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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 217 records · Page 12
All sourcesNASA NTRSDOE OSTI
NASA NTRS1974

Mean and fluctuating flow measurements in axisymmetric supersonic boundary layer flow subjected to distributed adverse pressure gradients

Measurements have been made of the mean flow properties and turbulent fluctuations in adiabatic turbulent boundary layer flows subjected to distributed adverse pressure gradients. In the freestream region upstream of the adverse pressure gradient the Mach number was 3.86, the unit Reynolds number 5.3 million per foot. The boundary layer developed on the wall of an axisymmetric nozzle and straight test section. In order to avoid the effects of streamwise surface curvature the adverse pressure gradients at the test section wall were induced by contoured centerbodies mounted on the wind tunnel centerline. The flow under study simulated that which might be found in an axially symmetric engine inlet of a supersonic aircraft.

Gootzait, E.↗
NASA NTRS1987

An LDA investigation of three-dimensional normal shock-boundary layer interactions in a corner

Nonintrusive, three-dimensional, measurements have been made of a normal shock wave-turbulent boundary layer interaction. The measurements were made in the corner of the test section of a continuous supersonic wind tunnel in which a normal shock wave had been stabilized. LDA, surface pressure measurement and flow visualization techniques were employed for two freestream Mach number test cases: 1.6 and 1.3. The former contained separated flow regions and a system of shock waves. The latter was found to be far less complicated. The reported results are believed to accurately define the flow physics of each case and may be used as benchmark data to verify three-dimensional computer codes.

Chriss, R. M.↗
NASA NTRS1987

Design and test of a natural laminar flow/large Reynolds number airfoil with a high design cruise lift coefficient

Research activity on an airfoil designed for a large airplane capable of very long endurance times at a low Mach number of 0.22 is examined. Airplane mission objectives and design optimization resulted in requirements for a very high design lift coefficient and a large amount of laminar flow at high Reynolds number to increase the lift/drag ratio and reduce the loiter lift coefficient. Natural laminar flow was selected instead of distributed mechanical suction for the measurement technique. A design lift coefficient of 1.5 was identified as the highest which could be achieved with a large extent of laminar flow. A single element airfoil was designed using an inverse boundary layer solution and inverse airfoil design computer codes to create an airfoil section that would achieve performance goals. The design process and results, including airfoil shape, pressure distributions, and aerodynamic characteristics are presented. A two dimensional wind tunnel model was constructed and tested in a NASA Low Turbulence Pressure Tunnel which enabled testing at full scale design Reynolds number. A comparison is made between theoretical and measured results to establish accuracy and quality of the airfoil design technique.

Kolesar, C. E.↗
NASA NTRS1976

Full-coverage film cooling heat transfer study: Summary of data for normal-hole injection and 30 deg slant-hole injection

Heat transfer to a full coverage film cooled turbulent boundary layer over a flat surface was studied. The surface consisted of a discrete hole test section containing 11 rows of holes spaced 5 diameters apart in a staggered array and an instrumented recovery region. Ten diameter spacing was also studied by plugging appropriate holes. Two test sections were used, one having holes normal to the surface and the other having holes angled 30 deg to the surface in the downstream direction. Stanton number data were obtained both in the full coverage region and in the downstream recovery region for a range of blowing ratios, or mass flux ratios, from 0 to 1.3. Initial conditions at the upstream edge of the blowing region were varied from 500 to 5000 for momentum thickness Reynolds number and from 100 to 1800 for enthalpy thickness Reynolds number. The range of Reynolds numbers based on hole diameter and mainstream velocity was 6000 to 22000. Initial boundary layer thicknesses range from 0.5 to 2.0 hole diameters. Air was used as the working fluid. The data were taken for the secondary injection temperature equal to the wall temperature and also equal to the mainstream temperature. Superposition was then used to obtain Stanton number as a continuous function of the injectant temperature. The heat transfer coefficient was defined on the basis of a mainstream-to-wall temperature difference. This definition permits direct comparison of performance between film cooling and transpiration cooling.

Crawford, M. E.↗
NASA NTRS2015

Numerical Study of the High-Speed Leg of a Wind Tunnel

The paper describes a numerical study of the high-speed leg of the NASA Langley 14 x 22-ft Low-Speed Wind Tunnel. The high-speed leg consists of the settling chamber, contraction, test section, and first diffuser. Results are shown comparing two different sources of surface geometry, and two different unstructured grid solvers for the flow characteristics. Numerical simulations of the flow on the tunnel centerline, boundary layer profiles on the floor, and wall static pressures have been compared with experiment. Flow angularities along the test section length have also been determined.

