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Sample records for 11-ft transonic wind

  1. Computational Modeling of the Ames 11-Ft Transonic Wind Tunnel in Conjunction with IofNEWT

    NASA Technical Reports Server (NTRS)

    Djomehri, M. Jahed; Buning, Pieter G.; Erickson, Larry L.; George, Michael W. (Technical Monitor)

    1995-01-01

    Technical advances in Computational Fluid Dynamics have now made it possible to simulate complex three-dimensional internal flows about models of various size placed in a Transonic Wind Tunnel. TWT wall interference effects have been a source of error in predicting flight data from actual wind tunnel measured data. An advantage of such internal CFD calculations is to directly compare numerical results with the actual tunnel data for code assessment and tunnel flow analysis. A CFD capability has recently been devised for flow analysis of the NASA/Ames 11-Ft TWT facility. The primary objectives of this work are to provide a CFD tool to study the NASA/Ames 11-Ft TWT flow characteristics, to understand the slotted wall interference effects, and to validate CFD codes. A secondary objective is to integrate the internal flowfield calculations with the Pressure Sensitive Paint data, a surface pressure distribution capability in Ames' production wind tunnels. The effort has been part of the Ames IofNEWT, Integration of Numerical and Experimental Wind Tunnels project, which is aimed at providing further analytical tools for industrial application. We used the NASA/Ames OVERFLOW code to solve the thin-layer Navier-Stokes equations. Viscosity effects near the model are captured by Baldwin-Lomax or Baldwin-Barth turbulence models. The solver was modified to model the flow behavior in the vicinity of the tunnel longitudinal slotted walls. A suitable porous type wall boundary condition was coded to account for the cross-flow through the test section. Viscous flow equations were solved in generalized coordinates with a three-factor implicit central difference scheme in conjunction with the Chimera grid procedure. The internal flow field about the model and the tunnel walls were descretized by the Chimera overset grid system. This approach allows the application of efficient grid generation codes about individual components of the configuration; separate minor grids were developed

  2. Pressure-Sensitive Paint Measurements on the NASA Common Research Model in the NASA 11-ft Transonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Bell, James H.

    2011-01-01

    The luminescence lifetime technique was used to make pressure-sensitive paint (PSP) measurements on a 2.7% Common Research Model in the NASA Ames 11ft Transonic Wind Tunnel. PSP data were obtained on the upper and lower surfaces of the wing and horizontal tail, as well as one side of the fuselage. Data were taken for several model attitudes of interest at Mach numbers between 0.70 and 0.87. Image data were mapped onto a three-dimensional surface grid suitable both for comparison with CFD and for integration of pressures to determine loads. Luminescence lifetime measurements were made using strobed LED (light-emitting diode) lamps to illuminate the PSP and fast-framing interline transfer cameras to acquire the PSP emission.

  3. Experimental Investigations of the NASA Common Research Model in the NASA Langley National Transonic Facility and NASA Ames 11-Ft Transonic Wind Tunnel (Invited)

    NASA Technical Reports Server (NTRS)

    Rivers, S. M.; Dittberner, Ashley

    2011-01-01

    Experimental aerodynamic investigations of the NASA Common Research Model have been conducted in the NASA Langley National Transonic Facility and the NASA Ames 11-ft wind tunnel. Data have been obtained at chord Reynolds numbers of 5 million for five different configurations at both wind tunnels. Force and moment, surface pressure and surface flow visualization data were obtained in both facilities but only the force and moment data are presented herein. Nacelle/pylon, tail effects and tunnel to tunnel variations have been assessed. The data from both wind tunnels show that an addition of a nacelle/pylon gave an increase in drag, decrease in lift and a less nose down pitching moment around the design lift condition of 0.5 and that the tail effects also follow the expected trends. Also, all of the data shown fall within the 2-sigma limits for repeatability. The tunnel to tunnel differences are negligible for lift and pitching moment, while the drag shows a difference of less than ten counts for all of the configurations. These differences in drag may be due to the variation in the sting mounting systems at the two tunnels.

  4. Comparison of the NASA Common Research Model European Transonic Wind Tunnel Test Data to NASA Test Data

    NASA Technical Reports Server (NTRS)

    Rivers, Melissa B.; Quest, Jurgen; Rudnik, Ralf

    2015-01-01

    Experimental aerodynamic investigations of the NASA Common Research Model have been conducted in the NASA Langley National Transonic Facility, the NASA Ames 11-ft wind tunnel, and the European Transonic Wind Tunnel. In the NASA Ames 11-ft wind tunnel, data have been obtained at only a chord Reynolds number of 5 million for a wing/body/tail = 0 degree incidence configuration. Data have been obtained at chord Reynolds numbers of 5, 19.8 and 30 million for the same configuration in the National Transonic Facility and in the European Transonic Facility. Force and moment, surface pressure, wing bending and twist, and surface flow visualization data were obtained in all three facilities but only the force and moment, surface pressure and wing bending and twist data are presented herein.

  5. Comparison of the NASA Common Research Model European Transonic Wind Tunnel Test Data to NASA Test Data

    NASA Technical Reports Server (NTRS)

    Rivers, Melissa; Quest, Juergen; Rudnik, Ralf

    2015-01-01

    Experimental aerodynamic investigations of the NASA Common Research Model have been conducted in the NASA Langley National Transonic Facility, the NASA Ames 11-ft wind tunnel, and the European Transonic Wind Tunnel. In the NASA Ames 11-ft wind tunnel, data have been obtained at only a chord Reynolds number of 5 million for a wing/body/tail = 0 degree incidence configuration. Data have been obtained at chord Reynolds numbers of 5, 19.8 and 30 million for the same configuration in the National Transonic Facility and in the European Transonic Facility. Force and moment, surface pressure, wing bending and twist, and surface flow visualization data were obtained in all three facilities but only the force and moment and surface pressure data are presented herein.

  6. Calibration of transonic and supersonic wind tunnels

    NASA Technical Reports Server (NTRS)

    Reed, T. D.; Pope, T. C.; Cooksey, J. M.

    1977-01-01

    State-of-the art instrumentation and procedures for calibrating transonic (0.6 less than M less than 1.4) and supersonic (M less than or equal to 3.5) wind tunnels were reviewed and evaluated. Major emphasis was given to transonic tunnels. Continuous, blowdown and intermittent tunnels were considered. The required measurements of pressure, temperature, flow angularity, noise and humidity were discussed, and the effects of measurement uncertainties were summarized. A comprehensive review of instrumentation currently used to calibrate empty tunnel flow conditions was included. The recent results of relevant research are noted and recommendations for achieving improved data accuracy are made where appropriate. It is concluded, for general testing purposes, that satisfactory calibration measurements can be achieved in both transonic and supersonic tunnels. The goal of calibrating transonic tunnels to within 0.001 in centerline Mach number appears to be feasible with existing instrumentation, provided correct calibration procedures are carefully followed. A comparable accuracy can be achieved off-centerline with carefully designed, conventional probes, except near Mach 1. In the range 0.95 less than M less than 1.05, the laser Doppler velocimeter appears to offer the most promise for improved calibration accuracy off-centerline.

  7. 4. VIEW NORTHWEST OF SUPERSONIC WIND TUNNEL BUILDING TO TRANSONIC ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    4. VIEW NORTHWEST OF SUPERSONIC WIND TUNNEL BUILDING TO TRANSONIC WIND TUNNEL BUILDING - Naval Surface Warfare Center, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD

  8. 1. VIEW SOUTHWEST OF SUBSONIC WIND TUNNEL BUILDING AND TRANSONIC ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    1. VIEW SOUTHWEST OF SUBSONIC WIND TUNNEL BUILDING AND TRANSONIC WIND TUNNEL BUILDING - Naval Surface Warfare Center, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD

  9. 5. VIEW NORTHWEST OF SUBSONIC WIND TUNNEL BUILDING TO TRANSONIC ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    5. VIEW NORTHWEST OF SUBSONIC WIND TUNNEL BUILDING TO TRANSONIC WIND TUNNEL BUILDING - Naval Surface Warfare Center, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD

  10. 3. VIEW SOUTHEAST OF TRANSONIC WIND TUNNEL BUILDING TO SUBSONIC ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    3. VIEW SOUTHEAST OF TRANSONIC WIND TUNNEL BUILDING TO SUBSONIC WIND TUNNEL BUILDING - Naval Surface Warfare Center, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD

  11. 2. VIEW SOUTH OF TRANSONIC WIND TUNNEL BUILDING AND SUPERSONIC ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    2. VIEW SOUTH OF TRANSONIC WIND TUNNEL BUILDING AND SUPERSONIC WIND TUNNEL BUILDING - Naval Surface Warfare Center, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD

  12. 7. VIEW NORTHWEST OF SUBSONIC WIND TUNNEL BUILDING TO TRANSONIC ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    7. VIEW NORTHWEST OF SUBSONIC WIND TUNNEL BUILDING TO TRANSONIC WIND TUNNEL BUILDING - Naval Surface Warfare Center, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD

  13. An experimental investigation of boundary layer and crossflow characteristics of the Ames 2 by 2 foot and 11 by 11 foot transonic wind-tunnel walls

    NASA Technical Reports Server (NTRS)

    Matyk, G.; Kobayashi, Y.

    1977-01-01

    The boundary layer and crossflow characteristics of 2- by 2-foot and 11- by 11-foot transonic wind-tunnel wall configurations have been studied for Mach numbers ranging from 0.5 to 1.2 and for various crossflow to free stream unit mass flow ratios. For the 2- by 2-ft and 11- by 11-ft wall configurations, these ratios ranged from 0 to 0.12 and from 0 to 0.07, respectively. Most notably, for both wall configurations, the pressure-drop coefficient across the wall was nonlinear with mass flow and invariant with Mach number.

  14. Emerging technology for transonic wind-tunnel-wall interference assessment and corrections

    NASA Technical Reports Server (NTRS)

    Newman, P. A.; Kemp, W. B., Jr.; Garriz, J. A.

    1988-01-01

    Several nonlinear transonic codes and a panel method code for wind tunnel/wall interference assessment and correction (WIAC) studies are reviewed. Contrasts between two- and three-dimensional transonic testing factors which affect WIAC procedures are illustrated with airfoil data from the NASA/Langley 0.3-meter transonic cyrogenic tunnel and Pathfinder I data. Also, three-dimensional transonic WIAC results for Mach number and angle-of-attack corrections to data from a relatively large 20 deg swept semispan wing in the solid wall NASA/Ames high Reynolds number Channel I are verified by three-dimensional thin-layer Navier-Stokes free-air solutions.

  15. The status of analytical preparation for 2-dimensional testing at high transonic speeds in the University of Southampton transonic self-streamlining wind tunnel

    NASA Technical Reports Server (NTRS)

    Lewis, M. C.

    1984-01-01

    Validation data from the Transonic Self-Streamlining Wind Tunnel has proved the feasibility of streamlining two dimensional flexible walls at low speeds and up to transonic speeds, the upper limit being the speed where the flexible walls are just supercritical. At this condition, breakdown of the wall setting strategy is evident in that convergence is neither as rapid nor as stable as for lower speeds, and wall streamlining criteria are not always completely satisfied. The only major step necessary to permit the extension of two dimensional testing into higher transonic speeds is the provision of a rapid algorithm to solve for mixed flow in the imagery flow fields. The status of two dimensional high transonic testing in the Transonic Self-Streamlining Wind Tunnel is outlined and, in particular, the progress of adapting an algorithm, which solves the Transonic Small Perturbation Equation, for predicting the imagery flow fields is detailed.

  16. Transonic wind tunnel test of a supersonic nozzle installation

    NASA Technical Reports Server (NTRS)

    Yetter, J. A.; Evelyn, G. B.; Mercer, C.

    1982-01-01

    The design of the propulsion system installation affects strongly the total drag and overall performance of an aircraft, and the concept, placement, and integration details of the exhaust nozzle are major considerations in the configuration definition. As part of the NASA Supersonic Cruise Research (SCR) program, a wind tunnel test program has been conducted to investigate exhaust nozzle-airframe interactions at transonic speeds. First phase testing is to establish guidelines for follow-on testing. A summary is provided of the results of first phase testing, taking into account the test approach, the effect of nozzle closure on aircraft aerodynamic characteristics, nozzle installation effects and nacelle interference drag, and an analytical study of the effects of nozzle closure on the aircraft.

  17. A survey of the three-dimensional high Reynolds number transonic wind tunnel

    NASA Technical Reports Server (NTRS)

    Takashima, K.; Sawada, H.; Aoki, T.

    1982-01-01

    The facilities for aerodynamic testing of airplane models at transonic speeds and high Reynolds numbers are surveyed. The need for high Reynolds number testing is reviewed, using some experimental results. Some approaches to high Reynolds number testing such as the cryogenic wind tunnel, the induction driven wind tunnel, the Ludwieg tube, the Evans clean tunnel and the hydraulic driven wind tunnel are described. The level of development of high Reynolds number testing facilities in Japan is discussed.

  18. The Use of Heavy Gas for Increased Reynolds Numbers in Transonic Wind Tunnels

    NASA Technical Reports Server (NTRS)

    Anders, J. B.; Anderson, W. K.; Murthy, A. V.

    1998-01-01

    The use of a high molecular weight test gas to increase the Reynolds number range of transonic wind tunnels is explored. Modifications to a small transonic wind tunnel are described and the real gas properties of the example heavy gas (sulfur hexafluoride) are discussed. Sulfur hexafluoride is shown to increase the test Reynolds number by a factor of more than 2 over air at the same Mach number. Experimental and computational pressure distributions on an advanced supercritical airfoil configuration at Mach 0.7 in both sulfur hexafluoride and nitrogen are presented. Transonic similarity theory is shown to be partially successful in transforming the heavy gas results to equivalent nitrogen (air) results, provided the correct definition of gamma is used.

  19. Static Aeroelastic Analysis of Transonic Wind Tunnel Models Using Finite Element Methods

    NASA Technical Reports Server (NTRS)

    Hooker, John R.; Burner, Alpheus W.; Valla, Robert

    1997-01-01

    A computational method for accurately predicting the static aeroelastic deformations of typical transonic transport wind tunnel models is described. The method utilizes a finite element method (FEM) for predicting the deformations. Extensive calibration/validation of this method was carried out using a novel wind-off wind tunnel model static loading experiment and wind-on optical wing twist measurements obtained during a recent wind tunnel test in the National Transonic Facility (NTF) at NASA LaRC. Further validations were carried out using a Navier-Stokes computational fluid dynamics (CFD) flow solver to calculate wing pressure distributions about several aeroelastically deformed wings and comparing these predictions with NTF experimental data. Results from this aeroelastic deformation method are in good overall agreement with experimentally measured values. Including the predicted deformations significantly improves the correlation between CFD predicted and experimentally measured wing & pressures.

  20. Modeling and control of transonic cryogenic wind tunnel: A summary report

    NASA Technical Reports Server (NTRS)

    Balakrishna, S.

    1980-01-01

    A representative, control compatible, mathematical model of a cryogenic wind tunnel was synthesized which describes the unsteady behavior of the tunnel. The performance of this model was validated by appropriate transient and quasisteady state tests on the 0.3 m transonic cryogenic tunnel. An understanding was developed of the cryogenic tunnel behavior by an appropriate interactive real time simulation of the developed model.

  1. User guide for WIACX: A transonic wind-tunnel wall interference assessment and correction procedure for the NTF

    NASA Technical Reports Server (NTRS)

    Garriz, Javier A.; Haigler, Kara J.

    1992-01-01

    A three dimensional transonic Wind-tunnel Interference Assessment and Correction (WIAC) procedure developed specifically for use in the National Transonic Facility (NTF) at NASA Langley Research Center is discussed. This report is a user manual for the codes comprising the correction procedure. It also includes listings of sample procedures and input files for running a sample case and plotting the results.

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

    NASA Technical Reports Server (NTRS)

    Wolf, S. W. D.

    1978-01-01

    Work was 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 (perhaps through changes in Reynold's number and freestream turbulence levels) 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, which will eventually allow on-line computer operation of the wind tunnel, was outlined.

  3. TWINTAN: A program for transonic wall interference assessment in two-dimensional wind tunnels

    NASA Technical Reports Server (NTRS)

    Kemp, W. B., Jr.

    1980-01-01

    A method for assessing the wall interference in transonic two dimensional wind tunnel test was developed and implemented in a computer program. The method involves three successive solutions of the transonic small disturbance potential equation to define the wind tunnel flow, the perturbation attriburable to the model, and the equivalent free air flow around the model. Input includes pressure distributions on the model and along the top and bottom tunnel walls which are used as boundary conditions for the wind tunnel flow. The wall induced perturbation fields is determined as the difference between the perturbation in the tunnel flow solution and the perturbation attributable to the model. The methodology used in the program is described and detailed descriptions of the computer program input and output are presented. Input and output for a sample case are given.

  4. Check-Standard Testing Across Multiple Transonic Wind Tunnels with the Modern Design of Experiments

    NASA Technical Reports Server (NTRS)

    Deloach, Richard

    2012-01-01

    This paper reports the result of an analysis of wind tunnel data acquired in support of the Facility Analysis Verification & Operational Reliability (FAVOR) project. The analysis uses methods referred to collectively at Langley Research Center as the Modern Design of Experiments (MDOE). These methods quantify the total variance in a sample of wind tunnel data and partition it into explained and unexplained components. The unexplained component is further partitioned in random and systematic components. This analysis was performed on data acquired in similar wind tunnel tests executed in four different U.S. transonic facilities. The measurement environment of each facility was quantified and compared.

  5. Experimental/analytical approach to understanding mistuning in a transonic wind tunnel compressor

    NASA Technical Reports Server (NTRS)

    Kaiser, Teri; Hansen, Reed S.; Nguyen, Nhan; Hampton, Roy W.; Muzzio, Doug; Chargin, Mladen K.; Guist, Roy; Hamm, Ken; Walker, Len

    1994-01-01

    This paper will briefly set forth some of the basic tenets of mistuned rotating bladed-disk assemblies. The experience with an existing three stage compressor in a transonic wind tunnel will be documented. The manner in which the theoretical properties manifest themselves in this non-ideal compressor will be described. A description of mistuning behaviors that can and cannot be accurately substantiated will be discussed.

  6. Within-Tunnel Variations in Pressure Data for Three Transonic Wind Tunnels

    NASA Technical Reports Server (NTRS)

    DeLoach, Richard

    2014-01-01

    This paper compares the results of pressure measurements made on the same test article with the same test matrix in three transonic wind tunnels. A comparison is presented of the unexplained variance associated with polar replicates acquired in each tunnel. The impact of a significance component of systematic (not random) unexplained variance is reviewed, and the results of analyses of variance are presented to assess the degree of significant systematic error in these representative wind tunnel tests. Total uncertainty estimates are reported for 140 samples of pressure data, quantifying the effects of within-polar random errors and between-polar systematic bias errors.

  7. Operating manual holographic interferometry system for 2 x 2 foot transonic wind tunnel

    NASA Technical Reports Server (NTRS)

    Craig, J. E.

    1981-01-01

    A holographic interferometer system was installed in a 2X2 foot transonic wind tunnel. The system incorporates a modern, 10 pps, Nd:YAG pulsed laser which provides reliable operation and is easy to align. The spatial filtering requirements of the unstable resonator beam are described as well as the integration of the system into the existing Schieren system. A two plate holographic interferometer is used to reconstruct flow field data. For static wind tunnel models the single exposure holograms are recorded in the usual manner; however, for dynamic models such as oscillating airfoils, synchronous laser hologram recording is used.

  8. Computational design of low aspect ratio wing-winglet configurations for transonic wind-tunnel tests

    NASA Technical Reports Server (NTRS)

    Kuhlman, John M.; Brown, Christopher K.

    1989-01-01

    Computational designs were performed for three different low aspect ratio wing planforms fitted with nonplanar winglets; one of the three configurations was selected to be constructed as a wind tunnel model for testing in the NASA LaRC 8-foot transonic pressure tunnel. A design point of M = 0.8, C(sub L) is approximate or = to 0.3 was selected, for wings of aspect ratio equal to 2.2, and leading edge sweep angles of 45 deg and 50 deg. Winglet length is 15 percent of the wing semispan, with a cant angle of 15 deg, and a leading edge sweep of 50 deg. Winglet total area equals 2.25 percent of the wing reference area. The design process and the predicted transonic performance are summarized for each configuration. In addition, a companion low-speed design study was conducted, using one of the transonic design wing-winglet planforms but with different camber and thickness distributions. A low-speed wind tunnel model was constructed to match this low-speed design geometry, and force coefficient data were obtained for the model at speeds of 100 to 150 ft/sec. Measured drag coefficient reductions were of the same order of magnitude as those predicted by numerical subsonic performance predictions.

  9. Optimization of transonic wind tunnel data acquisition and control systems for providing continuous mode tests

    NASA Astrophysics Data System (ADS)

    Petronevich, V. V.

    2016-10-01

    The paper observes the issues related to the increase of efficiency and information content of experimental research in transonic wind tunnels (WT). In particular, questions of optimizing the WT Data Acquisition and Control Systems (DACS) to provide the continuous mode test method are discussed. The problem of Mach number (M number) stabilization in the test section of the large transonic compressor-type wind tunnels at subsonic flow conditions with continuous change of the aircraft model angle of attack is observed on the example of T-128 wind tunnel. To minimize the signals distortion in T-128 DACS measurement channels the optimal MGCplus filter settings of the data acquisition system used in T-128 wind tunnel to measure loads were experimentally determined. As a result of the tests performed a good agreement of the results of balance measurements for pitch/pause and continuous test modes was obtained. Carrying out balance tests for pitch/pause and continuous test methods was provided by the regular data acquisition and control system of T-128 wind tunnel with unified software package POTOK. The architecture and functional abilities of POTOK software package are observed.

  10. Experimental parametric studies of transonic T-tail flutter. [wind tunnel tests

    NASA Technical Reports Server (NTRS)

    Ruhlin, C. L.; Sandford, M. C.

    1975-01-01

    Wind-tunnel tests of the T-tail of a wide-body jet airplane were made at Mach numbers up to 1.02. The model consisted of a 1/13-size scaled version of the T-tail, fuselage, and inboard wing of the airplane. Two interchangeable T-tails were tested, one with design stiffness for flutter-clearance studies and one with reduced stiffness for flutter-trend studies. Transonic antisymmetric-flutter boundaries were determined for the models with variations in: (1) fin-spar stiffness, (2) stabilizer dihedral angle (-5 deg and 0 deg), (3) wing and forward-fuselage shape, and (4) nose shape of the fin-stabilizer juncture. A transonic symmetric-flutter boundary and flutter trends were established for variations in stabilizer pitch stiffness. Photographs of the test configurations are shown.

  11. Optimization of wave cancellation in variable porosity transonic wind tunnel flows

    NASA Technical Reports Server (NTRS)

    Davis, J. W.

    1973-01-01

    A technique has been developed which is capable of determining the optimum wall configuration for a variable porosity perforated wall transonic wind tunnel. The technique is based on a mathematical model arrived at by considering the results of theory and past experimental investigations. A performance index was determined as a function of the significant wind tunnel parameters by comparing a formulation of this mathematical model, using MSFC 14 inch Trisonic Wind Tunnel experimental results, to interference free results. The resulting relationship was then used to determine the combination of wind tunnel parameters which should yield minimum reflected wave interference. A theoretical development of wall porosity requirements for thick wall inclined hole test sections is included which follows the trends and generally the magnitude of available experimental data. This theory is useful in studying the present variable porosity case, but also should be of value in studies concerning the wave cancellation process for fixed porosity walls.

  12. Comparison of the Aerodynamic Characteristics of Similar Models in Two Size Wind Tunnels at Transonic Speeds

    NASA Technical Reports Server (NTRS)

    Springer, Anthony M.

    1998-01-01

    The aerodynamic characteristics of two similar models of a lifting body configuration were run in two transonic wind tunnels, one a 16 foot the other a 14-inch and are compared. The 16 foot test used a 2% model while the 14-inch test used a 0.7% scale model. The wind tunnel model configurations varied only in vertical tail size and an aft sting shroud. The results from these two tests compare the effect of tunnel size, Reynolds number, dynamic pressure and blockage on the longitudinal aerodynamic characteristics of the vehicle. The data accuracy and uncertainty are also presented. It was concluded from these tests that the data resultant from a small wind tunnel compares very well to that of a much larger wind tunnel in relation to total vehicle aerodynamic characteristics.

  13. The status of two-dimensional testing at high transonic speeds in the University of Southampton transonic self-streamlining wind tunnel

    NASA Technical Reports Server (NTRS)

    Lewis, M. C.

    1985-01-01

    This report briefly outlines the progress made during the last 2 years in extending the operational range of the Transonic Self-Streamlining Wind Tunnel (at the University of Southampton) into high subsonic speeds. Analytical preparation completed in order to achieve such an extension is outlined and a summary of the preliminary model validation tests is presented. Future work necessary to allow further validation and development is discussed.

  14. Modernization and Activation of the NASA Ames 11- by 11-Foot Transonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Kmak, Frank J.

    2000-01-01

    The Unitary Plan Wind Tunnel (UPWT) was modernized to improve performance, capability, productivity, and reliability. Automation systems were installed in all three UPWT tunnel legs and the Auxiliaries facility. Major improvements were made to the four control rooms, model support systems, main drive motors, and main drive speed control. Pressure vessel repairs and refurbishment to the electrical distribution system were also completed. Significant changes were made to improve test section flow quality in the 11-by 11-Foot Transonic leg. After the completion of the construction phase of the project, acceptance and checkout testing was performed to demonstrate the capabilities of the modernized facility. A pneumatic test of the tunnel circuit was performed to verify the structural integrity of the pressure vessel before wind-on operations. Test section turbulence, flow angularity, and acoustic parameters were measured throughout the tunnel envelope to determine the effects of the tunnel flow quality improvements. The new control system processes were thoroughly checked during wind-off and wind-on operations. Manual subsystem modes and automated supervisory modes of tunnel operation were validated. The aerodynamic and structural performance of both the new composite compressor rotor blades and the old aluminum rotor blades was measured. The entire subsonic and supersonic envelope of the 11-by 11-Foot Transonic leg was defined up to the maximum total pressure.

  15. Empty test section streamlining of the transonic self-streamlining wind tunnel fitted with new walls

    NASA Technical Reports Server (NTRS)

    Lewis, M. C.

    1988-01-01

    The original flexible top and bottom walls of the Transonic Self-Streamlining Wind Tunnel (TSWT), at the University of Southampton, have been replaced with new walls featuring a larger number of static pressure tappings and detailed mechanical improvements. This report describes the streamling method, results, and conclusions of a series of tests aimed at defining sets of aerodynamically straight wall contours for the new flexible walls. This procedure is a necessary prelude to model testing. The quality of data obtained compares favorably with the aerodynamically straight data obtained with the old walls. No operational difficulties were experienced with the new walls.

  16. Application of Computational Fluid Dynamics (CFD) in transonic wind-tunnel/flight-test correlation

    NASA Technical Reports Server (NTRS)

    Murman, E. M.

    1982-01-01

    The capability for calculating transonic flows for realistic configurations and conditions is discussed. Various phenomena which were modeled are shown to have the same order of magnitude on the influence of predicted results. It is concluded that CFD can make the following contributions to the task of correlating wind tunnel and flight test data: some effects of geometry differences and aeroelastic distortion can be predicted; tunnel wall effects can be assessed and corrected for; and the effects of model support systems and free stream nonuniformities can be modeled.

  17. Transonic Flutter Suppression Control Law Design, Analysis and Wind Tunnel Results

    NASA Technical Reports Server (NTRS)

    Mukhopadhyay, Vivek

    1999-01-01

    The benchmark active controls technology and wind tunnel test program at NASA Langley Research Center was started with the objective to investigate the nonlinear, unsteady aerodynamics and active flutter suppression of wings in transonic flow. The paper will present the flutter suppression control law design process, numerical nonlinear simulation and wind tunnel test results for the NACA 0012 benchmark active control wing model. The flutter suppression control law design processes using (1) classical, (2) linear quadratic Gaussian (LQG), and (3) minimax techniques are described. A unified general formulation and solution for the LQG and minimax approaches, based on the steady state differential game theory is presented. Design considerations for improving the control law robustness and digital implementation are outlined. It was shown that simple control laws when properly designed based on physical principles, can suppress flutter with limited control power even in the presence of transonic shocks and flow separation. In wind tunnel tests in air and heavy gas medium, the closed-loop flutter dynamic pressure was increased to the tunnel upper limit of 200 psf The control law robustness and performance predictions were verified in highly nonlinear flow conditions, gain and phase perturbations, and spoiler deployment. A non-design plunge instability condition was also successfully suppressed.

  18. Transonic Flutter Suppression Control Law Design Using Classical and Optimal Techniques with Wind-Tunnel Results

    NASA Technical Reports Server (NTRS)

    Mukhopadhyay, Vivek

    1999-01-01

    The benchmark active controls technology and wind tunnel test program at NASA Langley Research Center was started with the objective to investigate the nonlinear, unsteady aerodynamics and active flutter suppression of wings in transonic flow. The paper will present the flutter suppression control law design process, numerical nonlinear simulation and wind tunnel test results for the NACA 0012 benchmark active control wing model. The flutter suppression control law design processes using (1) classical, (2) linear quadratic Gaussian (LQG), and (3) minimax techniques are described. A unified general formulation and solution for the LQG and minimax approaches, based on the steady state differential game theory is presented. Design considerations for improving the control law robustness and digital implementation are outlined. It was shown that simple control laws when properly designed based on physical principles, can suppress flutter with limited control power even in the presence of transonic shocks and flow separation. In wind tunnel tests in air and heavy gas medium, the closed-loop flutter dynamic pressure was increased to the tunnel upper limit of 200 psf. The control law robustness and performance predictions were verified in highly nonlinear flow conditions, gain and phase perturbations, and spoiler deployment. A non-design plunge instability condition was also successfully suppressed.

  19. Transonic Flutter Suppression Control Law Design, Analysis and Wind-Tunnel Results

    NASA Technical Reports Server (NTRS)

    Mukhopadhyay, Vivek

    1999-01-01

    The benchmark active controls technology and wind tunnel test program at NASA Langley Research Center was started with the objective to investigate the nonlinear, unsteady aerodynamics and active flutter suppression of wings in transonic flow. The paper will present the flutter suppression control law design process, numerical nonlinear simulation and wind tunnel test results for the NACA 0012 benchmark active control wing model. The flutter suppression control law design processes using (1) classical, (2) linear quadratic Gaussian (LQG), and (3) minimax techniques are described. A unified general formulation and solution for the LQG and minimax approaches, based on the steady state differential game theory is presented. Design considerations for improving the control law robustness and digital implementation are outlined. It was shown that simple control laws when properly designed based on physical principles, can suppress flutter with limited control power even in the presence of transonic shocks and flow separation. In wind tunnel tests in air and heavy gas medium, the closed-loop flutter dynamic pressure was increased to the tunnel upper limit of 200 psf. The control law robustness and performance predictions were verified in highly nonlinear flow conditions, gain and phase perturbations, and spoiler deployment. A non-design plunge instability condition was also successfully suppressed.

  20. Transonic Flutter Suppression Control Law Design, Analysis and Wind-Tunnel Results

    NASA Technical Reports Server (NTRS)

    Mukhopadhyay, Vivek

    1999-01-01

    The benchmark active controls technology and wind tunnel test program at NASA Langley Research Center was started with the objective to investigate the nonlinear, unsteady aerodynamics and active flutter suppression of wings in transonic flow. The paper will present the flutter suppression control law design process, numerical nonlinear simulation and wind tunnel test results for the NACA 0012 benchmark active control wing model. The flutter suppression control law design processes using classical, and minimax techniques are described. A unified general formulation and solution for the minimax approach, based on the steady state differential game theory is presented. Design considerations for improving the control law robustness and digital implementation are outlined. It was shown that simple control laws when properly designed based on physical principles, can suppress flutter with limited control power even in the presence of transonic shocks and flow separation. In wind tunnel tests in air and heavy gas medium, the closed-loop flutter dynamic pressure was increased to the tunnel upper limit of 200 psf. The control law robustness and performance predictions were verified in highly nonlinear flow conditions, gain and phase perturbations, and spoiler deployment. A non-design plunge instability condition was also successfully suppressed.

