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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 73 records · Page 4
All sourcesNASA NTRSDOE OSTI
NASA NTRS1973

Local skin friction coefficients and boundary layer profiles obtained in flight from the XB-70-1 airplane at Mach numbers up to 2.5

Boundary-layer and local friction data for Mach numbers up to 2.5 and Reynolds numbers up to 3.6 x 10 to the 8th power were obtained in flight at three locations on the XB-70-1 airplane: the lower forward fuselage centerline (nose), the upper rear fuselage centerline, and the upper surface of the right wing. Local skin friction coefficients were derived at each location by using (1) a skin friction force balance, (2) a Preston probe, and (3) an adaptation of Clauser's method which derives skin friction from the rake velocity profile. These three techniques provided consistent results that agreed well with the von Karman-Schoenherr relationship for flow conditions that are quasi-two-dimensional. At the lower angles of attack, the nose-boom and flow-direction vanes are believed to have caused the momentum thickness at the nose to be larger than at the higher angles of attack. The boundary-layer data and local skin friction coefficients are tabulated. The wind-tunnel-model surface-pressure distribution ahead of the three locations and the flight surface-pressure distribution ahead of the wing location are included.

Fisher, D. F.↗
NASA NTRS1986

Measurement of heat transfer and pressure drop in rectangular channels with turbulence promoters

Periodic rib turbulators were used in advanced turbine cooling designs to enhance the internal heat transfer. The objective of the present project was to investigate the combined effects of the rib angle of attack and the channel aspect ratio on the local heat transfer and pressure drop in rectangular channels with two opposite ribbed walls for Reynolds number varied from 10,000 to 60,000. The channel aspect ratio (W/H) was varied from 1 to 2 to 4. The rib angle of attack (alpha) was varied from 90 to 60 to 45 to 30 degree. The highly detailed heat transfer coefficient distribution on both the smooth side and the ribbed side walls from the channel sharp entrance to the downstream region were measured. The results showed that, in the square channel, the heat transfer for the slant ribs (alpha = 30 -45 deg) was about 30% higher that of the transverse ribs (alpha = 90 deg) for a constant pumping power. However, in the rectangular channels (W/H = 2 and 4, ribs on W side), the heat transfer at alpha = 30 -45 deg was only about 5% higher than 90 deg. The average heat transfer and friction correlations were developed to account for rib spacing, rib angle, and channel aspect ratio over the range of roughness Reynolds number.

Han, J. C.↗
NASA NTRS1987

Investigation of a delta-wing fighter model flow field at transonic speeds

The paper reports a flow-field investigation on a 7.52-percent scale model of an advanced fighter aircraft design conducted in the NASA-Langley 16-ft Transonic Tunnel. The effects of free-stream Mach number, angle-of-attack, angle of sideslip, and various vortex control devices on the local flow values were studied. The model was tested at Mach numbers of 0.6, 0.9, and 1.2 and the angles of sideslip of 0 and +/- 5 deg; the model angle-of-attack was varied from -4 to 30 deg. Results are presented in terms of contour plots of local total pressure recovery. The dominant influence on the over-wing flow field was found to be the wing leading-edge vortex which first appears in the survey region at an angle-of-attack of 8 deg and increases in strength and influence with increasing angle-of-attack, finally dominating the entire survey region at very high angles-of-attack.

Bare, E. Ann↗
NASA NTRS1956

Investigation of Heat Transfer From

The convective heat transfer from the surface of an ellipsoidal forebody of fineness ratio 3 and 20-inch maximum diameter was investigated in clear air for both stationary and rotating operation over a range of conditions including air speeds up to 240 knots, rotational speeds up to 1200 rpm, and angles of attack of 0 deg, 3 deg, and 6 deg. The results are presented in the form of heat-transfer coefficients and the correlation of Nusselt and Reynolds numbers. Both a uniform surface temperature and a uniform input heater density distribution were used. The experimental results agree well with theoretical predictions for uniform surface temperature distribution. Complete agreement was not obtained with uniform input heat density in the laminar-flow region because of conduction effects. No significant effects of rotation were obtained over the range of airstream and rotational speeds investigated. Operation at angle of attack had only minor effects on the local heat transfer. Transition from laminar to turbulent heat transfer occurred over a wide range of Reynolds numbers. The location of transition depended primarily on surface roughness and pressure and temperature gradients. Limited transient heating data indicate that the variation of surface temperature with time followed closely an exponential relation.

