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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 685 records · Page 38
All sourcesNASA NTRSDOE OSTI
NASA NTRS1977

F-8 supercritical wing flight pressure, Boundary layer, and wake measurements and comparisons with wind tunnel data

Data for speeds from Mach 0.50 to Mach 0.99 are presented for configurations with and without fuselage area-rule additions, with and without leading-edge vortex generators, and with and without boundary-layer trips on the wing. The wing pressure coefficients are tabulated. Comparisons between the airplane and model data show that higher second velocity peaks occurred on the airplane wing than on the model wing. The differences were attributed to wind tunnel wall interference effects that caused too much rear camber to be designed into the wing. Optimum flow conditions on the outboard wing section occurred at Mach 0.98 at an angle of attack near 4 deg. The measured differences in section drag with and without boundary-layer trips on the wing suggested that a region of laminar flow existed on the outboard wing without trips.

Montoya, L. C.↗
NASA NTRS1977

Efficient solution of unsteady transonic flows about airfoils

An implicit finite difference procedure was developed for the efficient solution of unsteady transonic flow fields. Sample computations illustrate applications of procedures to aerodynamic problems. Solutions are presented that illustrate three types of shock wave motion that can result from airfoil control surface oscillations. The significant effect of wind tunnel wall conditions on these shock wave motions is demonstrated. Solutions are also presented for a simple aeroelastic problem in which the flow field equations and the structural motion equations are integrated simultaneously in time. Both stable and unstable aeroelastic interactions are considered. The procedure is adapted to compute unsteady aerodynamic force coefficients by the indicial method.

Ballhaus, W. F.↗
NASA NTRS1981

Preliminary results of buffet tests in a cryogenic wind tunnel

Buffet tests of two wings with different leading-edge sweep show that it is feasible to use the standards wing root bending moment technique in a cryogenic wing tunnel. The results for the 65 deg sweep delta wing indicate the importance of matching the reduced frequency parameter in model tests for planforms which are sensitive to reduced frequency parameter if quantitative buffet measurements are required. The unique ability of a pressurized cryogenic wind tunnel to separate the effects of Reynolds number and of aeroelastic distortion by variations in the tunnel stagnation temperature and pressure was demonstrated.

Boyden, R. P.↗
NASA NTRS1981

Report of the panel on theoretical aerodynamics

Interactions between theoretical aerodynamics and the NTF are discussed. The development and validation of computational fluid dynamics computer codes, the determination of Reynolds number scaling laws, and extension of the data bases of entrainment type turbulence models to include high Reynolds number data are recommended areas of study. The major benefit theoretical aerodynamics could have on the NTF is in the quantitative description of wind tunnel wall interference effects.

South, J. C., Jr.↗
NASA NTRS1982

Experimental measurements of unsteady turbulent boundary layers near separation

Investigations conducted to document the behavior of turbulent boundary layers on flat surfaces that separate due to adverse pressure gradients are reported. Laser and hot wire anemometers measured turbulence and flow structure of a steady free stream separating turbulent boundary layer produced on the flow of a wind tunnel section. The effects of sinusoidal and unsteadiness of the free stream velocity on this separating turbulent boundary layer at a reduced frequency were determined. A friction gage and a thermal tuft were developed and used to measure the surface skin friction and the near wall fraction of time the flow moves downstream for several cases. Abstracts are provided of several articles which discuss the effects of the periodic free stream unsteadiness on the structure or separating turbulent boundary layers.

Simpson, R. L.↗
NASA NTRS1983

Transonic empirical configuration design process

This lecture describes some of the experimental research pertaining to transonic configuration development conducted by the Transonic Aerodynamics Branch of the NASA Langley Research Center. Discussions are presented of the following: use of florescent oil films for the study of surface boundary layer flows; the severe effect of wind tunnel wall interference on the measured configuration drag rise near the speed of sound as determined by a comparison between wind tunnel and free air results; the development of a near sonic transport configuration incorporating a supercritical wing and an indented fuselage, designed on the basis of the area rule with a modification to account for the presence of local supersonic flow above the wing; a device for improving the transonic pitch up of swept wings with very little added drag at the cruise condition; a means for reducing the large transonic aerodynamic interference between the wing, fuselage, nacelle and pylon for a for a fuselage mounted nacelle having the inlet above the wing; and methods for reducing the transonic interference between flows over a winglet and the wing.

Whitcomb, R. T.↗
NASA NTRS1985

A simple, accurate and efficient algorithm for unsteady transonic flow

The derivation of an algorithm for computing an airfoil's response to small unsteady perturbations is described. The analysis of the far-field boundary conditions for the flow field, and the application of the time-linearization technique to develop an accurate description of the flow field are discussed. The treatment of moving shock waves is examined. The alternating direction implicit scheme of Ballhaus and Steger (1975) is utilized for the discretization of the governing equations and boundary conditions. Numerical examples demonstrating the applicability of the time-linearization method to determine the correlation between airfoil and pressure distribution are provided. The effect of wind tunnel walls on unsteady transonic testing is studied.