  1. Multiple transonic solutions and a new class of shock transitions in solar and stellar winds

    NASA Technical Reports Server (NTRS)

    Habbal, S. R.; Tsinganos, K.; Rosner, R.

    1983-01-01

    The steady isothermal solar wind equations are shown to admit, under certain circumstances, mutliple transonic solutions when, for example, momentum deposition gives rise to multiplee critical points in the flow. These multiple solutions consist of a continuous solution and solutions which involve shock transitions between critical solutions. The ambiguity arising from the multiplicity of the solutions can be resolved by following the time evolution of a wind profile with one critical point. Results of the numerical integration of the time-dependent equations with momentum addition show that each of these multiple solutions is physically accessible and depends on the rate of change of momentum deposition. These results suggest that standing shocks are likely to be present in the inner solar wind flow.

  2. Adaptive wall technology for minimization of wall interferences in transonic wind tunnels

    NASA Technical Reports Server (NTRS)

    Wolf, Stephen W. D.

    1988-01-01

    Modern experimental techniques to improve free air simulations in transonic wind tunnels by use of adaptive wall technology are reviewed. Considered are the significant advantages of adaptive wall testing techniques with respect to wall interferences, Reynolds number, tunnel drive power, and flow quality. The application of these testing techniques relies on making the test section boundaries adjustable and using a rapid wall adjustment procedure. A historical overview shows how the disjointed development of these testing techniques, since 1938, is closely linked to available computer support. An overview of Adaptive Wall Test Section (AWTS) designs shows a preference for use of relatively simple designs with solid adaptive walls in 2- and 3-D testing. Operational aspects of AWTS's are discussed with regard to production type operation where adaptive wall adjustments need to be quick. Both 2- and 3-D data are presented to illustrate the quality of AWTS data over the transonic speed range. Adaptive wall technology is available for general use in 2-D testing, even in cryogenic wind tunnels. In 3-D testing, more refinement of the adaptive wall testing techniques is required before more widespread use can be planned.

  3. Flow Quality Measurements in the NASA Ames Upgraded 11-by 11-Foot Transonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Amaya, Max A.; Murthy, Sreedhara V.; George, M. W. (Technical Monitor)

    2000-01-01

    Among the many upgrades designed and implemented in the NASA Ames 11-by 11-Foot Transonic Wind Tunnel over the past few years, several directly affect flow quality in the test section: a turbulence reduction system with a honeycomb and two screens, a flow smoothing system in the back leg diffusers, an improved drive motor control system, and a full replacement set of composite blades for the compressor. Prior to the shut-down of the tunnel for construction activities, an 8-foot span rake populated with flow instrumentation was traversed in the test section to fully document the flow quality and establish a baseline against which the upgrades could be characterized. A similar set of measurements was performed during the recent integrated system test trials, but the scope was somewhat limited in accordance with the primary objective of such tests, namely to return the tunnel to a fully operational status. These measurements clearly revealed substantial improvements in flow angularity and significant reductions in turbulence level for both full-span and semi-span testing configurations, thus making the flow quality of the tunnel one of the best among existing transonic facilities.

  4. Transonic wind tunnel test of a 14 percent thick oblique wing

    NASA Technical Reports Server (NTRS)

    Kennelly, Robert A., Jr.; Kroo, Ilan M.; Strong, James M.; Carmichael, Ralph L.

    1990-01-01

    An experimental investigation was conducted at the ARC 11- by 11-Foot Transonic Wind Tunnel as part of the Oblique Wing Research Aircraft Program to study the aerodynamic performance and stability characteristics of a 0.087-scale model of an F-8 airplane fitted with an oblique wing designed by Rockwell International. The 10.3 aspect ratio, straight-tapered wing of 0.14 thickness/chord ratio was tested at two different mounting heights above the fuselage. Additional tests were conducted to assess low-speed behavior with and without flaps, aileron effectiveness at representative flight conditions, and transonic drag divergence with 0 degree wing sweep. Longitudinal stability data were obtained at sweep angles of 0, 30, 45, 60, and 65 degrees, at Mach numbers ranging from 0.25 to 1.40. Test Reynolds numbers varied from 3.2 to 6.6 x 10 exp 6/ft. and angle of attack ranged from -5 to +18 degrees. Most data were taken at zero sideslip, but a few runs were at sideslip angles of +/- 5 degrees. The raised wing position proved detrimental overall, although side force and yawing moment were reduced at some conditions. Maximum lift coefficient with the flaps deflected was found to fall short of the value predicted in the preliminary design document. The performance and trim characteristics of the present wing are generally inferior to those obtained for a previously tested wing designed at ARC.

  5. Multiple transonic solutions with a new class of shock transitions in steady isothermal solar and stellar winds

    NASA Technical Reports Server (NTRS)

    Habbal, S. R.; Tsinganos, K.

    1983-01-01

    It is shown that a new class of shock transitions arises in the transonic solutions of the steady isothermal solar wind equations when momentum deposition and/or nonradial flow tube divergence give rise to multiple critical points in the flow. These shock transitions between critical solutions occur for a certain range of the parameters which characterize the momentum deposition function. The isothermal wind equations allow multiple transonic solutions in the presence of such shock transitions, yielding a continuous solution passing through an inner critical point and solutions involving a shock transition between critical solutions. It is determined that these multiple transonic solutions have the same flow speed at the base but different supersonic flow speeds at infinity. It is found that the nonradial flow tube divergence and momentum addition are equivalent, which gives rise to multiple critical points and hence to multiple transonic solutions with shock transitions. In addition, the physical relevance of these properties are examined for astrophysical systems such as the inner solar wind, flows in extragalactic jets, and accretion discs.

  6. Aircraft wake flow effect and horizontal tail buffet. [pressure distribution and responses of fighter aircraft in transonic maneuvers

    NASA Technical Reports Server (NTRS)

    Hwang, C.; Pi, W. S.

    1979-01-01

    As part of a program to investigate the fluctuating pressure distribution and response behavior of a fighter aircraft in transonic maneuver, an F-5A scale model has previously been tested in an 11-ft transonic wind tunnel. The model, with a number of static and dynamic pressure transducers imbedded in the lifting surfaces was tested at various angles of attack up to 16 deg. In this paper, test results of particular interest to wake flow and horizontal tail buffet are described. It is shown that the dynamic pressure data on the tail surface at the specified flight conditions serve to determine the local dynamic loads. They also influence the control performance of the aircraft under maneuver conditions where buffet is encountered. The data presented demonstrate a number of contributing factors that affect the tail dynamic pressures in the transonic regime.

  7. Transonic wind tunnel tests of A.015 scale space shuttle orbiter model, volume 1

    NASA Technical Reports Server (NTRS)

    Struzynski, N. A.

    1975-01-01

    Transonic wind tunnel tests were run on a 0.015 scale model of the Space Shuttle Orbiter Vehicle in an eight-foot tunnel during August 1975. The purpose of the program was to obtain basic shuttle aerodynamic data through a full range of elevon and aileron deflections, verification of data obtained at other facilities, and effects of Reynolds numbers. The first part of a discussion of test procedures and results in both tabular and graphical form were presented. Tests were performed at Mach numbers from 0.35 to 1.20, and at Reynolds numbers for 3.5 million to 8.2 million per foot. The angle of attack was varied from -1 to +20 degrees at sideslip angles of -2, 0, +2 degrees. Sideslip was varied from -6 to +8 degrees at constant angles of attack from 0 to +20 degrees. Various aileron and ailevon settings were tested for various angles of attack.

  8. Direct Validation of the Wall Interference Correction System of the Ames 11-Foot Transonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Ulbrich, Norbert; Boone, Alan R.

    2003-01-01

    Data from the test of a large semispan model was used to perform a direct validation of a wall interference correction system for a transonic slotted wall wind tunnel. At first, different sets of uncorrected aerodynamic coefficients were generated by physically changing the boundary condition of the test section walls. Then, wall interference corrections were computed and applied to all data points. Finally, an interpolation of the corrected aerodynamic coefficients was performed. This interpolation made sure that the corrected Mach number of a given run would be constant. Overall, the agreement between corresponding interpolated lift, drag, and pitching moment coefficient sets was very good. Buoyancy corrections were also investigated. These studies showed that the accuracy goal of one drag count may only be achieved if reliable estimates of the wall interference induced buoyancy correction are available during a test.

  9. TWINTN4: A program for transonic four-wall interference assessment in two-dimensional wind tunnels

    NASA Technical Reports Server (NTRS)

    Kemp, W. B., Jr.

    1984-01-01

    A method for assessing the wall interference in transonic two-dimensional wind tunnel tests including the effects of the tunnel sidewall boundary layer was developed and implemented in a computer program named TWINTN4. The method involves three successive solutions of the transonic small disturbance potential equation to define the wind tunnel flow, the equivalent free air flow around the model, and the perturbation attributable to the model. Required input includes pressure distributions on the model and along the top and bottom tunnel walls which are used as boundary conditions for the wind tunnel flow. The wall-induced perturbation field is determined as the difference between the perturbation in the tunnel flow solution and the perturbation attributable to the model. The methodology used in the program is described and detailed descriptions of the computer program input and output are presented. Input and output for a sample case are given.

  10. Using transonic small disturbance theory for predicting the aeroelastic stability of a flexible wind-tunnel model

    NASA Technical Reports Server (NTRS)

    Silva, Walter A.; Bennett, Robert M.

    1990-01-01

    The CAP-TSD (Computational Aeroelasticity Program - Transonic Small Disturbance) code, developed at the NASA - Langley Research Center, is applied to the Active Flexible Wing (AFW) wind tunnel model for prediction of the model's transonic aeroelastic behavior. Static aeroelastic solutions using CAP-TSD are computed. Dynamic (flutter) analyses are then performed as perturbations about the static aeroelastic deformations of the AFW. The accuracy of the static aeroelastic procedure is investigated by comparing analytical results to those from previous AFW wind tunnel experiments. Dynamic results are presented in the form of root loci at different Mach numbers for a heavy gas and air. The resultant flutter boundaries for both gases are also presented. The effects of viscous damping and angle-of-attack, on the flutter boundary in air, are presented as well.

  11. A fast-response aspirating probe for measurements of total temperature and pressure in transonic cryogenic wind tunnel

    NASA Technical Reports Server (NTRS)

    Ng, W.-F.; Rosson, J. C.

    1986-01-01

    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. The probe consists of two coplanar constant temperature hot wires at different overheat ratios operating in a 1.5-mm-diam channel with a choked exit. Thus, the constant Mach number flow by the wires is influenced only by free-stream stagnation temperature and pressure. Diffusion of the free-stream Mach number to a lower value in the channel reduces the dynamic drag on the hot-wire. Frequency response of the present design is dc to 20 kHz. The probe was used to measure the unsteady wake shed from an oscillating airfoil tested in the 0.3-m Transonic Cryogenic Tunnel at NASA-Langley Research Center. The hot-wire lasted for more than ten hours before breaking, proving the ruggedness of the probe and the usefulness of the technique in a high dynamic pressure, transonic cryogenic wind tunnel. Typical data obtained from the experiment are presented after reduction to stagnation pressure and temperature.

  12. Investigation of transonic flow over segmented slotted wind tunnel wall with mass transfer

    NASA Technical Reports Server (NTRS)

    Bhat, M. K.; Vakili, A. D.; Wu, J. M.

    1990-01-01

    The flowfield on a segmented multi-slotted wind tunnel wall was studied at transonic speeds by measurements in and near the wall layer using five port cone probes. The slotted wall flowfield was observed to be three-dimensional in nature for a relatively significant distance above the slot. The boundary layer characteristics measured on the single slotted wall were found to be very sensitive to the applied suction through the slot. The perturbation in the velocity components generated due to the flow through the slot decay rapidly in the transverse direction. A vortex-like flow existed on the single slotted wall for natural ventilation but diminished with increased suction flow rate. For flow on a segmented multi-slotted wall, the normal velocity component changes were found to be maximum for measurement points located between the segmented slots atop the active chamber. The lateral influence due to applied suction and blowing, through a compartment, exceeded only slightly that in the downstream direction. Limited upstream influence was observed. Influence coefficients were determined from the data in the least-square sense for blowing and suction applied through one and two compartments. This was found to be an adequate determination of the influence coefficients for the range of mass flows considered.

  13. Transonic wind-tunnel investigation of the Galileo Probe parachute configuration

    NASA Technical Reports Server (NTRS)

    Corridan, R. E.; Givens, J. G.; Kepley, B. M.

    1984-01-01

    Wind-tunnel tests have been made of 1/4-scale and 1/2-scale models of the Galileo Probe conical ribbon-parachute at flight conditions. For the Galileo mission, the parachute is deployed behind a blunt shape at transonic speed. The investigation was initiated to verify the solutions to the delayed main-parachute opening with squidding that was observed in the 1982 Galileo Probe system balloon drop test. In addition to varying model scale, the forebody shape, angle of attack, dynamic pressure, Mach number, and canopy trailing distance behind the forebody were varied to determine their effect on parachute performance. Both steady-state and deployment tests were conducted. Parachute drag was measured and was seen to degrade severely at canopy trailing distances of 5.5 and 7 forebody diameters. Performance was shown to be good at the trailing distances of 9 and 11 diameters. A second balloon drop test demonstrated the desired parachute performance at the system level.

  14. Further investigations of the aeroelastic behavior of the AFW wind-tunnel model using transonic small disturbance theory

    NASA Technical Reports Server (NTRS)

    Silva, Walter A.; Bennett, Robert M.

    1992-01-01

    The Computational Aeroelasticity Program-Transonic Small Disturbance (CAP-TSD) code, developed at LaRC, is applied to the active flexible wing wind-tunnel model for prediction of transonic aeroelastic behavior. A semi-span computational model is used for evaluation of symmetric motions, and a full-span model is used for evaluation of antisymmetric motions, and a full-span model is used for evaluation of antisymmetric motions. Static aeroelastic solutions using CAP-TSD are computed. Dynamic deformations are presented as flutter boundaries in terms of Mach number and dynamic pressure. Flutter boundaries that take into account modal refinements, vorticity and entropy corrections, antisymmetric motion, and sensitivity to the modeling of the wing tip ballast stores are also presented with experimental flutter results.

  15. An Experimental Evaluation of Advanced Rotorcraft Airfoils in the NASA Ames Eleven-foot Transonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Flemming, Robert J.

    1984-01-01

    Five full scale rotorcraft airfoils were tested in the NASA Ames Eleven-Foot Transonic Wind Tunnel for full scale Reynolds numbers at Mach numbers from 0.3 to 1.07. The models, which spanned the tunnel from floor to ceiling, included two modern baseline airfoils, the SC1095 and SC1094 R8, which have been previously tested in other facilities. Three advanced transonic airfoils, designated the SSC-A09, SSC-A07, and SSC-B08, were tested to confirm predicted performance and provide confirmation of advanced airfoil design methods. The test showed that the eleven-foot tunnel is suited to two-dimensional airfoil testing. Maximum lift coefficients, drag coefficients, pitching moments, and pressure coefficient distributions are presented. The airfoil analysis codes agreed well with the data, with the Grumman GRUMFOIL code giving the best overall performance correlation.

  16. Wind tunnel investigation of the interaction and breakdown characteristics of slender wing vortices at subsonic, transonic, and supersonic speeds

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.

    1991-01-01

    The vortex dominated aerodynamic characteristics of a generic 65 degree cropped delta wing model were studied in a wind tunnel at subsonic through supersonic speeds. The lee-side flow fields over the wing-alone configuration and the wing with leading edge extension (LEX) added were observed at M (infinity) equals 0.40 to 1.60 using a laser vapor screen technique. These results were correlated with surface streamline patterns, upper surface static pressure distributions, and six-component forces and moments. The wing-alone exhibited vortex breakdown and asymmetry of the breakdown location at the subsonic and transonic speeds. An earlier onset of vortex breakdown over the wing occurred at transonic speeds due to the interaction of the leading edge vortex with the normal shock wave. The development of a shock wave between the vortex and wing surface caused an early separation of the secondary boundary layer. With the LEX installed, wing vortex breakdown asymmetry did not occur up to the maximum angle of attack in the present test of 24 degrees. The favorable interaction of the LEX vortex with the wing flow field reduced the effects of shock waves on the wing primary and secondary vortical flows. The direct interaction of the wing and LEX vortex cores diminished with increasing Mach number. The maximum attainable vortex-induced pressure signatures were constrained by the vacuum pressure limit at the transonic and supersonic speeds.

  17. Experimental studies of transonic flow field near a longitudinally slotted wind tunnel wall. Ph.D. Thesis - George Washington Univ., 1988

    NASA Technical Reports Server (NTRS)

    Everhart, Joel L.; Bobbitt, Percy J.

    1994-01-01

    The results of detailed parametric experiments are presented for the near-wall flow field of a longitudinally slotted transonic wind tunnel. Existing data are reevaluated and new data obtained in the Langley 6- by 19-inch Transonic Wind Tunnel are presented and analyzed. In the experiments, researchers systematically investigate many pertinent wall-geometry variables such as the wall openness and the number of slots along with the free stream Mach number and model angle of attack. Flow field surveys on the plane passing through the centerline of the slot were conducted and are presented. The effects of viscosity on the slot flow are considered in the analysis. The present experiments, combined with those of previous investigations, give a more complete physical characterization of the flow near and through the slotted wall of a transonic wind tunnel.

  18. Some anomalies observed in wind-tunnel tests of a blunt body at transonic and supersonic speeds

    NASA Technical Reports Server (NTRS)

    Brooks, J. D.

    1976-01-01

    An investigation of anomalies observed in wind tunnel force tests of a blunt body configuration was conducted at Mach numbers from 0.20 to 1.35 in the Langley 8-foot transonic pressure tunnel and at Mach numbers of 1.50, 1,80, and 2.16 in the Langley Unitary Plan wind tunnel. At a Mach number of 1.35, large variations occurred in axial force coefficient at a given angle of attack. At transonic and low supersonic speeds, the total drag measured in the wind tunnel was much lower than that measured during earlier ballistic range tests. Accurate measurements of total drag for blunt bodies will require the use of models smaller than those tested thus far; however, it appears that accurate forebody drag results can be obtained by using relatively large models. Shock standoff distance is presented from experimental data over the Mach number range from 1.05 to 4.34. Theory accurately predicts the shock standoff distance at Mach numbers up to 1.75.

  19. Wind tunnel investigations of forebody strakes for yaw control on F/A-18 model at subsonic and transonic speeds

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.; Murri, Daniel G.

    1993-01-01

    Wind tunnel investigations have been conducted of forebody strakes for yaw control on 0.06-scale models of the F/A-18 aircraft at free-stream Mach numbers of 0.20 to 0.90. The testing was conducted in the 7- by 10-Foot Transonic Tunnel at the David Taylor Research Center and the Langley 7- by 10-Foot High-Speed Tunnel. The principal objectives of the testing were to determine the effects of the Mach number and the strake plan form on the strake yaw control effectiveness and the corresponding strake vortex induced flow field. The wind tunnel model configurations simulated an actuated conformal strake deployed for maximum yaw control at high angles of attack. The test data included six-component forces and moments on the complete model, surface static pressure distributions on the forebody and wing leading-edge extensions, and on-surface and off-surface flow visualizations. The results from these studies show that the strake produces large yaw control increments at high angles of attack that exceed the effect of conventional rudders at low angles of attack. The strake yaw control increments diminish with increasing Mach number but continue to exceed the effect of rudder deflection at angles of attack greater than 30 degrees. The character of the strake vortex induced flow field is similar at subsonic and transonic speeds. Cropping the strake planform to account for geometric and structural constraints on the F-18 aircraft has a small effect on the yaw control increments at subsonic speeds and no effect at transonic speeds.

  20. Analysis of Fluctuating Static Pressure Measurements in a Large High Reynolds Number Transonic Cryogenic Wind Tunnel. Ph.D. Thesis

    NASA Technical Reports Server (NTRS)

    Igoe, William B.

    1991-01-01

    Dynamic measurements of fluctuating static pressure levels were made using flush mounted high frequency response pressure transducers at eleven locations in the circuit of the National Transonic Facility (NTF) over the complete operating range of this wind tunnel. Measurements were made at test section Mach numbers from 0.2 to 1.2, at pressure from 1 to 8.6 atmospheres and at temperatures from ambient to -250 F, resulting in dynamic flow disturbance measurements at the highest Reynolds numbers available in a transonic ground test facility. Tests were also made independently at variable Mach number, variable Reynolds number, and variable drivepower, each time keeping the other two variables constant thus allowing for the first time, a distinct separation of these three important variables. A description of the NTF emphasizing its flow quality features, details on the calibration of the instrumentation, results of measurements with the test section slots covered, downstream choke, effects of liquid nitrogen injection and gaseous nitrogen venting, comparisons between air and nitrogen, isolation of the effects of Mach number, Reynolds number, and fan drive power, and identification of the sources of significant flow disturbances is included. The results indicate that primary sources of flow disturbance in the NTF may be edge-tones generated by test section sidewall re-entry flaps and the venting of nitrogen gas from the return leg of the tunnel circuit between turns 3 and 4 in the cryogenic mode of operation. The tests to isolate the effects of Mach number, Reynolds number, and drive power indicate that Mach number effects predominate. A comparison with other transonic wind tunnels shows that the NTF has low levels of test section fluctuating static pressure especially in the high subsonic Mach number range from 0.7 to 0.9.

  1. Adjoint Method and Predictive Control for 1-D Flow in NASA Ames 11-Foot Transonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Nguyen, Nhan; Ardema, Mark

    2006-01-01

    This paper describes a modeling method and a new optimal control approach to investigate a Mach number control problem for the NASA Ames 11-Foot Transonic Wind Tunnel. The flow in the wind tunnel is modeled by the 1-D unsteady Euler equations whose boundary conditions prescribe a controlling action by a compressor. The boundary control inputs to the compressor are in turn controlled by a drive motor system and an inlet guide vane system whose dynamics are modeled by ordinary differential equations. The resulting Euler equations are thus coupled to the ordinary differential equations via the boundary conditions. Optimality conditions are established by an adjoint method and are used to develop a model predictive linear-quadratic optimal control for regulating the Mach number due to a test model disturbance during a continuous pitch

  2. Application of a transonic similarity rule to correct the effects of sidewall boundary layers in two-dimensional transonic wind tunnels. M.S. Thesis - George Washington Univ.

    NASA Technical Reports Server (NTRS)

    Sewall, W. G.

    1982-01-01

    A transonic similarity rule which accounts for the effects of attached sidewall boundary layers is presented and evaluated by comparison with the characteristics of airfoils tested in a two dimensional transonic tunnel with different sidewall boundary layer thicknesses. The rule appears valid provided the sidewall boundary layer both remains attached in the vicinity of the model and occupies a small enough fraction of the tunnel width to preserve sufficient two dimensionality in the tunnel.

  3. FLPP IXV Re-entry Vehicle, Transonic Characterisation Based on FOI T1500 Wind Tunnel Tests and CFD

    NASA Astrophysics Data System (ADS)

    Torngren, L.; Chiarelli, C.; Mareschi, V.; Tribot, J.-P.; Binetti, P.; Walloschek, T.

    2009-01-01

    The European Space Agency ESA, has engaged in 2004, the IXV project (Intermediate eXperimental Vehicle) which is part of the FLPP (Future Launcher Preparatory Programme) aiming at answering to critical technological issues, while supporting the future generation launchers and to improve in general European capabilities in the strategic field of atmospheric re-entry for space transportation, exploration and scientific applications. The IXV key mission and system objectives are the design, development, manufacturing, assembling and on-ground to in-flight verification of an autonomous European lifting and aerodynamically controlled re-entry system, integrating the critical re-entry technologies at the system level. The current IXV vehicle is a slender body type exhibiting rounded shape, thick body controlled by means of two control surfaces. The current mission is to perform an atmospheric re- entry ended by a safe recovery in supersonic regime. A potential extension of the flight domain down to the transonic regime was proposed to be analyzed. The objectives were to study the capability of the IXV for flying autonomously enabling a recovery of the vehicle by means of a subsonic parachute based DRS. The vehicle designed for the hypersonic speeds integrating a large base with only two control surfaces located close to the plane of symmetry is definitively not tuned for transonic ones. CFD done by Thales Alenia Space and wind tunnel activities involving FOI T1500 facility contributed to built up an Aerodynamic Data Base (AEDB) to be used as inputs for flying qualities analysis and re-entry simulations. The paper presents the main objectives of the transonic activities with emphasis on CFD and WTT including a description of the different prediction tools and discussing the main outcomes of the current data comparisons.

  4. Wind tunnel investigation of vortex flows on F/A-18 configuration at subsonic through transonic speed

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.

    1991-01-01

    A wind tunnel experiment was conducted in the David Taylor Research Center 7- by 10-Foot Transonic Tunnel of the wing leading-edge extension (LEX) and forebody vortex flows at subsonic and transonic speeds about a 0.06-scale model of the F/A-18. The primary goal was to improve the understanding and control of the vortical flows, including the phenomena of vortex breakdown and vortex interactions with the vertical tails. Laser vapor screen flow visualizations, LEX, and forebody surface static pressures, and six-component forces and moments were obtained at angles of attack of 10 to 50 degrees, free-stream Mach numbers of 0.20 to 0.90, and Reynolds numbers based on the wing mean aerodynamic chord of 0.96 x 10(exp 6) to 1.75 x 10(exp 6). The wind tunnel results were correlated with in-flight flow visualizations and handling qualities trends obtained by NASA using an F-18 High-Alpha Research Vehicle (HARV) and by the Navy and McDonnell Douglas on F-18 aircraft with LEX fences added to improve the vertical tail buffet environment. Key issues that were addressed include the sensitivity of the vortical flows to the Reynolds number and Mach number; the reduced vertical tail excitation, and the corresponding flow mechanism, in the presence of the LEX fence; the repeatability of data obtained during high angle-of-attack wind tunnel testing of F-18 models; the effects of particle seeding for flow visualization on the quantitative model measurements; and the interpretation of off-body flow visualizations obtained using different illumination and particle seeding techniques.

  5. Wind Tunnel Application of a Pressure-Sensitive Paint Technique to a Faceted Missile Model at Subsonic and Transonic Speeds

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.

    2004-01-01

    A pressure-sensitive paint (PSP) technique was applied in a wind tunnel experiment in the NASA Langley Research Center 8-Foot Transonic Pressure Tunnel to quantify the vortex-induced surface static pressures on a slender, faceted missile model at subsonic and transonic speeds. Global PSP calibrations were obtained using an in-situ method featuring the simultaneous electronically-scanned pressures (ESP) measurements. Both techniques revealed the significant influence leading-edge vortices on the surface pressure distributions. The mean error in the PSP measurements relative to the ESP data was approximately 0.6 percent at M(sub infinity)=0.70 and 2.6 percent at M(sub infinity)=0.90 and 1.20. The vortex surface pressure signatures obtained from the PSP and ESP techniques were correlated with the off-surface vortex cross-flow structures obtained using a laser vapor screen (LVS) flow visualization technique. The on-surface and off-surface techniques were complementary, since each provided details of the vortex-dominated flow that were not clear or apparent in the other.

  6. Wind Tunnel Application of a Pressure-Sensitive Paint Technique to a Faceted Missile Model at Subsonic and Transonic Speeds

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.

    2004-01-01

    A pressure-sensitive paint (PSP) technique was applied in a wind tunnel experiment in the NASA Langley Research Center 8-Foot Transonic Pressure Tunnel to quantify the vortex-induced surface static pressures on a slender, faceted missile model at subsonic and transonic speeds. Satisfactory global calibrations of the PSP were obtained at =0.70, 0.90, and 1.20, angles of attack from 10 degrees to 20 degrees, and angles of sideslip of 0 and 2.5 degrees using an in-situ method featuring the simultaneous acquisition of electronically-scanned pressures (ESP) at 57 discrete locations on the model. Both techniques clearly revealed the significant influence on the surface pressure distributions of the vortices shed from the sharp, chine-like leading edges. The mean error in the PSP measurements relative to the ESP data was approximately 0.6 percent at M infinity =0.70 and 2.6 percent at M infinity =0.90 and 1.20. The vortex surface pressure signatures obtained from the PSP and ESP techniques were correlated with the off-surface vortex cross-flow structures obtained using a laser vapor screen (LVS) flow visualization technique. The on-surface and off-surface techniques were complementary, since each provided details of the vortex-dominated flow that were not clear or apparent in the other.

  7. Pressure coefficient evaluation on the surface of the SONDA III model tested in the TTP Pilot Transonic Wind Tunnel

    NASA Astrophysics Data System (ADS)

    Reis, M. L. C. C.; Falcao Filho, J. B. P.; Basso, E.; Caldas, V. R.

    2015-02-01

    A test campaign of the Brazilian sounding rocket Sonda III was carried out at the Pilot Transonic Wind Tunnel, TTP. The aim of the campaign was to investigate aerodynamic phenomena taking place at the connection region of the first and second stages. Shock and expansion waves are expected at this location causing high gradients in airflow properties around the vehicle. Pressure taps located on the surface of a Sonda III half model measure local static pressures. Other measured parameters were freestream static and total pressures of the airflow. Estimated parameters were pressure coefficients and Mach numbers. Uncertainties associated with the estimated parameters were calculated by employing the Law of Propagation of Uncertainty and the Monte Carlo method. It was found that both uncertainty evaluation methods resulted in similar values. A Computational Fluid Dynamics simulation code was elaborated to help understand the changes in the flow field properties caused by the disturbances.

  8. An evaluation in a modern wind tunnel of the transonic adaptive wall adjustment strategy developed by NPL in the 1940's

    NASA Technical Reports Server (NTRS)

    Lewis, M. C.

    1988-01-01

    The first documented wind tunnel employing a flexible walled test section for the purpose of eliminating wall interference was constructed in England by the National Physical Laboratory (NPL) during the late 1930's. The tunnel was transonic and designed for two-dimensional testing. In an attempt to eliminate the top and bottom wall interference effects on the model NPL developed a strategy to adjust two flexible walls to streamlined shapes. This report covers an evaluation of the NPL wall adjustment strategy in a modern wind tunnel, e.g., the Transonic Self-Streamlining Wind Tunnel (TSWT) at the University of Southampton, England. The evaluation took the form of performance comparisons with other modern strategies which have been developed for use in, and proven in, the TSWT.

  9. Design considerations of the national transonic facility

    NASA Technical Reports Server (NTRS)

    Baals, D. D.

    1976-01-01

    The inability of existing wind tunnels to provide aerodynamic test data at transonic speeds and flight Reynolds numbers was examined. The proposed transonic facility is a high Reynolds number transonic wind tunnel designed to meet the research and development needs of industry, and the scientific community. The facility employs the cryogenic approach to achieve high transonic Reynolds numbers at acceptable model loads and tunnel power. By using temperature as a test variable, a unique capability to separate scale effects from model aeroelastic effects is provided. The performance envelope of the facility is shown to provide a ten fold increase in transonic Reynolds number capability compared to currently available facilities.

  10. Revalidation of the NASA Ames 11-by 11-Foot Transonic Wind Tunnel with a Commercial Airplane Model

    NASA Technical Reports Server (NTRS)

    Kmak, Frank J.; Hudgins, M.; Hergert, D.; George, Michael W. (Technical Monitor)

    2001-01-01

    The 11-By 11-Foot Transonic leg of the Unitary Plan Wind Tunnel (UPWT) was modernized to improve tunnel performance, capability, productivity, and reliability. Wind tunnel tests to demonstrate the readiness of the tunnel for a return to production operations included an Integrated Systems Test (IST), calibration tests, and airplane validation tests. One of the two validation tests was a 0.037-scale Boeing 777 model that was previously tested in the 11-By 11-Foot tunnel in 1991. The objective of the validation tests was to compare pre-modernization and post-modernization results from the same airplane model in order to substantiate the operational readiness of the facility. Evaluation of within-test, test-to-test, and tunnel-to-tunnel data repeatability were made to study the effects of the tunnel modifications. Tunnel productivity was also evaluated to determine the readiness of the facility for production operations. The operation of the facility, including model installation, tunnel operations, and the performance of tunnel systems, was observed and facility deficiency findings generated. The data repeatability studies and tunnel-to-tunnel comparisons demonstrated outstanding data repeatability and a high overall level of data quality. Despite some operational and facility problems, the validation test was successful in demonstrating the readiness of the facility to perform production airplane wind tunnel%, tests.

