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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

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NASA NTRS1995

Analysis of Tangential Slot Blowing on F/A-18 Isolated Forebody

The generation of significant side forces and yawing moments on an F/A-18 fuselage through tangential slot blowing is analyzed using computational fluid dynamics. The effects of freestream Mach number, jet exit conditions, jet length, and jet location are studied. The effects of over- and underblowing on force and moment production are analyzed. Non-time-accurate solutions are obtained to determine the steady-state side forces, yawing moments, and surface pressure distributions generated by tangential slot blowing. Time-accurate solutions are obtained to study the force onset time lag of tangential slot blowing. Comparison with available experimental data from full-scale wind-tunnel and subscale wind-tunnel tests are made. This computational analysis complements the experimental results and provides a detailed understanding of the effects of tangential slot blowing on the flowfield about the isolated F/A-18 forebody. Additionally, it extends the slot-blowing database to transonic maneuvering Mach numbers.

Gee, Ken↗

NASA NTRS2000

Effects of Hole Length, Supply Plenum Geometry, and Freestream Turbulence on Film Cooling Performance

Experimental measurements are presented in this report to document the sensitivity of film cooling performance to the hole length and coolant delivery plenum geometry. Measurements with hot-wire anemometry detail velocity, local turbulence, and spectral distributions over the exit plane of film cooling holes and downstream of injection in the coolant-freestream interaction zone. Measurements of discharge coefficients and adiabatic effectiveness are also provided. Coolant is supplied to the film cooling holes by means of a large, open plenum and through plenums which force the coolant to approach the holes either co-current or counter-current to the freestream. A single row of film cooling holes with 35 degree-inclined streamwise at two coolant-to-freestream velocity ratios, 0.5 and 1.0, is investigated. The coolant-to-freestream density ratio is maintained in the range 0.96 to 1.0. Measurements were taken under high-freestream (FSTI = 12%) and low-freestream turbulence intensity (FSTI = 0.5%) conditions. The results document the effects of the hole L/D, coolant supply plenum geometry, velocity ratio, and FSTI. In general, hole L/D and the supply plenum geometry play influential roles in the film cooling performance. Hole L/D effects, however, are more pronounced. Film cooling performance is also dependent upon the velocity ratio and FSTI.

Burd, Steven W.↗

NASA NTRSDate not supplied

Low-Noise Propeller Design with the Vortex Lattice Method and Gradient-Based Optimization

This work aims to combine an aerodynamic model based on the unsteady vortex lattice method with an acoustic model provided by Farassat’s formulation 1A to perform gradient-based propeller optimizations with aerodynamic and acoustic constraints. Two optimization cases are attempted and successfully converged: an isolated propeller and a propeller-wing configuration, both operating in a 0.11 Mach freestream, representative of a cruise condition. The results show that simultaneously decreasing the RPM and increasing the collective angle of a propeller is an effective noise reduction technique, as expected, and little difference is observed between the two cases. A vortex dissipation model was found to be needed to stabilize the derivatives of the vortex lattice outputs, and a constraint on the propeller blade’s chord distribution concavity was needed to produce realistic-looking blade chord distributions.

Propellers↗

NASA NTRS2013

Numerical Study of Pressure Fluctuations due to a Mach 6 Turbulent Boundary Layer

Direct numerical simulations (DNS) are used to examine the pressure fluctuations generated by a Mach 6 turbulent boundary layer with nominal freestream Mach number of 6 and Reynolds number of Re(sub t) approx. =. 464. The emphasis is on comparing the primarily vortical pressure signal at the wall with the acoustic freestream signal under higher Mach number conditions. Moreover, the Mach-number dependence of pressure signals is demonstrated by comparing the current results with those of a supersonic boundary layer at Mach 2.5 and Re(sub t) approx. = 510. It is found that the freestream pressure intensity exhibits a strong Mach number dependence, irrespective of whether it is normalized by the mean wall shear stress or by the mean pressure, with the normalized fluctuation amplitude being significantly larger for the Mach 6 case. Spectral analysis shows that both the wall and freestream pressure fluctuations of the Mach 6 boundary layer have enhanced energy content at high frequencies, with the peak of the premultiplied frequency spectrum of freestream pressure fluctuations being at a frequency of omega(delta)/U(sub infinity) approx. = 3.1, which is more than twice the corresponding frequency in the Mach 2.5 case. The space-time correlations indicate that the pressure-carrying eddies for the higher Mach number case are of smaller size, less elongated in the spanwise direction, and convect with higher convection speeds relative to the Mach 2.5 case. The demonstrated Mach-number dependence of the pressure field, including radiation intensity, directionality, and convection speed, is consistent with the trend exhibited in experimental data and can be qualitatively explained by the notion of "eddy Mach wave" radiation.