Lewis, James P.↗
NASA NTRS2004

Application of a Momentum Source Model to the RAH-66 Comanche FANTAIL

A Momentum Source Model has been revised and implemented in the flow solver OVERFLOW-D. In this approach, the fan forces are evaluated from two-dimensional airfoil tables as a function of local Mach number and angle-of-attack and applied as source terms in the discretized Navier-Stokes equations. The model revisions include a new model for forces in the tip region and axial distribution of the source terms. The model revisions improve the results significantly. The Momentum Source Model agrees well with a discrete blade model for all computed collective pitch angles. The two models agree well with experimental data for thrust vs. torque. The Momentum Source Model is a good complement to Discrete Blade Models for ducted fan computations. The lower computational and labor costs make parametric studies, optimization studies and interactional aerodynamics studies feasible for cases beyond what is practical with a Discrete Blade Model today.

Nygaard, Tor A.↗
NASA NTRS1957

Experimental Droplet Impingement on Four Bodies of Revolution

The rate and. area of cloud droplet impingement on four bodies of revolution were obtained experimentally in the NACA Lewis icing tunnel with a dye-tracer technique. The study included spheres, ellipsoidal forebodies of fineness ratios of 2.5 and 3.0, and a conical forebody of 300 included angle and covered a range of angles of attack from 0? to 60 and rotational speeds up to 1200 rpm. The data were obtained at an airspeed of 157 knots and are correlated by dimensionless impingement parameters. In general, the experimental data show that the local and total impingement rates and impingement limits of bodies of revolution are primarily functions of the modified inertia parameters, the body shape, and fineness ratio. Both the local impingement rate and impingement limits depend upon the angle of attack. Rotation of the bodies had a negligible effect on the impingement characteristics except for an averaging effect at angle of attack. For comparable diameters the bluffer bodies had the largest total impingement efficiency, but the finer and sharper bodies had the largest values of maximum local impingement efficiency and, in most cases, the largest limits of impingement. In most cases, the impingement characteristics were less than those calculated from theoretical trajectories; in general, however, fairly good agreement was obtained between the experimental and theoretical impingement characteristics.

Lewis, James P.↗
NASA NTRS2001

Flight, Wind-Tunnel, and Computational Fluid Dynamics Comparison for Cranked Arrow Wing (F-16XL-1) at Subsonic and Transonic Speeds

Geometrical, flight, computational fluid dynamics (CFD), and wind-tunnel studies for the F-16XL-1 airplane are summarized over a wide range of test conditions. Details are as follows: (1) For geometry, the upper surface of the airplane and the numerical surface description compare reasonably well. (2) For flight, CFD, and wind-tunnel surface pressures, the comparisons are generally good at low angles of attack at both subsonic and transonic speeds, however, local differences are present. In addition, the shock location at transonic speeds from wind-tunnel pressure contours is near the aileron hinge line and generally is in correlative agreement with flight results. (3) For boundary layers, flight profiles were predicted reasonably well for attached flow and underneath the primary vortex but not for the secondary vortex. Flight data indicate the presence of an interaction of the secondary vortex system and the boundary layer and the boundary-layer measurements show the secondary vortex located more outboard than predicted. (4) Predicted and measured skin friction distributions showed qualitative agreement for a two vortex system. (5) Web-based data-extraction and computational-graphical tools have proven useful in expediting the preceding comparisons. (6) Data fusion has produced insightful results for a variety of visualization-based data sets.

John E Lamar↗
NASA NTRS1975

Flight transition data for angles of attack at Mach 22 with correlations of the data

Boundary-layer transition data for angles of attack from 2.5 to 47 deg from a flight experiment with a cone that reentered at angles of attack up to 75 deg were analyzed and their local flow conditions are presented. The transition data were obtained from both acoustic and electrostatic sensors. The data from the acoustic and electrostatic sensors were correlated by use of three different sets of correlating parameters. For each set of correlating parameters, the transitional and turbulent data from the acoustic and electrostatic sensors were plotted separately. Each of the correlations was compared with a linear curve fit of previous cone-flight data at near zero angle of attack. Results indicate that the data from acoustic sensors falls below the linear curve fits to the flight data and indicates the sensitivity of the sensors. It is shown that the transition data from the electrostatic sensors tends to scatter much more than the data from the acoustic sensors. However, the data from the electrostatic sensors tends to scatter about the linear curve fits based on previous flight data.

