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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 1,171 records · Page 65
All sourcesNASA NTRSDOE OSTI
NASA NTRS1982

Effects of installation of F101 DFE exhaust nozzles on the afterbody-nozzle characteristics of the F-14 airplane

A wind-tunnel investigation was conducted to determine the effects of F101 DFE (derivative fighter engine) nozzle axial positioning on the afterbody-nozzle longitudinal aerodynamic characteristics of the F-14 airplane. The model was tested in the Langley 16-Foot Transonic Tunnel at Mach numbers from 0.7 to 1.25 and angles of attack from about -2 to 6 degrees. Compressed air was used to simulate nozzle exhaust flow at jet total-pressure ratios from 1 (jet off) to about 8. The results of the investigation show that for subsonic Mach numbers the intermediate cruise nozzle position of the three positions tested resulted in the lowest drag.

Reubush, D. E.↗
NASA NTRSDate not supplied

NASA Langley FUN3D Analyses in Support of the 1st AIAA Stability and Control Prediction Workshop

This work summarizes the results of FUN3D analyses conducted for the 1st AIAA Stability and Control Workshop on behalf of participants from the NASA Langley Research Center. The workshop was created to establish best practices for the prediction of stability and control derivatives using computational fluid dynamics and assess the limitations of these methods when those best practices are applied. The inaugural workshop considered the ONERA version of the NASA/Boeing Common Research Model, which includes the wing, body, horizontal tail, and a vertical tail designed by ONERA. Wind tunnel data at small sideslip angles remain unpublished and served as ‘blind’ data for computational comparisons. The present research generated workshop test case data using the NASA FUN3D code, which is a parallelized, unstructured, node-based, finite-volume discretization, Reynolds-averaged Navier-Stokes flow solver. Steady- state numerical simulations were conducted for workshop test cases investigating the following: grid convergence, Mach number effect on static stability, wind tunnel sting increments, static stability-derivative calculations, and a sideslip angle sweep. Results were generated for two series of unstructured, mixed-element grids, one set provided by the workshop and another set created using the HeldenMesh grid generation software. The results provided include total- and component-level breakdowns of the force and moment coefficients, in addition to sectional pressure distributions for the wing and tail components for comparisons to wind tunnel data.

CFD↗
NASA NTRS2023

Turbulence Model Validation Through Joint Experimental /Computational Studies of Separated Flow Over A Three-Dimensional Tapered Bump: Part I - Experimental Investigation

The results of a three-year experimental campaign aimed at comprehensively documenting the separated flow over a three-dimensional bump are presented with the purpose of generating a benchmark experimental database useful in validating computational fluid dynamics (CFD) flow simulations. The bump model geometry was designed to provide well-defined and repeatable smooth-body flow separation conditions that were suitable for both experiments and simulations. The bump had a Gaussian streamwise profile with a constant height equal to 8.5% of its width over the central 60% of the test section width. The remaining 40% were outboard spanwise portions that gradually taper to zero using an error function profile to minimize side-wall boundary layer effects. The model was immersed in a turbulent boundary layer that was developed on a suspended flat plate in the Notre Dame Mach 0.6 Wind Tunnel. In order to document the effect of the incoming boundary layer thickness on the flow separation, the bump model could be located at two streamwise positions. The mean velocity and turbulence intensity of the wind tunnel freestream flow field and approaching turbulent boundary layer were fully documented. The measurements of the flow separation region included surface visualization, wall shear stress using oil-film interferometry, mean and dynamic surface pressure, and planar and stereoscopic particle image velocimetry. The experiments were conducted over a range of Mach numbers from 0.05 to 0.2 corresponding to a range of Reynolds numbers based on the test section spanwise dimension (0.914m) of 1.0 × 10 6 ≤ Re L = U ∞ L/ν ≤ 4.0 × 10 6 . The bulk of the results are presented for the higher Mach number conditions of 0.1 and 0.2 with Re L = 2.0 × 10 6 and 4.0 × 10 6 , respectively. Extensive uncertainty analysis of the data was performed. The data is archived in the NASA Langley Turbulence Modeling Resource website at https://turbmodels.larc.nasa.gov/Other_exp_Data/speedbump_sep_exp.html. In addition to the experiments, a computational effort was made in parallel by the CFD group at Boeing Research & Technology highlighting the usefulness of the data set, which is outlined in the accompanying CFD report.

