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All sourcesNASA NTRSDOE OSTI
NASA NTRS1981

The development of cryogenic wind tunnels and their application to maneuvering aircraft technology

The cryogenic wind tunnel and its potential for advancing maneuvering aircraft technology is discussed. A brief overview of the cryogenic wind tunnel concept and the capabilities and status of the Langley cryogenic facilities is given, as is a review of the considerations leading to the selection of the cryogenic concept such as capital and operating costs of the tunnel, model and balance construction implications, and test condition. Typical viscous, compressibility and aeroelastic effects encountered by maneuvering aircraft are illustrated and the unique ability of the cryogenic wind tunnels to isolate and investigate these parameters while simulating full scale conditions is discussed. The status of the Langley cryogenic wind tunnel facilities is reviewed and their operating envelopes described in relation to maneuvering aircraft research and development requirements. The status of cryogenic testing technology specifically related to aircraft maneuverability studies including force balances and buffet measurement techniques is discussed. Included are examples of research carried out in the Langley 0.3 meter transonic cryogenic wind tunnel to verify the various techniques.

Polhamus, E. C.↗
NASA NTRS1980

Real-gas effects 1: Simulation of ideal gas flow by cryogenic nitrogen and other selected gases

The thermodynamic properties of nitrogen gas do not thermodynamically approximate an ideal, diatomic gas at cryogenic temperatures. Choice of a suitable equation of state to model its behavior is discussed and the equation of Beattie and Bridgeman is selected as best meeting the needs for cryogenic wind tunnel use. The real gas behavior of nitrogen gas is compared to an ideal, diatomic gas for the following flow processes: isentropic expansion; normal shocks; boundary layers; and shock wave boundary layer interactions. The only differences in predicted pressure ratio between nitrogen and an ideal gas that may limit the minimum operating temperatures of transonic cryogenic wind tunnels seem to occur at total pressures approaching 9atmospheres and total temperatures 10 K below the corresponding saturation temperature, where the differences approach 1 percent for both isentropic expansions and normal shocks. Several alternative cryogenic test gases - air, helium, and hydrogen - are also analyzed. Differences in air from an ideal, diatomic gas are similar in magnitude to those of nitrogen. Differences for helium and hydrogen are over an order of magnitude greater than those for nitrogen or air. Helium and hydrogen do not approximate the compressible flow of an ideal, diatomic gas.

Hall, R. M.↗
NASA NTRS1986

Cryogenic tunnel measurement of total temperature and pressure

A newly developed, 3-mm-diam, dual hot-wire aspirating probe was used to measure the time-resolved stagnation temperature and pressure in a transonic cryogenic wind tunnel. Measurements were taken in the freestream of the settling chamber and test section. Data were also obtained in the unsteady wake shed from an airfoil oscillating at 5 Hz. The investigation revealed the presence of large fluctuations in the settling chamber occuring at the blade passing frequency of the driving fan of the tunnel. These fluctuations decrease at the test section. The rms value of the fluctuating stagnation pressure decreased from 17.5 percent in the settling chamber to 3.7 percent in the test section. Fluctuating stagnation temperature decreased from 12.3 percent to 8.4 percent. Measurements in the wake of the oscillating airfoil showed a fluctuating stagnation temperature of as much as 42 K in rms value.

Ng, W.-F.↗
NASA NTRS2000

Development of X-33/X-34 Aerothermodynamic Data Bases: Lessons Learned and Future Enhancements

