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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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All sourcesNASA NTRSDOE OSTI
NASA NTRS2005

Statistical Calibration and Validation of a Homogeneous Ventilated Wall-Interference Correction Method for the National Transonic Facility

Wind tunnel experiments will continue to be a primary source of validation data for many types of mathematical and computational models in the aerospace industry. The increased emphasis on accuracy of data acquired from these facilities requires understanding of the uncertainty of not only the measurement data but also any correction applied to the data. One of the largest and most critical corrections made to these data is due to wall interference. In an effort to understand the accuracy and suitability of these corrections, a statistical validation process for wall interference correction methods has been developed. This process is based on the use of independent cases which, after correction, are expected to produce the same result. Comparison of these independent cases with respect to the uncertainty in the correction process establishes a domain of applicability based on the capability of the method to provide reasonable corrections with respect to customer accuracy requirements. The statistical validation method was applied to the version of the Transonic Wall Interference Correction System (TWICS) recently implemented in the National Transonic Facility at NASA Langley Research Center. The TWICS code generates corrections for solid and slotted wall interference in the model pitch plane based on boundary pressure measurements. Before validation could be performed on this method, it was necessary to calibrate the ventilated wall boundary condition parameters. Discrimination comparisons are used to determine the most representative of three linear boundary condition models which have historically been used to represent longitudinally slotted test section walls. Of the three linear boundary condition models implemented for ventilated walls, the general slotted wall model was the most representative of the data. The TWICS code using the calibrated general slotted wall model was found to be valid to within the process uncertainty for test section Mach numbers less than or equal to 0.60. The scatter among the mean corrected results of the bodies of revolution validation cases was within one count of drag on a typical transport aircraft configuration for Mach numbers at or below 0.80 and two counts of drag for Mach numbers at or below 0.90.

Walker, Eric Lee↗
NASA NTRS2005

Statistical Calibration and Validation of a Homogeneous Ventilated Wall-Interference Correction Method for the National Transonic Facility

Wind tunnel experiments will continue to be a primary source of validation data for many types of mathematical and computational models in the aerospace industry. The increased emphasis on accuracy of data acquired from these facilities requires understanding of the uncertainty of not only the measurement data but also any correction applied to the data. One of the largest and most critical corrections made to these data is due to wall interference. In an effort to understand the accuracy and suitability of these corrections, a statistical validation process for wall interference correction methods has been developed. This process is based on the use of independent cases which, after correction, are expected to produce the same result. Comparison of these independent cases with respect to the uncertainty in the correction process establishes a domain of applicability based on the capability of the method to provide reasonable corrections with respect to customer accuracy requirements. The statistical validation method was applied to the version of the Transonic Wall Interference Correction System (TWICS) recently implemented in the National Transonic Facility at NASA Langley Research Center. The TWICS code generates corrections for solid and slotted wall interference in the model pitch plane based on boundary pressure measurements. Before validation could be performed on this method, it was necessary to calibrate the ventilated wall boundary condition parameters. Discrimination comparisons are used to determine the most representative of three linear boundary condition models which have historically been used to represent longitudinally slotted test section walls. Of the three linear boundary condition models implemented for ventilated walls, the general slotted wall model was the most representative of the data. The TWICS code using the calibrated general slotted wall model was found to be valid to within the process uncertainty for test section Mach numbers less than or equal to 0.60. The scatter among the mean corrected results of the bodies of revolution validation cases was within one count of drag on a typical transport aircraft configuration for Mach numbers at or below 0.80 and two counts of drag for Mach numbers at or below 0.90.

Walker, Eric L.↗
NASA NTRS1984

Asymptotic methods for wind tunnel wall corrections at transonic speed

The effort to develop classical methods to compute wall interference at transonic speeds is outlined. The two-dimensional theory and three-dimensional development are discussed. Also, some numerical application of the two-dimensional work are indicated. The basic advantages of the asymptotic theory are noted.

Malmuth, N. D.↗
NASA NTRS1986

Transonic airfoil calculations including wind tunnel wall-interference effects

The results of Reynolds-averaged time-dependent inviscid and turbulent compressible Navier-Stokes computations using the implicit finite-difference approach of Steger (1978), modified by incorporating a pressure boundary condition, (PBC) to account for wall interference are compared with experimental data on a NACA 64A010 airfoil (Johnson and Bachalo, 1980) in graphs and briefly characterized. The computational approach is the same as that used by King and Johnson (1980), but a 137 x 50 mesh is used instead of a 97 x 35 mesh, and special care is taken in resolving the nose, shock, and trailing-edge regions. Imposition of PBC is shown to improve significantly the accuracy of the computations for the flowfield on the upper surface of the airfoil, shifting the shock forward to its experimentally measured position in the case of turbulent flow. The failure of the method, even with PBC, to match the experimental shock location in the case of a flow with a separation bubble is attributed to inadequacies in the algebraic turbulence model employed (Baldwin and Lomax, 1978).