  11. Overview of advanced wing design. [Ames 12-Foot Pressure Tunnel and 11-Foot Transonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Hicks, R. M.

    1981-01-01

    Examples of experiment theory correlation are presented to give an indication of the capabilities and limitations of wing design and analysis for transonic applications by potential flow theory. The examples include correlations of experimental pressure distributions with theoretical results from isolated wing codes and wing-body codes. Both conservative and non conservative differencing as well as body and boundary layer corrections are considered. A full potential isolated wing code correlates well with data from an isolated wing test but may give poor prediction of the aerodynamic characteristics of some wing-body configurations. Potential flow wing body codes were found to improve the correlation for the wing-body configurations considered.

  12. High angle canard missile test in the Ames 11-foot transonic wind tunnel

    NASA Technical Reports Server (NTRS)

    Schwind, R. G.

    1978-01-01

    Four blunted ogive-cylinder missile models with a length-to-diameter ratio of 10.4 were tested at transonic speeds and large angles of attack. The configurations are: body, body with tail panels, body with canards, and body with canards and tails. Forces and moments from the entire model and each of the eight fins were measured over the pitch range of 20 deg to 50 deg and 0 deg to 45 deg roll. Canard deflection angles between 0 deg and 15 deg were tested. Exploratory vapor screen flow visualization testing was also performed. Sample force and moment data are reported along with observations from the vapor screen tests.

  13. Development of a Large Field of View Shadowgraph System for a 16 Ft. Transonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Talley, Michael A.; Jones, Stephen B.; Goodman, Wesley L.

    2000-01-01

    A large field of view shadowgraph flow visualization system for the Langley 16 ft. Transonic Tunnel (16 ft.TT) has been developed to provide fast, low cost, aerodynamic design concept evaluation capability to support the development of the next generation of commercial and military aircraft and space launch vehicles. Key features of the 16 ft. TT shadowgraph system are: (1) high resolution (1280 X 1024) digital snap shots and sequences; (2) video recording of shadowgraph at 30 frames per second; (3) pan, tilt, & zoom to find and observe flow features; (4) one microsecond flash for freeze frame images; (5) large field of view approximately 12 X 6 ft; and (6) a low maintenance, high signal/noise ratio, retro-reflective screen to allow shadowgraph imaging while test section lights are on.

  14. Study on needs for a magnetic suspension system operating with a transonic wind tunnel

    NASA Technical Reports Server (NTRS)

    Martindale, W. R.; Butler, R. W.; Starr, R. F.

    1985-01-01

    The U.S. aeronautical industry was surveyed to determine if current and future transonic testing requirements are sufficient to justify continued development work on magnetic suspension and balance systems (MSBS) by NASA. The effort involved preparation of a brief technical description of magnetic suspension and balance systems, design of a survey form asking specific questions about the role of the MSBS in satisfying future testing requirements, selecting nine major aeronautics companies to which the description and survey forms were sent, and visiting the companies and discussing the survey to obtain greater insight to their response to the survey. Evaluation and documentation of the survey responses and recommendations which evolved from the study are presented.

  15. HSCT Ref-H Transonic Flap Data Base: Wind-Tunnel Test and Comparison with Theory

    NASA Technical Reports Server (NTRS)

    Vijgen, Paul M.

    1999-01-01

    In cooperation with personnel from the Boeing ANP Laboratory and NASA Langley, a performance test was conducted using the Reference-H 1.675% model ("NASA Modular Model") without nacelles at the NASA Langley 16-Ft Transonic Tunnel. The main objective of the test was to determine the drag reduction achievable with leading-edge and trailing-edge flaps deflected along the outboard wing span at transonic Mach numbers (M = 0.9 to 1.2) for purpose of preliminary design and for comparison with computational predictions. The obtained drag data with flap deflections for Mach numbers of 1.07 to 1.20 are unique for the Reference H wing. Four leading-edge and two trailing-edge flap deflection angles were tested at a mean-wing chord-Reynolds number of about 5.7 million. An outboard-wing leading-edge flap deflection of 81 provides a 4.5 percent drag reduction at M = 1.2 A = 0.2), and much larger values at lower Mach numbers with larger flap deflections. The present results for the baseline (no flaps deflected) compare reasonably well with previous Boeing and NASA Ref-H tunnel tests, including high-Reynolds number NTF results. Viscous CFD simulations using the OVERFLOW thin-layer N.S. method properly predict the observed trend in drag reduction at M = 1.2 as function of leading-edge flap deflection. Modified linear theory properly predicts the flap effects on drag at subsonic conditions (Aero2S code), and properly predicts the absolute drag for the 40 and 80 leading-edge deflection at M = 1.2 (A389 code).

  16. Experimental Wind-Tunnel Investigation of the Transonic Damping-in-Pitch Characteristics of Two Wing-Body Combinations

    NASA Technical Reports Server (NTRS)

    Emerson, Horace F.; Robinson, Robert C.

    1958-01-01

    The results of an experimental wind-tunnel investigation of the damping in pitch of two wing-body combinations are presented. The tests were conducted in the Ames 14-foot transonic wind tunnel over a Mach number range from 0.60 to 1.18. Reynolds numbers varied from 2.3 million to 5.5 million. One model with a triangular wing of aspect ratio 2 having NACA 0003-63 sections was oscillated at an amplitude of 1.5 and a frequency of 17 cycles per second. The second model with a straight, tapered wing of aspect ratio 3 having 3-percent biconvex circular-arc sections was oscillated at an amplitude of 1.0 deg and a frequency of 21 cycles per second. The tests were made with the models at a mean angle of attack of 0 deg. The models were oscillated with a dynamic balance that was actuated by an electrohydraulic servo valve. The results of this investigation indicate the usefulness of this new apparatus. The experimental results of a previous damping-in-pitch investigation conducted in the Ames 6- by 6-foot supersonic wind tunnel at Mach numbers from 1.2 to 1.7 are included along with the theoretical results for this Mach number range. In the region of Mach numbers available for comparison, good agreement is shown to exist between the data obtained in the two facilities, except for some inconsistency in the slopes of the curves at M = 1.2 for the triangular wing. The results of this investigation clearly show that for the models tested the maximum values of the damping in pitch occur at Mach numbers very close to 1.0, and that abrupt changes in the pitch damping are encountered near sonic velocity.

  17. Self streamlining wind tunnel: Low speed testing and transonic test section design

    NASA Technical Reports Server (NTRS)

    Wolf, S. W. D.; Goodyer, M. J.

    1977-01-01

    Comprehensive aerodynamic data on an airfoil section were obtained through a wide range of angles of attack, both stalled and unstalled. Data were gathered using a self streamlining wind tunnel and were compared to results obtained on the same section in a conventional wind tunnel. The reduction of wall interference through streamline was demonstrated.

  18. Wind tunnel tests of modified cross, hemisflo, and disk-gap-band parachutes with emphasis in the transonic range

    NASA Technical Reports Server (NTRS)

    Foughner, J. T., Jr.; Alexander, W. C.

    1974-01-01

    Transonic wind-tunnel studies were conducted with modified cross, hemisflo, and disk-gap-band parachute models in the wake of a cone-cylinder shape forebody. The basic cross design was modified with the addition of a circumferential constraining band at the lower edge of the canopy panels. The tests covered a Mach number range of 0.3 to 1.2 and a dynamic pressure range from 479 Newtons per square meter to 5746 Newtons per square meter. The parachute models were flexible textile-type structures and were tethered to a rigid forebody with a single flexible riser. Different size models of the modified cross and disk-gap-band canopies were tested to evaluate scale effects. Model reference diameters were 0.30, 0.61, and 1.07 meters (1.0, 2.0, and 3.5 ft) for the modified cross; and nominal diameters of 0.25 and 0.52 meter (0.83 and 1.7 ft) for the disk-gap-band; and 0.55 meter (1.8 ft) for the hemisflo. Reefing information is presented for the 0.61-meter-diameter cross and the 0.52-meter-diameter disk-gap-band. Results are presented in the form of the variation of steady-state average drag coefficient with Mach number. General stability characteristics of each parachute are discussed. Included are comments on canopy coning, spinning, and fluttering motions.

  19. Procedures for Computing Transonic Flows for Control of Adaptive Wind Tunnels. Ph.D. Thesis - Technische Univ., Berlin, Mar. 1986

    NASA Technical Reports Server (NTRS)

    Rebstock, Rainer

    1987-01-01

    Numerical methods are developed for control of three dimensional adaptive test sections. The physical properties of the design problem occurring in the external field computation are analyzed, and a design procedure suited for solution of the problem is worked out. To do this, the desired wall shape is determined by stepwise modification of an initial contour. The necessary changes in geometry are determined with the aid of a panel procedure, or, with incident flow near the sonic range, with a transonic small perturbation (TSP) procedure. The designed wall shape, together with the wall deflections set during the tunnel run, are the input to a newly derived one-step formula which immediately yields the adapted wall contour. This is particularly important since the classical iterative adaptation scheme is shown to converge poorly for 3D flows. Experimental results obtained in the adaptive test section with eight flexible walls are presented to demonstrate the potential of the procedure. Finally, a method is described to minimize wall interference in 3D flows by adapting only the top and bottom wind tunnel walls.

  20. Simulation of ideal-gas flow by nitrogen and other selected gases at cryogenic temperatures. [transonic flow in cryogenic wind tunnels

    NASA Technical Reports Server (NTRS)

    Hall, R. M.; Adcock, J. B.

    1981-01-01

    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.

  1. Canard-body-tail missile test at angles of attack to 50 deg in the Ames 11-foot transonic wind tunnel

    NASA Technical Reports Server (NTRS)

    Allen, C. Q.; Schwind, R. G.; Malcolm, G. N.

    1978-01-01

    Blunted ogive cylinder missile models with a length-to-diameter ratio of 10.4 were tested at transonic speeds and large angles of attack in an 11 foot transonic wind tunnel. The configurations are: body, body with tail panels, body with canards, and body with canards and tails. Forces and moments from the entire model and each of the eight fins were measured over the pitch range of 20 deg to 50 deg, and roll angles of 0 deg to 45 deg and canard deflection angles between 0 deg and 15 deg. The Reynolds number ranged from 3.9 x 10 to the 6th power per meter. Large side forces and yawing moments were observed for some of the test cases involving a symmetric geometry.

  2. Transonic Wind-Tunnel Tests of an Error-Compensated Static-Pressure Probe

    NASA Technical Reports Server (NTRS)

    Capone, Francis J.

    1961-01-01

    An investigation of the pressure-sensing characteristics of an error-compensated static-pressure probe mounted on the nose section of a missile body has been conducted in the Langley 16-foot transonic tunnel. The probe was free to rotate about its roll axis and was equipped with a vane so that the crossflow velocity component due to angles of attack or sideslip was always alined with the probe's vertical plane of symmetry. The probe was tested in five axial positions with respect to the missile nose at Mach numbers from 0.30 to 1.08 and at angles of attack from -2.7 to 15.3 deg. The test Reynolds number per foot varied from 1.79 x 10(exp 6) to 4.05 x 10(exp 6). Results showed that at a Mach number of 1.00 the static-pressure error decreased from 3.5 percent to 0.8 percent of the free-stream static pressure, as a result of a change in orifice location from 0.15 maximum missile diameter to 0.20 maximum missile diameter forward of the missile nose. Although compensation for pressure-sensing errors due to angles of attack up to 15.3 was maintained at Mach numbers from M = 0.30 to M = 0.50, there was an increase in error with an increase in angle of attack for Mach numbers between M 0.50 and M = 1.08.

  3. Correlation of transonic-cone preston-tube data and skin friction

    NASA Technical Reports Server (NTRS)

    Abu-Mostafa, A. S.; Reed, T. D.

    1984-01-01

    Preston-tube measurements obtained on the Arnold Engineering Development Center (AEDC) Transition Cone have been correlated with theoretical skin friction coefficients in transitional and turbulent flow. This has been done for the NASA Ames 11-Ft Transonic Wind Tunnel (11 TWT) and flight tests. The developed semi-empirical correlations of Preston-tube data have been used to derive a calibration procedure for the 11 TWT flow quality. This procedure has been applied to the corrected laminar data, and an effective freestream unit Reynolds number is defined by requiring a matching of the average Preston-tube pressure in flight and in the tunnel. This study finds that the operating Reynolds number is below the effective value required for a match in laminar Preston-tube data. The distribution of this effective Reynolds number with Mach number correlates well with the freestream noise level in this tunnel. Analyses of transitional and turbulent data, however, did not result in effective Reynolds numbers that can be correlated with background noise. This is a result of the fact that vorticity fluctuations present in transitional and turbulent boundary layers dominate Preston-tube pressure fluctuations and, therefore, mask the tunnel noise eff ects. So, in order to calibrate the effects of noise on transonic wind tunnel tests only laminar data should be used, preferably at flow conditions similar to those in flight tests. To calibrate the effects of transonic wind-tunnel noise on drag measurements, however, the Preston-tube data must be supplemented with direct measurements of skin friction.

  4. Fiber-optic-based laser vapor screen flow visualization system for aerodynamic research in larger scale subsonic and transonic wind tunnels

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.; Inenaga, Andrew S.

    1994-01-01

    Laser vapor screen (LVS) flow visualization systems that are fiber-optic based were developed and installed for aerodynamic research in the Langley 8-Foot Transonic Pressure Tunnel and the Langley 7- by 10-Foot High Speed Tunnel. Fiber optics are used to deliver the laser beam through the plenum shell that surrounds the test section of each facility and to the light-sheet-generating optics positioned in the ceiling window of the test section. Water is injected into the wind tunnel diffuser section to increase the relative humidity and promote condensation of the water vapor in the flow field about the model. The condensed water vapor is then illuminated with an intense sheet of laser light to reveal features of the flow field. The plenum shells are optically sealed; therefore, video-based systems are used to observe and document the flow field. Operational experience shows that the fiber-optic-based systems provide safe, reliable, and high-quality off-surface flow visualization in smaller and larger scale subsonic and transonic wind tunnels. The design, the installation, and the application of the Langley Research Center (LaRC) LVS flow visualization systems in larger scale wind tunnels are highlighted. The efficiency of the fiber optic LVS systems and their insensitivity to wind tunnel vibration, the tunnel operating temperature and pressure variations, and the airborne contaminants are discussed.

  5. Status of the KTH-NASA Wind-Tunnel Test for Acquisition of Transonic Nonlinear Aeroelastic Data

    NASA Technical Reports Server (NTRS)

    Silva, Walter A.; Ringertz, Ulf; Stenfelt, Gloria; Eller, David; Keller, Donald F.; Chwalowski, Pawel

    2016-01-01

    This paper presents a status report on the collaboration between the Royal Institute of Technology (KTH) in Sweden and the NASA Langley Research Center regarding the design, fabrication, modeling, and testing of a full-span lighter configuration in the Transonic Dynamics Tunnel (TDT). The goal of the test is to acquire transonic limit-cycle- oscillation (LCO) data, including accelerations, strains, and unsteady pressures. Finite element models (FEMs) and aerodynamic models are presented and discussed along with results obtained to date.

  6. Unsteady transonic aerodynamics

    SciTech Connect

    Nixon, D.

    1989-01-01

    Various papers on unsteady transonic aerodynamics are presented. The topics addressed include: physical phenomena associated with unsteady transonic flows, basic equations for unsteady transonic flow, practical problems concerning aircraft, basic numerical methods, computational methods for unsteady transonic flows, application of transonic flow analysis to helicopter rotor problems, unsteady aerodynamics for turbomachinery aeroelastic applications, alternative methods for modeling unsteady transonic flows.

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

    NASA Technical Reports Server (NTRS)

    Lamar, John E.; Obara, Clifford J.; Fisher, Bruce D.; Fisher, David F.

    2001-01-01

    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.

  8. Study of design and analysis methods for transonic flow

    NASA Technical Reports Server (NTRS)

    Murman, E. M.

    1977-01-01

    An airfoil design program and a boundary layer analysis were developed. Boundary conditions were derived for ventilated transonic wind tunnels and performing transonic windtunnel wall calculations. A computational procedure for rotational transonic flow in engine inlet throats was formulated. Results and conclusions are summarized.

  9. Wind Tunnel Investigation of Passive Vortex Control and Vortex-Tail Interactions on a Slender Wing at Subsonic and Transonic Speeds

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.

    2013-01-01

    A wind tunnel experiment was conducted in the NASA Langley 8-Foot Transonic Pressure Tunnel to determine the effects of passive porosity on vortex flow interactions about a slender wing configuration at subsonic and transonic speeds. Flow-through porosity was applied in several arrangements to a leading-edge extension, or LEX, mounted to a 65-degree cropped delta wing as a longitudinal instability mitigation technique. Test data were obtained with LEX on and off in the presence of a centerline vertical tail and twin, wing-mounted vertical fins to quantify the sensitivity of the aerodynamics to tail placement and orientation. A close-coupled canard was tested as an alternative to the LEX as a passive flow control device. Wing upper surface static pressure distributions and six-component forces and moments were obtained at Mach numbers of 0.50, 0.85, and 1.20, unit Reynolds number of 2.5 million, angles of attack up to approximately 30 degrees, and angles of sideslip to +/-8 degrees. The off-surface flow field was visualized in cross planes on selected configurations using a laser vapor screen flow visualization technique. Tunnel-to-tunnel data comparisons and a Reynolds number sensitivity assessment were also performed. 15.

  10. Effects of flaps on buffet characteristics and wind-rock onset of an F-8C airplane at subsonic and transonic speeds

    NASA Technical Reports Server (NTRS)

    Monaghan, R. C.; Friend, E. L.

    1973-01-01

    Wind-up-turn maneuvers were performed to establish the values of airplane normal force coefficient for buffet onset, wing-rock onset, and buffet loads with various combinations of leading- and trailing-edge flap deflections. Data were gathered at both subsonic and transonic speeds covering a range from Mach 0.64 to Mach 0.92. Buffet onset and buffet loads were obtained from wingtip acceleration and wing-root bending-moment data, and wing-rock onset was obtained from airplane roll rate data. Buffet onset, wing-rock onset, and buffet loads were similarly affected by the various combinations of leading- and training-edge flaps. Subsonically, the 12 deg leading-edge-flap and trailing-edge-flap combination was most effective in delaying buffet onset, wing-rock onset, and equivalent values of buffet loads to a higher value of airplane normal force coefficient. This was the maximum flap deflection investigated. Transonically, however, the optimum leading-edge flap position was generally less than 12 deg.

  11. A two-dimensional adaptive-wall test section with ventilated walls in the Ames 2- by 2-foot transonic wind tunnel

    NASA Technical Reports Server (NTRS)

    Schairer, Edward T.; Lee, George; Mcdevitt, T. Kevin

    1989-01-01

    The first tests conducted in the adaptive-wall test section of the Ames Research Center's 2- by 2-Foot Transonic Wind Tunnel are described. A procedure was demonstrated for reducing wall interference in transonic flow past a two-dimensional airfoil by actively controlling flow through the slotted walls of the test section. Flow through the walls was controlled by adjusting pressures in compartments of plenums above and below the test section. Wall interference was assessed by measuring (with a laser velocimeter) velocity distributions along a contour surrounding the model, and then checking those measurements for their compatibility with free-air far-field boundary conditions. Plenum pressures for minimum wall interference were determined from empirical influence coefficients. An NACA 0012 airfoil was tested at angles of attach of 0 and 2, and at Mach numbers between 0.70 and 0.85. In all cases the wall-setting procedure greatly reduced wall interference. Wall interference, however, was never completely eliminated, primarily because the effect of plenum pressure changes on the velocities along the contour could not be accurately predicted.

  12. Calculations Of Transonic Flow About A Wing

    NASA Technical Reports Server (NTRS)

    Holst, Terry L.; Gundy, Karen L.; Flores, Jolen; Chaderjian, Neal; Kaynak, Univer; Thomas, Scott D.

    1988-01-01

    Report describes calculations of transonic airflows about wing in wind tunnel. Basic equations of flow used in study are Reynolds-averaged Navier-Stokes equations in strong conservation-law form. Equations of flow incorporated into finite-difference computer code called TNS (Transonic Navier-Stokes). Computational grid generated by solution of partial differential equations yielding smooth meshes conforming to surfaces of wing and wind tunnel.

  13. Transonic empirical configuration design process

    NASA Technical Reports Server (NTRS)

    Whitcomb, R. T.

    1983-01-01

    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.

  14. Transonic Symposium: Theory, Application, and Experiment, volume 1, part 2

    NASA Technical Reports Server (NTRS)

    Foughner, Jerome T., Jr. (Compiler)

    1989-01-01

    In order to assess the state of the art in transonic flow disciplines and to glimpse at future directions, NASA-Langley held a Transonic Symposium. Emphasis was placed on steady, three dimensional external, transonic flow and its simulation, both numerically and experimentally. The symposium included technical sessions on wind tunnel and flight experiments; computational fluid dynamic applications; inviscid methods and grid generation; viscous methods and boundary layer stability; and wind tunnel techniques and wall interference. This, being volume 1, is unclassified.

  15. Results of wind tunnel tests of an ASRM configured 0.03 scale Space Shuttle integrated vehicle model (47-OTS) in the AEDC 16-foot Transonic wind tunnel (IA613A), volume 1

    NASA Astrophysics Data System (ADS)

    Marroquin, J.; Lemoine, P.

    1992-10-01

    An experimental Aerodynamic and Aero-Acoustic loads data base was obtained at transonic Mach numbers for the Space Shuttle Launch Vehicle configured with the ASRM Solid Rocket Boosters as an increment to the current flight configuration (RSRB). These data were obtained during transonic wind tunnel tests (IA 613A) conducted in the Arnold Engineering Development Center 16-Foot transonic propulsion wind tunnel from March 27, 1991 through April 12, 1991. This test is the first of a series of two tests covering the Mach range from 0.6 to 3.5. Steady state surface static and fluctuating pressure distributions over the Orbiter, External Tank and Solid Rocket Boosters of the Shuttle Integrated Vehicle were measured. Total Orbiter forces, Wing forces and Elevon hinge moments were directly measured as well from force balances. Two configurations of Solid Rocket Boosters were tested, the Redesigned Solid Rocket Booster (RSRB) and the Advanced Solid Rocket Motor (ASRM). The effects of the position (i.e. top, bottom, top and bottom) of the Integrated Electronics Assembly (IEA) box, mounted on the SRB attach ring, were obtained on the ASRM configured model. These data were obtained with and without Solid Plume Simulators which, when used, matched as close as possible the flight derived pressures on the Orbiter and External Tank base. Data were obtained at Mach numbers ranging from 0.6 to 1.55 at a Unit Reynolds Number of 2.5 million per foot through model angles of attack from -8 to +4 degrees at sideslip angles of 0, +4 and -4 degrees.

  16. Results of wind tunnel tests of an ASRM configured 0.03 scale Space Shuttle integrated vehicle model (47-OTS) in the AEDC 16-foot transonic wind tunnel, volume 2

    NASA Astrophysics Data System (ADS)

    Marroquin, J.; Lemoine, P.

    1992-10-01

    An experimental Aerodynamic and Aero-Acoustic loads data base was obtained at transonic Mach numbers for the Space Shuttle Launch Vehicle configured with the ASRM Solid Rocket Boosters as an increment to the current flight configuration (RSRB). These data were obtained during transonic wind tunnel tests (IA 613A) conducted in the Arnold Engineering Development Center 16-Foot transonic propulsion wind tunnel from March 27, 1991 through April 12, 1991. This test is the first of a series of two tests covering the Mach range from 0.6 to 3.5. Steady state surface static and fluctuating pressure distributions over the Orbiter, External Tank and Solid Rocket Boosters of the Shuttle Integrated Vehicle were measured. Total Orbiter forces, Wing forces and Elevon hinge moments were directly measured as well from force balances. Two configurations of Solid Rocket Boosters were tested, the Redesigned Solid Rocket Booster (RSRB) and the Advanced Solid Rocket Motor (ASRM). The effects of the position (i.e., top, bottom, top and bottom) of the Integrated Electronics Assembly (IEA) box, mounted on the SRB attach ring, were obtained on the ASRM configured model. These data were obtained with and without Solid Plume Simulators which, when used, matched as close as possible the flight derived pressures on the Orbiter and External Tank base. Data were obtained at Mach numbers ranging from 0.6 to 1.55 at a Unit Reynolds Number of 2.5 million per foot through model angles of attack from -8 to +4 degrees at sideslip angles of 0, +4 and -4 degrees.

  17. Results of wind tunnel tests of an ASRM configured 0.03 scale Space Shuttle integrated vehicle model (47-OTS) in the AEDC 16-foot transonic wind tunnel, volume 2

    NASA Technical Reports Server (NTRS)

    Marroquin, J.; Lemoine, P.

    1992-01-01

    An experimental Aerodynamic and Aero-Acoustic loads data base was obtained at transonic Mach numbers for the Space Shuttle Launch Vehicle configured with the ASRM Solid Rocket Boosters as an increment to the current flight configuration (RSRB). These data were obtained during transonic wind tunnel tests (IA 613A) conducted in the Arnold Engineering Development Center 16-Foot transonic propulsion wind tunnel from March 27, 1991 through April 12, 1991. This test is the first of a series of two tests covering the Mach range from 0.6 to 3.5. Steady state surface static and fluctuating pressure distributions over the Orbiter, External Tank and Solid Rocket Boosters of the Shuttle Integrated Vehicle were measured. Total Orbiter forces, Wing forces and Elevon hinge moments were directly measured as well from force balances. Two configurations of Solid Rocket Boosters were tested, the Redesigned Solid Rocket Booster (RSRB) and the Advanced Solid Rocket Motor (ASRM). The effects of the position (i.e., top, bottom, top and bottom) of the Integrated Electronics Assembly (IEA) box, mounted on the SRB attach ring, were obtained on the ASRM configured model. These data were obtained with and without Solid Plume Simulators which, when used, matched as close as possible the flight derived pressures on the Orbiter and External Tank base. Data were obtained at Mach numbers ranging from 0.6 to 1.55 at a Unit Reynolds Number of 2.5 million per foot through model angles of attack from -8 to +4 degrees at sideslip angles of 0, +4 and -4 degrees.

  18. Results of wind tunnel tests of an ASRM configured 0.03 scale Space Shuttle integrated vehicle model (47-OTS) in the AEDC 16-foot Transonic wind tunnel (IA613A), volume 1

    NASA Technical Reports Server (NTRS)

    Marroquin, J.; Lemoine, P.

    1992-01-01

    An experimental Aerodynamic and Aero-Acoustic loads data base was obtained at transonic Mach numbers for the Space Shuttle Launch Vehicle configured with the ASRM Solid Rocket Boosters as an increment to the current flight configuration (RSRB). These data were obtained during transonic wind tunnel tests (IA 613A) conducted in the Arnold Engineering Development Center 16-Foot transonic propulsion wind tunnel from March 27, 1991 through April 12, 1991. This test is the first of a series of two tests covering the Mach range from 0.6 to 3.5. Steady state surface static and fluctuating pressure distributions over the Orbiter, External Tank and Solid Rocket Boosters of the Shuttle Integrated Vehicle were measured. Total Orbiter forces, Wing forces and Elevon hinge moments were directly measured as well from force balances. Two configurations of Solid Rocket Boosters were tested, the Redesigned Solid Rocket Booster (RSRB) and the Advanced Solid Rocket Motor (ASRM). The effects of the position (i.e. top, bottom, top and bottom) of the Integrated Electronics Assembly (IEA) box, mounted on the SRB attach ring, were obtained on the ASRM configured model. These data were obtained with and without Solid Plume Simulators which, when used, matched as close as possible the flight derived pressures on the Orbiter and External Tank base. Data were obtained at Mach numbers ranging from 0.6 to 1.55 at a Unit Reynolds Number of 2.5 million per foot through model angles of attack from -8 to +4 degrees at sideslip angles of 0, +4 and -4 degrees.

  19. Geared-elevator flutter study. [wind tunnel tests of transonic flutter effects on control surfaces of supersonic transport tail assemblies, conducted in a NASA-Langley transonic wind tunnel

    NASA Technical Reports Server (NTRS)

    Ruhlin, C. L.; Doggett, R. V., Jr.; Gregory, R. A.

    1976-01-01

    An experimental and analytical study was made of the transonic flutter characteristics of a supersonic transport tail assembly model having an all-movable, horizontal tail with a geared elevator. Two model configurations, namely, one with a gear-elevator (2.8 to 1.0 gear ratio) and one with locked-elevator (1.0 to 1.0 gear ratio), were flutter tested in the Langley transonic dynamics tunnel with an empennage cantilever-mounted on a sting. The geared-elevator configuration fluttered experimentally at about 20% higher dynamic pressures than the locked-elevator configuration. The experimental flutter dynamic pressure boundaries for both configurations were nearly flat over a Mach number range from 0.9 to 1.1. Flutter calculations (mathematical models) were made for the geared-elevator configuration using three subsonic lifting-surface methods. In one method, the elevator was treated as a discrete surface, and in the other two methods, the stabilizer and elevator were treated as a single warped-surface with the primary difference between these two methods being in the mathematical implementation used. A comparison of the experimental and analytical results shows that the discrete-elevator method predicted best the experimental flutter dynamic pressure level. However, the single warped-surface methods predicts more closely the experimental flutter frequencies and Mach number trends.

  20. Boundary layer separation on isolated boattail nozzles. M.S. Thesis; [conducted in the Langley 16-foot transonic wind tunnel

    NASA Technical Reports Server (NTRS)

    Abeyounis, W. K.

    1977-01-01

    The phenomenon of separated flow on a series of circular-arc afterbodies was investigated using the Langley 16-foot transonic tunnel at free-stream Mach numbers from 0.40 to 0.95 at 0 deg angle of attack. Both high-pressure air and solid circular cylinders with a diameter equal to the nozzle exit diameter were used to simulate jet exhausts. A detailed data base of boundary layer separation locations was obtained using oil-flow techniques. The results indicate that boundary layer separation is most extensive on steep boattails at high Mach numbers.

  1. Transonic and Supersonic Wind-Tunnel Tests of Wing-Body Combinations Designed for High Efficiency at a Mach Number of 1.41

    NASA Technical Reports Server (NTRS)

    Grant, Frederick C.; Sevier, John R., Jr.

    1960-01-01

    Wind-tunnel force tests of a number of wing-body combinations designed for high lift-drag ratio at a Mach number of 1.41 are reported. Five wings and six bodies were used in making up the various wing-body combinations investigated. All the wings had the same highly swept dis- continuously tapered plan form with NACA 65A-series airfoil sections 4 percent thick at the root tapering linearly to 3 percent thick at the tip. The bodies were based on the area distribution of a Sears-Haack body of revolution for minimum drag with a given length and volume. These wings and bodies were used to determine the effects of wing twist., wing twist and camber, wing leading-edge droop, a change from circular to elliptical body cross-sectional shape, and body indentation by the area-rule and streamline methods. The supersonic test Mach numbers were 1.41 and 2.01. The transonic test Mach number range was from 0.6 to 1.2. For the transition-fixed condition and at a Reynolds number of 2.7 x 10(exp 6) based on the mean aerodynamic chord, the maximum value of lift- drag ratio at a Mach number of 1.41 was 9.6 for a combination with a twisted wing and an indented body of elliptical cross section. The tests indicated that the transonic rise in minimum drag was low and did not change appreciably up to the highest test Mach number of 2.01. The lower values of lift-drag ratio obtained at a Mach number of 2.01 can be attributed to the increase of drag due to lift with Mach number.

  2. Wind Tunnel Application of a Pressure-Sensitive Paint Technique to a Double Delta Wing Model at Subsonic and Transonic Speeds

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.; Gonzalez, Hugo A.