Duan, Lian↗

NASA NTRS2001

A Wind Tunnel Study on the Mars Pathfinder (MPF) Lander Descent Pressure Sensor

The primary focus of this study was to determine the accuracy of the Mars Pathfinder lander local pressure readings in accordance with the actual ambient atmospheric pressures of Mars during parachute descent. In order to obtain good measurements, the plane of the lander pressure sensor opening should ideally be situated so that it is parallel to the freestream. However, due to two unfavorable conditions, the sensor was positioned in locations where correction factors are required. One of these disadvantages is due to the fact that the parachute attachment point rotated the lander's center of gravity forcing the location of the pressure sensor opening to be off tangent to the freestream. The second and most troublesome factor was that the lander descends with slight oscillations that could vary the amplitude of the sensor readings. In order to accurately map the correction factors required at each sensor position, an experiment simulating the lander descent was conducted in the Martian Surface Wind Tunnel at NASA Ames Research Center. Using a 115 scale model at Earth ambient pressures, the test settings provided the necessary Reynolds number conditions in which the actual lander was possibly subjected to during the descent. In the analysis and results of this experiment, the readings from the lander sensor were converted to the form of pressure coefficients. With a contour map of pressure coefficients at each lander oscillatory position, this report will provide a guideline to determine the correction factors required for the Mars Pathfinder lander descent pressure sensor readings.

Soriano, J. Francisco↗

NASA NTRSDate not supplied

FLEET and PLIF Velocimetry Within A Mach 10 Hypersonic Air Flow

Femtosecond laser electronic excitation tagging (FLEET) and planar laser-induced fluorescence (PLIF) velocity measurements utilizing molecular tagging velocity (MTV) methods from three recent test campaigns conducted at the 31-in Mach 10 Air Tunnel at the NASA Langley Research Center are highlighted within. The FLEET measurements reported here include the first direct measurement of freestream velocity at this hypersonic wind tunnel facility. Measurement challenges were exasperated by the low gas density of the Mach 10 air freestream (~0.4% of standard temperature and pressure conditions) and even lower gas densities within the hypersonic wake of a 70-degree sphere-cone model. In addition, the hypersonic freestream and very low speed velocities in the wake also tested the measurement dynamic range. To complement the FLEET measurements in the wake of the sphere-cone model, PLIF velocimetry using seeded nitric oxide was also performed. While NO-PLIF velocimetry has been performed at this facility several times by previous researchers, the use of a 1D diffractive optical element for NO-PLIF velocimetry is reported here for the first time. The 1D DOE enabled the generation of up to 75 laser lines simultaneously and improved the spatial extent of the measurement three times compared to previous work. This enabled a wide velocity measurement plane of approximately 130 mm x 130 mm. The velocimetry methods demonstrated here are expected to improve wind tunnel characterization, provide critical data to validate CFD codes, and improve the design of flight vehicles for planetary entry.

FLEET↗

NASA NTRSDate not supplied

FLEET and PLIF Velocimetry Within A Mach 10 Hypersonic Air Flow

FLEET↗

NASA NTRS1991

The NASA Lewis Research Center Internal Fluid Mechanics Facility

An experimental facility specifically designed to investigate internal fluid duct flows is described. It is built in a modular fashion so that a variety of internal flow test hardware can be installed in the facility with minimal facility reconfiguration. The facility and test hardware interfaces are discussed along with design constraints of future test hardware. The plenum flow conditioning approach is also detailed. Available instrumentation and data acquisition capabilities are discussed. The incoming flow quality was documented over the current facility operating range. The incoming flow produces well behaved turbulent boundary layers with a uniform core. For the calibration duct used, the boundary layers approached 10 percent of the duct radius. Freestream turbulence levels at the various operating conditions varied from 0.64 to 0.69 percent of the average freestream velocity.