A synoptic of programmatic and technical lessons learned in the development of aerothermodynamic data bases for the X-33 and X-34 programs is presented in general terms and from the perspective of the NASA Langley Research Center Aerothermodynamics Branch. The format used is that of the "aerothermodynamic chain," the links of which are personnel, facilities, models/test articles, instrumentation, test techniques, and computational fluid dynamics (CFD). Because the aerodynamic data bases upon which the X-33 and X-34 vehicles will fly are almost exclusively from wind tunnel testing, as opposed to CFD, the primary focus of the lessons learned is on ground-based testing. The period corresponding to the development of X-33 and X-34 aerothermodynamic data bases was challenging, since a number of other such programs (e.g., X-38, X-43) competed for resources at a time of downsizing of personnel, facilities, etc., outsourcing, and role changes as NASA Centers served as subcontractors to industry. The impact of this changing environment is embedded in the lessons learned. From a technical perspective, the relatively long times to design and fabricate metallic force and moment models, delays in delivery of models, and a lack of quality assurance to determine the fidelity of model outer mold lines (OML) prior to wind tunnel testing had a major negative impact on the programs. On the positive side, the application of phosphor thermography to obtain global, quantitative heating distributions on rapidly fabricated ceramic models revolutionized the aerothermodynamic optimization of vehicle OMLs, control surfaces, etc. Vehicle designers were provided with aeroheating information prior to, or in conjunction with, aerodynamic information early in the program, thereby allowing trades to be made with both sets of input; in the past only aerodynamic data were available as input. Programmatically, failure to include transonic aerodynamic wind tunnel tests early in the assessment phase led to delays in the optimization phase, as OMLs required modification to provide adequate transonic aerodynamic performance without sacrificing subsonic and hypersonic performance. Funding schedules for industry, based on technical milestones, also presented challenges to aerothermodynamics seeking optimum flying characteristics across the subsonic to hypersonic speed regimes and minimum aeroheating. This paper is concluded with a brief discussion of enhancements in ground-based testing/CFD capabilities necessary to partially/fully satisfy future requirements.

Miller, C. G.↗
NASA NTRS2002

Investigation of Reynolds Number Effects on a Generic Fighter Configuration in the National Transonic Facility

A wind tunnel test of a generic fighter configuration was tested in the National Transonic Facility through a cooperative agreement between NASA Langley Research Center and McDonnell Douglas. The primary purpose of the test was to assess Reynolds number scale effects on a thin-wing, fighter-type configuration up to full-scale flight conditions (that is, Reynolds numbers of the order of 60 million). The test included longitudinal and lateral/directional studies at subsonic and transonic conditions across a range of Reynolds numbers from that available in conventional wind tunnels to flight conditions. Results are presented for three Mach numbers (0.6, 0.8, and 0.9) and three configurations: (1) Fuselage/Wing; (2) Fuselage/Wing/Centerline Vertical Tail/Horizontal Tail; and (3) Fuselage/Wing/Trailing-Edge Extension/Twin Vertical Tails. Reynolds number effects on the longitudinal aerodynamic characteristics are presented herein.

Tomek, W. G.↗
NASA NTRS2017

Computational Results for the KTH-NASA Wind-Tunnel Model Used for Acquisition of Transonic Nonlinear Aeroelastic Data

A status report is provided on the collaboration between the Royal Institute of Technology (KTH) in Sweden and the NASA Langley Research Center regarding the aeroelastic analyses of a full-span fighter configuration wind-tunnel model. This wind-tunnel model was tested in the Transonic Dynamics Tunnel (TDT) in the summer of 2016. Large amounts of data were acquired including steady/unsteady pressures, accelerations, strains, and measured dynamic deformations. The aeroelastic analyses presented include linear aeroelastic analyses, CFD steady analyses, and analyses using CFD-based reduced-order models (ROMs).

Silva, Walter A.↗
NASA NTRS1983

The performance of a circulation control airfoil at transonic speeds

A two-dimensional wind tunnel test has been performed on a small circulation control airfoil section equipped with trailing edge blowing. The model was mounted in the NASA-Ames 2 x 2 ft transonic wind tunnel and tested at speeds up to free-stream Mach equals 0.75 for a range of incidences. Jet pressure ratios up to 3 (relative to tunnel static conditions) were evaluated together with the effects of Reynolds number. Normal force and pitching moment coefficients were calculated from surface pressures using a Scanivalve pressure measuring system. Drag force coefficients were calculated from wake rake pressures. The results obtained indicated that this airfoil was capable of producing useful lift at high subsonic Mach numbers. Some changes in the stall characteristics were apparent at above free-stream Mach equals 0.4 and some dependence between lift augmentation and incidence was observed. There also appeared to be a significant Reynolds number effect on the airfoil drag performance.