Nayani, Sudheer↗
NASA NTRS2001

Initial Circulation and Peak Vorticity Behavior of Vortices Shed from Airfoil Vortex Generators

An extensive parametric study of vortices shed from airfoil vortex generators has been conducted to determine the dependence of initial vortex circulation and peak vorticity on elements of the airfoil geometry and impinging flow conditions. These elements include the airfoil angle of attack, chord length, span, aspect ratio, local boundary layer thickness, and free stream Mach number. In addition, the influence of airfoil-to-airfoil spacing on the circulation and peak vorticity has been examined for pairs of co-rotating and counter-rotating vortices. The vortex generators were symmetric airfoils having a NACA-0012 cross-sectional profile. These airfoils were mounted either in isolation, or in pairs, on the surface of a straight pipe. The turbulent boundary layer thickness to pipe radius ratio was about 17 percent. The circulation and peak vorticity data were derived from cross-plane velocity measurements acquired with a seven-hole probe at one chord-length downstream of the airfoil trailing edge location. The circulation is observed to be proportional to the free-stream Mach number, the angle-of-attack, and the span-to-boundary layer thickness ratio. With these parameters held constant, the circulation is observed to fall off in monotonic fashion with increasing airfoil aspect ratio. The peak vorticity is also observed to be proportional to the free-stream Mach number, the airfoil angle-of-attack, and the span-to-boundary layer thickness ratio. Unlike circulation, however, the peak vorticity is observed to increase with increasing aspect ratio, reaching a peak value at an aspect ratio of about 2.0 before falling off again at higher values of aspect ratio. Co-rotating vortices shed from closely spaced pairs of airfoils have values of circulation and peak vorticity under those values found for vortices shed from isolated airfoils of the same geometry. Conversely, counter-rotating vortices show enhanced values of circulation and peak vorticity when compared to values obtained in isolation. The circulation may be accurately modeled with an expression based on Prandtl's relationship between finite airfoil circulation and airfoil geometry. A correlation for the peak vorticity has been derived from a conservation relationship equating the moment at the airfoil tip to the rate of angular momentum production of the shed vortex, modeled as a Lamb (ideal viscous) vortex. This technique provides excellent qualitative agreement to the observed behavior of peak vorticity for low aspect ratio airfoils typically used as vortex generators.

Wendt, Bruce J.↗
NASA NTRS2009

Boundary Layer Transition Protuberance Tests at NASA JSC Arc-Jet Facility

A series of arc-jet tests in support of the Shuttle Orbiter Boundary Layer Transition flight experiment was conducted in the Channel Nozzle of the NASA Johnson Space Center Atmospheric Reentry Materials and Structures Facility. The boundary layer trip was a protrusion of a certain height and geometry fabricated as part of a 6"x6" tile insert, a special test article made of the Boeing Rigid Insulation tile material and coated with the Reaction Cured Glass used for the bottom fuselage tiles of the Space Shuttle Orbiter. A total of five such tile inserts were manufactured: four with the 0.25-in. trip height, and one with the 0.35-in. trip height. The tile inserts were interchangeably installed in the center of the 24"x24" variable configuration tile array mounted in the 24"x24" test section of the channel nozzle. The objectives of the test series were to demonstrate that the boundary layer trip can safely withstand the Space Shuttle Orbiter flight-like re-entry environments and provide temperature data on the protrusion surface, surfaces of the nearby tiles upstream and downstream of the trip, as well as the bond line between the tiles and the structure. The targeted test environments were defined for the tip of the protrusion, away from the nominal surface of the tile array. The arc jet test conditions were approximated in order to produce the levels of the free stream total enthalpy at the protrusion height similar to those expected in flight. The test articles were instrumented with surface, sidewall and bond line thermocouples. Additionally, Tempilaq temperature-indicating paint was applied to the nominal tiles of the tile array in locations not interfering with the protrusion trip. Five different grades of paint were used that disintegrate at different temperatures between 1500 and 2000 deg F. The intent of using the paint was to gauge the RCG-coated tile surface temperature, as well as determine its usefulness for a flight experiment. This paper provides an overview of the channel nozzle arc jet, test articles and test conditions, as well as the results of the arc-jet tests including the measured temperature response of the test articles, their pre- and post-test surface scans, condition of the thermal paint, and continents on the protrusion tip heating achieved in tests compared to the computational fluid dynamics predictions.