    2006-01-01

    A pressure-sensitive paint (PSP) technique was applied in a wind tunnel experiment in the NASA Langley Research Center 8-Foot Transonic Pressure Tunnel to study the effect of wing fillets on the global vortex induced surface static pressure field about a sharp leading-edge 76 deg./40 deg. double delta wing, or strake-wing, model at subsonic and transonic speeds. Global calibrations of the PSP were obtained at M(sub infinity) = 0.50, 0.70, 0.85, 0.95, and 1.20, a Reynolds number per unit length of 2.0 million, and angles of attack from 10 degrees to 20 degrees using an insitu method featuring the simultaneous acquisition of electronically scanned pressures (ESP) at discrete locations on the model. The mean error in the PSP measurements relative to the ESP data was approximately 2 percent or less at M(sub infinity) = 0.50 to 0.85 but increased to several percent at M(sub infinity) =0.95 and 1.20. The PSP pressure distributions and pseudo-colored, planform-view pressure maps clearly revealed the vortex-induced pressure signatures at all Mach numbers and angles of attack. Small fillets having parabolic or diamond planforms situated at the strake-wing intersection were respectively designed to manipulate the vortical flows by removing the leading-edge discontinuity or introducing additional discontinuities. The fillets caused global changes in the vortex-dominated surface pressure field that were effectively captured in the PSP measurements. The vortex surface pressure signatures were compared to available off-surface vortex cross-flow structures obtained using a laser vapor screen (LVS) flow visualization technique. The fillet effects on the PSP pressure distributions and the observed leading-edge vortex flow characteristics were consistent with the trends in the measured lift, drag, and pitching moment coefficients.

  3. Results of flutter test OS7 obtained using the 0.14-scale space shuttle orbiter fin/rudder model number 55-0 in the NASA LaRC 16-foot transonic dynamics wind tunnel

    NASA Technical Reports Server (NTRS)

    Berthold, C. L.

    1977-01-01

    A 0.14-scale dynamically scaled model of the space shuttle orbiter vertical tail was tested in a 16-foot transonic dynamic wind tunnel to determine flutter, buffet, and rudder buzz boundaries. Mach numbers between .5 and 1.11 were investigated. Rockwell shuttle model 55-0 was used for this investigation. A description of the test procedure, hardware, and results of this test is presented.

  4. Results of flutter test OS6 obtained using the 0.14-scale wing/elevon model (54-0) in the NASA LaRC 16-foot transonic dynamics wind tunnel

    NASA Technical Reports Server (NTRS)

    Berthold, C. L.

    1977-01-01

    A 0.14-scale dynamically scaled model of the space shuttle orbiter wing was tested in the Langley Research Center 16-Foot Transonic Dynamics Wind Tunnel to determine flutter, buffet, and elevon buzz boundaries. Mach numbers between 0.3 and 1.1 were investigated. Rockwell shuttle model 54-0 was utilized for this investigation. A description of the test procedure, hardware, and results of this test is presented.

  5. Transonic Symposium: Theory, Application and Experiment, volume 2

    NASA Technical Reports Server (NTRS)

    Foughner, Jerome T., Jr. (Compiler)

    1989-01-01

    Papers presented at the Transonic Symposium are compiled. The following subject areas are covered: National Transonic Facility status; transonic aerodynamics of slender wing-body configuration; laminar flow flight experiments; laminar flow wind tunnel experiments; computational support of X-29A flight experiment; transition location on a clean-up glove installed on a F-14 aircraft; and design studies for a laminar glove for the X-29 aircraft.

  6. Transonic Airfoil Development

    NASA Technical Reports Server (NTRS)

    Whitcomb, R. T.

    1983-01-01

    This lecture consists of three parts, in which discussions are presented of the current state of development of transonic or supercritical airfoils designed for fully turbulent boundary layers on the surfaces, previous research on subcritical airfoils designed to achieve laminar boundary layers on all or parts of the surfaces, and current research on supercritical airfoils designed to achieve laminar boundary layers. In the first part the use of available two dimensional computer codes in the development of supercritical airfoils and the general trends in the design of such airfoils with turbulent boundary layers are discussed. The second part provides the necessary background on laminar boundary layer phenomena. The last part, which constitutes the major portion of the lecture, covers research by NASA on supercritical airfoils utilizing both decreasing pressure gradients and surface suction for stabilizing the laminar boundary layer. An investigation of the former has been recently conducted in fight using gloves on the wing panels of the U.S. Air Force F111 TACT airplane, research on the later is currently being conducted in a transonic wind tunnel which has been modified to greatly reduce the stream turbulence and noise levels in the tests section.

  7. Report on tests of a CAST 10 airfoil with fixed transition in the T2 transonic cryogenic wind tunnel with self-adaptive walls

    NASA Technical Reports Server (NTRS)

    Seraudie, A.; Blanchard, A.; Breil, J. F.

    1985-01-01

    Described are tests on the CAST 10 airfoil in tripped-transition, carried out in the cryogenic and transonic wind-tunnel T2 fitted with self-adaptive walls. These tests follow those which were performed in natural transition and were presented in a previous note. Firstly, a complement was realized to pinpoint the location of the natural transition on the upper surface of the airfoil; this was done by a longitudinal exploration in the boundary layer. Secondly, in a first stage, the transition was only tripped on the lower surface with a carborundum strip of 0.045 mm thickness, situated at 5% of chord (T 1/2 D). These tests were performed here to separate the phenomena in relation to the lower surface and those in relation to the upper surface which occur in natural transition (TN). In a second stage, the transition was normally tripped on both sides of the profile (TD), likewise at x/c = 5% and h = 0.045 mm. The test configurations of the previous serial were experimented again and results obtained in the three cases (TN), (T 1/2 N) and (TD) were compared, in particular those concerned with the effect of the Reynolds number on aerodynamic coefficients of the airfoil. The gathering of the experimental values around a Reynolds number of 20 millions is observed; but before this number, the evolutions of the curves in the three cases tested are different.

  8. Time-Accurate Unsteady Pressure Loads Simulated for the Space Launch System at Wind Tunnel Conditions

    NASA Technical Reports Server (NTRS)

    Alter, Stephen J.; Brauckmann, Gregory J.; Kleb, William L.; Glass, Christopher E.; Streett, Craig L.; Schuster, David M.

    2015-01-01

    A transonic flow field about a Space Launch System (SLS) configuration was simulated with the Fully Unstructured Three-Dimensional (FUN3D) computational fluid dynamics (CFD) code at wind tunnel conditions. Unsteady, time-accurate computations were performed using second-order Delayed Detached Eddy Simulation (DDES) for up to 1.5 physical seconds. The surface pressure time history was collected at 619 locations, 169 of which matched locations on a 2.5 percent wind tunnel model that was tested in the 11 ft. x 11 ft. test section of the NASA Ames Research Center's Unitary Plan Wind Tunnel. Comparisons between computation and experiment showed that the peak surface pressure RMS level occurs behind the forward attach hardware, and good agreement for frequency and power was obtained in this region. Computational domain, grid resolution, and time step sensitivity studies were performed. These included an investigation of pseudo-time sub-iteration convergence. Using these sensitivity studies and experimental data comparisons, a set of best practices to date have been established for FUN3D simulations for SLS launch vehicle analysis. To the author's knowledge, this is the first time DDES has been used in a systematic approach and establish simulation time needed, to analyze unsteady pressure loads on a space launch vehicle such as the NASA SLS.

  9. The transonic Reynolds number problem. [limitations of transonic aerodynamic test facilities

    NASA Technical Reports Server (NTRS)

    Jones, J. L.

    1977-01-01

    Problems in modeling the complex interacting flow fields in the transonic speed regime are reviewed. The limitations of wind tunnel test capabilities are identified, and options for resolving the deficiency are examined. The evolution of the National Transonic Facility, and the various needs for research investigations to be done there are discussed. The relative priorities that should be given within and across subdisciplines for guidance in planning for the most effective use of the facility are considered.

  10. Transonic control effectiveness for full and partial span elevon configurations on a 0.0165 scale model space shuttle orbiter tested in the LaRC 8-foot transonic wind tunnel (LA48)

    NASA Technical Reports Server (NTRS)

    1977-01-01

    A transonic pressure tunnel test is reported on an early version of the space shuttle orbiter (designated 089B-139) 0.0165 scale model to systematically determine both longitudinal and lateral control effectiveness associated with various combinations of inboard, outboard, and full span wing trailing edge controls. The test was conducted over a Mach number range from 0.6 to 1.08 at angles of attack from -2 deg to 23 deg at 0 deg sideslip.

  11. Preliminary results of buffet tests in a cryogenic wind tunnel. [conducted in Langley 0.3 m transonic cryogenic tunnel

    NASA Technical Reports Server (NTRS)

    Boyden, R. P.; Johnson, W. G., Jr.

    1981-01-01

    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.

  12. National Transonic Facility: A review of the operational plan

    NASA Technical Reports Server (NTRS)

    Liepmann, H. W.; Black, R. E.; Dietz, R. O.; Kirchner, M. E.; Sears, W. R.

    1980-01-01

    The proposed National Transonic Facility (NTF) operational plan is reviewed. The NTF will provide an aerodynamic test capability significantly exceeding that of other transonic regime wind tunnels now available. A limited number of academic research program that might use the NTF are suggested. It is concluded that the NTF operational plan is useful for management, technical, instrumentation, and model building techniques available in the specialized field of aerodynamic analysis and simulation. It is also suggested that NASA hold an annual conference to discuss wind tunnel research results and to report on developments that will further improve the utilization and cost effectiveness of the NTF and other wind tunnels.

  13. Results of tests OA63 and IA29 on an 0.015 scale model of the space shuttle configuration 140 A/B in the NASA/ARC 6- by 6-foot transonic wind tunnel, volume 1

    NASA Technical Reports Server (NTRS)

    Spangler, R. H.; Thornton, D. E.

    1974-01-01

    Tests were conducted in the NASA/ARC 6- by 6-foot transonic wind tunnel from September 12 to September 28, 1973 on an 0.015-scale model of the space shuttle configuration 140 A/B. Surface pressure data were obtained for the orbiter for both launch and entry configuration at Mach numbers from 0.6 to 2.0. The surface pressures were obtained in the vicinity of the cargo bay door hinge and parting lines, the side of the fuselage at the crew compartment and below the OMS pods at the aft compartment. Data were obtained at angles of attack and sideslip consistent with the expected divergencies along the nominal trajectory. These tests were first in a series of tests supporting the orbiter venting analysis. The series will include tests in three facilities covering a total Mach number range from 0.6 to 10.4.

  14. Results of test IA137 in the NASA/ARC 14 foot transonic wind tunnel of the 0.07 scale external tank forebody (model 68-T) to determine auxiliary aerodynamic data system feasibility

    NASA Technical Reports Server (NTRS)

    Thornton, D. E.

    1976-01-01

    Tests were conducted in a 14 foot transonic wind tunnel to examine the feasibility of the auxiliary aerodynamic data system (AADS) for determining angles of attack and sideslip during boost flight. The model used was a 0.07 scale replica of the external tank forebody consisting of the nose portion and a 60 inch (full scale) cylindrical section of the ogive cylinder tangency point. The model terminated in a blunt base with a 320.0 inch diameter at external tank (ET) station 1120.37. Pressure data were obtained from five pressure orifices (one total and four statics) on the nose probe, and sixteen surface static pressure orifices along the ET forebody.

  15. Results of the space shuttle vehicle ascent air data system probe calibration test using a 0.07-scale external tank forebody model (68T) in the AEDC 16-foot transonic wind tunnel (IA-310), volume 1

    NASA Technical Reports Server (NTRS)

    Collette, J. G. R.

    1991-01-01

    A recalibration of the Space Shuttle Vehicle Ascent Air Data System probe was conducted in the Arnold Engineering Development Center (AEDC) transonic wind tunnel. The purpose was to improve on the accuracy of the previous calibration in order to reduce the existing uncertainties in the system. A probe tip attached to a 0.07-scale External Tank Forebody model was tested at angles of attack of -8 to +4 degrees and sideslip angles of -4 to +4 degrees. High precision instrumentation was used to acquire pressure data at discrete Mach numbers ranging from 0.6 to 1.55. Pressure coefficient uncertainties were estimated at less than 0.0020. Data is given in graphical and tabular form.

  16. Results of the space shuttle vehicle ascent air data system probe calibration test using a 0.07-scale external tank forebody model (68T) in the AEDC 16-foot transonic wind tunnel (IA-310), volume 2

    NASA Technical Reports Server (NTRS)

    Collette, J. G. R.

    1991-01-01

    A recalibration of the Space Shuttle Vehicle Ascent Air Data System probe was conducted in the Arnold Engineering and Development Center (AEDC) transonic wind tunnel. The purpose was to improve on the accuracy of the previous calibration in order to reduce the existing uncertainties in the system. A probe tip attached to a 0.07-scale External Tank Forebody model was tested at angles of attack of -8 to +4 degrees and sideslip angles of -4 to +4 degrees. High precision instrumentation was used to acquire pressure data at discrete Mach numbers ranging from 0.6 to 1.55. Pressure coefficient uncertainties were estimated at less than 0.0020. Additional information is given in tabular form.

  17. Axisymmetric & non-axisymmetric exhaust jet induced-effects on a V/STOL vehicle design. Part 1: Data presentation. [conducted in Ames 11-foot transonic tunnel

    NASA Technical Reports Server (NTRS)

    Schnell, W. C.; Ordonez, G. W.

    1981-01-01

    A 1/8 scale jet-effects model was tested in the NASA Ames 11 ft transonic tunnel at static conditions and over a range of Mach numbers from 0.4 to 1.4. The data presented show that significant differences in aeropropulsion performance can be expected by varying the exhaust nozzle type and its geometric parameters on a V/STOL underwing nacelle installation.

  18. Investigations for Supersonic Transports at Transonic and Supersonic Conditions

    NASA Technical Reports Server (NTRS)

    Rivers, S. Melissa B.; Owens, Lewis R.; Wahls, Richard A.

    2007-01-01

    Several computational studies were conducted as part of NASA s High Speed Research Program. Results of turbulence model comparisons from two studies on supersonic transport configurations performed during the NASA High-Speed Research program are given. The effects of grid topology and the representation of the actual wind tunnel model geometry are also investigated. Results are presented for both transonic conditions at Mach 0.90 and supersonic conditions at Mach 2.48. A feature of these two studies was the availability of higher Reynolds number wind tunnel data with which to compare the computational results. The transonic wind tunnel data was obtained in the National Transonic Facility at NASA Langley, and the supersonic data was obtained in the Boeing Polysonic Wind Tunnel. The computational data was acquired using a state of the art Navier-Stokes flow solver with a wide range of turbulence models implemented. The results show that the computed forces compare reasonably well with the experimental data, with the Baldwin-Lomax with Degani-Schiff modifications and the Baldwin-Barth models showing the best agreement for the transonic conditions and the Spalart-Allmaras model showing the best agreement for the supersonic conditions. The transonic results were more sensitive to the choice of turbulence model than were the supersonic results.

  19. A Transonic Wind-Tunnel Investigation of a Seaplane Configuration having a 40 Deg Sweptback Wing, TED No. NACA DE 387

    NASA Technical Reports Server (NTRS)

    Hieser, Gerald; Kudlacik, Louis; Gray, W. H.

    1956-01-01

    During the course of an aerodynamic loads investigation of a model of the Martin XP6M-1 flying boat in the.Langley 16-foot transonic tunnel, longitudinal-aerodynamic-performance information was obtained. Data were obtained at speeds up to and exceeding those anticipated for the seaplane in level flight and included the Mach number range from 0.84. to 1.09. The angle of attack was varied from -2deg to 6deg and the average Reynolds number, based on wing mean aerodyn&ic chord, was about 3.7 x 10(exp 6). This seaplane, although not designed to maintain level flight at Mach numbers beyond the force break, was found to have a transonic drag-rise coefficient of 0.0728, with an accompanying drag-rise Mach number of about 0.85. A large portion of the.drag rise and the relatively low value of drag-rise Mach number result from the axial coincidence of the maximum areas of the principal airplane components.

  20. Results of tests CS4 and CS5 to investigate dynamic loads and pressures on 0.03-scale models (Ax1319-3/4 and 45-0) of mated 747 cam and space shuttle orbiter in the Boeing transonic wind tunnel

    NASA Technical Reports Server (NTRS)

    1976-01-01

    A 0.03-scale model of the 747 CAM/Orbiter was tested in an 8 x 12 foot transonic wind tunnel. Dynamic loads, pressure, and empennage flow field data were obtained using pressure transducers, strain gages, and a split film anemometer. The test variables included Mach number, angle of attack, sideslip angle, orbiter tailcone on and off, orbiter partial tailcone, orbiter nozzle air scoops, orbiter body flap angle, and orbiter elevon angle.

  1. Transonic Wind-Tunnel Investigation of the Fin Loads on a 1/8-Scale Model Simulating the First Stage of the Scout Research Vehicle

    NASA Technical Reports Server (NTRS)

    Kelly, Thomas C.

    1961-01-01

    An investigation to determine the fin loads on a 1/8-scale model simulating the first stage of the Scout research vehicle was made in the Langley 8-foot transonic tunnel at Mach numbers from 0.40 to 1.20. Tests were conducted over an angle-of-attack range from about -10 to 10 deg and at a Reynolds number per foot of approximately 3.5 x 10(exp 6). Results of the tests indicate that for a given angle of attack, negative tip-control deflections caused decreases in normal-force and fin-bending-moment coefficients and increases in pitching-moment coefficient, as would be expected. The effects were slight at a model angle of attack of -10 deg where tip-control stall had probably occurred but increased with an increase in angle of attack.

  2. Wind-tunnel tests of a 1/4 scale model of the Bell XS-1 transonic airplane. 1: Longitudinal stability and control characteristics

    NASA Technical Reports Server (NTRS)

    Donlan, C. J.; Kemp, W. B., Jr.; Polhamus, E. C.

    1976-01-01

    A 1/4 scale model of the Bell XS-1 transonic aircraft was tested in the Langley 300 mile-per-hour 7 by 10 foot tunnel to determine its low speed longitudinal stability and control characteristics. Pertinent longitudinal flying qualities expected of the XS-1 research airplane were estimated from the results of these tests including the effects of compressibility likely to be encountered at speeds below the force break. It appears that the static longitudinal stability and elevator control power will be adequate, but that the elevator control force gradient in steady flight will be undesirably low for all configurations. It is suggested that a centering spring be incorporated in the elevator control system of the airplane in order to increase the control force gradient in steady flight and in maneuvers.

  3. Transonic and supersonic ground effect aerodynamics

    NASA Astrophysics Data System (ADS)

    Doig, G.

    2014-08-01

    A review of recent and historical work in the field of transonic and supersonic ground effect aerodynamics has been conducted, focussing on applied research on wings and aircraft, present and future ground transportation, projectiles, rocket sleds and other related bodies which travel in close ground proximity in the compressible regime. Methods for ground testing are described and evaluated, noting that wind tunnel testing is best performed with a symmetry model in the absence of a moving ground; sled or rail testing is ultimately preferable, though considerably more expensive. Findings are reported on shock-related ground influence on aerodynamic forces and moments in and accelerating through the transonic regime - where force reversals and the early onset of local supersonic flow is prevalent - as well as more predictable behaviours in fully supersonic to hypersonic ground effect flows.

  4. Analysis of a theoretically optimized transonic airfoil

    NASA Technical Reports Server (NTRS)

    Lores, M. E.; Burdges, K. P.; Shrewsbury, G. D.

    1978-01-01

    Numerical optimization was used in conjunction with an inviscid, full potential equation, transonic flow analysis computer code to design an upper surface contour for a conventional airfoil to improve its supercritical performance. The modified airfoil was tested in a compressible flow wind tunnel. The modified airfoil's performance was evaluated by comparison with test data for the baseline airfoil and for an airfoil developed by optimization of leading edge of the baseline airfoil. While the leading edge modification performed as expected, the upper surface re-design did not produce all of the expected performance improvements. Theoretical solutions computed using a full potential, transonic airfoil code corrected for viscosity were compared to experimental data for the baseline airfoil and the upper surface modification. These correlations showed that the theory predicted the aerodynamics of the baseline airfoil fairly well, but failed to accurately compute drag characteristics for the upper surface modification.

  5. Kuechemann Carrots for transonic drag reduction.

    NASA Astrophysics Data System (ADS)

    Bechert, D. W.; Hage, W.; Stanewsky, E.

    1999-11-01

    Wave drag reduction bodies on the suction side of transonic wings are investigated. Following the original invention by O. Frenzl (1942), subsequently, such bodies have been suggested by Kuechemann and Whitcomb. These devices have been used sucessfully on various TUPOLEV aircraft and on the CONVAIR 990 airliner. New transonic wind tunnel data from an unswept wing with an array of Kuechemann Carrots are presented (airfoil: CAST 10/DOA-2). In a certain parameter range (M= 0.765-0.86) the measurements exhibit a significant reduction of the shock strength on a wing between the Kuechemann Carrots. This entails a dramatic reduction of drag, in a certain Mach number and angular regime up to 50-60%.

  6. Buffet test in the National Transonic Facility

    NASA Technical Reports Server (NTRS)

    Young, Clarence P., Jr.; Hergert, Dennis W.; Butler, Thomas W.; Herring, Fred M.

    1992-01-01

    A buffet test of a commercial transport model was accomplished in the National Transonic Facility at the NASA Langley Research Center. This aeroelastic test was unprecedented for this wind tunnel and posed a high risk to the facility. This paper presents the test results from a structural dynamics and aeroelastic response point of view and describes the activities required for the safety analysis and risk assessment. The test was conducted in the same manner as a flutter test and employed onboard dynamic instrumentation, real time dynamic data monitoring, automatic, and manual tunnel interlock systems for protecting the model. The procedures and test techniques employed for this test are expected to serve as the basis for future aeroelastic testing in the National Transonic Facility. This test program was a cooperative effort between the Boeing Commercial Airplane Company and the NASA Langley Research Center.

  7. Transonic compressor technology advancements

    NASA Technical Reports Server (NTRS)

    Benser, W. A.

    1974-01-01

    The highlights of the NASA program on transonic compressors are presented. Effects of blade shape and throat area on losses and flow range are discussed. Some effects of casing treatment on stall margin are presented. Results of tests with varying solidity are also presented. High Mach number, highly loaded stators are discussed and some results of stator hub slit suction are presented.

  8. Selected data from a transonic flexible walled test section

    NASA Technical Reports Server (NTRS)

    Wolf, S. W. D.

    1980-01-01

    Twenty four test runs of the Transonic Self-Streamlining Wind Tunnel were performed with the flexible walls 'streamlined' around a two dimensional section of four inch chord, over the Mach number range 0.3 to 0.89. Relevant wall and model data for the streamlined cases are presented.

  9. Subsonic and transonic hinge moment and wing bending/torsion characteristics of .015 scale space shuttle models 49-0 and 67-TS in the Rockwell International trisonic wind tunnel (IA70), volume 1

    NASA Technical Reports Server (NTRS)

    Hughes, M. T.; Mennell, R. C.

    1974-01-01

    Experimental aerodynamic investigations were conducted on an 0.015-scale representation of the integrated space shuttle launch vehicle in the trisonic wind tunnel. The primary test objective was to obtain subsonic and transonic elevon and bodyflap hinge moments and wing bending-torsion moments in the presence of the launch vehicle. Wing pressures were also recorded for the upper and lower right wing surfaces at two spanwise stations. The hinge moment, wing bending/torsion moments and wing pressure data were recorded over an angle-of-attack (alpha) range from -8 deg to +8 deg, and angle-of-sideslip (beta) range from -8 deg to +8 deg and at Mach numbers of 0.90, 1.12, 1.24 and 1.50. Tests were also conducted to determine the effects of the orbiter rear attach cross beam and the forward attach wedge and strut diameter. The orbiter alone was tested at 0.90 and 1.24 Mach number only.

  10. Comparison of the Drag of a Fin-Stabilized Body of Revolution and of a Complete Airplane Configuration as Obtained at Transonic Speeds in a Slotted Wind Tunnel and in Free Flight

    NASA Technical Reports Server (NTRS)

    Howell, Robert R.; Braslow, Albert L.

    1955-01-01

    A comparison of the zero-lift drag coefficients at Mach numbers from 0.81 to 1.41 of a fin-stabilized parabolic body of revolution as measured in the Langley transonic blowdown tunnel has been made with measurements obtained in free-flight on a larger but geometrically similar model. The absolute values of drag coefficient obtained in the slotted wind tunnel were equivalent to the free-flight drag-coefficient values up to a Mach number of 1.4 when adjustments were made for the effect on viscous drag of differences in Reynolds number between the two test conditions. Excellent agreement was obtained between the two tests for the pressure-drag variation with Mach number, regardless of whether the scale effect on skin friction was considered. Favorable agreement was also obtained between the pressure-drag increments due t o the presence of the stabilizing fins as determined in the wine tunnel from fins-on and fins-off tests and as obtained by a different method in free flight. Tests of a specific airplane configuration to obtain an indication of the problems involved in the construction and tests of small-scale (approximately 7-inch span) complete airplane configuration with internal air flow indicated that reliable zero-lift drag-coefficient measurements at Mach numbers up to 1.4 can be attained with such models, provided the model is constructed with a high but not an unreasonable degree of accuracy.

  11. Users Guide for the National Transonic Facility Research Data System

    NASA Technical Reports Server (NTRS)

    Foster, Jean M.; Adcock, Jerry B.

    1996-01-01

    The National Transonic Facility is a complex cryogenic wind tunnel facility. This report briefly describes the facility, the data systems, and the instrumentation used to acquire research data. The computational methods and equations are discussed in detail and many references are listed for those who need additional technical information. This report is intended to be a user's guide, not a programmer's guide; therefore, the data reduction code itself is not documented. The purpose of this report is to assist personnel involved in conducting a test in the National Transonic Facility.

  12. Contributions of Transonic Dynamics Tunnel Testing to Airplane Flutter Clearance

    NASA Technical Reports Server (NTRS)

    Rivera, Jose A.; Florance, James R.

    2000-01-01

    The Transonic Dynamics Tunnel (TDT) became in operational in 1960, and since that time has achieved the status of the world's premier wind tunnel for testing large in aeroelastically scaled models at transonic speeds. The facility has many features that contribute to its uniqueness for aeroelastic testing. This paper will briefly describe these capabilities and features, and their relevance to aeroelastic testing. Contributions to specific airplane configurations and highlights from the flutter tests performed in the TDT aimed at investigating the aeroelastic characteristics of these configurations are presented.

  13. Buffet test in the National Transonic Facility

    NASA Technical Reports Server (NTRS)

    Young, Clarence P., Jr.; Hergert, Dennis W.; Butler, Thomas W.; Herring, Fred M.

    1992-01-01

    A buffet test of a commercial transport model was accomplished in the National Transonic Facility at the NASA Langley Research Center. This aeroelastic test was unprecedented for this wind tunnel and posed a high risk for the facility. Presented here are the test results from a structural dynamics and aeroelastic response point of view. The activities required for the safety analysis and risk assessment are described. The test was conducted in the same manner as a flutter test and employed on-board dynamic instrumentation, real time dynamic data monitoring, and automatic and manual tunnel interlock systems for protecting the model.

  14. Investigations of transonic buffet control on civil aircraft wing with the use of tangential jet blowing

    NASA Astrophysics Data System (ADS)

    Abramova, K. A.; Petrov, A. V.; Potapchick, A. V.; Soudakov, V. G.

    2016-10-01

    Numerical and experimental investigations of transonic buffet control by tangential jet blowing are presented. To suppress the shock-induced boundary layer separation and the buffet at transonic speeds, compressed air jet is blown through a small slot nozzle tangentially to the upper surface of the supercritical airfoil. Numerical simulations were carried out on the basis of the unsteady Reynolds averaged Navier-Stokes (URANS) equations. Experimental studies of the tangential jet blowing were performed in the transonic wind tunnel T-112 of TsAGI. Results show that the jet moves the shock downstream, increases lift, suppresses flow separation under shock foot and delays buffet onset.

  15. Numerical design of streamlined tunnel walls for a two-dimensional transonic test

    NASA Technical Reports Server (NTRS)

    Newman, P. A.; Anderson, E. C.

    1978-01-01

    An analytical procedure is discussed for designing wall shapes for streamlined, nonporous, two-dimensional, transonic wind tunnels. It is based upon currently available 2-D inviscid transonic and boundary layer analysis computer programs. Predicted wall shapes are compared with experimental data obtained from the NASA Langley 6 by 19 inch Transonic Tunnel where the slotted walls were replaced by flexible nonporous walls. Comparisons are presented for the empty tunnel operating at a Mach number of 0.9 and for a supercritical test of an NACA 0012 airfoil at zero lift. Satisfactory agreement is obtained between the analytically and experimentally determined wall shapes.

  16. Terminal area energy management regime investigations utilizing an 0.030-scale model (47-0) of the space shuttle vehicle orbiter configuration 140A/B/C/R in the Ames Research Center 11 x 11 foot transonic wind tunnel (OA148), volume 5

    NASA Technical Reports Server (NTRS)

    Hawthorne, P. J.

    1976-01-01

    Data obtained in wind tunnel test OA148 are presented. The objectives of the test series were to: (1) obtain pressure distributions, forces and moments over the vehicle 5 orbiter in the thermal area energy management (TAEM) and approach phases of flight; (2) obtain elevon and rudder hinge moments in the TAEM and approach phases of flight; (3) obtain body flap and elevon loads for verification of loads balancing with integrated pressure distributions; and (4) obtain pressure distributions near the short OMS pods in the high subsonic, transonic and low supersonic Mach number regimes.

  17. Comparison of wind tunnel and flight test afterbody and nozzle pressures for a twin-jet fighter aircraft at transonic speeds

    NASA Technical Reports Server (NTRS)

    Nugent, Jack; Pendergraft, Odis C., Jr.

    1987-01-01

    Afterbody and nozzle pressures measured on a 1/12-scale model and in flight on a twin-jet fighter aircraft were compared as Mach number varied from 0.6 to 1.2, Reynolds number from 17.5 million to 302.5 million, and angle of attack from 1 to 7 deg. At Mach 0.6 and 0.8, nozzle pressure coefficient distributions and nozzle axial force coefficients agreed and showed good recompression. At Mach 0.9 and 1.2, flow complexity caused a loss in recompression for both flight and wind tunnel nozzle data. The flight data exhibited less negative values of pressure coefficient and lower axial force coefficients than did the wind tunnel data. Reynolds number effects were noted only at these Mach numbers. Jet temperature and mass flux ratio did not affect the comparisons of nozzle axial flow coefficient. At subsonic speeds, the levels of pressure coefficient distributions on the upper fuselage and lower nacelle surfaces for flight were less negative than those for the model. The model boundary layer thickness at the aft rake station exceeded that for the forward rake station and increased with increasing angle of attack. The flight boundary layer thickness at the aft rake station was less than that for the forward rake station and decreased with increasing angle of attack.

  18. Numerical calculation of transonic flow about slender bodies of revolution

    NASA Technical Reports Server (NTRS)

    Bailey, F. R.

    1971-01-01

    A relaxation method is described for the numerical solution of the transonic small disturbance equation for flow about a slender body of revolution. Results for parabolic arc bodies, both with and without an attached sting, are compared with wind-tunnel measurements for a free-stream Mach number range from 0.90 to 1.20. The method is also used to show the effects of wind-tunnel wall interference by including boundary conditions representing porous-wall and open-jet wind-tunnel test sections.

  19. Space Shuttle Model In The 16 Foot Transonic Tunnel

    NASA Technical Reports Server (NTRS)

    1978-01-01

    What may appear at first glance to be a swimming shark is a wind tunnel model of the Space Shuttle Orbiter, being tested at NASA's Langley Research Center in Hampton,VA. The Orbiter model is 5.5 feet long (1/20th of the real Orbiter's length) and has remotely operated control surfaces. Inside Langley's 16 foot Transonic Wind Tunnel, the model simulated Orbiter re-entry into the Earth's atmosphere, when it must fly through the transonic speed range (the range that crosses the sound barrier). Information on Orbiter stability and control, collected and analyzed during the tests, were integrated with other data to become part of computerized flight simulation programs.