King, L. S.↗
NASA NTRS1982

High Reynolds number tests of a Boeing BAC I airfoil in the Langley 0.3-meter transonic cryogenic tunnel

A wind tunnel investigation of an advanced-technology airfoil was conducted in the Langley 0.3-Meter Transonic Cryogenic Tunnel (TCT). This investigation represents the first in a series of NASA/U.X. industry two dimensional airfoil studies to be completed in the Advanced Technology Airfoil Test program. Test temperature was varied from ambient to about 100 K at pressures ranging from about 1.2 to 6.0 atm. Mach number was varied from about 0.40 to 0.80. These variables provided a Reynolds number (based on airfoil chord) range from about .0000044 to .00005. This investigation was specifically designed to: (1) test a Boeing advanced airfoil from low to flight-equivalent Reynolds numbers; (2) provide the industry participant (Boeing) with experience in cryogenic wind-tunnel model design and testing techniques; and (3) demonstrate the suitability of the 0.3-m TCT as an airfoil test facility. All the objectives of the cooperative test were met. Data are included which demonstrate the effects of fixed transition, Mach number, and Reynolds number on the aerodynamic characteristics of the airfoil. Also included are remarks on the model design, the model structural integrity, and the overall test experience.

Johnson, W. G., Jr.↗
NASA NTRS1984

Qualification of the T2 wind tunnel in cryogenic operation. B: Flow fluctuations, particle detection and qualification

This report presents part of the tests for verification of the T2 transonic induction wind tunnel in cryogenic operation. The first part of the results presented concerns fluctuations in pressure and temperature at ambient temperature and in cryogenic regulation. The second part presents the condensation phenomena which could be observed in the cryogenic flow by means of an optical particle detection system in the test section.

Dor, J. B.↗
NASA NTRSDate not supplied

Overview of the High Reynolds Number Ascent Wind Tunnel Test of the Space Launch System at the National Transonic Facility

Transonic, high Reynolds number wind tunnel testing for the Space Launch System (SLS) ascent flight environment was conducted in the National Transonic Facility (NTF) at the NASA Langley Research Center from December 2019 to April 2021. The test was sponsored by the SLS Program and the NASA Engineering & Safety Center with the ob- jective of assessing and characterizing Reynolds number effects on the ascent aerodynamics of the SLS launch vehicle. The cryogenic test article was a 1.75%-scale representation of the SLS Block 1 Cargo configuration that was fabricated mostly out of stainless steel, but with several additive-manufactured components including the Solid Rocket Booster (SRB) attach brackets and the SRB nozzles. This was the first use of additive manufacturing on structural components of a cryogenic model in the NTF, therefore post-fabrication material testing and inspections were performed to satisfy safety requirements. Force & moment and surface pressure data were acquired for Mach numbers between 0.50 and 0.95 over a range of Reynolds numbers based on core stage diameter between 2x10^6 and a maximum of 40x10^6 corresponding to 45% of flight Reynolds number. Additionally, flow visualization data using Pressure Sensitive Paint were acquired with a focus on the SRB forward attach area, which is susceptible to Reynolds number sensitivity at transonic speeds. This pa- per provides an overview of the test campaign including details on the unique test article and the experimental setup and test execution. General test findings and observations are also presented, but the majority of the test results and data analyses are provided in a companion paper.

Space Launch System↗
NASA NTRSDate not supplied

Overview of the High Reynolds Number Ascent Wind Tunnel Test of the Space Launch System at the National Transonic Facility

Transonic, high Reynolds number wind tunnel testing for the Space Launch System (SLS) ascent flight environment was conducted in the National Transonic Facility (NTF) at the NASA Langley Research Center from December 2019 to April 2021. The test was sponsored by the SLS Program and the NASA Engineering & Safety Center with the ob- jective of assessing and characterizing Reynolds number effects on the ascent aerodynamics of the SLS launch vehicle. The cryogenic test article was a 1.75%-scale representation of the SLS Block 1 Cargo configuration that was fabricated mostly out of stainless steel, but with several additive-manufactured components including the Solid Rocket Booster (SRB) attach brackets and the SRB nozzles. This was the first use of additive manufacturing on structural components of a cryogenic model in the NTF, therefore post-fabrication material testing and inspections were performed to satisfy safety requirements. Force & moment and surface pressure data were acquired for Mach numbers between 0.50 and 0.95 over a range of Reynolds numbers based on core stage diameter between 2x10^6 and a maximum of 40x10^6 corresponding to 45% of flight Reynolds number. Additionally, flow visualization data using Pressure Sensitive Paint were acquired with a focus on the SRB forward attach area, which is susceptible to Reynolds number sensitivity at transonic speeds. This paper provides an overview of the test campaign including details on the unique test article and the experimental setup and test execution. General test findings and observations are also presented, but the majority of the test results and data analyses are provided in a companion paper.