Porro, A. R.↗

NASA NTRS2016

Measurements in a Transitioning Cone Boundary Layer at Freestream Mach 3.5

An experimental study was conducted in the Supersonic Low-Disturbance Tunnel to investigate naturally-occurring instabilities in a supersonic boundary layer on a 7 deg half- angle cone. All tests were conducted with a nominal freestream Mach number of M(sub infinity) = 3:5, total temperature of T(sub 0) = 299:8 K, and unit Reynolds numbers of Re(sub infinity) x 10(exp -6) = 9:89, 13.85, 21.77, and 25.73 m(exp -1). Instability measurements were acquired under noisy- ow and quiet- ow conditions. Measurements were made to document the freestream and the boundary-layer edge environment, to document the cone baseline flow, and to establish the stability characteristics of the transitioning flow. Pitot pressure and hot-wire boundary- layer measurements were obtained using a model-integrated traverse system. All hot- wire results were single-point measurements and were acquired with a sensor calibrated to mass ux. For the noisy-flow conditions, excellent agreement for the growth rates and mode shapes was achieved between the measured results and linear stability theory (LST). The corresponding N factor at transition from LST is N 3:9. The stability measurements for the quiet-flow conditions were limited to the aft end of the cone. The most unstable first-mode instabilities as predicted by LST were successfully measured, but this unstable first mode was not the dominant instability measured in the boundary layer. Instead, the dominant instabilities were found to be the less-amplified, low-frequency disturbances predicted by linear stability theory, and these instabilities grew according to linear theory. These low-frequency unstable disturbances were initiated by freestream acoustic disturbances through a receptivity process that is believed to occur near the branch I locations of the cone. Under quiet-flow conditions, the boundary layer remained laminar up to the last measurement station for the largest Re1, implying a transition N factor of N greater than 8:5.

King, Rudolph A.↗

NASA NTRS1976

Investigations of the 0.020-scale 88-OTS Integrated Space Shuttle Vehicle Jet-Plume Model in the NASA/Ames Research Center 11 by11-Foot Unitary Plan Wind Tunnel (IA80)

The results are documented of jet plume effects wind tunnel test of the 0.020-scale 88-OTS launch configuration space shuttle vehicle model in the 11 x 11 foot leg of the NASA/Ames Research Center Unitary Plan Wind Tunnel. This test involved cold gas main propulsion system (MPS) and solid rocket motor (SRB) plume simulations at Mach numbers from 0.6 to 1.4. Integrated vehicle surface pressure distributions, elevon and rudder hinge moments, and wing and vertical tail root bending and torsional moments due to MPS and SRB plume interactions were determined. Nozzle power conditions were controlled per pretest nozzle calibrations. Model angle of attack was varied from -4 deg to +4 deg; model angle of sideslip was varied from -4 deg to +4 deg. Reynolds number was varied for certain test conditions and configurations, with the nominal freestream total pressure being 14.69 psia. Plotted force and pressure data are presented.

Nichols, M. E.↗

NASA NTRS1983

Heat-transfer distributions on biconics at incidence in hypersonic-hypervelocity He, N2, air, and CO2 flows

Laminar heat transfer rates were measured on spherically blunted, 13 deg/7 deg on axis and bent biconics (fore cone bent 7 deg upward relative to aft cone) at hypersonic hypervelocity flow conditions in the Langley Expansion Tube. Freestream velocities from 4.5 to 6.9 km/sec and Mach numbers from 6 to 9 were generated using helium, nitrogen, air, and carbon dioxide test gases, resulting in normal shock density ratios from 4 to 19. Angle of attack, referenced to the axis of the aft cone, was varied from 0 to 20 deg in 4 deg increments. The effect of nose bend, angle of attack, and real gas phenomena on heating distributions are presented along with comparisons of measurement to prediction from a code which solves the three dimensional parabolized Navier-Stokes equations.

Miller, C. G.↗

NASA NTRS1983

Heat-transfer distributions on biconics at incidence in hypersonic-hypervelocity He, N2, air, and CO2 flows

Laminar heat-transfer rates were measured on spherically blunted, 13 degrees/F degrees on-axis and bent biconics (fore cone bent 7 degrees upward relative to aft cone) at hypersonic-hypervelocity flow conditions in the Langley Expansion Tube. Freestream velocities from 4.5 to 6.9 km/sec and Mach numbers from 6 to 9 were generated using helium, nitrogen, air, and carbon dioxide test gases, resulting in normal shock density ratios from 4 to 19. Angle of attack, referenced to the axis of the aft cone, was varied from zero to 20 degrees in 4 degree increments. The effect of nose bend, angle of attack, and real-gas phenomena on heating distributions are presented along with comparisons of measurement to prediction from a code which solves the three-dimensional 'parabolized Navier-Stokes' equations.

Miller, C. G.↗

NASA NTRS1985

Numerical simulation of viscous supersonic flow over a generic fighter configuration

A procedure is presented, as well as some results, to calculate the flow over a generic fighter configuration. A parabolized marching Navier-Stokes code is used to obtain the solution over a wing-canopy body. The flow conditions simulate supersonic cruise with a freestream Mach number of 2.169 and angles of attack of 4 and 10 deg. The body surface is considered to be an adiabatic wall and the flow is assumed to be turbulent for the given Reynolds number.