  20. Laser velocimeter optical traverse scheme: An investigation of a proposed optical scanning technique for Arnold Engineering and Development Center's four-foot transonic wind tunnel

    NASA Astrophysics Data System (ADS)

    Krajci, G. S.

    1983-12-01

    This investigation analyzed a nonstandard laser velocimeter setup proposed for use in AEDC Wind Tunnel 4T. The setup uses a gimballed mirror to move the probe volume from point to point, and the translation of a lens to control the distance in the tunnel the probe volume reaches. Results show that for equal indices of refraction inside and outside the tunnel, the laser beams of a converging pair do not totally converge with its associated beam except under certain conditions, and the probe volumes created by each pair of overlapping laser beams do not always coincide. This work then provides the conditions necessary for total convergence of a pair of laser beams for this setup. A solution is then proposed to insure convergence of each laser beam pair and overlap of the two probe volumes. More than a solution to the above problems, a method is given to determine the azimuth and elevation angles for a mirror such that the reflected beam off the mirror passes through a given point in the tunnel after traversing a window. To carry out these investigations, a computer code was written to simulate the nonstandard laser velocimeter setup, and a second code was written to determine the azimuth and elevation angles for a mirror such that the reflected beam off the mirror passes through a given point in the tunnel after traversing a window. Both codes were written in FORTRAN 77, implemented on a CDC 6000-CYBER 74.

  1. Transonic-Small-Disturbance and Linear Analyses for the Active Aeroelastic Wing Program

    NASA Technical Reports Server (NTRS)

    Wiesman, Carol D.; Silva, Walter A.; Spain, Charles V.; Heeg, Jennifer

    2005-01-01

    Analysis serves many roles in the Active Aeroelastic Wing (AAW) program. It has been employed to ensure safe testing of both a flight vehicle and wind tunnel model, has formulated models for control law design, has provided comparison data for validation of experimental methods and has addressed several analytical research topics. Aeroelastic analyses using mathematical models of both the flight vehicle and the wind tunnel model configurations have been conducted. Static aeroelastic characterizations of the flight vehicle and wind tunnel model have been produced in the transonic regime and at low supersonic Mach numbers. The flight vehicle has been analyzed using linear aerodynamic theory and transonic small disturbance theory. Analyses of the wind-tunnel model were performed using only linear methods. Research efforts conducted through these analyses include defining regions of the test space where transonic effects play an important role and investigating transonic similarity. A comparison of these aeroelastic analyses for the AAW flight vehicle is presented in this paper. Results from a study of transonic similarity are also presented. Data sets from these analyses include pressure distributions, stability and control derivatives, control surface effectiveness, and vehicle deflections.

  2. Design optimization of transonic airfoils

    NASA Technical Reports Server (NTRS)

    Joh, C.-Y.; Grossman, B.; Haftka, R. T.

    1991-01-01

    Numerical optimization procedures were considered for the design of airfoils in transonic flow based on the transonic small disturbance (TSD) and Euler equations. A sequential approximation optimization technique was implemented with an accurate approximation of the wave drag based on the Nixon's coordinate straining approach. A modification of the Euler surface boundary conditions was implemented in order to efficiently compute design sensitivities without remeshing the grid. Two effective design procedures producing converged designs in approximately 10 global iterations were developed: interchanging the role of the objective function and constraint and the direct lift maximization with move limits which were fixed absolute values of the design variables.

  3. Transonic wall interference effects on bodies of revolution.

    NASA Technical Reports Server (NTRS)

    Couch, L. M.

    1972-01-01

    Efforts to develop a near sonic transport have placed renewed emphasis on obtaining accurate aerodynamic force and pressure data in the near sonic speed range. Comparison of wind-tunnel and flight data obtained for a blunt-nose body of revolution showed significant discrepancies in drag levels near Mach 1 - apparently due to wind-tunnel wall interference. Subsequent tests of geometrically similar bodies of revolution showed that increasing the model-to-test-section blockage ratio from 0.00017 to 0.0043 resulted in altered drag curve shapes, delayed drag divergence, and 'transonic creep' from subsonic drag levels due to increased wall interference.

  4. Turbulence Model Comparisons for Supersonic Transports at Transonic and Supersonic Conditions

    NASA Technical Reports Server (NTRS)

    Rivers, S. M. B.; Wahls, R. A.

    2003-01-01

    Results of turbulence model comparisons from two studies on supersonic transport configurations performed during the NASA High-speed Research program are given. Results are presented for both transonic conditions at Mach 0.90 and supersonic conditions at Mach 2.48. A feature of these two studies was the availability of higher Reynolds number wind tunnel data with which to compare the computational results. The transonic wind tunnel data was obtained in the National Transonic Facility at NASA Langley, and the supersonic data was obtained in the Boeing Polysonic Wind Tunnel. The computational data was acquired using a state of the art Navier-Stokes flow solver with a wide range of turbulence models implemented. The results show that the computed forces compare reasonably well with the experimental data, with the Baldwin- Lomax with Degani-Schiff modifications and the Baldwin-Barth models showing the best agreement for the transonic conditions and the Spalart-Allmaras model showing the best agreement for the supersonic conditions. The transonic results were more sensitive to the choice of turbulence model than were the supersonic results.

  5. Initial Assesment of Space Launch System Transonic Unsteady Pressure Environment

    NASA Technical Reports Server (NTRS)

    Sekula, Martin K.; Piatak, David J.; Rausch, Russ D.; Florance, James R.; Ramey, James M.

    2015-01-01

    A series of wind tunnel tests were conducted at the NASA Langley Research Center Transonic Dynamics Tunnel to assess the transonic buffet environment for the Space Launch System (SLS) launch vehicle. An initial test, conducted in 2012, indicated an elevated buffet environment prompting a second test to provide further insight into the buffet phenomena and assess potential solutions to reduce the response levels of these environments. During the course of the test program, eight variants of the SLS-10000 configuration were examined. The effect of these configuration variants on the coefficient of the root-mean-square fluctuation of pressure about the mean as a function of test condition indicates that the maximum fluctuating pressure levels are extremely sensitive to the geometry of the forward attachment of the solid rocket boosters (SRBs) to the SLS Core. The addition of flow fences or changes to the SRB nose cone geometry can alleviate the unsteady pressure environment.

  6. National Transonic Facility model and model support vibration problems

    NASA Technical Reports Server (NTRS)

    Young, Clarence P., Jr.; Popernack, Thomas G., Jr.; Gloss, Blair B.

    1990-01-01

    Vibrations of models and model support system were encountered during testing in the National Transonic Facility. Model support system yaw plane vibrations have resulted in model strain gage balance design load limits being reached. These high levels of vibrations resulted in limited aerodynamic testing for several wind tunnel models. The yaw vibration problem was the subject of an intensive experimental and analytical investigation which identified the primary source of the yaw excitation and resulted in attenuation of the yaw oscillations to acceptable levels. This paper presents the principal results of analyses and experimental investigation of the yaw plane vibration problems. Also, an overview of plans for development and installation of a permanent model system dynamic and aeroelastic response measurement and monitoring system for the National Transonic Facility is presented.

  7. Reynolds Number Effects on a Supersonic Transport at Transonic Conditions

    NASA Technical Reports Server (NTRS)

    Wahls, R. N.; Owens, L. R.; Rivers, S. M. B.

    2001-01-01

    A High Speed Civil Transport configuration was tested in the National Transonic Facility at the NASA Langley Research Center as part of NASA's High Speed Research Program. The primary purposes of the tests were to assess Reynolds number scale effects and the high Reynolds number aerodynamic characteristics of a realistic, second generation supersonic transport while providing data for the assessment of computational methods. The tests included longitudinal and lateral/directional studies at low speed high-lift and transonic conditions across a range of Reynolds numbers from that available in conventional wind tunnels to near flight conditions. Results are presented which focus on both the Reynolds number and static aeroelastic sensitivities of longitudinal characteristics at Mach 0.90 for a configuration without an empennage.

  8. Experiment and analysis on the flow process dynamics of the NASA-Langley eight foot transonic pressure tunnel

    NASA Technical Reports Server (NTRS)

    Tcheng, P.

    1977-01-01

    A dynamic response test performed in a eight foot transonic pressure tunnel is described. The dynamics of the flow process of the wind tunnel at transonic conditions were obtained. Descriptions of the test conditions, instrumentation, presentation of raw data, analysis of data, and finally, based on experimental evidences, an attempt to construct an input output relationship of the flow process from the viewpoints of control engineering are included.

  9. Design and physical characteristics of the Transonic Aircraft Technology (TACT) research aircraft

    NASA Technical Reports Server (NTRS)

    Painter, W. D.; Caw, L. J.

    1978-01-01

    The Transonic Aircraft Technology (TACT) research program provided data necessary to verify aerodynamic concepts, such as the supercritical wing, and to gain the confidence required for the application of such technology to advanced high performance aircraft. An F-111A aircraft was employed as the flight test bed to provide full scale data. The data were correlated extensively with predictions based on data obtained from wind tunnel tests. An assessment of the improvement afforded at transonic speeds in drag divergence, maneuvering performance, and airplane handling qualities by the use of the supercritical wing was included in the program. Transonic flight and wind tunnel testing techniques were investigated, and specific research technologies evaluated were also summarized.

  10. Terminal area energy management regime investigations utilizing an 0.030-scale model (47-0) of the space shuttle vehicle orbiter configuration 140A/B/C/R in the Ames Research Center 11 x 11 foot transonic wind tunnel (0A148), volume 1

    NASA Technical Reports Server (NTRS)

    Hawthorne, P. J.

    1976-01-01

    Data obtained in wind tunnel tests are presented. The objectives of the tests were to: (1) obtain pressure distributions, forces and moments over the vehicle 5 Orbiter in the terminal area energy management (TAEM) and approach phases of flight; (2) obtain elevon and rudder hinge moments in the TAEM and approach phases of flight; (3) obtain body flap and elevon loads for verification of loads balancing with integrated pressure distributions; and (4) obtain pressure distributions near the short OMS pods in the high subsonic, transonic and low supersonic Mach number regimes. Testing was conducted over a Mach number range from 0.6 to 1.4 with Reynolds number variations from 4.57 million to 2.74 million per foot. Model angle-of-attack was varied from -4 to 16 degrees and angles of side slip ranged from -8 to 8 degrees.

  11. Some remarks on the design of transonic tunnels with low levels of flow unsteadiness

    NASA Technical Reports Server (NTRS)

    Mabey, D. G.

    1976-01-01

    The principal sources of flow unsteadiness in the circuit of a transonic wind tunnel are presented. Care must be taken to avoid flow separations, acoustic resonances and large scale turbulence. Some problems discussed are the elimination of diffuser separations, the aerodynamic design of coolers and the unsteadiness generated in ventilated working sections.

  12. Calibration of the Langley 16-foot transonic tunnel with test section air removal

    NASA Technical Reports Server (NTRS)

    Corson, B. W., Jr.; Runckel, J. F.; Igoe, W. B.

    1974-01-01

    The Langley 16-foot transonic tunnel with test section air removal (plenum suction) was calibrated to a Mach number of 1.3. The results of the calibration, including the effects of slot shape modifications, test section wall divergence, and water vapor condensation, are presented. A complete description of the wind tunnel and its auxiliary equipment is included.

  13. TAIR: A transonic airfoil analysis computer code

    NASA Technical Reports Server (NTRS)

    Dougherty, F. C.; Holst, T. L.; Grundy, K. L.; Thomas, S. D.

    1981-01-01

    The operation of the TAIR (Transonic AIRfoil) computer code, which uses a fast, fully implicit algorithm to solve the conservative full-potential equation for transonic flow fields about arbitrary airfoils, is described on two levels of sophistication: simplified operation and detailed operation. The program organization and theory are elaborated to simplify modification of TAIR for new applications. Examples with input and output are given for a wide range of cases, including incompressible, subcritical compressible, and transonic calculations.

  14. 12. VIEW EAST, BUILDING 12 INTERIOR, WIND TUNNEL 157 ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    12. VIEW EAST, BUILDING 12 INTERIOR, WIND TUNNEL 157 - Naval Surface Warfare Center, Transonic Wind Tunnel Building, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD

  15. 2. VIEW SOUTH OF WIND TUNNEL 157, NORTH ELEVATION ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    2. VIEW SOUTH OF WIND TUNNEL 157, NORTH ELEVATION - Naval Surface Warfare Center, Transonic Wind Tunnel Building, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD

  16. 13. VIEW NORTHEAST, BUILDING 12 INTERIOR, WIND TUNNEL FAN ASSEMBLY ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    13. VIEW NORTHEAST, BUILDING 12 INTERIOR, WIND TUNNEL FAN ASSEMBLY - Naval Surface Warfare Center, Transonic Wind Tunnel Building, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD

  17. Parametric Evaluation of Thin, Transonic Circulation-Control Airfoils

    NASA Technical Reports Server (NTRS)

    Schlecht, Robin; Anders, Scott

    2007-01-01

    Wind-tunnel tests were conducted in the NASA Langley Transonic Dynamics Tunnel on a 6 percent-thick, elliptical circulation-control airfoil with upper-surface and lower-surface blowing capability. Results for elliptical Coanda trailing-edge geometries, biconvex Coanda trailing-edge geometries, and leading-edge geometries are reported. Results are presented at subsonic and transonic Mach numbers of 0.3 and 0.8, respectively. When considering one fixed trailing-edge geometry, for both the subsonic and transonic conditions it was found that the [3.0:1] ratio elliptical Coanda surface with the most rounded leading-edge [03] performed favorably and was determined to be the best compromise between comparable configurations that took advantage of the Coanda effect. This configuration generated a maximum. (Delta)C(sub 1) = 0.625 at a C(sub mu) = 0.06 at M = 0.3, alpha = 6deg. This same configuration generated a maximum (Delta)C(sub 1) = 0.275 at a C(sub mu) = 0.0085 at M = 0.8, alpha = 3deg.

  18. Multiple Solutions of Transonic Flow over NACA0012 Airfoil

    NASA Astrophysics Data System (ADS)

    Xiong, Juntao; Liu, Ya; Liu, Feng; Luo, Shijun; Zhao, Zijie; Ren, Xudong; Gao, Chao

    2012-11-01

    Multiple solutions of the small-disturbance potential equation and full potential equation were known for the NACA0012 airfoil in a certain range of transonic Mach numbers and at zero angle of attack. However the multiple solutions for this airfoil were not observed using Euler or Navier-Stokes equations under the above flow conditions. In the present work, both the Unsteady Reynolds-Averaged Navier-Stokes (URANS) computations and transonic wind tunnel experiments are performed under certain Reynolds numbers to further study the problem. The results of the two methods reveal that buffet appears in a narrow Mach number range where the potential flow methods predict multiple solutions. Boundary layer displacement thickness computed from URANS at the same flow condition is used to modify the geometry of the airfoil. Euler equations are then solved for the modified geometry. The results show that the addition of the boundary layer displacement thickness creates multiple solutions for the NACA0012 airfoil. Global linear stability analysis is also performed on the original and the modified airfoils. This shows a close relationship between the viscous unsteady shock buffet phenomenon of transonic airfoil flow and the existence of multiple solutions of the external inviscid flow. Postdoctoral Research Assistant.

  19. MAVRIC Flutter Model Transonic Limit Cycle Oscillation Test

    NASA Technical Reports Server (NTRS)

    Edwards, John W.; Schuster, David M.; Spain, Charles V.; Keller, Donald F.; Moses, Robert W.

    2001-01-01

    The Models for Aeroelastic Validation Research Involving Computation semi-span wind-tunnel model (MAVRIC-I), a business jet wing-fuselage flutter model, was tested in NASA Langley's Transonic Dynamics Tunnel with the goal of obtaining experimental data suitable for Computational Aeroelasticity code validation at transonic separation onset conditions. This research model is notable for its inexpensive construction and instrumentation installation procedures. Unsteady pressures and wing responses were obtained for three wingtip configurations clean, tipstore, and winglet. Traditional flutter boundaries were measured over the range of M = 0.6 to 0.9 and maps of Limit Cycle Oscillation (LCO) behavior were made in the range of M = 0.85 to 0.95. Effects of dynamic pressure and angle-of-attack were measured. Testing in both R134a heavy gas and air provided unique data on Reynolds number, transition effects, and the effect of speed of sound on LCO behavior. The data set provides excellent code validation test cases for the important class of flow conditions involving shock-induced transonic flow separation onset at low wing angles, including Limit Cycle Oscillation behavior.

  20. MAVRIC Flutter Model Transonic Limit Cycle Oscillation Test

    NASA Technical Reports Server (NTRS)

    Edwards, John W.; Schuster, David M.; Spain, Charles V.; Keller, Donald F.; Moses, Robert W.

    2001-01-01

    The Models for Aeroelastic Validation Research Involving Computation semi-span wind-tunnel model (MAVRIC-I), a business jet wing-fuselage flutter model, was tested in NASA Langley's Transonic Dynamics Tunnel with the goal of obtaining experimental data suitable for Computational Aeroelasticity code validation at transonic separation onset conditions. This research model is notable for its inexpensive construction and instrumentation installation procedures. Unsteady pressures and wing responses were obtained for three wingtip configurations of clean, tipstore, and winglet. Traditional flutter boundaries were measured over the range of M = 0.6 to 0.9 and maps of Limit Cycle Oscillation (LCO) behavior were made in the range of M = 0.85 to 0.95. Effects of dynamic pressure and angle-of-attack were measured. Testing in both R134a heavy gas and air provided unique data on Reynolds number, transition effects, and the effect of speed of sound on LCO behavior. The data set provides excellent code validation test cases for the important class of flow conditions involving shock-induced transonic flow separation onset at low wing angles, including LCO behavior.

  1. Transonic CFD applications at Boeing

    NASA Technical Reports Server (NTRS)

    Tinoco, E. N.

    1989-01-01

    The use of computational methods for three dimensional transonic flow design and analysis at the Boeing Company is presented. A range of computational tools consisting of production tools for every day use by project engineers, expert user tools for special applications by computational researchers, and an emerging tool which may see considerable use in the near future are described. These methods include full potential and Euler solvers, some coupled to three dimensional boundary layer analysis methods, for transonic flow analysis about nacelle, wing-body, wing-body-strut-nacelle, and complete aircraft configurations. As the examples presented show, such a toolbox of codes is necessary for the variety of applications typical of an industrial environment. Such a toolbox of codes makes possible aerodynamic advances not previously achievable in a timely manner, if at all.

  2. Semidirect computations for transonic flow

    NASA Technical Reports Server (NTRS)

    Swisshelm, J. M.; Adamczyk, J. J.

    1983-01-01

    A semidirect method, driven by a Poisson solver, was developed for inviscid transonic flow computations. It is an extension of a recently introduced algorithm for solving subsonic rotational flows. Shocks are captured by implementing a form of artificial compressibility. Nonisentropic cases are computed using a shock tracking procedure coupled with the Rankine-Hugoniot relationships. Results are presented for both subsonic and transonic flows. For the test geometry, an unstaggered cascade of 20 percent thick circular arc airfoils at zero angle of attack, shocks are crisply resolved in supercritical situations and the algorithm converges rapidly. In addition, the convergence rate appears to be nearly independent of the entropy and vorticity production at the shock.

  3. Unsteady transonic flow in cascades

    NASA Technical Reports Server (NTRS)

    Surampudi, S. P.; Adamczyk, J. J.

    1984-01-01

    There is a need for methods to predict the unsteady air loads associated with flutter of turbomachinery blading at transonic speeds. The results of such an analysis in which the steady relative flow approaching a cascade of thin airfoils is assumed to be transonic, irrotational, and isentropic is presented. The blades in the cascade are allowed to undergo a small amplitude harmonic oscillation which generates a small unsteady flow superimposed on the existing steady flow. The blades are assumed to oscillate with a prescribed motion of constant amplitude and interblade phase angle. The equations of motion are obtained by linearizing about a uniform flow the inviscid nonheat conducting continuity and momentum equations. The resulting equations are solved by employing the Weiner Hopf technique. The solution yields the unsteady aerodynamic forces acting on the cascade at Mach number equal to 1. Making use of an unsteady transonic similarity law, these results are compared with the results obtained from linear unsteady subsonic and supersonic cascade theories. A parametric study is conducted to find the effects of reduced frequency, solidity, stagger angle, and position of pitching axis on the flutter.

  4. Transonic unsteady airloads on an energy efficient transport wing with oscillating control surfaces

    NASA Technical Reports Server (NTRS)

    Sandford, M. C.; Ricketts, R. H.; Cazier, F. W., Jr.; Cunningham, H. J.

    1980-01-01

    An aspect-ratio 10.8 supercritical wing with oscillating control surfaces is described. The wing is instrumented with 252 static orifices and 164 in situ dynamic pressure transducers for studying the effects of control surface deflection on steady and unsteady pressures at transonic speeds. Selected results from initial wind-tunnel tests conducted in the Langley Transonic Dynamics Tunnel are discussed. Unsteady pressure results are presented for two trailing-edge control surfaces oscillating separately at the design Mach number of 0.78. Some experimental results are compared with analytical results obtained by using linear lifting-surface theory.

  5. An experimental investigation of internal area ruling for transonic and supersonic channel flow

    NASA Technical Reports Server (NTRS)

    Roberts, W. B.; Vanrintel, H. L.; Rizvi, G.

    1982-01-01

    A simulated transonic rotor channel model was examined experimentally to verify the flow physics of internal area ruling. Pressure measurements were performed in the high speed wind tunnel at transonic speeds with Mach 1.5 and Mach 2 nozzle blocks to get an indication of the approximate shock losses. The results showed a reduction in losses due to internal area ruling with the Mach 1.5 nozzle blocks. The reduction in total loss coefficient was of the order of 17 percent for a high blockage model and 7 percent for a cut-down model.

  6. Transonic unsteady airloads on an energy efficient transport wing with oscillating control surfaces

    NASA Technical Reports Server (NTRS)

    Sandford, M. C.; Ricketts, R. H.; Cazier, F. W., Jr.; Cunningham, H. J.

    1980-01-01

    An aspect ratio 10.8 supercritical wing with oscillating control surfaces is described. The wing is instrumental with 252 static orifices and 164 in situ dynamic pressure transducers for studying the effects of control surface deflection on steady and unsteady pressures at transonic speeds. Results from initial wind tunnel tests conducted in the Langley Transonic Dynamics Tunnel are discussed. Unsteady pressure results are presented for two trailing edge control surfaces oscillating separately at the design Mach number of 0.78. Some experimental results are compared with analytical results obtained by using linear lifting surface theory.

  7. Langley 16- Ft. Transonic Tunnel Pressure Sensitive Paint System

    NASA Technical Reports Server (NTRS)

    Sprinkle, Danny R.; Obara, Clifford J.; Amer, Tahani R.; Leighty, Bradley D.; Carmine, Michael T.; Sealey, Bradley S.; Burkett, Cecil G.

    2001-01-01

    This report describes the NASA Langley 16-Ft. Transonic Tunnel Pressure Sensitive Paint (PSP) System and presents results of a test conducted June 22-23, 2000 in the tunnel to validate the PSP system. The PSP system provides global surface pressure measurements on wind tunnel models. The system was developed and installed by PSP Team personnel of the Instrumentation Systems Development Branch and the Advanced Measurement and Diagnostics Branch. A discussion of the results of the validation test follows a description of the system and a description of the test.

  8. Operational manual for two-dimensional transonic code TSFOIL

    NASA Technical Reports Server (NTRS)

    Stahara, S. S.

    1978-01-01

    This code solves the two-dimensional, transonic, small-disturbance equations for flow past lifting airfoils in both free air and various wind-tunnel environments by using a variant of the finite-difference method. A description of the theoretical and numerical basis of the code is provided, together with complete operating instructions and sample cases for the general user. In addition, a programmer's manual is also presented to assist the user interested in modifying the code. Included in the programmer's manual are a dictionary of subroutine variables in common and a detailed description of each subroutine.

  9. Numerical study of porous airfoils in transonic flow

    NASA Technical Reports Server (NTRS)

    Chen, C. L.; Chow, C. Y.; Holst, T. L.; Vandalsem, W. R.

    1985-01-01

    A numerical study was made to examine the effect of a porous surface on the aerodynamic performance of a transonic airfoil. The pressure jump across the normal shock wave on the upper surface of the airfoil was reduced by making the surface below the shock porous. The weakened shock is preceded by an oblique shock at the upstream end of the porous surface where air is blown out of the cavity. The lambda shock structure shown in the numerical result qualitatively agrees with that observed in the wind tunnel. According to the present analysis, the porous airfoil has a smaller drag and a higher lift than the solid airfoil.

  10. Introductory remarks. [fluid mechanics research for the National Transonic Facility: theoretical aerodynamics

    NASA Technical Reports Server (NTRS)

    Gessow, A.

    1977-01-01

    Suggested fluid mechanics research to be conducted in the National Transonic Facility include: wind tunnel calibration; flat plate skin friction, flow visualization and measurement techniques; leading edge separation; high angle of attack separation; shock-boundary layer interaction; submarine shapes; low speed studies of cylinder normal to flow; and wall interference effects. These theoretical aerodynamic investigations will provide empirical inputs or validation data for computational aerodynamics, and increase the usefulness of existing wind tunnels.

  11. Transonic aeroelastic analysis of the B-1 wing

    NASA Technical Reports Server (NTRS)

    Guruswamy, G. P.; Goorjian, P. M.; Ide, H.; Miller, G. D.

    1986-01-01

    The flow over the B-1 wing is studied computationally, including the aeroelastic response of the wing. Computed results are compared with results from wind tunnel and flight tests for both low- and high-sweep cases, at 25.0 and 67.5 deg, respectively, for selected transonic Mach numbers. The aerodynamic and aeroelastic computations are made by using the transonic unsteady code ATRAN3S. Steady aerodynamic computations compare well with wind tunnel results for the 25.0 deg sweep case and also for small angles of attack at 67.5 deg sweep case. The aeroelastic response results show that the wing is stable at the low-sweep angle for the calculation at the Mach number at which there is a shock wave. In the higher-sweep case, for the higher angle of attack at which oscillations were observed in the flight and wind tunnel tests, the calculations do not show any shock waves. Their absence lends support to the hypothesis that the observed oscillations are due to the presence of leading-edge separation vortices and not to shock wave motion, as was previously proposed.

  12. Transonic aerodynamic and aeroelastic characteristics of a variable sweep wing

    NASA Technical Reports Server (NTRS)

    Goorjian, P. M.; Guruswamy, G. P.; Ide, H.; Miller, G.

    1985-01-01

    The flow over the B-1 wing is studied computationally, including the aeroelastic response of the wing. Computed results are compared with results from wind tunnel and flight tests for both low-sweep and high-sweep cases, at 25.0 and 67.5 deg., respectively, for selected transonic Mach numbers. The aerodynamic and aeroelastic computations are made by using the transonic unsteady code ATRAN3S. Steady aerodynamic computations compare well with wind tunnel results for the 25.0 deg sweep case and also for small angles of attack at the 67.5 deg sweep case. The aeroelastic response results show that the wing is stable at the low sweep angle for the calculation at the Mach number at which there is a shock wave. In the higher sweep case, for the higher angle of attack at which oscillations were observed in the flight and wind tunnel tests, the calculations do not show any shock waves. Their absence lends support to the hypothesis that the observed oscillations are due to the presence of leading edge separation vortices and are not due to shock wave motion as was previously proposed.

  13. Transonic aerodynamic and aeroelastic characteristics of a variable sweep wing

    NASA Technical Reports Server (NTRS)

    Goorjian, P. M.; Guruswamy, G. P.; Ide, H.; Miller, G.

    1985-01-01

    The flow over the B-1 wing is studied computationally, including the aeroelastic response of the wing. Computed results are compared with results from wind tunnel and flight tests for both low-sweep and high-sweep cases, at 25.0 deg. and 67.5 deg., respectively, for selected transonic Mach numbers. The aerodynamic and aeroelastic computations are made by using the transonic unsteady code ATRAN3S. Steady aerodynamic computations compare well with wind tunnel results for the 25.0 deg. sweep case and also for small angles of attack at the 67.5 deg. sweep case. The aeroelastic response results show that the wing is stable at the low sweep angle for the calculation at the Mach number at which there is a shock wave. In the higher sweep case, for the higher angle of attack at which oscillations were observed in the flight and wind tunnel tests, the calculations do not show any shock waves. Their absence lends support to the hypothesis that the observed oscillations are due to the presence of leading edge separation vortices and are not due to shock wave motion as was previously proposed.

  14. Ares Launch Vehicle Transonic Buffet Testing and Analysis Techniques

    NASA Technical Reports Server (NTRS)

    Piatak, David J.; Sekula, Martin K.; Rausch, Russ D.

    2010-01-01

    It is necessary to define the launch vehicle buffet loads to ensure that structural components and vehicle subsystems possess adequate strength, stress, and fatigue margins when the vehicle structural dynamic response to buffet forcing functions are considered. In order to obtain these forcing functions, the accepted method is to perform wind-tunnel testing of a rigid model instrumented with hundreds of unsteady pressure transducers designed to measure the buffet environment across the desired frequency range. The buffet wind-tunnel test program for the Ares Crew Launch Vehicle employed 3.5 percent scale rigid models of the Ares I and Ares I-X launch vehicles instrumented with 256 unsteady pressure transducers each. These models were tested at transonic conditions at the Transonic Dynamics Tunnel at NASA Langley Research Center. The ultimate deliverable of the Ares buffet test program are buffet forcing functions (BFFs) derived from integrating the measured fluctuating pressures on the rigid wind-tunnel models. These BFFs are then used as input to a multi-mode structural analysis to determine the vehicle response to buffet and the resulting buffet loads and accelerations. This paper discusses the development of the Ares I and I-X rigid buffet model test programs from the standpoint of model design, instrumentation system design, test implementation, data analysis techniques to yield final products, and presents normalized sectional buffet forcing function root-mean-squared levels.

  15. An airfoil flutter model suspension system to accommodate large static transonic airloads

    NASA Technical Reports Server (NTRS)

    Reed, W. H., III

    1985-01-01

    A pitch/plunge flutter model suspension system and associated two-dimensional MBB-A3 airfoil models is described. The system is designed for installation in the Langley 6-by-19-inch and 6-by-18-inch transonic blowdown wind tunnels to enable systematic study of the transonic flutter characteristics and static pressure distributions of supercritical airfoils at transonic Mach numbers. A compound spring suspension concept is introduced which simultaneously meets requirements for low plunge-mode stiffness, lightweight suspended model, and large steady lift due to angle of attack without the need for excessive static deflections of the plunge spring. The system features variable pitch and plunge frequencies, changeable airfoil rotation axes, and a self aligning control system to maintain a constant mean position of the model with changing airload.

  16. Investigation of Transonic Reynolds Number Scaling on a Twin-Engine Transport

    NASA Technical Reports Server (NTRS)

    Curtin, M. M.; Bogue, D. R.; Om, D.; Rivers, S. M. B.; Pendergraft, O. C., Jr.; Wahls, R. A.

    2002-01-01

    This paper discusses Reynolds number scaling for aerodynamic parameters including force and wing pressure measurements. A full-span model of the Boeing 777 configuration was tested at transonic conditions in the National Transonic Facility (NTF) at Reynolds numbers (based on mean aerodynamic chord) from 3.0 to 40.0 million. Data was obtained for a tail-off configuration both with and without wing vortex generators and flap support fairings. The effects of aeroelastics were separated from Reynolds number effects by varying total pressure and temperature independently. Data from the NTF at flight Reynolds number are compared with flight data to establish the wind tunnel/flight correlation. The importance of high Reynolds number testing and the need for developing a process for transonic Reynolds number scaling is discussed. This paper also identifies issues that need to be worked for Boeing Commercial to continue to conduct future high Reynolds number testing in the NTF.