Wood, N. J.↗
NASA NTRS1981

Simulation of ideal-gas flow by nitrogen and other selected gases at cryogenic temperatures

The real gas behavior of nitrogen, the gas normally used in transonic cryogenic tunnels, is reported for the following flow processes: isentropic expansion, normal shocks, boundary layers, and interactions between shock waves and boundary layers. The only difference in predicted pressure ratio between nitrogen and an ideal gas which may limit the minimum operating temperature of transonic cryogenic wind tunnels occur at total pressures approaching 9 atm and total temperatures 10 K below the corresponding saturation temperature. These pressure differences approach 1 percent for both isentropic expansions and normal shocks. Alternative cryogenic test gases were also analyzed. Differences between air and an ideal diatomic gas are similar in magnitude to those for nitrogen and should present no difficulty. However, differences for helium and hydrogen are over an order of magnitude greater than those for nitrogen or air. It is concluded that helium and cryogenic hydrogen would not approximate the compressible flow of an ideal diatomic gas.

Hall, R. M.↗
NASA NTRS1985

The joined wing - An overview

The joined wing is a new type of aircraft configuration which employs tandem wings arranged to form diamond shapes in plan view and front view. Wind-tunnel tests and finite-element structural analyses have shown that the joined wing provides the following advantages over a comparable wing-plus-tail system; lighter weight and higher stiffness, higher span-efficiency factor, higher trimmed maximum lift coefficient, lower wave drag, plus built-in direct lift and direct sideforce control capability. A summary is given of research performed on the joined wing. Calculated joined wing weights are correlated with geometric parameters to provide simple weight estimation methods. The results of low-speed and transonic wind-tunnel tests are summarized, and guidelines for design of joined-wing aircraft are given. Some example joined-wing designs are presented and related configurations having connected wings are reviewed.

Wolkovitch, J.↗
NASA NTRS2017

Transonic Semispan Aerodynamic Testing of the Hybrid Wing Body with Over Wing Nacelles in the National Transonic Facility

A wind tunnel investigation of a 0.04-scale model of the Lockheed Martin Hybrid Wing Body (HWB) with Over Wing Nacelles (OWN) air mobility transport configuration was conducted in the National Transonic Facility at the NASA Langley Research Center under a collaborative partnership between NASA, the Air Force Research Laboratory, and Lockheed Martin Aeronautics Company. The wind tunnel test sought to validate the transonic aerodynamic performance of the HWB and to validate the efficiency benefits of the OWN installation as compared to the traditional under-wing installation. The semispan HWB model was tested in a clean wing configuration and also tested with two different nacelles representative of a modern turbofan engine and a future advanced high bypass ratio engine. The nacelles were installed in three different locations with two over-wing positions and one under-wing position. Five-component force and moment data, surface static pressure data, and aeroelastic deformation data were acquired. For the cruise configuration, the model was tested in an angle-of-attack range between -2 and 10 degrees at free-stream Mach numbers from 0.3 to 0.9 and at unit Reynolds numbers between 8 and 39 million per foot, achieving a maximum of 80% of flight Reynolds numbers across the Mach number range. The test results validated pretest computational fluid dynamic (CFD) simulations of the HWB performance including the OWN benefit and the results also exhibited excellent transonic drag data repeatability to within +/-1 drag count. This paper details the experimental setup and model overview, presents some sample data results, and describes the facility improvements that led to the success of the test.

Chan, David T.↗
NASA NTRS2025

Subsonic Ultra Green Aircraft Research: Phase V – Buffet Test Report

This test report summarizes work performed by the Boeing Subsonic Ultra-Green Aircraft Research (SUGAR) team in the Phase V Transonic Truss-Braced Wing contract task. The task was awarded in September 2020, and ended at the end of September 2022. As a part of the SUGAR Phase V contract task, Boeing developed a wind tunnel test plan, model requirements, then fabricated and conducted a dedicated transonic buffet wind tunnel test. The test focused on the investigation of buffet onset, and the ability to accurately predict buffet onset for the TTBW configuration. This report summarizes the results and post-test analysis of the 9% scale, semispan, Transonic Truss-Braced Wing (TTBW) transonic buffet model tested at the NASA Ames Research Center Unitary Plan Wind Tunnel (UPWT) 11-Foot Transonic Wind Tunnel (11-Ft TWT) facility located at Moffett Field, CA. It was tested from January 24, 2022, through February 18, 2022. The wind tunnel test was successful in capturing buffet onset at the model design dynamic pressure. Multiple configurations were tested starting in the wing-body-strut configuration and ending with the full configuration including wing-body-strut-nacelle/pylon, and flap hinge. The test also measured the effect of control surface deflections (ailerons and spoilers), as well as the effect of an outboard strut flap. Steady data collected during the test included standard force and moment measurements, as well as wing, strut, nacelle, and body (fuselage) static pressures. These data were collected in a standard pitch-pause mode. Unsteady data captured in the test, collected in continuous pitch mode, included wing and strut dynamic pressures, accelerations, and strains. According to test data, previous concerns regarding critical buffet onset on the strut (at low vehicle lift coefficient) and in the wing-strut channel have been mitigated. Buffet onset on the TTBW appears to follow conventional wing-induced buffet onset mechanisms. However, the critical station where buffet onset first occurs has moved inboard to the wing-strut juncture span location (55-60% span) from the typical ~70% span location. Test data obtained by unsteady pressure transducers provided a good match to expected buffet onset modes. An assessment of different tools and methodologies for predicting buffet onset was conducted. Reasonable agreement was found for buffet boundary predictions between most of the methods examined when using a forward trip location, with the greatest discrepancies from the pitching moment break and trailing edge pressure divergence methods. When an aft trip location was used the buffet boundary predictions diverged more significantly from unsteady pressure measurements. In addition, good agreement between the buffet onset trends were found when the Reynolds number, configuration, or spanload was changed.