Fung, K.-Y.↗
NASA NTRS1987

Estimation of blade airloads from rotor blade bending moments

A method is developed to estimate the blade normal airloads by using measured flap bending moments; that is, the rotor blade is used as a force balance. The blade's rotation is calculated in vacuum modes and the airloads are then expressed as an algebraic sum of the mode shapes, modal amplitudes, mass distribution, and frequency properties. The modal amplitudes are identified from the blade bending moments using the Strain Pattern Analysis Method. The application of the method is examined using simulated flap bending moment data that have been calculated for measured airloads for a full-scale rotor in a wind tunnel. The estimated airloads are compared with the wind tunnel measurements. The effects of the number of measurements, the number of modes, and errors in the measurements and the blade properties are examined, and the method is shown to be robust.

Bousman, William G.↗
NASA NTRS1989

Comparison of two- and three-dimensional Navier-Stokes solutions with NASA experimental data for CAST-10 airfoil

The two-dimensional (2-D) and three-dimensional Navier-Stokes equations are solved for flow over a NAE CAST-10 airfoil model. Recently developed finite-volume codes that apply a multistage time stepping scheme in conjunction with steady state acceleration techniques are used to solve the equations. Two-dimensional results are shown for flow conditions uncorrected and corrected for wind tunnel wall interference effects. Predicted surface pressures from 3-D simulations are compared with those from 2-D calculations. The focus of the 3-D computations is the influence of the sidewall boundary layers. Topological features of the 3-D flow fields are indicated. Lift and drag results are compared with experimental measurements.

Swanson, R. Charles↗
NASA NTRS1993

Casting Using A Polystyrene Pattern

New technique for making metal aircraft models saves significant amount of time and effort in comparison with conventional lost-wax method. Produces inexpensive, effective wind-tunnel models. Metal wind-tunnel model cast by use of polystyrene pattern.

Vasquez, Peter↗
NASA NTRS1950

Altitude Performance Characteristics of Tail-pipe Burner with Variable-area Exhaust Nozzle

An investigation was conducted in the NACA Lewis altitude wind tunnel to determine effect of altitude and flight Mach number on performance of tail-pipe burner equipped with variable-area exhaust nozzle and installed on full-scale turbojet engine. At a given flight Mach number, with constant exhaust-gas and turbine-outlet temperatures, increasing altitude lowered the tail-pipe combustion efficiency and raised the specific fuel consumption while the augmented thrust ratio remained approximately constant. At a given altitude, increasing flight Mach number raised the combustion efficiency and augmented thrust ratio and lowered the specific fuel consumption.

TURBINES, GAS - AFTERBURNING↗
NASA NTRS1954

Thrust and drag characteristics of a convergent-divergent nozzle with various exhaust jet temperatures

An investigation was conducted in the 8-by-6 foot supersonic wind tunnel on the effect of exhaust-gas temperatures on the external and internal characteristics of a convergent-divergent nozzle having an area expansion ratio of 1.83. Data were obtained over a pressure-ratio range from 1 to 20 at free-stream Mach numbers of 1.6 and 2.0 for exhaust temperatures of 860 degrees, 1650 degrees, and 2000 degrees R. Results of this investigation indicated that generally both the internal and external performance characteristics were only slightly affected by a large change in jet temperature. The small differences in performance which did occur were predicted satisfactorily from theoretical considerations.

Hearth, Donald P↗
NASA NTRS1981

Calculation of Unsteady Transonic Flow over an Airfoil

An implicit finite-difference solver for either the Euler equations or the "thin-layer" Navier-Stokes equations was used to calculate a transonic flow over the NACA 64A010 airfoil pitching about its one-quarter chord. An unsteady automatic grid-generation procedure that will improve significantly the computational efficiency of various unsteady flow problems is described. The calculated results for both inviscid and viscous flows at Much number 0.8 over the airfoil oscillating with reduced frequency referenced to one-half chord, 0.2, are compared with experimental data measured in the Ames 11 x 11-ft Transonic Wind Tunnel. Nonlinear, unsteady effects of the flow on the surface pressure variations, shock-wave excursions, and overall airloads are examined. Good agreements between the results of computations and experiments were obtained. In the shock-wave region, however, the results of the viscous-flow computations showed closer agreement with the experimental data.