  17. Euler solvers for transonic applications

    NASA Technical Reports Server (NTRS)

    Vanleer, Bram

    1989-01-01

    The 1980s may well be called the Euler era of applied aerodynamics. Computer codes based on discrete approximations of the Euler equations are now routinely used to obtain solutions of transonic flow problems in which the effects of entropy and vorticity production are significant. Such codes can even predict separation from a sharp edge, owing to the inclusion of artificial dissipation, intended to lend numerical stability to the calculation but at the same time enforcing the Kutta condition. One effect not correctly predictable by Euler codes is the separation from a smooth surface, and neither is viscous drag; for these some form of the Navier-Stokes equation is needed. It, therefore, comes as no surprise to observe that the Navier-Stokes has already begun before Euler solutions were fully exploited. Moreover, most numerical developments for the Euler equations are now constrained by the requirement that the techniques introduced, notably artificial dissipation, must not interfere with the new physics added when going from an Euler to a full Navier-Stokes approximation. In order to appreciate the contributions of Euler solvers to the understanding of transonic aerodynamics, it is useful to review the components of these computational tools. Space discretization, time- or pseudo-time marching and boundary procedures, the essential constituents are discussed. The subject of grid generation and grid adaptation to the solution are touched upon only where relevant. A list of unanswered questions and an outlook for the future are covered.

  18. A user's guide to the Langley 16-foot transonic tunnel complex. Revision 1

    NASA Technical Reports Server (NTRS)

    1990-01-01

    The operational characteristics and equipment associated with the Langley 16-foot transonic tunnel complex which is located in buildings 1146 and 1234 at the Langley Research Center are described in detail. This complex consists of the 16-foot transonic wind tunnel, the static test facility, and the 16- by 24-inch water tunnel research facilities. The 16-foot transonic tunnel is a single-return atmospheric wind tunnel with a 15.5 foot diameter test section and a Mach number capability from 0.20 to 1.30. The emphasis for research conducted in this research complex is on the integration of the propulsion system into advanced aircraft concepts. In the past, the primary focus has been on the integration of nozzles and empennage into the afterbody of fighter aircraft. During the last several years this experimental research has been expanded to include developing the fundamental data base necessary to verify new theoretical concepts, inlet integration into fighter aircraft, nozzle integration for supersonic and hypersonic transports, nacelle/pylon/wing integration for subsonic transport configurations, and the study of vortical flows (in the 16- by 24-inch water tunnel). The purpose here is to provide a comprehensive description of the operational characteristics of the research facilities of the 16-foot transonic tunnel complex and their associated systems and equipments.

  19. Wing Twist Measurements at the National Transonic Facility

    NASA Technical Reports Server (NTRS)

    Burner, Alpheus W.; Wahls, Richard A.; Goad, William K.

    1996-01-01

    A technique for measuring wing twist currently in use at the National Transonic Facility is described. The technique is based upon a single camera photogrammetric determination of two dimensional coordinates with a fixed (and known) third dimensional coordinate. The wing twist is found from a conformal transformation between wind-on and wind-off 2-D coordinates in the plane of rotation. The advantages and limitations of the technique as well as the rationale for selection of this particular technique are discussed. Examples are presented to illustrate run-to-run and test-to-test repeatability of the technique in air mode. Examples of wing twist in cryogenic nitrogen mode are also presented.

  20. X-29 High Alpha Test in the National Transonic Facility

    NASA Technical Reports Server (NTRS)

    Underwood, Pamela J.; Owens, Lewis R.; Wahls, Richard A.; Williams, Susan

    2003-01-01

    This paper describes the X-29A research program at the National Transonic Facility. This wind tunnel test leveraged the X-29A high alpha flight test program by enabling ground-to-flight correlation studies with an emphasis on Reynolds number effects. The background and objectives of this test program, as well as the comparison of high Reynolds number wind tunnel data to X-29A flight test data are presented. The effects of Reynolds number on the forebody pressures at high angles of attack are also presented. The purpose of this paper is to document this test and serve as a reference for future ground-to-flight correlation studies, and high angle-of-attack investigations. Good ground-to-flight correlations were observed for angles of attack up to 50 deg, and Reynolds number effects were also observed.

  1. A Compliant Casing for Transonic Axial Compressors

    NASA Technical Reports Server (NTRS)

    Bloch, Gregory S.; Hah, Chunill

    2003-01-01

    A viewgraph presentation on the concept of compliant casing for transonic axial compressors is shown. The topics include: 1) Concept for compliant casing; 2) Rig and facility details; 3) Experimental results; and 4) Numerical results.

  2. Inviscid transonic flow computations with shock fitting

    NASA Technical Reports Server (NTRS)

    Yu, N. J.; Seebass, A. R.

    1975-01-01

    First-and second-order numerical procedures are presented for calculating two-dimensional transonic flows that treat shock waves as discontinuities. Their application to a simple but nontrivial problem for which there are limited theoretical results is discussed.

  3. Transonic-Wind-Tunnel Tests of the Aerodynamic Characteristics of a 0.15-Scale Model of the North American Aviation 255-Inch Fin-Stabilized External Store, Coord No. AF-AM-4

    NASA Technical Reports Server (NTRS)

    Fischetti, Thomas L.

    1958-01-01

    An investigation has been made in the Langley 8-foot transonic tunnels on the aerodynamic characteristics of a 0.15-scale model of the North American Aviation 255-inch fin-stabilized external store over a maximum Mach number range of 0.60 to 1.2 and on the effects of mounting lugs, of fin orientation, of fin aspect ratio, and of fixed-transition. The Reynolds number (based on a body length of 37.50 inches) varied from 9.8 x 10(exp 6) to 13.1 x 10(exp 6). The results indicate that the static margin of the finned store at low lift coefficients was only 9 percent of body length at subsonic Mach numbers and was reduced to zero at a Mach number of 1.0, Increasing the fin aspect ratio from 1.82 to 2.41 increased the subsonic static margin to 18 percent and provided a minimum margin of 9 percent near a Mach number of l.O. Store mounting lugs or fin orientation had only small effects on the aerodynamic characteristics of the basic store.

  4. Transonic rotor noise: Theoretical and experimental comparisons

    NASA Technical Reports Server (NTRS)

    Schmitz, F. H.; Yu, Y. H.

    1980-01-01

    Two complementary methods of describing the high speed rotor noise problem are discussed. The first method uses the second order transonic potential equation to define and characterize the nature of the aerodynamic and acoustic fields and to explain the appearance of radiating shock waves. The second employs the Ffowcs Williams and Hawkings equation to successfully calculate the acoustic far field. Good agreement between theoretical and experimental waveforms is shown for transonic hover tip Mach numbers from 0.8 to 0.9.

  5. Control software for two dimensional airfoil tests using a self-streamlining flexible walled transonic test section

    NASA Technical Reports Server (NTRS)

    Wolf, S. W. D.; Goodyer, M. J.

    1982-01-01

    Operation of the Transonic Self-Streamlining Wind Tunnel (TSWT) involved on-line data acquisition with automatic wall adjustment. A tunnel run consisted of streamlining the walls from known starting contours in iterative steps and acquiring model data. Each run performs what is described as a streamlining cycle. The associated software is presented.

  6. Fifteen Years of Operation at NASA's National Transonic Facility with the World's Largest Adjustable Speed Drive

    NASA Technical Reports Server (NTRS)

    Sydnor, George H.; Bhatia, Ram; Krattiger, Hansueli; Mylius, Justus; Schafer, D.

    2012-01-01

    In September 1995, a project was initiated to replace the existing drive line at NASA's most unique transonic wind tunnel, the National Transonic Facility (NTF), with a single 101 MW synchronous motor driven by a Load Commutated Inverter (LCI). This Adjustable Speed Drive (ASD) system also included a custom four-winding transformer, harmonic filter, exciter, switch gear, control system, and feeder cable. The complete system requirements and design details have previously been presented and published [1], as well as the commissioning and acceptance test results [2]. The NTF was returned to service in December 1997 with the new drive system powering the fan. Today, this installation still represents the world s largest horizontal single motor/drive combination. This paper describes some significant events that occurred with the drive system during the first 15 years of service. These noteworthy issues are analyzed and root causes presented. Improvements that have substantially increased the long term viability of the system are given.

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

    NASA Technical Reports Server (NTRS)

    Tomek, W. G.; Hall, R. M.; Wahls, R. A.; Luckring, J. M.; Owens, L. R.

    2002-01-01

    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.

  8. Transonic flow analysis for rotors. Part 3: Three-dimensional, quasi-steady, Euler calculation

    NASA Technical Reports Server (NTRS)

    Chang, I-Chung

    1990-01-01

    A new method is presented for calculating the quasi-steady transonic flow over a lifting or non-lifting rotor blade in both hover and forward flight by using Euler equations. The approach is to solve Euler equations in a rotor-fixed frame of reference using a finite volume method. A computer program was developed and was then verified by comparison with wind-tunnel data. In all cases considered, good agreement was found with published experimental data.

  9. Data reduction formulas for the 16-foot transonic tunnel: NASA Langley Research Center, revision 2

    NASA Technical Reports Server (NTRS)

    Mercer, Charles E.; Berrier, Bobby L.; Capone, Francis J.; Grayston, Alan M.

    1992-01-01

    The equations used by the 16-Foot Transonic Wind Tunnel in the data reduction programs are presented in nine modules. Each module consists of equations necessary to achieve a specific purpose. These modules are categorized in the following groups: (1) tunnel parameters; (2) jet exhaust measurements; (3) skin friction drag; (4) balance loads and model attitudes calculations; (5) internal drag (or exit-flow distribution); (6) pressure coefficients and integrated forces; (7) thrust removal options; (8) turboprop options; and (9) inlet distortion.

  10. Transonic aerodynamic characteristics of a proposed wing-body reusable launch vehicle concept

    NASA Technical Reports Server (NTRS)

    Springer, A. M.

    1995-01-01

    A proposed wing-body reusable launch vehicle was tested in the NASA Marshall Space Flight Center's 14 x 14-inch trisonic wind tunnel during the winter of 1994. This test resulted in the vehicle's subsonic and transonic, Mach 0.3 to 1.96, longitudinal and lateral aerodynamic characteristics. The effects of control surface deflections on the basic vehicle's aerodynamics, including a body flap, elevons, ailerons, and tip fins, are presented.

  11. Transonic Aerodynamic Characteristics of a Proposed Wing Body Reusable Launch Vehicle Concept

    NASA Technical Reports Server (NTRS)

    Springer, Anthony M.

    1996-01-01

    A proposed wing body reusable launch vehicle was tested in the Marshall Space Flight Center (MSFC) 14 x 14 inch trisonic wind tunnel during the winter of 1994. this test resulted in the vehicle's subsonic and transonic (Mach 03 to 1.96) longitudinal and lateral aerodynamic characteristics. The effects of control surface deflections on the basic vehicle aerodynamics including a body flap, elevons, ailerons, and tip fins are presented.

  12. Upgrades at the NASA Langley Research Center National Transonic Facility

    NASA Technical Reports Server (NTRS)

    Paryz, Roman W.

    2012-01-01

    Several projects have been completed or are nearing completion at the NASA Langley Research Center (LaRC) National Transonic Facility (NTF). The addition of a Model Flow-Control/Propulsion Simulation test capability to the NTF provides a unique, transonic, high-Reynolds number test capability that is well suited for research in propulsion airframe integration studies, circulation control high-lift concepts, powered lift, and cruise separation flow control. A 1992 vintage Facility Automation System (FAS) that performs the control functions for tunnel pressure, temperature, Mach number, model position, safety interlock and supervisory controls was replaced using current, commercially available components. This FAS upgrade also involved a design study for the replacement of the facility Mach measurement system and the development of a software-based simulation model of NTF processes and control systems. The FAS upgrades were validated by a post upgrade verification wind tunnel test. The data acquisition system (DAS) upgrade project involves the design, purchase, build, integration, installation and verification of a new DAS by replacing several early 1990's vintage computer systems with state of the art hardware/software. This paper provides an update on the progress made in these efforts. See reference 1.

  13. Investigation of Transonic Wake Dynamics for Mechanically Deployable Entry Systems

    NASA Technical Reports Server (NTRS)

    Stern, Eric; Barnhardt, Michael; Venkatapathy, Ethiraj; Candler, Graham; Prabhu, Dinesh

    2012-01-01

    A numerical investigation of transonic flow around a mechanically deployable entry system being considered for a robotic mission to Venus has been performed, and preliminary results are reported. The flow around a conceptual representation of the vehicle geometry was simulated at discrete points along a ballistic trajectory using Detached Eddy Simulation (DES). The trajectory points selected span the low supersonic to transonic regimes with freestream Mach numbers from 1:5 to 0:8, and freestream Reynolds numbers (based on diameter) between 2:09 x 10(exp 6) and 2:93 x 10(exp 6). Additionally, the Mach 0:8 case was simulated at angles of attack between 0 and 5 . Static aerodynamic coefficients obtained from the data show qualitative agreement with data from 70deg sphere-cone wind tunnel tests performed for the Viking program. Finally, the effect of choices of models and numerical algorithms is addressed by comparing the DES results to those using a Reynolds Averaged Navier-Stokes (RANS) model, as well as to results using a more dissipative numerical scheme.

  14. Enhancements to the FAST-MAC Circulation Control Model and Recent High-Reynolds Number Testing in the National Transonic Facility

    NASA Technical Reports Server (NTRS)

    Milholen, William E., II; Jones, Gregory S.; Chan, David T.; Goodliff, Scott L.; Anders, Scott G.; Melton, Latunia P.; Carter, Melissa B.; Allan, Brian G.; Capone, Francis J.

    2013-01-01

    A second wind tunnel test of the FAST-MAC circulation control model was recently completed in the National Transonic Facility at the NASA Langley Research Center. The model was equipped with four onboard flow control valves allowing independent control of the circulation control plenums, which were directed over a 15% chord simple-hinged flap. The model was configured for low-speed high-lift testing with flap deflections of 30 and 60 degrees, along with the transonic cruise configuration with zero degree flap deflection. Testing was again conducted over a wide range of Mach numbers up to 0.88, and Reynolds numbers up to 30 million based on the mean chord. The first wind tunnel test had poor transonic force and moment data repeatability at mild cryogenic conditions due to inadequate thermal conditioning of the balance. The second test demonstrated that an improvement to the balance heating system significantly improved the transonic data repeatability, but also indicated further improvements are still needed. The low-speed highlift performance of the model was improved by testing various blowing slot heights, and the circulation control was again demonstrated to be effective in re-attaching the flow over the wing at off-design transonic conditions. A new tailored spanwise blowing technique was also demonstrated to be effective at transonic conditions with the benefit of reduced mass flow requirements.

  15. Blended-Wing-Body Transonic Aerodynamics: Summary of Ground Tests and Sample Results

    NASA Technical Reports Server (NTRS)

    Carter, Melissa B.; Vicroy, Dan D.; Patel, Dharmendra

    2009-01-01

    The Blended-Wing-Body (BWB) concept has shown substantial performance benefits over conventional aircraft configuration with part of the benefit being derived from the absence of a conventional empennage arrangement. The configuration instead relies upon a bank of trailing edge devices to provide control authority and augment stability. To determine the aerodynamic characteristics of the aircraft, several wind tunnel tests were conducted with a 2% model of Boeing's BWB-450-1L configuration. The tests were conducted in the NASA Langley Research Center's National Transonic Facility and the Arnold Engineering Development Center s 16-Foot Transonic Tunnel. Characteristics of the configuration and the effectiveness of the elevons, drag rudders and winglet rudders were measured at various angles of attack, yaw angles, and Mach numbers (subsonic to transonic speeds). The data from these tests will be used to develop a high fidelity simulation model for flight dynamics analysis and also serve as a reference for CFD comparisons. This paper provides an overview of the wind tunnel tests and examines the effects of Reynolds number, Mach number, pitch-pause versus continuous sweep data acquisition and compares the data from the two wind tunnels.

  16. Static Aeroelastic Predictions for a Transonic Transport Model Using an Unstructured-Grid Flow Solver Coupled With a Structural Plate Technique

    NASA Technical Reports Server (NTRS)

    Allison, Dennis O.; Cavallo, Peter A.

    2003-01-01

    An equivalent-plate structural deformation technique was coupled with a steady-state unstructured-grid three-dimensional Euler flow solver and a two-dimensional strip interactive boundary-layer technique. The objective of the research was to assess the extent to which a simple accounting for static model deformations could improve correlations with measured wing pressure distributions and lift coefficients at transonic speeds. Results were computed and compared to test data for a wing-fuselage model of a generic low-wing transonic transport at a transonic cruise condition over a range of Reynolds numbers and dynamic pressures. The deformations significantly improved correlations with measured wing pressure distributions and lift coefficients. This method provided a means of quantifying the role of dynamic pressure in wind-tunnel studies of Reynolds number effects for transonic transport models.

  17. Airframe-propulsion system aerodynamic interference predictions at high transonic Mach numbers including off-design engine airflow effects

    NASA Technical Reports Server (NTRS)

    Kulfan, R. M.; Sigalla, A.

    1981-01-01

    The transonic speed regime for airplanes at conditions where inlet spillage takes place is discussed. A wind tunnel test program to evaluate aerodynamic performance penalties associated with propulsion system installation and operation at subsonic through low supersonic speeds was conducted. The accuracy of analytic methods for predicting transonic engine airframe interference effects was assessed. Study variables included Mach number, angle of attack, relative nacelle location, and nacelle mass flow ratio. Results include test theory comparisons of forces as well as induced pressure fields. Prediction capability of induced shock wave strength and locations is assessed. It was found that large interference forces due to engine location and flow spillage occur at transonic speeds, that theory explains these effects; and that theory can predict quantitatively these effects.

  18. An experimental investigation of nacelle-pylon installation on an unswept wing at subsonic and transonic speeds

    NASA Technical Reports Server (NTRS)

    Carlson, J. R.; Compton, W. B., III

    1984-01-01

    A wind tunnel investigation was conducted to determine the aerodynamic interference associated with the installation of a long duct, flow-through nacelle on a straight unswept untapered supercritical wing. Experimental data was obtained for the verification of computational prediction techniques. The model was tested in the 16-Foot Transonic Tunnel at Mach numbers from 0.20 to 0.875 and at angles of attack from about 0 deg to 5 deg. The results of the investigation show that strong viscous and compressibility effects are present at the transonic Mach numbers. Numerical comparisons show that linear theory is adequate for subsonic Mach number flow prediction, but is inadequate for prediction of the extreme flow conditions that exist at the transonic Mach numbers.

  19. Recent National Transonic Facility Test Process Improvements (Invited)

    NASA Technical Reports Server (NTRS)

    Kilgore, W. A.; Balakrishna, S.; Bobbitt, C. W., Jr.; Adcock, J. B.

    2001-01-01

    This paper describes the results of two recent process improvements; drag feed-forward Mach number control and simultaneous force/moment and pressure testing, at the National Transonic Facility. These improvements have reduced the duration and cost of testing. The drag feedforward Mach number control reduces the Mach number settling time by using measured model drag in the Mach number control algorithm. Simultaneous force/moment and pressure testing allows simultaneous collection of force/moment and pressure data without sacrificing data quality thereby reducing the overall testing time. Both improvements can be implemented at any wind tunnel. Additionally the NTF is working to develop and implement continuous pitch as a testing option as an additional method to reduce costs and maintain data quality.

  20. Guide for users of the National Transonic Facility

    NASA Technical Reports Server (NTRS)

    Fuller, D. E.; Gloss, B. B.; Nystrom, D.

    1981-01-01

    The National Transonic Facility (NTF) is a fan-driven, closed-circuit, continuous flow, pressurized wind tunnel. The test section is 2.5 m x 2.5 m and 7.62 m long with a slotted-wall configuration. The NTF will have a Mach number range from 0.2 to 1.2, with Reynolds number up to 120 10 to the sixth power at Mach 1 (based on a reference length of 0.25 m). The pressure range for the facility will be from 1 to about 9 bars (1 ban = 100 kPa), and the temperature can be varied from 340 to 78 K. This report provides potential users of the NTF with the information required for preliminary planning to test programs and for preliminary layout of models and model supports which may be used in such programs.

  1. Heavy Gas Conversion of the NASA Langley Transonic Dynamics Tunnel

    NASA Technical Reports Server (NTRS)

    Corliss, James M.; Cole, Stanley, R.

    1998-01-01

    The heavy gas test medium has recently been changed in the Transonic Dynamics Tunnel (TDT) at the NASA Langley Research Center. A NASA Construction of Facilities project has converted the TDT heavy gas from dichlorodifluoromethane (R12) to 1,1,1,2 tetrafluoroethane (R134a). The facility s heavy gas processing system was extensively modified to implement the conversion to R134a. Additional system modifications have improved operator interfaces, hardware reliability, and quality of the research data. The facility modifications included improvements to the heavy gas compressor and piping, the cryogenic heavy gas reclamation system, and the heavy gas control room. A series of wind tunnel characterization and calibration tests are underway. Results of the flow characterization tests show the TDT operating envelope in R134a to be very similar to the previous operating envelope in R12.

  2. Experimental uncertainty and drag measurements in the national transonic facility

    NASA Technical Reports Server (NTRS)

    Batill, Stephen M.

    1994-01-01

    This report documents the results of a study which was conducted in order to establish a framework for the quantitative description of the uncertainty in measurements conducted in the National Transonic Facility (NTF). The importance of uncertainty analysis in both experiment planning and reporting results has grown significantly in the past few years. Various methodologies have been proposed and the engineering community appears to be 'converging' on certain accepted practices. The practical application of these methods to the complex wind tunnel testing environment at the NASA Langley Research Center was based upon terminology and methods established in the American National Standards Institute (ANSI) and the American Society of Mechanical Engineers (ASME) standards. The report overviews this methodology.

  3. A free flight investigation of transonic sting interference

    NASA Technical Reports Server (NTRS)

    Jaffe, P.

    1975-01-01

    Transonic sting interference has been studied in a supersonic wind tunnel to obtain free flight and sting support data on identical models. The two principal configurations, representing fuselage bodies, were cigar shaped with tail fins. The others were a sharp 10-deg cone, a sphere, and a blunt entry body. Comparative data indicated that the sting had an appreciable effect on drag for the fuselage-like configurations; drag rise occurred 0.02 Mach number earlier in free flight, and drag level was 15% greater. The spheres and the blunt bodies were insensitive to the presence of stings regardless of their size. The 10-deg cones were in between, experiencing no drag difference with a minimum diameter sting, but a moderate difference with the largest diameter sting tested. All data tend to confirm the notion that for the more slender bodies the sting not only affects flow but the forebody flow as well.

  4. Recent National Transonic Facility Test Process Improvements (Invited)

    NASA Technical Reports Server (NTRS)

    Kilgore, W. A.; Balakrishna, S.; Bobbitt, C. W., Jr.; Adcock, J. B.

    2001-01-01

    This paper describes the results of two recent process improvements; drag feed-forward Mach number control and simultaneous force/moment and pressure testing, at the National Transonic Facility. These improvements have reduced the duration and cost of testing. The drag feed-forward Mach number control reduces the Mach number settling time by using measured model drag in the Mach number control algorithm. Simultaneous force/moment and pressure testing allows simultaneous collection of force/moment and pressure data without sacrificing data quality thereby reducing the overall testing time. Both improvements can be implemented at any wind tunnel. Additionally the NTF is working to develop and implement continuous pitch as a testing option as an additional method to reduce costs and maintain data quality.

  5. Recent Productivity Improvements to the National Transonic Facility

    NASA Technical Reports Server (NTRS)

    Popernack, Thomas G., Jr.; Sydnor, George H.

    1998-01-01

    Productivity gains have recently been made at the National Transonic Facility wind tunnel at NASA Langley Research Center. A team was assigned to assess and set productivity goals to achieve the desired operating cost and output of the facility. Simulations have been developed to show the sensitivity of selected process productivity improvements in critical areas to reduce overall test cycle times. The improvements consist of an expanded liquid nitrogen storage system, a new fan drive, a new tunnel vent stack heater, replacement of programmable logic controllers, an increased data communications speed, automated test sequencing, and a faster model changeout system. Where possible, quantifiable results of these improvements are presented. Results show that in most cases, improvements meet the productivity gains predicted by the simulations.

  6. Transonic aeroelasticity analysis for rotor blades

    NASA Technical Reports Server (NTRS)

    Chow, Chuen-Yen; Chang, I-Chung; Gea, Lie-Mine

    1989-01-01

    A numerical method is presented for calculating the unsteady transonic rotor flow with aeroelasticity effects. The blade structural dynamic equations based on beam theory were formulated by FEM and were solved in the time domain, instead of the frequency domain. For different combinations of precone, droop, and pitch, the correlations are very good in the first three flapping modes and the first twisting mode. However, the predicted frequencies are too high for the first lagging mode at high rotational speeds. This new structure code has been coupled into a transonic rotor flow code, TFAR2, to demonstrate the capability of treating elastic blades in transonic rotor flow calculations. The flow fields for a model-scale rotor in both hover and forward flight are calculated. Results show that the blade elasticity significantly affects the flow characteristics in forward flight.

  7. Experience with transonic unsteady aerodynamic calculations

    NASA Technical Reports Server (NTRS)

    Edwards, J. W.; Bland, S. R.; Seidel, D. A.

    1985-01-01

    Comparisons of calculated and experimental transonic unsteady pressures and airloads for four of the AGARD Two Dimensional Aeroelastic Configurations and for a rectangular supercritical wing are presented. The two dimensional computer code, XTRAN2L, implementing the transonic small perturbation equation was used to obtain results for: (1) pitching oscillations of the NACA 64A010A; NLR 7301 and NACA 0012 airfoils; (2) flap oscillations for the NACA 64A006 and NRL 7301 airfoils; and (3) transient ramping motions for the NACA 0012 airfoils. Results from the three dimensional code XTRAN3S are compared with data from a rectangular supercritical wing oscillating in pitch. These cases illustrate the conditions under which the transonic inviscid small perturbation equation provides reasonable predictions.

  8. Inverse transonic airfoil design including viscous interaction

    NASA Technical Reports Server (NTRS)

    Carlson, L. A.

    1976-01-01

    A numerical technique was developed for the analysis of specified transonic airfoils or for the design of airfoils having a prescribed pressure distribution, including the effect of weak viscous interaction. The method uses the full potential equation, a stretched Cartesian coordinate system, and the Nash-MacDonald turbulent boundary layer method. Comparisons with experimental data for typical transonic airfoils show excellent agreement. An example shows the application of the method to design a thick aft-cambered airfoil, and the effects of viscous interaction on its performance are discussed.

  9. A vapor generator for transonic flow visualization

    NASA Technical Reports Server (NTRS)

    Bruce, Robert A.; Hess, Robert W.; Rivera, Jose A., Jr.

    1989-01-01

    A vapor generator was developed for use in the NASA Langley Transonic Dynamics Tunnel (TDT). Propylene glycol was used as the vapor material. The vapor generator system was evaluated in a laboratory setting and then used in the TDT as part of a laser light sheet flow visualization system. The vapor generator provided satisfactory seeding of the air flow with visible condensate particles, smoke, for tests ranging from low subsonic through transonic speeds for tunnel total pressures from atmospheric pressure down to less than 0.1 atmospheric pressure.

  10. Unsteady transonic flow over cascade blades

    NASA Technical Reports Server (NTRS)

    Surampudi, S. P.; Adamczyk, J. J.

    1986-01-01

    An attempt is made to develop an efficient staggered cascade blade unsteady aerodynamics model for the neighborhood of March 1, representing the blade row by a rectilinear two-dimensional cascade of thin, flat plate airfoils. The equations of motion are derived on the basis of linearized transonic small perturbation theory, and an analytical solution is obtained by means of the Wiener-Hopf procedure. Making use of the transonic similarity law, the results obtained are compared with those of other linearized cascade analyses. A parametric study is conducted to find the effects of reduced frequency, stagger angle, solidity, and the location of the pitching axis on cascade stability.

  11. Characteristics of the Langley 8-foot Transonic Tunnel with Slotted Test Section

    NASA Technical Reports Server (NTRS)

    Wright, Ray H; Ritchie, Virgil S; Pearson, Albin O

    1958-01-01

    A large wind tunnel, approximately 8 feet in diameter, has been converted to transonic operation by means of slots in the boundary extending in the direction of flow. The usefulness of such a slotted wind tunnel, already known with respect to the reduction of the subsonic blockage interference and the production of continuously variable supersonic flows, has been augmented by devising a slot shape with which a supersonic test region with excellent flow quality could be produced. Experimental locations of detached shock waves ahead of axially symmetric bodies at low supersonic speeds in the slotted test section agreed satisfactorily with predictions obtained by use of existing approximate methods.

  12. An evaluation of three experimental processes for two-dimensional transonic tests

    NASA Technical Reports Server (NTRS)

    Zuppardi, Gennaro

    1989-01-01

    The aerodynamic measurements in conventional wind tunnels usually suffer from the interference effects of the sting supporting the model and the test section walls. These effects are particularly severe in the transonic regime. Sting interference effects can be overcome through the Magnetic Suspension technique. Wall effects can be alleviated by: testing airfoils in conventional, ventilated tunnels at relatively small model to tunnel size ratios; treatment of the tunnel wall boundary layers; or by utilization of the Adaptive Wall Test Section (AWTS) concept. The operating capabilities and results from two of the foremost two-dimensional, transonic, AWTS facilities in existence are assessed. These facilities are the NASA 0.3-Meter Transonic Cryogenic Tunnel and the ONERA T-2 facility located in Toulouse, France. In addition, the results derived from the well known conventional facility, the NAE 5 ft x 5 ft Canadian wind tunnel will be assessed. CAST10/D0A2 Airfoil results will be used in all of the evaluations.

  13. Boundary-layer measurements on a transonic low-aspect ratio wing

    NASA Technical Reports Server (NTRS)

    Keener, Earl R.

    1985-01-01

    Tabulations and plots are presented of boundary-layer velocity and flow-direction surveys from wind-tunnel tests of a large-scale (0.90 m semi-span) model of the NASA/Lockheed Wing C. This wing is a generic, transonic, supercritical, highly three-dimensional, low-aspect-ratio configuration designed with the use of a three-dimensional, transonic full-potential-flow wing code (FLO22). Tests were conducted at the design angle of attack of 5 deg over a Mach number range from 0.25 to 0.96 and a Reynolds number range of 3.4x10 to the 6th power. Wing pressures were measured at five span stations, and boundary-layer surveys were measured at the midspan station. The data are presented without analysis.

  14. Transonic shock-induced dynamics of a flexible wing with a thick circular-arc airfoil

    NASA Technical Reports Server (NTRS)

    Bennett, Robert M.; Dansberry, Bryan E.; Farmer, Moses G.; Eckstrom, Clinton V.; Seidel, David A.; Rivera, Jose A., Jr.

    1991-01-01

    Transonic shock boundary layer oscillations occur on rigid models over a small range of Mach numbers on thick circular-arc airfoils. Extensive tests and analyses of this phenomena have been made in the past but essentially all of them were for rigid models. A simple flexible wing model with an 18 pct. circular arc airfoil was constructed and tested in the Langley Transonic Dynamics Tunnel to study the dynamic characteristics that a wing might have under these circumstances. In the region of shock boundary layer oscillations, buffeting of the first bending mode was obtained. This mode was well separated in frequency from the shock boundary layer oscillations. A limit cycle oscillation was also measured in a third bending like mode, involving wind vertical bending and splitter plate motion, which was in the frequency range of the shock boundary layer oscillations. Several model configurations were tested, and a few potential fixes were investigated.

  15. Analysis of Fluctuating Static Pressure Measurements in the National Transonic Facility

    NASA Technical Reports Server (NTRS)

    Igoe, William B.

    1996-01-01

    Dynamic measurements of fluctuating static pressure levels were taken with flush-mounted, high-frequency response pressure transducers at 11 locations in the circuit of the National Transonic Facility (NTF) across the complete operating range of this wind tunnel. Measurements were taken at test-section Mach numbers from 0.1 to 1.2, at pressures from 1 to 8.6 atm, and at temperatures from ambient to -250 F, which resulted in dynamic flow disturbance measurements at the highest Reynolds numbers available in a transonic ground test facility. Tests were also made by independent variation of the Mach number, the Reynolds number, or the fan drive power while the other two parameters were held constant, which for the first time resulted in a distinct separation of the effects of these three important parameters.

  16. Application of shock tubes to transonic airfoil testing at high Reynolds numbers

    NASA Technical Reports Server (NTRS)

    Cook, W. J.; Chaney, M. J.; Presley, L. L.; Chapman, G. T.