strut↗
NASA NTRS1977

Self streamlining wind tunnel: Further low speed testing and final design studies for the transonic facility

Work has continued with the low speed self streamlining wind tunnel (SSWT) using the NACA 0012-64 airfoil in an effort to explain the discrepancies between the NASA Langley low turbulence pressure tunnel (LTPT) and SSWT results obtained with the airfoil stalled. Conventional wind tunnel corrections were applied to straight wall SSWT airfoil data, to illustrate the inadequacy of standard correction techniques in circumstances of high blockage. Also one SSWT test was re-run at different air speeds to investigate the effects of such changes on airfoil data and wall contours. Mechanical design analyses for the transonic self streamlining wind tunnel (TSWT) were completed by the application of theoretical airfoil flow field data to the elastic beam and streamline analysis. The control system for the transonic facility is outlined.

Wolf, S. W. D.↗
NASA NTRS1985

Assessment of wind tunnel corrections for multielement airfoils at transonic speeds

A finite volume formulation of the Euler equations using Cartesian grids is used to calculate the transonic flow over multielement airfoils and to use the resulting solutions to assess wall interference effects in wind tunnels. Available methods and recommendations for evaluating such effects, which are based on shifts in Mach number and angle of attack, are examined and the results are compared with measurements using the flapped supercritical SKF 1.1 airfoil. Based on the calculations, it is concluded that shifts in Mach number and angle of attack cannot by themselves account for viscous and wall effects on multielement airfoils at transonic speeds.

Gaffney, R. L., Jr.↗
NASA NTRS1978

Condensation and its growth down the test-section of the Langley 0.3-m transonic cryogenic tunnel

Four total pressure probes were used to measure the growth of condensation down the test section of the Langley 0.3-m tunnel, and the condensation data were employed to verify a mathematical model which assumes condensation results from heterogeneous nucleation on preexisting seed particles. The onset of effects occurs throughout the test section at the same total temperature but the magnitude of the effects increases with increasing length down the test section. Condensation is important because it determines the minimum operating temperature of transonic cryogenic wind tunnels.

Hall, R. M.↗
NASA NTRS1984

A flying superconducting magnet and cryostat for magnetic suspension of wind-tunnel models

The engineering practicality of a persistent high-field superconducting solenoid cryostat as a magnetic suspension and balance system (MSBS) for wind-tunnel testing of aircraft and missile models is examined. The test apparatus is a simple solenoid of filamentary NbTi superconductor with a cupronickel matrix. The apparatus, with a length-to-diameter ratio of 6 to 1 and a radius of 32 mm, used a 0.25 mm wire with a critical current of 27 A in an external field of 6 T. The total heat inleak of 150 mW was achieved. Helium boiloff rates were tested over a range of operating conditions, including pitch attitudes from 10 deg nose down to 90 deg nose up; the rate was estimated as low, but the aerodynamic acceptability of venting gaseous helium has not been determined. It is shown that the effectiveness of the concept increases with increasing scale, and performance in excess of that of conventional ferromagnets is achievable with reduction in size and costs, and with aptness to transonic wind-tunnel testing. Detailed specifications and schematics are included.