Larin, M. E.↗
NASA NTRS1974

A boundary-layer analysis of atmospheric motion over a semi-elliptical surface obstruction

Flow over surface obstructions can produce adverse flying conditions for helicopters, V/STOL vehicles, etc. The disturbed boundary-layer concept is applied in approximating the localized flow field induced around a surface obstruction (modeled by a two-dimensional cylinder with elliptical cross section) by an impinging wind. The analysis concludes that: (1) localized wind-speed maxima occur at the top of a surface obstruction, which are expected in physically real flows; (2) increased elliptical aspect ratio decreases with speed within the boundary layer at the top of the ellipse; (3) increased surface roughness decreases velocity in the boundary layer; (4) Reynolds number has a negligible effect on the overall flow for the Re range considered; (5) decreased elliptical aspect ratio and increased surface roughness cause larger separation regions.

Frost, W.↗
NASA NTRS2004

Design and Evaluation of Modifications to the NASA Langley Flow Impedance Tube

The need to minimize fan noise radiation from commercial aircraft engine nacelles continues to provide an impetus for developing new acoustic liner concepts. If the full value of such concepts is to be attained, an understanding of grazing flow effects is crucial. Because of this need for improved understanding of grazing flow effects, the NASA Langley Research Center Liner Physics Group has invested a large effort over the past decade into the development of a 2-D finite element method that characterizes wave propagation through a lined duct. The original test section in the Langley Grazing IncidenceTube was used to acquire data needed for implementation of this finite element method. This test section employed a stepper motor-driven axial-traversing bar, embedded in the wall opposite the test liner, to position a flush-mounted microphone at pre-selected locations. Complex acoustic pressure data acquired with this traversing microphone were used to educe the acoustic impedance of test liners using this 2-D finite element method and a local optimization technique. Results acquired in this facility have been extensively reported, and were compared with corresponding results from various U.S. aeroacoustics laboratories in the late 1990 s. Impedance data comparisons acquired from this multi-laboratory study suggested that it would be valuable to incorporate more realistic 3-D aeroacoustic effects into the impedance eduction methodology. This paper provides a description of modifications that have been implemented to facilitate studies of 3-D effects. The two key features of the modified test section are (1) the replacement of the traversing bar and its flush-mounted microphone with an array of 95 fixed-location microphones that are flush-mounted in all four walls of the duct, and (2) the inclusion of a suction device to modify the boundary layer upstream of the lined portion of the duct. The initial results achieved with the modified test section are provided in this report, and a comparison of these results with those achieved using the original test section is used to demonstrate that the data acquisition and analysis with the new test section can be confidently used for impedance eduction.

Jones, Michael G.↗
NASA NTRS1988

Wind tunnel-sidewall-boundary-layer effects in transonic airfoil testing-some correctable, but some not

The need to correct transonic airfoil wind tunnel test data for the influence of the tunnel sidewall boundary layers, in addition to the wall accepted corrections for the analytical investigation was carried out in order to evaluate sidewall boundary layer effects on transonic airfoil characteristics, and to validate proposed correction and the limit to their applications. This investigation involved testing of modern airfoil configurations in two different transonic airfoil test facilities, the 15 x 60 inch two-dimensional insert of the National Aeronautical Establishment (NAE) 5 foot tunnel in Ottawa, Canada, and the two-dimensional test section of the NASA Langley 0.3 m Transonic Cryogenic Tunnel (TCT). Results presented included effects of variations in sidewall-boundary layer bleed in both facilities, different sidewall boundary layer correction procedures, tunnel-to tunnel comparisons of correcte results, and flow conditions with and without separation.

Lynch, F. T.↗
NASA NTRS1985

Flow control in a diffusing S-Duct

Accurate measurements have been made of secondary flow in a 1.51 area ratio diffusing 30 deg - 30 deg S-Duct with circulair cross section. Turbulent flow was entering the duct at Mach number of 0.6, the boundary layer thickness at the duct entrance was ten percent of the duct inlet diameter. Through measurements made, local flow velocity vector as well as static and total pressures mapping of the flow at several stations were obtained. Strong secondary flow was measured in the first bend which continued into the second bend with new vorticity produced in there in the opposite direction. Surface oil flow visualization and wall pressures indicated a region of separated flow starting at theta approximately equal to 22 deg on the inside of the first bend up to theta approximately equal to 44 deg on the outside of the second bend. The flow separated in 'cyclone' form and never reattached in the duct. As a result of the secondary flow and the flow separation, significant total pressure distortion was observed at the exit of the duct. Using flow control devices the separation was eliminated while the exit distortion was improved.