    1978-01-01

    Performance analysis of a gas-driven shock tube shows that transonic airfoil flows with chord Reynolds numbers of the order of 100 million can be produced, with limitations being imposed by the structural integrity of the facility or the model. A study of flow development over a simple circular arc airfoil at zero angle of attack was carried out in a shock tube at low and intermediate Reynolds numbers to assess the testing technique. Results obtained from schlieren photography and airfoil pressure measurements show that steady transonic flows similar to those produced for the same airfoil in a wind tunnel can be generated within the available testing time in a shock tube with properly contoured test section walls.

  17. Results of transonic wind tunnel tests on an 0.015-scale space shuttle mated vehicle model (67-ots) in the LaRC 8 foot TPT (IA41)

    NASA Technical Reports Server (NTRS)

    Hardin, R.; Burrows, R. R.

    1974-01-01

    Wind tunnel tests were conducted to obtain aerodynamic force data for Mach numbers from 0.60 to 1.20. Data were obtained for an alpha range of -10 deg to +10 deg (beta = 0 deg beta = 5 deg) and beta range of -10 deg to +10 deg (alpha = 0 deg). Longitudinal and lateral-directional stability and control data were obtained for tank alone, tank plus SRB's, tank plus Orbiter, and mated configuration of tank + Orbiter + SRB's. Also, single-component rudder hinge moment data were obtained at rudder deflections of 0 and -20 deg for each Mach number tested. Plots of aerodynamic coefficients vs. Mach number are presented, using data from both test IA41 and tests LRC-UPWT-1056, 1073 (IA42A/B) for Mach numbers of 1.60 to 4.63. The model tested in IA42A/B was the same model as tested in IA41.

  18. Aerodynamic interactions from reaction controls for lateral control of the M2-F2 lifting-body entry configuration at transonic and supersonic and supersonic Mach numbers. [wind tunnel tests

    NASA Technical Reports Server (NTRS)

    Bailey, R. O.; Brownson, J. J.

    1979-01-01

    Tests were conducted in the Ames 6 by 6 foot wind tunnel to determine the interaction of reaction jets for roll control on the M2-F2 lifting-body entry vehicle. Moment interactions are presented for a Mach number range of 0.6 to 1.7, a Reynolds number range of 1.2 x 10 to the 6th power to 1.6 x 10 to the 6th power (based on model reference length), an angle-of-attack range of -9 deg to 20 deg, and an angle-of-sideslip range of -6 deg to 6 deg at an angle of attack of 6 deg. The reaction jets produce roll control with small adverse yawing moment, which can be offset by horizontal thrust component of canted jets.

  19. Investigation of Seal-to-Floor Effects on Semi-Span Transonic Models

    NASA Technical Reports Server (NTRS)

    Sleppy, Mark A.; Engel, Eric A.; Watson, Kevin T.; Atler, Douglas M.

    2009-01-01

    In an effort to achieve the maximum possible Reynolds number (Re) when conducting production testing for flight loads aerodynamic databases, it has been the preferred practice of The Boeing Company / Commercial Airplanes (BCA) -- Loads and Dynamics Group since the early 1990's to test large scale semi-span models in the 11- By 11-Foot Transonic Wind Tunnel (TWT) leg of the Unitary Plan Wind Tunnel (UPWT) at the NASA Ames Research Center (ARC). There are many problems related to testing large scale semi-span models of high aspect ratio flexible transport wings, such as; floor boundary layer effects, wing spanwise wall effects, solid blockage buoyancy effects, floor mechanical interference effects, airflow under the model effects, or tunnel flow gradient effects. For most of these issues, BCA has developed and implemented either standard testing methods or numerical correction schemes and these will not be discussed in this document. Other researchers have reported on semi-span transonic testing correction issues, however most of the reported research has been for low Mach testing. Some of the reports for low Mach testing address the difficult problem of preventing undesirable airflow under a semi-span model while ensuring unrestricted main balance functionality, however, for transonic models this issue has gone unresolved. BCA has been cognizant for sometime that there are marked differences in wing pressure distributions from semi-span transonic model testing than from full model or flight testing. It has been suspected that these differences are at least in part due to airflow under the model. Previous efforts by BCA to address this issue have proven to be ineffective or inconclusive and in one situation resulted in broken hardware. This paper reports on a Boeing-NASA collaborative investigation based on a series of small tests conducted between June 2006 and November 2007 in the 11 by 11 foot Transonic Wind Tunnel at NASA Ames on three large commercial jet

  20. User's guide for a revised computer program to analyze the LRC 16 foot transonic dynamics tunnel active cable mount system. [computer techniques - aircraft models

    NASA Technical Reports Server (NTRS)

    Chin, J.; Barbero, P.

    1975-01-01

    The revision of an existing digital program to analyze the stability of models mounted on a two-cable mount system used in a transonic dynamics wind tunnel is presented. The program revisions and analysis of an active feedback control system to be used for controlling the free-flying models are treated.

  1. A New Forced Oscillation Capability for the Transonic Dynamics Tunnel

    NASA Technical Reports Server (NTRS)

    Piatak, David J.; Cleckner, Craig S.

    2002-01-01

    A new forced oscillation system has been installed and tested at NASA Langley Research Center's Transonic Dynamics Tunnel (TDT). The system is known as the Oscillating Turntable (OTT) and has been designed for the purpose of oscillating, large semispan models in pitch at frequencies up to 40 Hz to acquire high-quality unsteady pressure and loads data. Precisely controlled motions of a wind-tunnel model on the OTT can yield unsteady aerodynamic phenomena associated with flutter, limit cycle oscillations, shock dynamics, and non-linear aerodynamic effects on many vehicle configurations. This paper will discuss general design and components of the OTT and will present test data from performance testing and from research tests on two rigid semispan wind-tunnel models. The research tests were designed to challenge the OTT over a wide range of operating conditions while acquiring unsteady pressure data on a small rectangular supercritical wing and a large supersonic transport wing. These results will be presented to illustrate the performance capabilities, consistency of oscillations, and usefulness of the OTT as a research tool.

  2. Computer model to simulate testing at the National Transonic Facility

    NASA Technical Reports Server (NTRS)

    Mineck, Raymond E.; Owens, Lewis R., Jr.; Wahls, Richard A.; Hannon, Judith A.

    1995-01-01

    A computer model has been developed to simulate the processes involved in the operation of the National Transonic Facility (NTF), a large cryogenic wind tunnel at the Langley Research Center. The simulation was verified by comparing the simulated results with previously acquired data from three experimental wind tunnel test programs in the NTF. The comparisons suggest that the computer model simulates reasonably well the processes that determine the liquid nitrogen (LN2) consumption, electrical consumption, fan-on time, and the test time required to complete a test plan at the NTF. From these limited comparisons, it appears that the results from the simulation model are generally within about 10 percent of the actual NTF test results. The use of actual data acquisition times in the simulation produced better estimates of the LN2 usage, as expected. Additional comparisons are needed to refine the model constants. The model will typically produce optimistic results since the times and rates included in the model are typically the optimum values. Any deviation from the optimum values will lead to longer times or increased LN2 and electrical consumption for the proposed test plan. Computer code operating instructions and listings of sample input and output files have been included.

  3. Transonic pressure distributions on a rectangular supercritical wing oscillating in pitch

    NASA Technical Reports Server (NTRS)

    Ricketts, R. H.; Sandford, M. C.; Seidel, D. A.; Watson, J. J.

    1983-01-01

    Steady and unsteady aerodynamic data were measured on a rectangular wing with a 12 percent thick supercritical airfoil mounted in the NASA Langley Transonic Dynamics Tunnel. The wing was oscillated in pitch to generate the unsteady aerodynamic data. The purpose of the wind-tunnel test was to measure data for use in the development and assessment of transonic analytical codes. The effects on the wing pressure distributions of Mach number, mean angle of attack, and oscillation frequency and amplitude were measured. Results from the newly-developed XTRAN3S program (a nonlinear transonic small disturbance code) and from the RHOIV program (a linear lifting surface kernel function code) were compared to measured data for a Mach number of 0.7 and for oscillation frequencies ranging from 0 to 20 Hz. The XTRAN3S steady and unsteady results agreed fairly well with the measured results. The RHOIV unsteady-result agreement was fair but, of course, did not predict shock effects.

  4. Transonic pressure distributions on a rectangular supercritical wing oscillating in pitch

    NASA Technical Reports Server (NTRS)

    Ricketts, R. H.; Sandford, M. C.; Seidel, D. A.; Watson, J. J.

    1983-01-01

    Steady and unsteady aerodynamic data were measured on a rectangular wing with a 12 percent thick supercritical airfoil mounted in the NASA Langley Transonic Dynamics Tunnel. The wing was oscillated in pitch to generate the unsteady aerodynamic data. The purpose of the wind-tunnel test was to measure data for use in the development and assessment of transonic analytical codes. The effects on the wing pressure distributions of Mach number, mean angle of attack, and oscillation frequency and amplitude were measured. Results from the newly-developed XTRAN3S program (a non-linear transonic small disturbance code) and from the RHOIV program (a linear lifting surface kernel function code) were compared to measured data for a Mach number of 0.7 and for oscillation frequencies ranging from 0 to 20 Hz. The XTRAN3S steady and unsteady results agreed fairly well with the measured data. The RHOIV unsteady-result agreement was fair but, of course, did not predict shock effects.

  5. Free-To-Roll Analysis of Abrupt Wing Stall on Military Aircraft at Transonic Speeds

    NASA Technical Reports Server (NTRS)

    Owens, D. Bruce; Capone, Francis J.; Brandon, Jay M.; Cunningham, Kevin; Chambers, Joseph R.

    2003-01-01

    Transonic free-to-roll and static wind tunnel tests for four military aircraft - the AV-8B, the F/A-18C, the preproduction F/A-18E, and the F-16C - have been analyzed. These tests were conducted in the NASA Langley 16-Foot Transonic Tunnel as a part of the NASA/Navy/Air Force Abrupt Wing Stall Program. The objectives were to evaluate the utility of the free-to-roll test technique as a tool for predicting areas of significant uncommanded lateral motions and for gaining insight into the wing-drop and wing-rock behavior of military aircraft at transonic conditions. The analysis indicated that the free-to-roll results had good agreement with flight data on all four models. A wide range of motions - limit cycle wing rock, occasional and frequent damped wing drop/rock and wing rock divergence - were observed. The analysis shows the effects that the static and dynamic lateral stability can have on the wing drop/rock behavior. In addition, a free-to-roll figure of merit was developed to assist in the interpretation of results and assessment of the severity of the motions.

  6. Vector processor algorithms for transonic flow calculations

    NASA Technical Reports Server (NTRS)

    South, J. C., Jr.; Keller, J. D.; Hafez, M. M.

    1979-01-01

    This paper discusses a number of algorithms for solving the transonic full-potential equation in conservative form on a vector computer, such as the CDC STAR-100 or the CRAY-1. Recent research with the 'artificial density' method for transonics has led to development of some new iteration schemes which take advantage of vector-computer architecture without suffering significant loss of convergence rate. Several of these more promising schemes are described and 2-D and 3-D results are shown comparing the computational rates on the STAR and CRAY vector computers, and the CYBER-175 serial computer. Schemes included are: (1) Checkerboard SOR, (2) Checkerboard Leapfrog, (3) odd-even vertical line SOR, and (4) odd-even horizontal line SOR.

  7. Transonic turbine blade cascade testing facility

    NASA Technical Reports Server (NTRS)

    Verhoff, Vincent G.; Camperchioli, William P.; Lopez, Isaac

    1992-01-01

    NASA LeRC has designed and constructed a new state-of-the-art test facility. This facility, the Transonic Turbine Blade Cascade, is used to evaluate the aerodynamics and heat transfer characteristics of blade geometries for future turbine applications. The facility's capabilities make it unique: no other facility of its kind can combine the high degree of airflow turning, infinitely adjustable incidence angle, and high transonic flow rates. The facility air supply and exhaust pressures are controllable to 16.5 psia and 2 psia, respectively. The inlet air temperatures are at ambient conditions. The facility is equipped with a programmable logic controller with a capacity of 128 input/output channels. The data acquisition system is capable of scanning up to 1750 channels per sec. This paper discusses in detail the capabilities of the facility, overall facility design, instrumentation used in the facility, and the data acquisition system. Actual research data is not discussed.

  8. Complex configuration analysis at transonic speeds

    NASA Technical Reports Server (NTRS)

    Boppe, C. W.; Aidala, P. V.

    1980-01-01

    Advanced performance requirements of new combat and transport aircraft together with design time constraints intensify the development and application of three dimensional computational analyses. A computational method which was developed for the specific purpose of providing an engineering analysis of complex aircraft configurations at transonic speeds. Particular attention is given to the recently incorporated wing viscous interaction and canard capabilities. The treatment of fuselage fairings, nacelles, and pylons is reviewed. The means for keeping computing resources at reasonable levels are identified. Three configurations were selected for correlations with experimental data. Taken together, the comparisons illustrate the full extent of current analysis capabilities. The configurations include: (1) a wing fuselage canard fighter; (2) a transport with fuselage fairings, four nacelles, four pylons; and (3) a space vehicle which includes an external fuel tank and rocket boosters (transonic launch configuration).

  9. Results of the AFRSI detailed-environment test of the 0.035-scale SSV pressure-loads model 84-0 in the Ames 11x11 ft. TWT and the Lewis 8x6 ft. and 10x10 ft. SWT (OA-310A, B, C), volume 1

    NASA Technical Reports Server (NTRS)

    Marshall, B. A.; Marroquin, J.

    1984-01-01

    Detailed orbiter aerodynamic and aeroacoustic pressure data were obtained in a three-part experimental investigation (OA-310A, B and C). The tests were conducted in three NASA facilities: OA-310A in the Ames 11x11-foot Transonic Wind Tunnel; OA-310B in the Lewis 8x6-foot Supersonic Wind Tunnel; and OA-310C in the Lewis 10x10-foot Supersonic Wind Tunnel. Test data were obtained to support analysis of the Space Transportation System (STS)-6 advanced flexible reusable surface insulation (AFRSI) anomaly using the 0.035-scale space shuttle vehicle pressure-loads Model 84-0. Data were obtained in the areas of the orbiter where AFRSI is to be applied to OV-099 and OV-103. Emphasis was placed on acquiring detailed aeroacoustic data and time-averaged pressure distributions on five affected areas: (1) canopy; (2) side of fuselage; (3) upper surface of wing; (4) OMS pods; and (5) vertical tail. Data were obtained at nominal ascent and entry atmospheric flight trajectory conditions between M=0.6 through M-3.5. Sample plotted data are given. aba M.G.

  10. Flow instabilities in transonic small disturbance theory

    NASA Technical Reports Server (NTRS)

    Williams, M. H.; Bland, S. R.; Edwards, J. W.

    1985-01-01

    The dynamics of unsteady transonic small disturbance flows about two-dimensional airfoils is examined, with emphasis on the behavior in the region where the steady state flow is nonunique. It is shown that nonuniqueness results from an extremely long time scale instability which occurs in a finite Mach number and angle of attack range. The similarity scaling rules for the instability are presented and the possibility of similar behavior in the Euler equations is discussed.

  11. Studying Transonic Gases With a Hydraulic Analog

    NASA Technical Reports Server (NTRS)

    Wagner, W.; Lepore, F.

    1986-01-01

    Water table for hydraulic-flow research yields valuable information about gas flow at transonic speeds. Used to study fuel and oxidizer flow in high-pressure rocket engines. Method applied to gas flows in such equipment as furnaces, nozzles, and chemical lasers. Especially suitable when wall contours nonuniform, discontinuous, or unusually shaped. Wall shapes changed quickly for study and evaluated on spot. Method used instead of computer simulation when computer models unavailable, inaccurate, or costly to run.

  12. Magnus effects on spinning transonic missiles

    NASA Technical Reports Server (NTRS)

    Seginer, A.; Rosenwasser, I.

    1983-01-01

    Magnus forces and moments were measured on a basic-finner model spinning in transonic flow. Spin was induced by canted fins or by full-span or semi-span, outboard and inboard roll controls. Magnus force and moment reversals were caused by Mach number, reduced spin rate, and angle of attack variations. Magnus center of pressure was found to be independent of the angle of attack but varied with the Mach number and model configuration or reduced spin rate.

  13. Transonic airfoil and axial flow rotary machine

    SciTech Connect

    Nagai, Naonori; Iwatani, Junji

    2015-09-01

    Sectional profiles close to a tip 124 and a part between a midportion 125 and a hub 123 are shifted to the upstream of an operating fluid flow in a sweep direction. Accordingly, an S shape is formed in which the tip 124 and the part between the midportion 125 and the hub 123 protrude. As a result, it is possible reduce various losses due to shook, waves, thereby forming a transonic airfoil having an excellent aerodynamic characteristic.

  14. Recent Enhancements to the National Transonic Facility

    NASA Technical Reports Server (NTRS)

    Kilgore, W. A.; Balakrishna, S.; Bobbitt, C. W.; Underwood, P.

    2003-01-01

    The National Transonic Facility continues to make enhancements to provide quality data in a safe, efficient and cost effective method for aerodynamic ground testing. Recent enhancements discussed in this paper include the restoration of reliability and improved performance of the heat exchanger systems resulting in the expansion of the NTF air operations envelope. Additionally, results are presented from a continued effort to reduce model dynamics through the use of a new stiffer balance and sting

  15. Recent Enhancements to the National Transonic Facility

    NASA Technical Reports Server (NTRS)

    Kilgore, W. A.; Balakrishna, S.; Bobbitt, C. W.; Underwood, P.

    2003-01-01

    The National Transonic Facility continues to make enhancements to provide quality data in a safe, efficient and cost effective method for aerodynamic ground testing. Recent enhancements discussed in this paper include the restoration of reliability and improved performance of the heat exchanger systems resulting in the expansion of the NTF air operations envelope. Additionally, results are presented from a continued effort to reduce model dynamics through the use of a new stiffer balance and sting.

  16. Geometric and structural properties of a rectangular supercritical wing oscillated in pitch for measurement of unsteady transonic pressure distributions

    NASA Technical Reports Server (NTRS)

    Ricketts, R. H.; Watson, J. J.; Sandford, M. C.; Seidel, D. A.

    1983-01-01

    Wind-tunnel tests to measure unsteady aerodynamic data in the transonic region have been completed on an aspect ratio 2.0 rectangular wing with a supercritical airfoil. The geometric and structural properties of the wing are presented. (Other references contain the measured aerodynamic data.) Both measured and design airfoil coordinates are presented and compared. In addition, measured wing bending and torsional stiffness distributions and some trailing-edge flexibility influence coefficients are presented.

  17. Development of a code for wall contour design in the transonic region of axisymmetric and square nozzles

    NASA Technical Reports Server (NTRS)

    Alcenius, Timothy; Schneider, Steven P.

    1994-01-01

    Nozzle design codes developed earlier under NAG1-1133 were modified and used in order to design a supersonic wind tunnel nozzle with square cross sections. As part of the design process, a computer code was written to implement the Hopkins and Hill perturbation solution for the flow in the transonic region of axisymmetric nozzles. This technique is used to design the bleed slot of quiet-flow nozzles. This new design code is documented in this report.

  18. Installed Transonic 2D Nozzle Nacelle Boattail Drag Study

    NASA Technical Reports Server (NTRS)

    Malone, Michael B.; Peavey, Charles C.

    1999-01-01

    The Transonic Nozzle Boattail Drag Study was initiated in 1995 to develop an understanding of how external nozzle transonic aerodynamics effect airplane performance and how strongly those effects are dependent on nozzle configuration (2D vs. axisymmetric). MDC analyzed the axisymmetric nozzle. Boeing subcontracted Northrop-Grumman to analyze the 2D nozzle. AU participants analyzed the AGARD nozzle as a check-out and validation case. Once the codes were checked out and the gridding resolution necessary for modeling the separated flow in this region determined, the analysis moved to the installed wing/body/nacelle/diverter cases. The boat tail drag validation case was the AGARD B.4 rectangular nozzle. This test case offered both test data and previous CFD analyses for comparison. Results were obtained for test cases B.4.1 (M=0.6) and B.4.2 (M=0.938) and compared very well with the experimental data. Once the validation was complete a CFD grid was constructed for the full Ref. H configuration (wing/body/nacelle/diverter) using a combination of patched and overlapped (Chimera) grids. This was done to ensure that the grid topologies and density would be adequate for the full model. The use of overlapped grids allowed the same grids from the full configuration model to be used for the wing/body alone cases, thus eliminating the risk of grid differences affecting the determination of the installation effects. Once the full configuration model was run and deemed to be suitable the nacelle/diverter grids were removed and the wing/body analysis performed. Reference H wing/body results were completed for M=0.9 (a=0.0, 2.0, 4.0, 6.0 and 8.0), M=1.1 (a=4.0 and 6.0) and M=2.4 (a=0.0, 2.0, 4.4, 6.0 and 8.0). Comparisons of the M=0.9 and M=2.4 cases were made with available wind tunnel data and overall comparisons were good. The axi-inlet/2D nozzle nacelle was analyzed isolated. The isolated nacelle data coupled with the wing/body result enabled the interference effects of the

  19. Historical Review of Uncommanded Lateral-Directional Motions At Transonic Conditions (Invited)

    NASA Technical Reports Server (NTRS)

    Chambers, Joseph R.; Hall, Robert M.

    2003-01-01

    This paper presents the results of a survey of past experiences with uncommanded lateral-directional motions at transonic speeds during specific military aircraft programs. The effort was undertaken to provide qualitative and quantitative information on past airplane programs that might be of use to the participants in the joint NASA/Navy/Air Force Abrupt Wing Stall (AWS) Program. The AWS Program was initiated because of the experiences of the F/A-18E/F development program, during which unexpected, severe wing-drop motions were encountered by preproduction aircraft at transonic conditions. These motions were judged to be significantly degrading to the primary mission requirements of the aircraft. Although the problem was subsequently solved for the production version of the F/A-l8E/F, a high-level review panel emphasized the poor understanding of such phenomena and issued a strong recommendation to: Initiate a national research effort to thoroughly and systematically study the wing drop phenomena. A comprehensive, cooperative NASA/Navy/Air Force AWS Program was designed to respond to provide the required technology requirements. As part of the AWS Program, a work element was directed at a historical review of wing-drop experiences in past aircraft development programs at high subsonic and transonic speeds. In particular, information was requested regarding: specific aircraft configurations that exhibited uncommanded motions and the nature of the motions; geometric characteristics of the air- planes; flight conditions involved in occurrences; relevant data, including wind-tunnel, computational, and flight sources; figures of merit used for analyses; and approaches used to alleviate the problem. An attempt was also made to summarize some of the more important lessons learned from past experiences, and to recommend specific research efforts. In addition to providing technical information to assist the AWS research objectives, the study produced fundamental information

  20. Historical Review of Uncommanded Lateral-Directional Motions at Transonic Conditions

    NASA Technical Reports Server (NTRS)

    Chambers, Joseph R.; Hall, Robert M.

    2003-01-01

    This paper presents the results of a survey of past experiences with uncommanded lateral-directional motions at transonic speeds during specific military aircraft programs. The effort was undertaken to provide qualitative and quantitative information on past airplane programs that might be of use to the participants in the joint NASA/Navy/Air Force Abrupt Wing Stall (AWS) Program. The AWS Program was initiated because of the experiences of the F/A-l8E/F development program, during which unexpected, severe wing-drop motions were encountered by preproduction aircraft at transonic conditions. These motions were judged to be significantly degrading to the primary mission requirements of the aircraft. Although the problem was subsequently solved for the production version of the F/A-l8E/F, a high-level review panel emphasized the poor understanding of such phenomena and issued a strong recommendation to: "Initiate a national research effort to thoroughly and systematically study the wing drop phenomena." A comprehensive, cooperative NASA/Navy/Air Force AWS Program was designed to respond to provide the required technology requirements. As part of the AWS Program, a work element was directed at a historical review of wing-drop experiences in past aircraft development programs at high subsonic and transonic speeds. In particular, information was requested regarding: specific aircraft configurations that exhibited uncommanded motions and the nature of the motions; geometric characteristics of the air- planes; flight conditions involved in occurrences; relevant data, including wind-tunnel, computational, and flight sources; figures of merit used for analyses; and approaches used to alleviate the problem. An attempt was also made to summarize some of the more important lessons learned from past experiences, and to recommend specific research efforts. In addition to providing technical information to assist the AWS research objectives, the study produced fundamental

  1. Orbiter Model in Wind Tunnel

    NASA Technical Reports Server (NTRS)

    1980-01-01

    Marshall Space Flight Center (MSFC) engineer holding a replica of the proposed Liquid Booster Module, observes the testing of a small Space Shuttle orbiter model at 14 Wind Tunnel at MSFC. 14 Wind Tunnel is a trisonic wind tunnel, which is capable of running subsonic, transonic, and supersonic. It is used to test the integrity of rockets and launch vehicles in launch and reentry environments. The Wind Tunnel was used to test rockets and launch vehicles from the Jupiter C through the Saturn family up to the current Space Shuttle and will be used to test future advanced launch vehicles.

  2. Numerical calculation of the three dimensional transonic flow over a yawed wing.

    NASA Technical Reports Server (NTRS)

    Jameson, A.

    1973-01-01

    Results are presented of calculations of the three dimensional steady transonic flow over a finite yawed wing. The full potential flow equation is solved in a transformed coordinate system which permits the boundary conditions to be satisfied exactly. The correct differential properties are enforced by rotating the difference scheme to conform with the flow direction, and fast convergence is assured by simulating a time dependent equation designed to settle quickly to a steady state. Computed lift drag ratios are consistent with the results of wind tunnel tests of a yawed wing conducted by R. T. Jones (1972).-

  3. Description of 0.186-scale model of high-speed duct of national transonic facility

    NASA Technical Reports Server (NTRS)

    Gentry, C. L., Jr.; Igoe, W. B.; Fuller, D. E.

    1981-01-01

    The National Transonic Facility (NTF) is a pressurized cryogenic wind tunnel with a 2.5 m square test section. A 0.186-scale model of the NTF was used to simulate the aerodynamic performance of the components of the high-speed duct of the NTF. These components consist of a wide-angle diffuser, settling chamber, contraction section, test section, model support section, and high-speed diffuser. The geometry of the model tunnel, referred to as the diffuser flow apparatus is described, and some of its operating characteristics are presented.

  4. Sensitivity analysis of cool-down strategies for a transonic cryogenic tunnel

    NASA Technical Reports Server (NTRS)

    Thibodeaux, J. J.

    1982-01-01

    Guidelines and suggestions substantiated by real-time simulation data to ensure optimum time and energy use of injected liquid nitrogen for cooling the Langley 0.3-Meter Transonic Cryogenic Tunnel (TCT) are presented. It is directed toward enabling operators and researchers to become cognizant of criteria for using the 0.3-m TCT in an energy- or time-efficient manner. Operational recommendations were developed based on information collected from a validated simulator of the 0.3-m TCT and experimental data from the tunnel. Results and trends, however, can be extrapolated to other similarly constructed cryogenic wind tunnels.

  5. Control of Interacting Vortex Flows at Subsonic and Transonic Speeds Using Passive Porosity

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.

    2003-01-01

    A wind tunnel experiment was conducted in the NASA Langley Research Center (LaRC) 8-foot Transonic Pressure Tunnel (TPT) to determine the effects of passive surface porosity on vortex flow interactions about a general research fighter configuration at subsonic and transonic speeds. Flow- through porosity was applied to a wind leading-edge extension (LEX) mounted to a 65 deg cropped delta wind model to promote large nose-down pitching moment increments at high angles of attack. Porosity decreased the vorticity shed from the LEX, which weakened the LEX vortex and altered the global interactions of the LEX and wing vortices at high angles of attack. Six-component forces and moments and wing upper surface static pressure distributions were obtained at free- stream Mach numbers of 0.50, 0.85, and 1.20, Reynolds number of 2.5(10(exp-6) per foot, angles of attack up to 30 deg and angles of sideslip to plus or minus 8 deg. The off-surface flow field was visualized in selected cross-planes using a laser vapor screen flow visualization technique. Test data were obtained with a centerline vertical tail and with alternate twin, wing-mounted vertical fins having 0 deg and 30 deg cant angles. In addition, the porosity of the LEX was compartmentalized to determine the sensitivity of the vortex- dominated aerodynamics to the location and level of porosity applied to the LEX.

  6. Numerical solution of transonic wing flows using an Euler/Navier-Stokes zonal approach

    NASA Technical Reports Server (NTRS)

    Holst, T. L.; Gundy, K. L.; Thomas, S. D.; Chaderjian, N. M.; Flores, J.

    1985-01-01

    Transonic flow fields about wing geometries are computed using an Euler/Navier-Stokes approach in which the flow field is divided into several zones. The grid zones immediately adjacent to the wing surface are suitably clustered and solved with the Navier-Stokes equations. Grid zones removed from the wing are less finely clustered and are solved with the Euler equations. Wind tunnel wall effects are easily and accurately modeled with the new grid-zoning algorithm because the wind tunnel grid is constructed as an exact subset of the corresponding free-air grid. Solutions are obtained that are in good agreement with experiment, including cases with significant wind tunnel wall effects and shock-induced separation on the upper wing surface.

  7. ATRAN3S: An unsteady transonic code for clean wings

    NASA Technical Reports Server (NTRS)

    Guruswamy, G. P.; Goorjian, P. M.; Merritt, F. J.

    1985-01-01

    The development and applications of the unsteady transonic code ATRAN3S for clean wings are discussed. Explanations of the unsteady, transonic small-disturbance aerodynamic equations that are used and their solution procedures are discussed. A detailed user's guide, along with input and output for a sample case, is given.

  8. Transonic cryogenic test section for the Goettingen tube facility

    NASA Technical Reports Server (NTRS)

    Hornung, H.; Hefer, G.; Krogmann, P.; Stanewsky, E.

    1983-01-01

    The design of modern aircraft requires the solution of problems related to transonic flow at high Reynolds numbers. To investigate these problems experimentally, it is proposed to extend the Ludwieg tube facility by adding a transonic cryogenic test section. After stating the requirements for such a test section, the technical concept is briefly explained and a preliminary estimate of the costs is given.

  9. Numerical solution techniques for unsteady transonic aerodynamics problems

    NASA Technical Reports Server (NTRS)

    Ballhaus, W. F.; Bridgeman, J. O.

    1980-01-01

    Basic concepts of finite difference solution techniques for unsteady transonic flows are presented. The hierarchy of mathematical forumulations that approximate the Navier-Stokes equations are reviewed. The basic concepts involved in constructing numerical algorthms to solve these formulations are given. Semi-implicit and implicit schemes are constructed and analyzed. The discussion focuses primarily on techniques for solving the low frequency transonic small disturbance equation. This is the simplest formulation that contains the essence of inviscid unsteady transonic flow physics. The low frequency formulation is emphasized here because codes based on this theory can be run in minutes of processor time on currently available computers. Furthermore, numerical techniques involved in solving this simple formulation also apply to the more complicated formulations. Extensions to these formulations are briefly described. An indication of the present capability for solving unsteady transonic flows is provided. Important areas of future research for the advancement of computational unsteady transonic aerodynamics are described.

  10. The NASA Langley 8-foot Transonic Pressure Tunnel calibration

    NASA Technical Reports Server (NTRS)

    Brooks, Cuyler W., Jr.; Harris, Charles D.; Reagon, Patricia G.

    1994-01-01

    The NASA Langley 8-Foot Transonic Pressure Tunnel is a continuous-flow, variable-pressure wind tunnel with control capability to independently vary Mach number, stagnation pressure, stagnation temperature, and humidity. The top and bottom walls of the test section are axially slotted to permit continuous variation of the test section Mach number from 0.2 to 1.2, the slot-width contour provides a gradient-free test section 50 in. long for Mach numbers equal to or greater than 1.0 and 100 in. long for Mach numbers less than 1.0. The stagnation pressure may be varied from 0.25 to 2.0 atm. The tunnel test section has been recalibrated to determine the relationship between the free-stream Mach number and the test chamber reference Mach number. The hardware was the same as that of an earlier calibration in 1972 but the pressure measurement instrumentation available for the recalibration was about an order of magnitude more precise. The principal result of the recalibration was a slightly different schedule of reentry flap settings for Mach numbers from 0.80 to 1.05 than that determined during the 1972 calibration. Detailed tunnel contraction geometry, test section geometry, and limited test section wall boundary layer data are presented.