Space Launch System↗
NASA NTRS1979

Design and wind tunnel tests of winglets on a DC-10 wing

Results are presented of a wind tunnel test utilizing a 4.7 percent scale semi-span model in the Langley Research Center 8-foot transonic pressure wind tunnel to establish the cruise drag improvement potential of winglets as applied to the DC-10 wide body transport aircraft. Winglets were investigated on both the DC-10 Series 10 (domestic) and 30/40 (intercontinental) configurations and compared with the Series 30/40 configuration. The results of the investigation confirm that for the DC-10 winglets provide approximately twice the cruise drag reduction of wing-tip extensions for about the same increase in bending moment at the wing fuselage juncture. Furthermore, the winglet configurations achieved drag improvements which were in close agreement to analytical estimates. It was observed that relatively small changes in wing-winglet tailoring effected large improvements in drag and visual flow characteristics. All final winglet configurations exhibited visual flow characteristics on the wing and winglets

Gilkey, R. D.↗
NASA NTRS1985

Pressure distribution from high Reynolds number tests of a NASA SC(3)-0712(B) airfoil in the Langley 0.3-meter transonic cryogenic tunnel

A wind tunnel investigation of a NASA 12-percent-thick, advanced-technology supercritical airfoil was conducted in the Langley 0.3-Meter Transonic Cryogenic Tunnel (TCT). This investigation represents another in the series of NASA/U.S. industry two-dimensional airfoil studies to be completed in the Advanced Technology Airfoil Tests program. Test temperature was varied from 220 K to 96 K at pressures ranging from 1.2 to 4.3 atm. Mach number was varied from 0.50 to 0.80. This investigation was designed to: (1) test a NASA advanced-technology airfoil from low to flight equivalent Reynolds numbers, (2) provide experience in cryogenic wind-tunnel model design and testing techniques, and (3) demonstrate the suitability of the 0.3-m TCT as an airfoil test facility. All the test objectives were met. The pressure data are presented without analysis in tabulated format and as plots of pressure coefficient versus position on the airfoil. This report was prepared for use in conjunction with the aerodynamic coefficient data published in NASA-TM-86371. Data are included which demonstrate the effects of fixed transition. Also included are remarks on the model design and fabrication.

Johnson, W. G., Jr.↗
NASA NTRS1985

High Reynolds number tests of a NASA SC(3)-0712(B) airfoil in the Langley 0.3-meter transonic cryogenic tunnel

A wind tunnel investigation of a NASA 12-percent-thick, advanced-technology supercritical airfoil was conducted in the Langley 0.3-Meter Transonic Cryogenic Tunnel (TCT). This investigation represents another in the series of NASA/U.S. industry two-dimensional airfoil studies to be completed in the Advanced Technology Airfoil Tests program. Test temperature was varied from 220 K to 96 K at pressures ranging from 1.2 to 4.3 atm. Mach number was varied from 0.60 to 0.80. These variables provided a Reynolds number range from 4,400,000 to 40,000,000 based on a 15.24-cm (6.0-in.) airfoil chord. This investigation was designed to test a NASA advanced-technology airfoil from low to flight-equivalent Reynolds numbers, provide experience in cryogenic wind tunnel model design and testing techniques, and demonstrate the suitability of the 0.3-m TCT as an airfoil test facility. The aerodynamic results are presented as integrated force and moment coefficients and pressure distributions. Data are included which demonstrate the effects of fixed transition, Mach number, and Reynolds number on the aerodynamic characteristics. Also included are remarks on the model design, the model structural integrity, and the overall test experience.

Johnson, W. G., Jr.↗
NASA NTRS2000

Active Control of Wind-Tunnel Model Aeroelastic Response Using Neural Networks

NASA Langley Research Center, Hampton, VA 23681 Under a joint research and development effort conducted by the National Aeronautics and Space Administration and The Boeing Company (formerly McDonnell Douglas) three neural-network based control systems were developed and tested. The control systems were experimentally evaluated using a transonic wind-tunnel model in the Langley Transonic Dynamics Tunnel. One system used a neural network to schedule flutter suppression control laws, another employed a neural network in a predictive control scheme, and the third employed a neural network in an inverse model control scheme. All three of these control schemes successfully suppressed flutter to or near the limits of the testing apparatus, and represent the first experimental applications of neural networks to flutter suppression. This paper will summarize the findings of this project.