Flow separation↗
NASA NTRS1975

The effects of body geometry on the structure in the near wake zone of bodies in a flowing plasma

Plasma wind tunnel experiments were performed to investigate the effects of body geometry and size on the structure of the ion flux in the near wake zone behind the body in a streaming plasma. The results are applicable to the development of wake and sheath experiments for Spacelab. The experiments indicate that the body's dimension in the flow direction does not alter the gross features of the near wake structure, but primarily determines the relative amplitude of the peak structure and the downstream distance at which it occurs for a given body potential relative to the plasma.

Oran, W. A.↗
NASA NTRS1984

Airfoil tip vortex formation noise

Spectral data are presented for the noise produced due to the turbulent three-dimensional vortex flow existing near the rounded tip of lifting airfoils. The results are obtained by the comparison of sets of two- and three-dimensional test data for different airfoil model sizes, angles of attack, and tunnel flow velocities. Microphone cross-correlation and cross-spectral methods were used to determine the radiated noise. Corrections were made for tunnel shear layer and source directivity effects. Interpretation of the results are aided by a three-dimensional flow analysis developed for this study which determines open tunnel and finite aspect ratio corrections heretofore neglected in tip vortex studies. Hot wire measurements were made in the tip vortex formation region for the specification of governing flow parameters. The spectral data is normalized in a format considered most useful for subsequent quantitative prediction of this noise mechanism for practical systems such as helicopter rotors. Comparison is made to the analysis of George and Chou. A recommended prediction method is given.

Brooks, T. F.↗
NASA NTRS1996

Effects of Winglets on the Drag of a Low-Aspect-Ratio Configuration

A wind-tunnel investigation has been performed to determine the effect of winglets on the induced drag of a low-aspect-ratio wing configuration at Mach numbers between 0.30 and 0.85 and a nominal angle-of-attack range from -2 deg to 20 deg. Results of the tests at the cruise lift coefficient showed significant increases in lift-drag ratio for the winglet configuration relative to a wing-alone configuration designed for the same lift coefficient and Mach number. Further, even larger increases in lift-drag ratio were observed at lift coefficients above the design value at all Mach numbers tested. The addition of these winglets had a negligible effect on the static lateral-directional stability characteristics of the configuration. No tests were made to determine the effect of these winglets at supersonic Mach numbers, where increases in drag caused by winglets might be more significant. Computational analyses were also performed for the two configurations studied. Linear and small-disturbance formulations were used. The codes were found to give reasonable performance estimates sufficient for predicting changes of this magnitude.

Smith, Leigh Ann↗
NASA NTRS1999

High Reynolds Number Effects on HSCT Stability and Control Characteristics

Two wind tunnel tests during 1995 in the National Transonic Facility (NTF 070 and 073) served to define Reynolds number effects on longitudinal and lateral-directional stability and control. Testing was completed at both high lift and transonic conditions. The effect of Reynolds number on the total airplane configuration, horizontal and vertical tail effectiveness, forebody chine performance, rudder control and model aeroelastics was investigated. This paper will present pertinent stability and control results from these two test entries. Note that while model aeroelastic effects are examined in this presentation, no corrections for these effects have been made to the data.

Elzey, Michael B.↗
NASA NTRS1972

An efficient algorithm using matrix methods to solve wind tunnel force-balance equations

An iterative procedure applying matrix methods to accomplish an efficient algorithm for automatic computer reduction of wind-tunnel force-balance data has been developed. Balance equations are expressed in a matrix form that is convenient for storing balance sensitivities and interaction coefficient values for online or offline batch data reduction. The convergence of the iterative values to a unique solution of this system of equations is investigated, and it is shown that for balances which satisfy the criteria discussed, this type of solution does occur. Methods for making sensitivity adjustments and initial load effect considerations in wind-tunnel applications are also discussed, and the logic for determining the convergence accuracy limits for the iterative solution is given. This more efficient data reduction program is compared with the technique presently in use at the NASA Langley Research Center, and computational times on the order of one-third or less are demonstrated by use of this new program.