Chaussee, D. S.↗

NASA NTRS1987

Simulation of transonic viscous flow over a fighter-like configuration including inlet

The simulation of transonic viscous flow over a modified F-16A including inlet is presented. A zonal approach is utilized which allows appropriate clustering suitable for viscous calculations on all solid surfaces. Computational efficiency is enhanced by solving the thin-layer Navier-Stokes equations in viscous zones adjacent to the aircraft and the Euler equations on those coarse zones away from the aircraft. The flow conditions for this transonic case are freestream Mach number of 0.9, angle of attack of 4.12 deg, and a Reynolds number (based on root chord) of 4.5 million. A total of 19 zones are utilized yielding a total of 350,000 grid points. This case required about 3000 iterations to reduce the residual by three orders, which takes about 10 hr of CPU time on the Cray X-MP/48 computer. Pressure distributions on the wing and on cross sections through the inlet region compare favorably with the experimental data for this transonic case.

Flores, Jolen↗

NASA NTRS1992

Real gas flow parameters for NASA Langley 22-inch Mach 20 helium tunnel

A computational procedure was developed which can be used to determine the flow properties in hypersonic helium wind tunnels in which real gas behavior is significant. In this procedure, a three-coefficient virial equation of state and the assumption of isentropic nozzle flow are employed to determine the tunnel reservoir, nozzle, throat, freestream, and post-normal shock conditions. This method was applied to a range of conditions which encompasses the operational capabilities of the LaRC 22-Inch Mach 20 Helium Tunnel. Results are presented graphically in the form of real gas correction factors which can be applied to perfect gas calculations. Important thermodynamic properties of helium are also plotted versus pressure and temperature. The computational scheme used to determine the real-helium flow parameters was incorporated into a FORTRAN code which is discussed.

Hollis, Brian R.↗

NASA NTRS2004

Performance and Flowfield Measurements on a 10-inch Ducted Rotor VTOL UAV

A ducted fan VTOL UAV with a 10-inch diameter rotor was tested in the US Army 7-by 10-Foot Wind Tunnel. The test conditions covered a range of angle of attack from 0 to 110 degrees to the freestream. The tunnel velocity was varied from 0 (simulating a hover condition) to 128 ft/sec in propeller mode. A six-component internal balance measured the aerodynamic loads for a range of model configurations. including the isolated rotor, the isolated duct, and the full configuration of the duct and rotor. For some conditions, hotwire velocity surveys were conducted along the inner and outer surface of the duct and across the downstream wake. In addition, fluorescent oil flow visualization allowed the flow separation patterns inside and outside of the duct to be mapped for a few test conditions. Two different duct shapes were tested to determine the performance effects of leading edge radius. For each duct, a range of rotor tip gap from 1%R to 4.5%R was tested to determine the performance penalty in hover and axial flight. Measured results are presented in terms of hover performance, hover performance in a crosswind, and high angle of attack performance in propeller mode. In each case, the effects of both tip gap and duct leading edge radius are illustrated using measurements. Some of the hover performance issues were also studied using a simple analytical method, and the results agreed with the measurements.

Martin, Preston↗

NASA NTRS1988

Vibration and flutter characteristics of the SR7L large-scale propfan

An investigation of the vibration characteristics and aeroelastic stability of the SR7L Large-Scale Advanced Propfan was performed using a finite element blade model and an improved aeroelasticity code. Analyses were conducted for different blade pitch angles, blade support conditions, number of blades, rotational speeds, and freestream Mach numbers. A finite element model of the blade was used to determine the blade's vibration behavior and sensitivity to support stiffness. The calculated frequencies and mode shape obtained with this model agreed well with the published experimental data. A computer code recently developed at NASA Lewis Research Center and based on three-dimensional, unsteady, lifting surface aerodynamic theory was used for the aeroelastic analysis to examine the blade's stability at a cruise condition of Mach 0.8 at 1700 rpm. The results showed that the blade is stable for that operating point. However, a flutter condition was predicted if the cruise Mach number was increased to 0.9.

August, Richard↗

NASA NTRS1989

Some characteristics of bypass transition in a heated boundary layer

Experimental measurements of both mean and conditionally sampled characteristics of laminar, transitional and low Reynolds number turbulent boundary layers on a heated flat plate are presented. Measurements were obtained in air over a range of freestream turbulence intensities from 0.3 percent to 6 percent with a freestream velocity of 30.5 m/s and zero pressure gradient. Conditional sampling performed in the transitional boundary layers indicate the existence of a near-wall drop in intermittency, especially pronounced at low intermittencies. Nonturbulent intervals were observed to possess large levels of low-frequency unsteadiness, and turbulent intervals had peak intensities as much as 50 percent higher than were measured at fully turbulent stations. Heat transfer results were consistent with results of previous researchers and Reynolds analogy factors were found to be well predicted by laminar and turbulent correlations which accounted for unheated starting length. A small dependence of the turbulent Reynolds analogy factors on freestream turbulence level was observed. Laminar boundary layer spectra indicated selective amplification of unstable frequencies. These instabilities appear to play a dominant role in the transition process only for the lowest freestream turbulence level studied, however.

Sohn, K. H.↗