Scott, Robert C.↗
NASA NTRSDate not supplied

On the Use of Liquid Nitrogen Droplets as Flow Tracers in Cryogenic Flow Facilities at NASA Langley Research Center

The injection of liquid nitrogen droplets to cool the gas temperature in cryogenic wind tunnels is discussed as a method of natural seeding for velocimetry-focused, particle-based, laser diagnostics. Historical observations and issues with seeding are presented for both ground-test facilities of interest in this work at NASA Langley: the 0.3-m Transonic Cryogenic Wind Tunnel (TCT) and the National Transonic Facility (NTF). Recent observations of natural seeding with a Rayleigh scattering instrument are presented in the two facilities, which motivated the purposeful use of a pulse-burst laser system to observe the particles directly with a sequentially operated laser sheet for the first time. Time-resolved image sequences of unevaporated liquid nitrogen droplets were readily acquired for tunnel total temperatures of 200 K and below. The preliminary results promote a discussion on the fitness of these natural particles as flow tracers for either a particle image velocimetry or a particle tracking velocimetry instrument in these high-Reynolds-number facilities.

cryogenic wind tunnels↗
NASA NTRS1987

Optical interferometry in fluid dynamics research

Optical interferometry techniques were applied to the investigation of transonic airfoil flow fields in large wind tunnels. Holographic interferometry techniques were used to study 2 dimensional symmetric NACA 64A010 and Douglas Aircraft Co. DSMA671 supercritical airfoil performance in the NASA Ames 2 x 2 ft transonic wind tunnel. Quantitative data obtained from the interferograms were compared to the surface pressure data. The agreement obtained verified the accuracy of the flow visualization and demonstrated the potential for acquiring quantitative scalar results. Measurements of the inviscid flow speed and the boundary layer and wake velocity profiles were extracted from the interferograms and compared to laser Doppler velocimeter measurements. These results were also in good agreement. A method for acquiring real time interferometric data in large scale facilities was developed. This method, based on the point diffraction interferometer, was successfully tested in the 2 x 2 ft transonic wind tunnel. The holographic and real time interferometry methods were applied to the investigations of circulation control airfoils utilizing the Coanda effect. These results reveals the details of the jet interacting with the trailing edge boundary layer and the other parameters affecting the lift augmentation.

Bachalo, W. D.↗
NASA NTRS1980

A theoretical analysis of simulated transonic boundary layers in cryogenic-nitrogen wind tunnels

A theoretical analysis was made to determine the real gas effects on simulation of transonic boundary layers in wind tunnels with cryogenic nitrogen as the test gas. The analysis included laminar and turbulent flat plate boundary layers and turbulent boundary layers on a two dimensional airfoil. The results indicate that boundary layers in such wind tunnels should not be substantially different from ideal gas boundary layers at standard conditions. At a pressure of 9.0 atm, two separate effects produce deviations of real gas values from ideal gas values which are in the opposite direction from deviations at 1.0 atm and are of the same insignificant order of magnitude. Results also show that nonadiabatic boundary layers should be adequately simulated if the enthalpy ratio is the correlating parameter rather than the temperature ratio.

Adcock, J. B.↗
NASA NTRS1971

Drag of a Supercritical Body of Revolution in Free Flight at Transonic Speeds and Comparison with Wind Tunnel Data

The forebody drag of a supercritical body of revolution was measured in free flight over a Mach number range of 0.85 to 1.05 and a Reynolds number range of 11.5 x 10 to the 6th power to 19.4 x 10 to the 6th power and was compared with wind-tunnel data. The forebody drag coefficient for a Mach number less than 0.96 was 0.111 compared with the wind-tunnel value of 0.103. A gradual increase in the drag occurred in the Langley 8-foot transonic pressure tunnel at a lower Mach number than in the Langley 16-foot transonic tunnel or in the free-flight test. The sharp drag rise occurred near Mach 0.98 in free flight whereas the rise occurred near Mach 0.99 in the Langley 16-foot transonic tunnel. The sharp rise was not as pronounced in the Langley 8-foot transonic pressure tunnel and was probably affected by tunnel-wall-interference effects. The increase occurred more slowly and at a higher Mach number. These results indicate that the drag measurements made in the wind tunnels near Mach 1 were significantly affected by the relative size of the model and the wind tunnel.

Usry, J. W.↗
NASA NTRS1974

Some tunnel-wall effects on transonic flutter

Significant effects of wind-tunnel walls were observed on the transonic flutter boundaries of wall-mounted models during two flutter model research studies. In these studies, flutter experiments with cantilevered SST-type wing models were conducted in three different wind tunnels. The experimental results are compared to flutter boundaries calculated for the models in free air. The results indicate that transonic flutter boundaries can be affected by tunnel-wall interference, tunnel resonances, and shock-wave reflections, and that flutter model data accuracy is a function of model/tunnel size and tunnel wall porosity. However, models within the recommended size limits should give accurate results in transonic tunnels with normal ventilation. A flutter trend analysis for a two-dimensional wing-demonstrating tunnel wall and resonance effects on flutter are also presented.

Ruhlin, C. L.↗