Vakili, A. D.↗
NASA NTRS1989

An experimental investigation of a low Reynolds number turbulent boundary layer subject to an adverse pressure gradient

A very low Reynolds number turbulent boundary layer subject to an adverse pressure gradient is studied. The aim is to obtain highly accurate mean-flow and turbulence measurements under conditions that can be closely related to the numerical simulations of Philippe Spalart for the purposes of CFD validation. Much of the Boundary Layer Wind Tunnel was completely rebuilt with a new wider contraction and working section which will improve compatibility with the simulations. A unique sophisticated high-speed computer controlled 3-D probe traversing mechanism was integrated into the test section. Construction of the tunnel and traverse is discussed in some detail. The hardware is now complete, and measurements are in progress. The mean-flow data indicate that a suitably two-dimensional base flow was established. Automation of the probe positioning and data acquistion have led to a decreased running time for total pressure measurements. However, the most significant benefits are expected to occur when using hot-wire probes. Calibrations can be performed automatically and there is no need to handle fragile probes when moving between measuring stations. Techniques are being developed which require sampling of the signals from moving hot-wire probes on the basis of their position in the flow. Measurements can be made in high intensity turbulence by flying probes upstream at high speed so that the relative magnitude of the turbulent velocity fluctuations are reduced. In regions, where the turbulence intensity is not too large, the probe can also be repetitively scanned across very dense spatial grids in other directions. With this technique, a complete profile can be measured in about 1/3 the time and with a spatial density about 50 times that obtainable using a stationary probe.

Watmuff, Jonathan H.↗
NASA NTRS1991

A laser velocimeter investigation of the normal shock-wave boundary layer interaction

Nonintrusive 3-D measurements were made of a normal shock wave/turbulent boundary layer interaction. The measurements were made through a quadrant of a square test section of a continuous flow supersonic wind tunnel in which a normal shock wave had been stabilized. Two dimensional measurements were made throughout the interaction region while 3-D measurements were made near the corner in the vicinity of the shock where 3-D flow effects were expected to be significant. Laser Doppler velocimetry, surface static pressure measurements, and flow visualization techniques were used for two freestream nominal Mach number test cases: 1.6 and 1.3. No turbulence information was obtained. The mean velocity measurements were converted to Mach number by recording the wind tunnel total temperature. Some shock oscillation was present during both of the test cases. After startup of the wind tunnel, the shock oscillated with an amplitude of approx. + or - 1 cm, however, after reaching steady condition, the shock oscillation amplitude was greatly reduced, as evidenced by the laser velocimeter results. The Mach 1.3 test case resulted in a nearly uniform flow without secondary shock waves and with no or very isolated corner separation.

Chriss, Randall M.↗
NASA NTRS1992

Unsteady pressure loads in a generic high speed engine model

Unsteady pressure loads were measured along the top interior wall of a generic high-speed engine (GHSE) model undergoing performance tests in the combustion-Heated Scramjet Test Facility at the Langley Research Center. Flow to the model inlet was simulated at 72000 ft and a flight Mach number of 4. The inlet Mach number was 3.5 with a total temperature and pressure of 1640 R and 92 psia. The unsteady pressure loads were measured with 5 piezoresistive gages, recessed into the wall 4 to 12 gage diameters to reduce incident heat flux to the diaphragms, and distributed from the inlet to the combustor. Contributors to the unsteady pressure loads included boundary layer turbulence, combustion noise, and transients generated by unstart loads. Typical turbulent boundary layer rms pressures in the inlet ranged from 133 dB in the inlet to 181 dB in the combustor over the frequency range from 0 to 5 kHz. Downstream of the inlet exist, combustion noise was shown to dominate boundary layer turbulence noise at increased heat release rates. Noise levels in the isolator section increased by 15 dB when the fuel-air ratio was increased from 0.37 to 0.57 of the stoichiometric ratio. Transient pressure disturbances associated with engine unstarts were measured in the inlet and have an upstream propagation speed of about 7 ft/sec and pressure jumps of at least 3 psia.

Parrott, Tony L.↗
NASA NTRS2001

Parametric Data from a Wind Tunnel Test on a Rocket-Based Combined-Cycle Engine Inlet