Johnson, C. B.↗
NASA NTRS1987

Local heat/mass transfer and pressure drop in a two-pass rib-roughened channel for turbine airfoil cooling

The heat transfer characteristics of turbulent air flow in a multipass channel were studied via the naphthalene sublimation technique. The naphthalene-coated test section, consisting of two straight, square channels joined by a 180 deg turn, resembled the internal cooling passages of gas turbine airfoils. The top and bottom surfaces of the test channel were roughened by rib turbulators. The rib height-to-hydraulic diameter ratio (e/D) were 0.063 and 0.094, and the rib pitch-to-height ratio (P/e) were 10 and 20. The local heat/mass transfer coefficients on the roughened top wall and on the smooth divider and side walls of the test channel were determined for three Reynolds numbers of 15, 30, and 60, thousand, and for three angles of attack (alpha) of 90, 60, and 45 deg. Results showed that the local Sherwood numbers on the ribbed walls were 1.5 to 6.5 times those for a fully developed flow in a smooth square duct. The average ribbed-wall Sherwood numbers were 2.5 to 3.5 times higher than the fully developed values, depending on the rib angle of attack and the Reynolds number. The results also indicated that, before the turn, the heat/mass transfer coefficients in the cases of alpha = 60 and 45 deg were higher than those in the case of alpha=90 deg. However, after the turn, the heat/mass transfer coefficients in the oblique-rib cases were lower than those in the transverse rib case. Correlations for the average Sherwood number ratios for individual channel surfaces and for the overall Sherwood number ratios are reported. Correlations for the fully developed friction factors and for the loss coefficients are also provided.

Han, J. C.↗
NASA NTRS1956

Experimental Droplet Impingement on Several Two-Dimensional Airfoils with Thickness Ratios of 6 to 16 Percent

The rate and area of cloud droplet impingement on several two-dimensional swept and unswept airfoils were obtained experimentally in the NACA Lewis icing tunnel with a dye-tracer technique. Airfoil thickness ratios of 6 to 16 percent; angles of attack from 0 deg to 12 deg, and chord sizes from 13 to 96 inches were included in the study. The data were obtained at 152 knots and are extended to other conditions by dimensionless impingement parameters. In general, the data show that the total and local collection efficiencies and impingement limits are primary functions of the modified inertia parameter (in which airspeed, droplet size, and body size are the most significant variables) and the airfoil thickness ratio. Local collection efficiencies and impingement limits also depend on angle of attack. Secondary factors affecting impingement characteristics are airfoil shape, camber, and sweep angle. The impingement characteristics obtained experimentally for the airfoils were within +/-10 percent on the average of the characteristics calculated from theoretical trajectories. Over the range of conditions studied, the experimental data demonstrate that a specific method can be used to predict the impingement characteristics of swept airfoils with large aspect ratios from the data for unswept airfoils of the same series.

Gelder, Thomas F.↗
NASA NTRS1953

Preliminary Results of Heat Transfer from a Stationary and Rotating Ellipsoidal Spinner

Convective heat-transfer coefficients in dry air were obtained for an ellipsoidal spinner of 30-inch maximum diameter for both stationary and rotating operation over a range of conditions including airspeeds up to 275 miles per hour, rotational speeds up to 1200 rpm, and angles of attack of zero and 40 The results are presented in terms of Nusselt numbers, Reynolds numbers, and convective heat-transfer coefficients. The studies included both uniform heating densities over the spinner and uniform surface temperatures.. In general, the results showed that rotation will increase the convective heat transfer from a spinner, especially in the turbulent-flow regions. Rotation of the spinner at 1200 rpm and at a free-stream velocity of 275 miles per hour increased the Nusselt number parameter in the turbulent-flow region by 32 percent over that obtained with a stationary spinner; whereas in the nose region, where the flow was laminar, an increase of only 18 percent was observed. Transition from laminar to turbulent flow occurred over a large range of Reynolds numbers primarily because of surface roughness of the spinner. Operation at an angle of attack of 40 had only small effects on the local convective heat transfer for the model studied.

vonGlahn, U.↗
NASA NTRS1947

Three-Component Force and Mass-Flow Measurement on a Jet Nacelle

The present report contains three-component forces and mass-flow measurements on a jet nacelle at small inlet-velocity ratios v(sub A)/v(sub o) < 1. The mass-flow measurement on two cross sections of the nacelle demonstrate the local-velocity distribution with varying flow and angle of attack.