Smith, D. L.↗
NASA NTRSDate not supplied

Upwind Buildings Effects on Test Section Performance of the NFAC 80- by 120-Foot Wind Tunnel

In 2013, a new development began that would put multi-story buildings within the exclusion zone of the 80- by 120- Foot Wind Tunnel (80x120) inlet that exceeded the then existing building height restriction. This raised questions concerning whether the tunnel could operate within the acceptable turbulence limits of ≤ 0.5% in the axial direction over 75% area in the wind tunnel test section. The goal of this testing was to determine the validity of this hypothesis. Testing was performed using the 1/50th-scale model of the 80- by 120-FootWind Tunnel within the full-scale facility test section to characterize the model’s test section turbulence levels as a function of wind direction with the presence of the upwind buildings that are currently being constructed. This paper reports the first testing performed to date on the 80x120 inlet as-built under controlled test conditions for wind magnitude and wind onset direction for a variety of buildings in the inlet external flow field, including no buildings (the control case). This testing determined that, at rated operational conditions the upwind blockages caused no increase in turbulence intensity levels measured in the test section and, in some limited cases, caused the turbulence levels to decrease. Lastly, looking at generic rectangular blockages showed that, to have a discernible effect on test section turbulence levels, scaled 100-ft tall blockages need to be within 600 ft of the existing inlet.

Upwind Buildings Effects↗
NASA NTRS1977

Effects of Simulated Flight on Fan Noise Suppression

Attenuation properties of three treated fan inlets were evaluated. Tunnel flow simulated the inflow clean-up effect on source noise observed in flight and allowed observation of the blade passage frequency tone cut-off phenomenon. Acoustic data consisted of isolated inlet noise measured in the far field at two fixed positions and with traverses at four frequencies. Attenuation and source noise properties with and without flight simulation are compared and discussed. Averaged attenuation properties showed relative agreement of the inlets with their design intent, however, tunnel flow significantly affected the attenuation spectra.

Heidmann, M. F.↗
NASA NTRS1979

Review of design and operational characteristics of the 0.3-meter transonic cryogenic tunnel

The fundamentals of cryogenic testing are validated both analytically and experimentally employing the 0.3-m transonic cryogenic tunnel. The tunnel with its unique Reynolds number capability has been used for a wide variety of aerodynamic tests. Techniques regarding real-gas effects have been developed and cryogenic tunnel conditions are set and maintained accurately. It is shown that cryogenic cooling, by injecting nitrogen directly into the tunnel circuit, imposes no problems with temperature distribution or dynamic response characteristics.

Ray, E. J.↗
NASA NTRS1986

Computer simulation of a wind tunnel test section with discrete finite-length wall slots

A computer simulation of a slotted wind tunnel test section which includes a discrete, finite-length wall slot representation with plenum chamber constraints and accounts for the nonlinear effects of the dynamic pressure of the slot outflow jet and of the low energy of slot inflow air was developed. The simulation features were selected to be those appropriate for the intended subsequent use of the simulation in a wall interference assessment procedure using sparsely located wall pressure measurements. It is demonstrated that accounting for slot discreteness is important in interpreting wall pressure measured between slots, and that accounting for nonlinear slot flow effects produces significant changes in tunnel-induced velocity distributions and, in particular, produces a longitudinal component of tunnel-induced velocity due to model lift. A characteristic mode of tunnel flow interaction with constraints imposed by the plenum chamber and diffuser entrance is apparent in simulation results and is derived analytically through a simplified analysis.

Kemp, W. B., Jr.↗
NASA NTRS1984

High Reynolds number tests of the cast 10-2/DOA2 transonic airfoil at ambient and cryogenic temperature conditions

The transonic airfoil CAST 10-2/DOA2 was investigated in several major transonic wind tunnels at Reynolds numbers ranging from Re=1 million six hundred thousand to forty five million at ambient and cryogenic temperature conditions. The main objective was to study the degree and extent of the effects of Reynolds on both the airfoil aerodynamic characteristics and the interference effects of various model-wind-tunnel systems. the initial analysis of the 10-2 airfoil results revealed appreciable real Reynolds number effects on this airfoil and, moreover, showed that wall interference, can be significantly affected by changes in Reynolds number thus appearing as true Reynolds number effects.