Chyu, W. J.↗
NASA NTRS1981

Calculation of Unsteady Transonic Flow Over an Airfoil

An implicit finite-difference solver for either the Euler equations or the "thin-layer" Navier-Stokes equations was used to calculate a transonic flow over the NACA 64A010 airfoil pitching about its one-quarter chord. An unsteady automatic grid-generation procedure that will improve significantly the computational efficiency of various unsteady flow problems is described. The calculated results for both inviscid and viscous flows at Mach number 0.8 over the airfoil oscillating with reduced frequency referenced to one-half chord, 0.2, are compared with experimental data measured in the Ames 11 x 11 ft Transonic Wind Tunnel. Nonlinear, unsteady effects of the flow on the surface pressure variations, shock-wave excursions, and overall airloads are examined. Good agreements between the results of computations and experiments were obtained. In the shock-wave region, however, the results of the viscous-flow computations showed closer agreement with the experimental data.

Chyu, W. J.↗
NASA NTRS2001

A Study of Ice Accretion Physics to Improve the Ice Accretion on Airfoils

This three-year grant began on November 7, 1996 and was no-cost extended to end on October 30, 2000. The objectives of the grant were: (1) To examine the effect of wind tunnel turbulence on ice accretion; (2) To determine the relationship between ice accretion geometry and airfoil performance; and (3) To determine if the wake-survey method was an appropriate experimental technique for iced-airfoil drag measurement. As specified in the grant the primary deliverables for this research were annual reports in the form of AIAA papers presented at national meetings each year. Masters theses and annual oral reports to be given at NASA Lewis (now Glenn) were also deliverables. Six AIAA papers documented the research findings from this study, Mr. Chad Henze's Masters thesis describes the wind tunnel turbulence work in detail, and a summary of the icing wind tunnel turbulence work was published in the archival AIAA Journal of Aircraft. A brief summary of the findings is given. Please refer to the reports for the details of the studies and findings.

Bragg, Michael B.↗
NASA NTRS2005

Design and Testing of a Blended Wing Body With Boundary Layer Ingestion Nacelles at High Reynolds Numbers

A knowledge-based aerodynamic design method coupled with an unstructured grid Navier-Stokes flow solver was used to improve the propulsion/airframe integration for a Blended Wing Body with boundary-layer ingestion nacelles. A new zonal design capability was used that significantly reduced the time required to achieve a successful design for each nacelle and the elevon between them. A wind tunnel model was built with interchangeable parts reflecting the baseline and redesigned configurations and tested in the National Transonic Facility (NTF). Most of the testing was done at the cruise design conditions (Mach number = 0.85, Reynolds number = 75 million). In general, the predicted improvements in forces and moments as well as the changes in wing pressures between the baseline and redesign were confirmed by the wind tunnel results. The effectiveness of elevons between the nacelles was also predicted surprisingly well considering the crudeness in the modeling of the control surfaces in the flow code. A novel flow visualization technique involving pressure sensitive paint in the cryogenic nitrogen environment used in high-Reynolds number testing in the NTF was also investigated.

Campbell, Richard L.↗
NASA NTRS2006

Designing and Testing a Blended Wing Body with Boundary Layer Ingestion Nacelles

A knowledge-based aerodynamic design method coupled with an unstructured grid Navier-Stokes flow solver was used to improve the propulsion/airframe integration for a Blended Wing Body with boundary-layer ingestion nacelles. A new zonal design capability was used that significantly reduced the time required to achieve a successful design for each nacelle and the elevon between them. A wind tunnel model was built with interchangeable parts reflecting the baseline and redesigned configurations and tested in the National Transonic Facility (NTF). Most of the testing was done at the cruise design conditions (Mach number = 0.85, Reynolds number = 75 million). In general, the predicted improvements in forces and moments as well as the changes in wing pressures between the baseline and redesign were confirmed by the wind tunnel results. The effectiveness of elevons between the nacelles was also predicted surprisingly well considering the crudeness in the modeling of the control surfaces in the flow code.

Carter, Melissa B.↗
NASA NTRS2012

Numerical Predictions of Sonic Boom Signatures for a Straight Line Segmented Leading Edge Model

A sonic boom wind tunnel test was conducted on a straight-line segmented leading edge (SLSLE) model in the NASA Langley 4- by 4- Foot Unitary Plan Wind Tunnel (UPWT). The purpose of the test was to determine whether accurate sonic boom measurements could be obtained while continuously moving the SLSLE model past a conical pressure probe. Sonic boom signatures were also obtained using the conventional move-pause data acquisition method for comparison. The continuous data acquisition approach allows for accurate signatures approximately 15 times faster than a move-pause technique. These successful results provide an incentive for future testing with greatly increased efficiency using the continuous model translation technique with the single probe to measure sonic boom signatures. Two widely used NASA codes, USM3D (Navier-Stokes) and CART3D-AERO (Euler, adjoint-based adaptive mesh), were used to compute off-body sonic boom pressure signatures of the SLSLE model at several different altitudes below the model at Mach 2.0. The computed pressure signatures compared well with wind tunnel data. The effect of the different altitude for signature extraction was evaluated by extrapolating the near field signatures to the ground and comparing pressure signatures and sonic boom loudness levels.

Elmiligui, Alaa A.↗