Britcher, C.↗
NASA NTRSDate not supplied

Comparison of Surface Pressure Fluctuations from Flight and Wind Tunnel Tests on the Orion Multi-Purpose Crew Vehicle

Transonic and supersonic wind-tunnel test of scaled-models are used to predict the surface pressure fluctuations on aerospace vehicles. An outstanding question is how good are such tests? The present paper attempts to answer this question for the Multi-Purpose Crew Vehicle by comparing data from the Ascent Abort-2 (AA2) flight test with those measured in two different scaled-model wind tunnel tests (WTT) 51AS and 134AS. A sensor by sensor and a Mach by Mach comparisons of spectra of surface pressure fluctuations are presented for nominal ascent part of flight. Spectra measured over a large part of the vehicle surface were found to be dominated by the wake vortices shed from the four Abort Motor nozzles situated on the upper part of the Launch Abort System. It was found that the shapes of the scaled-up spectra from WTTs were very close to those measured in the flight, but the levels, for the most part, were found to be lower than the flight data. The highest differences were from the regions of complex flows where wakes from the adjacent nozzles interacted with the free-stream flow. There the WTT predicted levels were consistently 2dB to 6dB lower than the flight data, across all spectral bands. Another region of significant under-prediction was the local separated flow region at the Fillet-Ogive junction. The accuracy of the sharp spectral peaks from regions along a nozzle axis was found to be dependent on the fidelity of the model. The model in the 51AS WTT used a simplified contour, which resulted in a large under prediction of the spectral peak by 1dB to 10dB. A limited number of data from 134AS, where the model accurately captured all features of the nozzles of the flight vehicle, however; showed that the under prediction was lower: 1-2dB. The fluctuation spectra on the large protuberance of the Umbilical Cover was found to be reasonably well-predicted by 134AS WTT, which meticulously reproduced a scaled-down shape of this protuberance. Such information will help to determine the margins to be applied to wind-tunnel data to create flight environments.

MPCV↗
DOE OSTI2020

Aerodynamic Data Fusion Toward the Digital Twin Paradigm

This paper considers the fusion of two aerodynamic data sets originating from differing types of physical or computer experiments. This paper specifically addresses the fusion of 1) noisy and in-complete fields from wind-tunnel measurements and 2) deterministic but biased fields from numerical simulations. These two data sources are fused in order to estimate the true field that best matches measured quantities that serve as the ground truth. For example, two sources of pressure fields about an aircraft are fused based on measured forces and moments from a wind-tunnel experiment. A fundamental challenge in this problem is that the true field is unknown and cannot be estimated with 100% certainty. A Bayesian framework is employed to infer the true fields conditioned on measured quantities of interest; essentially a statistical correction to the data is performed. The fused data may then be used to construct more accurate surrogate models suitable for early stages of aerospace design. An extension of the proper orthogonal decomposition with constraints is also introduced to solve the same problem. In this work, both methods are demonstrated on fusing the pressure distributions for flow past the RAE2822 airfoil and the Common Research Model wing at transonic conditions. Comparison of both methods reveals that the Bayesian method is more robust when data are scarce and capable of also accounting for uncertainties in the data. Furthermore, given adequate data, the proper-orthogonal-decomposition-based and Bayesian approaches lead to surprisingly similar results.

42 ENGINEERING↗
NASA NTRS1986

Transonic Navier-Stokes wing solution using a zonal approach. Part 1: Solution methodology and code validation

A fast diagonalized Beam-Warming algorithm is coupled with a zonal approach to solve the three-dimensional Euler/Navier-Stokes equations. The computer code, called Transonic Navier-Stokes (TNS), uses a total of four zones for wing configurations (or can be extended to complete aircraft configurations by adding zones). In the inner blocks near the wing surface, the thin-layer Navier-Stokes equations are solved, while in the outer two blocks the Euler equations are solved. The diagonal algorithm yields a speedup of as much as a factor of 40 over the original algorithm/zonal method code. The TNS code, in addition, has the capability to model wind tunnel walls. Transonic viscous solutions are obtained on a 150,000-point mesh for a NACA 0012 wing. A three-order-of-magnitude drop in the L2-norm of the residual requires approximately 500 iterations, which takes about 45 min of CPU time on a Cray-XMP processor. Simulations are also conducted for a different geometrical wing called WING C. All cases show good agreement with experimental data.

Flores, J.↗