Stanewsky, E.↗
NASA NTRS1946

The effect of wall interference upon the aerodynamic characteristics of an airfoil spanning a closed-throat circular wind tunnel

The results of a theoretical and experimental investigation of wall interference for an airfoil spanning a closed-throat circular wind tunnel are presented. Analytical equations are derived which relate the characteristics of an airfoil in the tunnel at subsonic speeds with the characteristics in free air. The analysis takes into consideration the effect of fluid compressibility and is based upon the assumption that the chord of the airfoil is small as compared with the diameter of the tunnel. The development is restricted to an untwisted, constant-chord airfoil spanning the middle of the tunnel. Brief theoretical consideration is also given to the problem of choking at high speeds. Results are then presented of tests to determine the low-speed characteristics of an NACA 4412 airfoil for two chord-diameter ratios. While, on the basis of these experiments, no appraisal is possible of the accuracy of the corrections at high speeds, the data indicate that at low Mach numbers the analytical results are valid, even for relatively large values of the chord-diameter ratio.

Vincenti, Walter G↗
NASA NTRS2020

Direct Numerical Simulations of Acoustic Disturbances in Various Rectangular Nozzle Configurations

We perform Direct Numerical Simulations (DNS) to study the acoustic freestream disturbances radiating from the turbulent boundary layers along the contoured nozzle walls of a hypersonic wind tunnel with a rectangular test section. To begin with, the effects of the spanwise end walls are suppressed by confining the spanwise computational domain to a finite segment of the overall nozzle cross section and by imposing periodic boundary conditions across that spanwise domain. Besides providing a building-block configuration to reveal par- tial effects of the enclosed acoustic environment within the wind tunnel, these computations serve as a stepping stone toward the goal of fully-3D computations including the nozzle end walls. Building upon the earlier simulations of Deegan et al. (2018), we show that the computed acoustic characteristics in the spanwise periodic simulations are insensitive to changes in the grid resolution parameters (e.g., x+ and z+). This is relative to the previous simulations that involved a coarser resolution in the streamwise and wall-normal directions, especially up- stream of the test section. Furthermore, we outline a comparison of the pressure fluctuations induced by the turbulent boundary layers over the contoured nozzle walls and several other calculations involving boundary layers over a single flat plate at nearly the same value of the edge Mach number. We also show the impact of having only one turbulent boundary layer instead of two within the computational setup. Various flow statistics, including the first and second moments of the unsteady flow field, are computed for the different flow configurations. Good comparisons of the statistics of the nozzle-wall boundary-layer turbulence and of the freestream acoustic disturbances between the simulations of coarser and finer grids confirms DNS procedure, but also the insensitivity of noise characteristics t section to the inflow turbulence generation technique.

Hildebrand, Nathaniel↗
NASA NTRS1976

Effects of forward velocity on noise for a J85 turbojet engine with multitube suppressor from wind tunnel and flight tests

Flight and wind tunnel noise tests were conducted using a J85 turbojet engine as a part of comprehensive programs to obtain an understanding of forward velocity effects on jet exhaust noise. Nozzle configurations of primary interest were a 104-tube suppressor with and without an acoustically-treated shroud. The installed configuration of the engine was as similar as possible in the flight and wind tunnel tests. Exact simultaneous matching of engine speed, exhaust velocity, and exhaust temperature was not possible, and the wind tunnel maximum Mach number was approximately 0.27, while the flight Mach number was approximately 0.37. The nominal jet velocity range was 450 to 640 m/sec. For both experiments, background noise limited the jet velocity range for which significant data could be obtained. In the present tests the observed directivity and forward velocity effects for the suppressor are more similar to predicted trends for internally-generated noise than unsuppressed jet noise.

Stone, J. R.↗
NASA NTRS1976

Effects of thickness on the aerodynamic characteristics of an initial low-speed family of airfoils for general aviation applications

Wind tunnel tests were conducted to determine the effects of airfoil thickness-ratio on the low speed aerodynamic characteristics of an initial family of airfoils. The results were compared with theoretical predictions obtained from a subsonic viscous method. The tests were conducted over a Mach number range from 0.10 to 0.28. Chord Reynolds numbers varied from about 2.0 x 1 million to 9.0 x 1 million.

Mcghee, R. J.↗