A 40-percent scale model of the inlet to a rocket-based combined-cycle (RBCC) engine was tested in the NASA Glenn Research Center 1- by 1-Foot Supersonic Wind Tunnel (SWT). The full-scale RBCC engine is scheduled for test in the Hypersonic Tunnel Facility (HTF) at NASA Glenn's Plum Brook Station at Mach 5 and 6. This engine will incorporate the configuration of this inlet model which achieved the best performance during the present experiment. The inlet test was conducted at Mach numbers of 4.0, 5.0, 5.5, and 6.0. The fixed-geometry inlet consists of an 8 deg.. forebody compression plate, boundary layer diverter, and two compressive struts located within 2 parallel sidewalls. These struts extend through the inlet, dividing the flowpath into three channels. Test parameters investigated included strut geometry, boundary layer ingestion, and Reynolds number (Re). Inlet axial pressure distributions and cross-sectional Pitot-pressure surveys at the base of the struts were measured at varying back-pressures. Inlet performance and starting data are presented. The inlet chosen for the RBCC engine self-started at all Mach numbers from 4 to 6. Pitot-pressure contours showed large flow nonuniformity on the body-side of the inlet. The inlet provided adequate pressure recovery and flow quality for the RBCC cycle even with the flow separation.

Fernandez, Rene↗
NASA NTRS2007

Preliminary Investigation of Curved Liner Sample in the NASA LaRC Curved Duct Test Rig

This viewgraph presentation reviews the preliminary investigation of the curved liner sample in the NASA LaRC Curved Duct Test Rig (CDTR). It reviews the purpose of the Curved Duct Test Rig. Its purpose is to develop capability to investigate acoustic and aerodynamic properties in ducts. It has several features to accomplish that purpose: (1) Large scale (2) Flow rate to M = 0.275 (3) Higher order mode control (4) Curved flow path (5) Adaptable test section (6) Flexible test configurations. The liner has minimal effect on turbulence or boundary layer growth in duct. The curved duct sample attenuation is affected by mode scattering. In conclusion, the CDTR is valid tool for aerodynamic and acoustic evaluation of duct treatment

Gerhold, Carl H.↗
NASA NTRS2012

Experimental Studies of Low-Pressure Turbine Flows and Flow Control

This report summarizes research performed in support of the NASA Glenn Research Center (GRC) Low-Pressure Turbine (LPT) Flow Physics Program. The work was performed experimentally at the U.S. Naval Academy faculties. The geometry corresponded to "Pak B" LPT airfoil. The test section simulated LPT flow in a passage. Three experimental studies were performed: (a) Boundary layer measurements for ten baseline cases under high and low freestream turbulence conditions at five Reynolds numbers of 25,000, 50,000, 100,000, 200,000, and 300,000, based on passage exit velocity and suction surface wetted length; (b) Passive flow control studies with three thicknesses of two-dimensional bars, and two heights of three-dimensional circular cylinders with different spanwise separations, at same flow conditions as the 10 baseline cases; (c) Active flow control with oscillating synthetic (zero net mass flow) vortex generator jets, for one case with low freestream turbulence and a low Reynolds number of 25,000. The Passive flow control was successful at controlling the separation problem at low Reynolds numbers, with varying degrees of success from case to case and varying levels of impact at higher Reynolds numbers. The active flow control successfully eliminated the large separation problem for the low Reynolds number case. Very detailed data was acquired using hot-wire anemometry, including single and two velocity components, integral boundary layer quantities, turbulence statistics and spectra, turbulent shear stresses and their spectra, and intermittency, documenting transition, separation and reattachment. Models were constructed to correlate the results. The report includes a summary of the work performed and reprints of the publications describing the various studies.

Volino, Ralph J.↗
NASA NTRS2023

Computational Study of Vortex-Induced Separation for a 5-Bladed Rotor

This work computationally investigated the rotor blade vortex-induced separation recently observed during an aerodynamic rotor test campaign in the NASA Langley Research Center 14- by 22-Foot Subsonic Tunnel. Two separate approaches (i.e., airfoil modification and blade tip modification) were studied to mitigate the vortex-induced separation. Low-fidelity tools based on blade element momentum theory were shown to mispredict the rotor inflow and were also shown to not capture the vortex-induced separation caused by perpendicular blade-vortex interaction. This misprediction was exploited to isolate the aerodynamic thrust deficit caused by the vortex-induced separation (i.e., 20%) from the thrust deficit due to inflow variation (i.e., 31%). High-fidelity tools were shown to reasonably predict aerodynamic forces within 13% and flow separation when compared to experimental results. The modified airfoil variant of the baseline rotor effectively mitigated the vortex-induced separation while the blade tip modified variant still showed separation, though the size and strength of the vortex was reduced. Acoustic predictions were underpredicted by 10 dB from preliminary measurements taken in the untreated wind tunnel. Broadband noise contributions from different rotor blade sections showed that self-noise due to flow separation and other turbulent boundary layer mechanisms was the dominant noise source for all three rotor cases, followed by blade-wake interaction noise caused by perpendicular blade-vortex interactions.

aerodynamics↗