator configurations. The actuator with the largest jet deflection angle, at the pressure ratios investigated, was the most efficient at controlling flow separation on the flap of the model. Oil flow visualization studies revealed that the flow field controlled by the sweeping jets was more three-dimensional than expected. The results presented also show that the actuator spacing was appropriate for the pressure ratios examined.

Melton, LaTunia Pack↗
NASA NTRS2015

Visualization of a Sweeping Jet by Laser Speckle Retro-reflective Background Oriented Schlieren

Sweeping jet actuators are currently being studied as a method of active flow control to increase the performance of wings and tails. Computational Fluid Dynamics simulations show that the jet does not sweep at a constant rate, instead spending more time at the minimum and maximum of its extent. As part of a test done to provide dynamic flow measurements for Computation Fluid Dynamics validation in NASA Ames Fluid Mechanics Lab, Particle Image Velocimetry (PIV) was performed on a single sweeping jet actuator designed by NASA and Boeing that was inset in a small airfoil. However, even at low jet pressures, the PIV could not be acquired at high enough frequencies to both spatially and temporally resolve the stickiness of the jet motion. Therefore, a small portion of the PIV laser was picked off and used to acquire flow visualization at higher frequencies through a new technique, laser speckle Retroreflective Background Oriented Schlieren (RBOS).

sweeping jet actuator↗
NASA NTRS1985

Velocity visualization in gaseous flows

Techniques are established for visualizing velocity in gaseous flows. Two approaches are considered, both of which are capable of yielding velocity simultaneously at a large number of flowfield locations, thereby providing images of velocity. The first technique employs a laser to mark specific fluid elements and a camera to track their subsequent motion. Marking is done by laser-induced phosphorescence of biacetyl, added as a tracer species in a flow of N2, or by laser-induced formation of sulfur particulates in SF6-H2-N2 mixtures. The second technique is based on the Doppler effect, and uses an intensified photodiode array camera and a planar form of laser-induced fluorescence to detect 2-d velocities of I2 (in I2-N2 mixtures) via Doppler-shifted absorption of narrow-linewidth laser radiation at 514.5 nm.

Hanson, R. K.↗
NASA NTRS1984

Vortex-generating coolant-flow-passage design for increased film-cooling effectiveness and surface coverage

The thermal film-cooling footprints observed by infrared imagery for three coolant-passage configurations embedded in adiabatic-test plates are discussed. The configurations included a standard round-hole cross section and two orientations of a vortex-generating flow passage. Both orientations showed up to factors of four increases in both film-cooling effectiveness and surface coverage over that obtained with the round coolant passage. The crossflow data covered a range of tunnel velocities from 15.5 to 45 m/sec with blowing rates from 0.20 to 2.05. A photographic streakline flow visualization technique supported the concept of the counterrotating apability of the flow passage design and gave visual credence to its role in inhibiting flow separation.

Papell, S. S.↗