Ilgmann, G.↗
NASA NTRSDate not supplied

Aeroacoustic Computations of a Generic Low Boom Concept in Landing Configuration: Part 1 - Aerodynamic Simulations

Computational results are presented for a 15%-scale, full-span, generic low-boom concept aircraft model simulated as installed in the NASA Langley Research Center 14- by 22-Foot Subsonic Tunnel and in free-air conditions. The simulations were conducted with the lattice Boltzmann solver PowerFLOW® to capture the time-accurate characteristics of the flow. The aerodynamic behavior of the aircraft was investigated with various combinations of flap and landing gear deployments. To replicate the flow speed used during tests of the model, a Mach number of approximately 0.20 was chosen while varying the angle of attack through a limited range. Analyses were performed on local flow quantities, global forces, as well as flow distributions between configurations. Aerodynamic quantities were shown to be sensitive to mesh resolution levels, especially with flap and landing gear systems retracted. When comparing wind tunnel to free-air simulations, the aerodynamic quantities sampled were found to be in good agreement, suggesting that tunnel blockage and model support effects did not adversely impact the aerodynamic behavior of the model during the wind tunnel test. The results of this study improve our understanding of the highly complex, vortical flow generated by configurations with swept, low aspect-ratio wings at speeds encountered during landing operations.

low boom↗
NASA NTRS1983

Aerodynamic characteristics of missile control fins in nonlinear flow fields

Recent experimental results show that the control effectiveness of a missile fin in supersonic flow at moderate-to-high angles of attack is a strong nonlinear function of free-stream Mach number, body incidence angle, fin bank angle and fin deflection angle. Analysis of the experimental results using an Euler finite-difference computer code with flow separation together with the equivalent angle-of-attack concept indicates that the observed nonlinearities are due to the variation of local dynamic pressure and local Mach number around the missile body alone. The nonlinearities are shown to be a strong source of control cross-coupling for high Mach number, high angle-of-attack combinations. The analysis suggests a relatively simple yet comprehensive approach for accurately accounting for these nonlinear effects.

Hemsch, M. J.↗
NASA NTRS1990

Application of localized surface heating to actively control the boundary layer separation

The use of localized surface heating to improve the performance of the airfoil of a fixed wing or rotating blade at high angle of attack is considered. Separation control by active means can be accomplished by acoustic perturbation of the separated area above the airfoil or by localized surface heating in a pressure gradient region. Active control by sound in the vicinity of the leading edge of an airfoil is shown to greatly reduce the area of separation. Therefore, the overall circulation is sufficiently changed to result in an increase in lift and a reduction in drag.

El-Hady, Nabit M.↗
NASA NTRS1977

Vortex lattice prediction of subsonic aerodynamics of hypersonic vehicle concepts

The vortex lattice method introduced by Lamar and Gloss (1975) was applied to the prediction of subsonic aerodynamic characteristics of hypersonic body-wing configurations. The reliability of the method was assessed through comparison of the calculated and observed aerodynamic performances of two National Hypersonic Flight Research Facility craft at Mach 0.2. The investigation indicated that a vortex lattice model involving 120 or more panel elements can give good results for the lift and induced drag coefficients of the craft, as well as for the pitching moment at angles of attack below 10 to 15 deg. Automated processes for calculating the local slopes of mean-camber surfaces may also render the method suitable for use in preliminary design phases.

Pittman, J. L.↗
NASA NTRS1989

Numerical simulation and comparison of symmetrical/supercritical airfoils for the near tip region of a helicopter in forward flight

Aerodynamic loads on a multi-bladed helicopter rotor in forward flight at transonic tip conditions are calculated. The unsteady, three-dimensional, time-accurate compressible Reynolds-averaged thin layer Navier-Stokes equations are solved in a rotating coordinate system on a body-conformed, curvilinear grid of C-H topology. Detailed boundary layer and global numerical comparisons of NACA-0012 symmetrical and CAST7-158 supercritical airfoils are made under identical forward flight conditions. The rotor wake effects are modeled by applying a correction to the geometric angle of attack of the blade. This correction is obtained by computing the local induced downwash velocity with a free wake analysis program. The calculations are performed on the Numerical Aerodynamic Simulation Cray 2 and the VPS32 (a derivative of a Cyber 205 at the Langley Research Center) for a model helicopter rotor in forward flight.

Badavi, F. F.↗