  11. Controlling Compressor Vane Flow Vectoring Angles at Transonic Speeds

    NASA Astrophysics Data System (ADS)

    Munson, Matthew; Rempfer, Dietmar; Williams, David; Acharya, Mukund

    2003-11-01

    The ability to control flow separation angles from compressor inlet guide vanes with a Coanda-type actuator is demonstrated using both wind tunnel experiments and finite element simulations. Vectoring angles up to 40 degrees from the uncontrolled baseline state were measured with helium schlieren visualization at transonic Mach numbers ranging from 0.1 to 0.6, and with airfoil chord Reynolds numbers ranging from 89,000 to 710,000. The magnitude of the vectoring angle is shown to depend upon the geometry of the trailing edge, and actuator slot size, and the momentum flux coefficient. Under certain conditions the blowing has no effect on the vectoring angle indicating that the Coanda effect is not present. DNS simulations with the finite element method investigated the effects of geometry changes and external flow. Continuous control of the vectoring angle is demonstrated, which has important implications for application to rotating machinery. The technique is shown to reduce the stall flow coefficient by 15 percent in an axial flow compressor.

  12. Numerical simulation of small perturbation transonic flows

    NASA Technical Reports Server (NTRS)

    Seebass, A. R.; Yu, N. J.

    1976-01-01

    The results of a systematic study of small perturbation transonic flows are presented. Both the flow over thin airfoils and the flow over wedges were investigated. Various numerical schemes were employed in the study. The prime goal of the research was to determine the efficiency of various numerical procedures by accurately evaluating the wave drag, both by computing the pressure integral around the body and by integrating the momentum loss across the shock. Numerical errors involved in the computations that affect the accuracy of drag evaluations were analyzed. The factors that effect numerical stability and the rate of convergence of the iterative schemes were also systematically studied.

  13. Parker Winds Revisited: An Extension to Disk Winds

    NASA Astrophysics Data System (ADS)

    Waters, T. R.; Proga, D.

    2012-08-01

    A simple, one-dimensional dynamical model of thermally driven disk winds, one in the spirit of the original Parker (1958) model, is presented. We consider two different axi-symmetric streamline geometries: geometry (i) is commonly used in kinematic models to compute synthetic spectra, while geometry (ii), which exhibits self-similarity and more closely resembles the geometry found by many numerical simulations of disk winds, is likely unused for this purpose, although it easily could be with existing kinematic models. We make the case that it should be, i.e., that geometry (ii) leads to transonic wind solutions with substantially different properties.

  14. Overview of the Space Launch System Transonic Buffet Environment Test Program

    NASA Technical Reports Server (NTRS)

    Piatak, David J.; Sekula, Martin K.; Rausch, Russ D.; Florance, James R.; Ivanco, Thomas G.

    2015-01-01

    Fluctuating aerodynamic loads are a significant concern for the structural design of a launch vehicle, particularly while traversing the transonic flight environment. At these trajectory conditions, unsteady aerodynamic pressures can excite the vehicle dynamic modes of vibration and result in high structural bending moments and vibratory environments. To ensure that vehicle structural components and subsystems possess adequate strength, stress, and fatigue margins in the presence of buffet and other environments, buffet forcing functions are required to conduct the coupled load analysis of the launch vehicle. The accepted method to obtain these buffet forcing functions is to perform wind-tunnel testing of a rigid model that is heavily instrumented with unsteady pressure transducers designed to measure the buffet environment within the desired frequency range. Two wind-tunnel tests of a 3 percent scale rigid buffet model have been conducted at the Langley Research Center Transonic Dynamics Tunnel (TDT) as part of the Space Launch System (SLS) buffet test program. The SLS buffet models have been instrumented with as many as 472 unsteady pressure transducers to resolve the buffet forcing functions of this multi-body configuration through integration of the individual pressure time histories. This paper will discuss test program development, instrumentation, data acquisition, test implementation, data analysis techniques, and several methods explored to mitigate high buffet environment encountered during the test program. Preliminary buffet environments will be presented and compared using normalized sectional buffet forcing function root-meansquared levels along the vehicle centerline.

  15. Investigation of Aerodynamics Scale Effects for a Generic Fighter Configuration in the National Transonic Facility (Invited)

    NASA Technical Reports Server (NTRS)

    Tomek, W. G.; Wahls, R. A.; Owens, L. R.; Burner, A. B.; Graves, S. S.; Luckring, J. M.

    2003-01-01

    Two wind tunnel tests of a generic fighter configuration have been completed in the National Transonic Facility. The primary purpose of the tests 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 tests 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 lateral-directional aerodynamic characteristics are presented herein, along with longitudinal data demonstrating the effects of fixing the boundary layer transition location for low Reynolds number conditions. In addition, an improved model videogrammetry system and results are discussed.

  16. Analysis of Ares Crew Launch Vehicle Transonic Alternating Flow Phenomenon

    NASA Technical Reports Server (NTRS)

    Sekula, Martin K.; Piatak, David J.; Rausch, Russ D.

    2012-01-01

    A transonic wind tunnel test of the Ares I-X Rigid Buffet Model (RBM) identified a Mach number regime where unusually large buffet loads are present. A subsequent investigation identified the cause of these loads to be an alternating flow phenomenon at the Crew Module-Service Module junction. The conical design of the Ares I-X Crew Module and the cylindrical design of the Service Module exposes the vehicle to unsteady pressure loads due to the sudden transition between a subsonic separated and a supersonic attached flow about the cone-cylinder junction as the local flow randomly fluctuates back and forth between the two flow states. These fluctuations produce a square-wave like pattern in the pressure time histories resulting in large amplitude, impulsive buffet loads. Subsequent testing of the Ares I RBM found much lower buffet loads since the evolved Ares I design includes an ogive fairing that covers the Crew Module-Service Module junction, thereby making the vehicle less susceptible to the onset of alternating flow. An analysis of the alternating flow separation and attachment phenomenon indicates that the phenomenon is most severe at low angles of attack and exacerbated by the presence of vehicle protuberances. A launch vehicle may experience either a single or, at most, a few impulsive loads since it is constantly accelerating during ascent rather than dwelling at constant flow conditions in a wind tunnel. A comparison of a windtunnel- test-data-derived impulsive load to flight-test-data-derived load indicates a significant over-prediction in the magnitude and duration of the buffet load. I. Introduction One

  17. The NASA Langley 16-Foot Transonic Tunnel: Historical Overview, Facility Description, Calibration, Flow Characteristics, and Test Capabilities

    NASA Technical Reports Server (NTRS)

    Capone, Francis J.; Bangert, Linda S.; Asbury, Scott C.; Mills, Charles T. L.; Bare, E. Ann

    1995-01-01

    The Langley 16-Foot Transonic Tunnel is a closed-circuit single-return atmospheric wind tunnel that has a slotted octagonal test section with continuous air exchange. The wind tunnel speed can be varied continuously over a Mach number range from 0.1 to 1.3. Test-section plenum suction is used for speeds above a Mach number of 1.05. Over a period of some 40 years, the wind tunnel has undergone many modifications. During the modifications completed in 1990, a new model support system that increased blockage, new fan blades, a catcher screen for the first set of turning vanes, and process controllers for tunnel speed, model attitude, and jet flow for powered models were installed. This report presents a complete description of the Langley 16-Foot Transonic Tunnel and auxiliary equipment, the calibration procedures, and the results of the 1977 and the 1990 wind tunnel calibration with test section air removal. Comparisons with previous calibrations showed that the modifications made to the wind tunnel had little or no effect on the aerodynamic characteristics of the tunnel. Information required for planning experimental investigations and the use of test hardware and model support systems is also provided.

  18. Asymptotic methods for internal transonic flows

    NASA Technical Reports Server (NTRS)

    Adamson, T. C., Jr.; Messiter, A. F.

    1989-01-01

    For many internal transonic flows of practical interest, some of the relevant nondimensional parameters typically are small enough that a perturbation scheme can be expected to give a useful level of numerical accuracy. A variety of steady and unsteady transonic channel and cascade flows is studied with the help of systematic perturbation methods which take advantage of this fact. Asymptotic representations are constructed for small changes in channel cross-section area, small flow deflection angles, small differences between the flow velocity and the sound speed, small amplitudes of imposed oscillations, and small reduced frequencies. Inside a channel the flow is nearly one-dimensional except in thin regions immediately downstream of a shock wave, at the channel entrance and exit, and near the channel throat. A study of two-dimensional cascade flow is extended to include a description of three-dimensional compressor-rotor flow which leads to analytical results except in thin edge regions which require numerical solution. For unsteady flow the qualitative nature of the shock-wave motion in a channel depends strongly on the orders of magnitude of the frequency and amplitude of impressed wall oscillations or fluctuations in back pressure. One example of supersonic flow is considered, for a channel with length large compared to its width, including the effect of separation bubbles and the possibility of self-sustained oscillations. The effect of viscosity on a weak shock wave in a channel is discussed.

  19. Transonic separated solutions for an augmentor-wing

    NASA Technical Reports Server (NTRS)

    Flores, J.; Van Dalsem, W. R.

    1985-01-01

    The viscous transonic flow about a multielement airfoil (augmentor-wing) is simulated by coupling full-potential and direct/inverse differential boundary-layer algorithms. Solutions have been obtained for a variety of conditions and are in fair agreement with available experimental data. Typical results from this transonic augmentor-wing code (TAUG-V) require approximately three minutes of CRAY-XMP CPU time. Since this viscous transonic code accounts for most of the important flow physics, yet is still economical, it is a practical tool for the design aerodynamicist.

  20. Control of the NASA Langley 16-Foot Transonic Tunnel with the Self-Organizing Map

    NASA Technical Reports Server (NTRS)

    Motter, Mark A.

    1999-01-01

    A predictive, multiple model control strategy is developed based on an ensemble of local linear models of the nonlinear system dynamics for a transonic wind tunnel. The local linear models are estimated directly from the weights of a self-organizing map (SOM). Multiple self-organizing maps collectively model the global response of the wind tunnel to a finite set of representative prototype controls. These prototype controls partition the control space and incorporate experiential knowledge gained from decades of operation. Each SOM models the combination of the tunnel with one of the representative controls, over the entire range of operation. The SOM based linear models are used to predict the tunnel response to a larger family of control sequences which are clustered on the representative prototypes. The control sequence which corresponds to the prediction that best satisfies the requirements on the system output is applied as the external driving signal.

  1. A study of test section configuration for shock tube testing of transonic airfoils

    NASA Technical Reports Server (NTRS)

    Cook, W. J.

    1978-01-01

    Two methods are investigated for alleviating wall interference effects in a shock tube test section intended for testing two-dimensional transonic airfoils. The first method involves contouring the test section walls to match approximate streamlines in the flow. Contours are matched to each airfoil tested to produce results close to those obtained in a conventional wind tunnel. Data from a previous study and the present study for two different airfoils demonstrate that useful results are obtained in a shock tube using a test section with contoured walls. The second method involves use of a fixed-geometry slotted-wall test section to provide automatic flow compensation for various airfoils. The slotted-wall test section developed exhibited the desired performance characteristics in the approximate Mach number range 0.82 to 0.89, as evidenced by good agreement obtained between shock tube and wind tunnel results for several airfoil flows.

  2. Contributions of the NASA Langley Transonic Dynamics Tunnel to Launch Vehicle and Spacecraft Development

    NASA Technical Reports Server (NTRS)

    Cole, Stanley R.; Keller, Donald F.; Piatak, David J.

    2000-01-01

    The NASA Langley Transonic Dynamics Tunnel (TDT) has provided wind-tunnel experimental validation and research data for numerous launch vehicles and spacecraft throughout its forty year history. Most of these tests have dealt with some aspect of aeroelastic or unsteady-response testing, which is the primary purpose of the TDT facility. However, some space-related test programs that have not involved aeroelasticity have used the TDT to take advantage of specific characteristics of the wind-tunnel facility. In general. the heavy gas test medium, variable pressure, relatively high Reynolds number and large size of the TDT test section have made it the preferred facility for these tests. The space-related tests conducted in the TDT have been divided into five categories. These categories are ground wind loads, launch vehicle dynamics, atmospheric flight of space vehicles, atmospheric reentry. and planetary-probe testing. All known TDT tests of launch vehicles and spacecraft are discussed in this report. An attempt has been made to succinctly summarize each wind-tunnel test, or in the case of multiple. related tests, each wind-tunnel program. Most summaries include model program discussion, description of the physical wind-tunnel model, and some typical or significant test results. When available, references are presented to assist the reader in further pursuing information on the tests.

  3. Introduction to cryogenic wind tunnels

    NASA Technical Reports Server (NTRS)

    Goodyer, M. J.

    1985-01-01

    The background to the evolution of the cryogenic wind tunnel is outlined, with particular reference to the late 60's/early 70's when efforts were begun to re-equip with larger wind tunnels. The problems of providing full scale Reynolds numbers in transonic testing were proving particularly intractible, when the notion of satisfying the needs with the cryogenic tunnel was proposed, and then adopted. The principles and advantages of the cryogenic tunnel are outlined, along with guidance on the coolant needs when this is liquid nitrogen, and with a note on energy recovery. Operational features of the tunnels are introduced with reference to a small low speed tunnel. Finally the outstanding contributions are highlighted of the 0.3-Meter Transonic Cryogenic Tunnel (TCT) at NASA Langley Research Center, and its personnel, to the furtherance of knowledge and confidence in the concept.

  4. Integral equations for flows in wind tunnels

    NASA Technical Reports Server (NTRS)

    Fromme, J. A.; Golberg, M. A.

    1979-01-01

    This paper surveys recent work on the use of integral equations for the calculation of wind tunnel interference. Due to the large number of possible physical situations, the discussion is limited to two-dimensional subsonic and transonic flows. In the subsonic case, the governing boundary value problems are shown to reduce to a class of Cauchy singular equations generalizing the classical airfoil equation. The theory and numerical solution are developed in some detail. For transonic flows nonlinear singular equations result, and a brief discussion of the work of Kraft and Kraft and Lo on their numerical solution is given. Some typical numerical results are presented and directions for future research are indicated.

  5. Aerodynamic Modeling of Transonic Aircraft Using Vortex Lattice Coupled with Transonic Small Disturbance for Conceptual Design

    NASA Technical Reports Server (NTRS)

    Chaparro, Daniel; Fujiwara, Gustavo E. C.; Ting, Eric; Nguyen, Nhan

    2016-01-01

    The need to rapidly scan large design spaces during conceptual design calls for computationally inexpensive tools such as the vortex lattice method (VLM). Although some VLM tools, such as Vorview have been extended to model fully-supersonic flow, VLM solutions are typically limited to inviscid, subcritical flow regimes. Many transport aircraft operate at transonic speeds, which limits the applicability of VLM for such applications. This paper presents a novel approach to correct three-dimensional VLM through coupling of two-dimensional transonic small disturbance (TSD) solutions along the span of an aircraft wing in order to accurately predict transonic aerodynamic loading and wave drag for transport aircraft. The approach is extended to predict flow separation and capture the attenuation of aerodynamic forces due to boundary layer viscosity by coupling the TSD solver with an integral boundary layer (IBL) model. The modeling framework is applied to the NASA General Transport Model (GTM) integrated with a novel control surface known as the Variable Camber Continuous Trailing Edge Flap (VCCTEF).

  6. TAIR- TRANSONIC AIRFOIL ANALYSIS COMPUTER CODE

    NASA Technical Reports Server (NTRS)

    Dougherty, F. C.

    1994-01-01

    The Transonic Airfoil analysis computer code, TAIR, was developed to employ a fast, fully implicit algorithm to solve the conservative full-potential equation for the steady transonic flow field about an arbitrary airfoil immersed in a subsonic free stream. The full-potential formulation is considered exact under the assumptions of irrotational, isentropic, and inviscid flow. These assumptions are valid for a wide range of practical transonic flows typical of modern aircraft cruise conditions. The primary features of TAIR include: a new fully implicit iteration scheme which is typically many times faster than classical successive line overrelaxation algorithms; a new, reliable artifical density spatial differencing scheme treating the conservative form of the full-potential equation; and a numerical mapping procedure capable of generating curvilinear, body-fitted finite-difference grids about arbitrary airfoil geometries. Three aspects emphasized during the development of the TAIR code were reliability, simplicity, and speed. The reliability of TAIR comes from two sources: the new algorithm employed and the implementation of effective convergence monitoring logic. TAIR achieves ease of use by employing a "default mode" that greatly simplifies code operation, especially by inexperienced users, and many useful options including: several airfoil-geometry input options, flexible user controls over program output, and a multiple solution capability. The speed of the TAIR code is attributed to the new algorithm and the manner in which it has been implemented. Input to the TAIR program consists of airfoil coordinates, aerodynamic and flow-field convergence parameters, and geometric and grid convergence parameters. The airfoil coordinates for many airfoil shapes can be generated in TAIR from just a few input parameters. Most of the other input parameters have default values which allow the user to run an analysis in the default mode by specifing only a few input parameters

  7. Transonic Wing Shape Optimization Using a Genetic Algorithm

    NASA Technical Reports Server (NTRS)

    Holst, Terry L.; Pulliam, Thomas H.; Kwak, Dochan (Technical Monitor)

    2002-01-01

    A method for aerodynamic shape optimization based on a genetic algorithm approach is demonstrated. The algorithm is coupled with a transonic full potential flow solver and is used to optimize the flow about transonic wings including multi-objective solutions that lead to the generation of pareto fronts. The results indicate that the genetic algorithm is easy to implement, flexible in application and extremely reliable.

  8. Separated transonic airfoil flow calculations with a nonequilibrium turbulence model

    NASA Technical Reports Server (NTRS)

    King, L. S.; Johnson, D. A.

    1985-01-01

    Navier-Stokes transonic airfoil calculations based on a recently developed nonequilibrium, turbulence closure model are presented for a supercritical airfoil section at transonic cruise conditions and for a conventional airfoil section at shock-induced stall conditions. Comparisons with experimental data are presented which show that this nonequilibrium closure model performs significantly better than the popular Baldwin-Lomax and Cebeci-Smith equilibrium algebraic models when there is boundary-layer separation that results from the inviscid-viscous interactions.

  9. Detailed transonic flow field measurements about a supercritical airfoil section

    NASA Technical Reports Server (NTRS)

    Hurley, F. X.; Spaid, F. W.; Roos, F. W.; Stivers, L. S., Jr.; Bandettini, A.

    1975-01-01

    The transonic flow field about a Whitcomb-type supercritical airfoil profile was measured in detail. In addition to the usual surface pressure distributions and wake surveys, schlieren photographs were taken and velocity vector profiles were determined in the upper surface boundary layer and in the near wake. Spanwise variations in the measured pressures were also determined. The data are analyzed with the aid of an inviscid transonic finite-difference computer program as well as with boundary layer modeling and calculation schemes.

  10. The National Transonic Facility: A Research Retrospective

    NASA Technical Reports Server (NTRS)

    Wahls, R. A.

    2001-01-01

    An overview of the National Transonic Facility (NTF) from a research utilization perspective is provided. The facility was born in the 1970s from an internationally recognized need for a high Reynolds number test capability based on previous experiences with preflight predictions of aerodynamic characteristics and an anticipated need in support of research and development for future aerospace vehicle systems. Selection of the cryogenic concept to meet the need, unique capabilities of the facility, and the eventual research utilization of the facility are discussed. The primary purpose of the paper is to expose the range of investigations that have used the NTF since being declared operational in late 1984; limited research results are included, though many more can be found in the references.

  11. Flutter Analysis of a Transonic Fan

    NASA Technical Reports Server (NTRS)

    Srivastava, R.; Bakhle, M. A.; Keith, T. G., Jr.; Stefko, G. L.

    2002-01-01

    This paper describes the calculation of flutter stability characteristics for a transonic forward swept fan configuration using a viscous aeroelastic analysis program. Unsteady Navier-Stokes equations are solved on a dynamically deforming, body fitted, grid to obtain the aeroelastic characteristics using the energy exchange method. The non-zero inter-blade phase angle is modeled using phase-lagged boundary conditions. Results obtained show good correlation with measurements. It is found that the location of shock and variation of shock strength strongly influenced stability. Also, outboard stations primarily contributed to stability characteristics. Results demonstrate that changes in blade shape impact the calculated aerodynamic damping, indicating importance of using accurate blade operating shape under centrifugal and steady aerodynamic loading for flutter prediction. It was found that the calculated aerodynamic damping was relatively insensitive to variation in natural frequency.

  12. Transonic Flow Computations Using Nonlinear Potential Methods

    NASA Technical Reports Server (NTRS)

    Holst, Terry L.; Kwak, Dochan (Technical Monitor)

    2000-01-01

    This presentation describes the state of transonic flow simulation using nonlinear potential methods for external aerodynamic applications. The presentation begins with a review of the various potential equation forms (with emphasis on the full potential equation) and includes a discussion of pertinent mathematical characteristics and all derivation assumptions. Impact of the derivation assumptions on simulation accuracy, especially with respect to shock wave capture, is discussed. Key characteristics of all numerical algorithm types used for solving nonlinear potential equations, including steady, unsteady, space marching, and design methods, are described. Both spatial discretization and iteration scheme characteristics are examined. Numerical results for various aerodynamic applications are included throughout the presentation to highlight key discussion points. The presentation ends with concluding remarks and recommendations for future work. Overall. nonlinear potential solvers are efficient, highly developed and routinely used in the aerodynamic design environment for cruise conditions. Published by Elsevier Science Ltd. All rights reserved.

  13. A transonic rectangular grid embedded panel method

    NASA Technical Reports Server (NTRS)

    Johnson, F. T.; Bussoletti, J. E.; James, R. M.; Young, D. P.; Woo, A. C.

    1982-01-01

    A method is presented that has the potential for solving transonic flow problems about the same complex aircraft configurations currently being analyzed by subsonic panel methods. This method does not require the generation of surface fitted grids. Instead it uses rectangular grids and subgrids together with embedded surface panels on which boundary conditions are imposed. Both the Euler and full potential equations are considered. The method of least squares is used to reduce the solution of these equations to the solution of a sequence of Poisson problems. The Poisson problems are solved using fast Fourier transforms and panel influence coefficient techniques. The overall method is still in its infancy but some two dimensional results are shown illustrating various key features.

  14. Transonic flow visualization using holographic interferometry

    NASA Technical Reports Server (NTRS)

    Bryanston-Cross, Peter J.

    1987-01-01

    An account is made of some of the applications of holographic interferometry to the visualization of transonic flows. In the case of the compressor shock visualization, the method is used regularly and has moved from being a research department invention to a design test tool. With the implementation of automatic processing and simple digitization systems, holographic vibrational analysis has also moved into routine nondestructive testing. The code verification interferograms were instructive, but the main turbomachinery interest is now in 3 dimensional flows. A major data interpretation effort will be required to compute tomographically the 3 dimensional flow around the leading or the trailing edges of a rotating blade row. The bolt on approach shows the potential application to current unsteady flows of interest. In particular that of the rotor passing and vortex interaction effects is experienced by the new generation of unducted fans. The turbocharger tests presents a new area for the application of holography.

  15. Laser velocimetry measurement in a transonic tunnel

    NASA Astrophysics Data System (ADS)

    Ng, T. T.; Mueller, T. J.

    1985-10-01

    Some preliminary velocity measurements were carried out inside the transonic tunnel using the laser velocimeter (LV) system in association with the smoke generator. Pressure measurements were also performed using a pressure tap located on the side wall of the test section slightly upstream of the windows. Though the pressure measurements and the LV measurements were not taken at exactly the same location, extrapolation of the pressure data into the location of the LV measurements indicated a very close agreement between the velocity values obtained using the two different methods. Thus it is believed that the smoke particle is following the air flow with little or no velocity slip. Velocity measurements with airfoil at various angles of attack are now being carried out in conjuction with schlieren flow visualization. In the near future pressure distribution around and on the airfoil will be obtained by putting pressure taps on the side windows and using a pressure tap model (currently under construction) of the airfoil.

  16. Laser velocimetry measurement in a transonic tunnel

    NASA Technical Reports Server (NTRS)

    Ng, T. T.; Mueller, T. J.

    1985-01-01

    Some preliminary velocity measurements were carried out inside the transonic tunnel using the laser velocimeter (LV) system in association with the smoke generator. Pressure measurements were also performed using a pressure tap located on the side wall of the test section slightly upstream of the windows. Though the pressure measurements and the LV measurements were not taken at exactly the same location, extrapolation of the pressure data into the location of the LV measurements indicated a very close agreement between the velocity values obtained using the two different methods. Thus it is believed that the smoke particle is following the air flow with little or no velocity slip. Velocity measurements with airfoil at various angles of attack are now being carried out in conjuction with schlieren flow visualization. In the near future pressure distribution around and on the airfoil will be obtained by putting pressure taps on the side windows and using a pressure tap model (currently under construction) of the airfoil.

  17. Wind Tunnel Wall Interference Assessment and Correction, 1983

    NASA Technical Reports Server (NTRS)

    Newman, P. A. (Editor); Barnwell, R. W. (Editor)

    1984-01-01

    Technical information focused upon emerging wall interference assessment/correction (WIAC) techniques applicable to transonic wind tunnels with conventional and passively or partially adapted walls is given. The possibility of improving the assessment and correction of data taken in conventional transonic wind tunnels by utilizing simultaneously obtained flow field data (generally taken near the walls) appears to offer a larger, nearer-term payoff than the fully adaptive wall concept. Development of WIAC procedures continues, and aspects related to validating the concept need to be addressed. Thus, the scope of wall interference topics discussed was somewhat limited.

  18. Subsonic Transonic Applied Refinements By Using Key Strategies - STARBUKS In the NASA Langley Research Center National Transonic Facility

    NASA Technical Reports Server (NTRS)

    Paryz, Roman W.

    2014-01-01

    Several upgrade projects have been completed at the NASA Langley Research Center National Transonic Facility over the last 1.5 years in an effort defined as STARBUKS - Subsonic Transonic Applied Refinements By Using Key Strategies. This multi-year effort was undertaken to improve NTF's overall capabilities by addressing Accuracy and Validation, Productivity, and Reliability areas at the NTF. This presentation will give a brief synopsis of each of these efforts.

  19. Interference-free wind-tunnel flows by adaptive-wall technology

    NASA Technical Reports Server (NTRS)

    Sears, W. R.; Vidal, R. J.; Erickson, J. C., Jr.; Ritter, A.

    1976-01-01

    The adaptive-wall or self-correcting wind tunnel has been proposed for such regimes as transonic and V/STOL where wall effects are large and cannot be corrected for. The power and generality of the concept are pointed out. In a two-dimensional transonic embodiment in the Calspan One-Foot Tunnel, the scheme has been shown to work at lower transonic Mach numbers. Several practical problems are cited, including instrumentation, the nature of the wall modification, and convergence of the iterative procedure. Moreover, questions of shock-wave neutralization at the wall and probable configuration of three-dimensional embodiments are discussed.

  20. High Reynolds number tests of the CAST-10-2/DOA 2 transonic airfoil at ambient and cryogenic temper ature conditions

    NASA Technical Reports Server (NTRS)

    Stanewsky, E.; Demurie, F.; Ray, Edward J.; Johnson, C. B.

    1989-01-01

    The transonic airfoil CAST 10-2/DOA 2 was investigated in several major transonic wind tunnels at Reynolds numbers ranging from Re=1.3 x 10(exp 6) to 45 x 10(exp 6) at ambient and cryogenic temperature conditions. The main objective was to study the degree and extent of the effects of Reynolds number on both the airfoil aerodynamic characteristics and the interference effects of various model-wind-tunnel systems. The initial analysis of the CAST 10-2 airfoil results revealed appreciable real Reynolds number effects on this airfoil and showed that wall interference can be significantly affected by changes in Reynolds number thus appearing as true Reynolds number effects.

  1. Hot-wire calibration in subsonic/transonic flow regimes

    NASA Technical Reports Server (NTRS)

    Nagabushana, K. A.; Ash, Robert L.

    1995-01-01

    A different approach for calibrating hot-wires, which simplifies the calibration procedure and reduces the tunnel run-time by an order of magnitude was sought. In general, it is accepted that the directly measurable quantities in any flow are velocity, density, and total temperature. Very few facilities have the capability of varying the total temperature over an adequate range. However, if the overheat temperature parameter, a(sub w), is used to calibrate the hot-wire then the directly measurable quantity, voltage, will be a function of the flow variables and the overheat parameter i.e., E = f(u,p,a(sub w), T(sub w)) where a(sub w) will contain the needed total temperature information. In this report, various methods of evaluating sensitivities with different dependent and independent variables to calibrate a 3-Wire hot-wire probe using a constant temperature anemometer (CTA) in subsonic/transonic flow regimes is presented. The advantage of using a(sub w) as the independent variable instead of total temperature, t(sub o), or overheat temperature parameter, tau, is that while running a calibration test it is not necessary to know the recovery factor, the coefficients in a wire resistance to temperature relationship for a given probe. It was deduced that the method employing the relationship E = f (u,p,a(sub w)) should result in the most accurate calibration of hot wire probes. Any other method would require additional measurements. Also this method will allow calibration and determination of accurate temperature fluctuation information even in atmospheric wind tunnels where there is no ability to obtain any temperature sensitivity information at present. This technique greatly simplifies the calibration process for hot-wires, provides the required calibration information needed in obtaining temperature fluctuations, and reduces both the tunnel run-time and the test matrix required to calibrate hotwires. Some of the results using the above techniques are presented

  2. Trailing Edge Blowing on a Two-Dimensional Six-Percent Thick Elliptical Circulation Control Airfoil Up to Transonic Conditions

    NASA Technical Reports Server (NTRS)

    Alexander, Michael G.; Anders, Scott G.; Johnson, Stuart K.; Florance, Jennifer P.; Keller, Donald F.

    2005-01-01

    A wind tunnel test was conducted in the NASA Langley Transonic Dynamics Tunnel (TDT) on a six percent thick slightly cambered elliptical circulation control airfoil with both upper and lower surface blowing capability. Parametric evaluations of jet slot heights and Coanda surface shapes were conducted at momentum coefficients (Cm) from 0.0 to 0.12. Test data were acquired at Mach numbers of 0.3, 0.5, 0.7, 0.8, and 0.84 at Reynolds numbers per foot of 2.43 x 105 to 1.05 x 106. For a transonic condition, (Mach = 0.8 at alpha = 3 degrees), it was generally found the smaller slot and larger Coanda surface combination was overall more effective than other slot/Coanda surface combinations. Lower surface blowing was not as effective as the upper surface blowing over the same range of momentum coefficients. No appreciable Coanda surface, slot height, or slot blowing position preference was indicated transonically with the dual slot blowing.

  3. Thrust Removal Scheme for the FAST-MAC Circulation Control Model Tested in the National Transonic Facility

    NASA Technical Reports Server (NTRS)

    Chan, David T.; Milholen, William E., II; Jones, Gregory S.; Goodliff, Scott L.

    2014-01-01

    A second wind tunnel test of the FAST-MAC circulation control semi-span model was recently completed in the National Transonic Facility at the NASA Langley Research Center. The model allowed independent control of four circulation control plenums producing a high momentum jet from a blowing slot near the wing trailing edge that was directed over a 15% chord simple-hinged flap. The model was configured for transonic testing of the cruise configuration with 0deg flap deflection to determine the potential for drag reduction with the circulation control blowing. Encouraging results from analysis of wing surface pressures suggested that the circulation control blowing was effective in reducing the transonic drag on the configuration, however this could not be quantified until the thrust generated by the blowing slot was correctly removed from the force and moment balance data. This paper will present the thrust removal methodology used for the FAST-MAC circulation control model and describe the experimental measurements and techniques used to develop the methodology. A discussion on the impact to the force and moment data as a result of removing the thrust from the blowing slot will also be presented for the cruise configuration, where at some Mach and Reynolds number conditions, the thrust-removed corrected data showed that a drag reduction was realized as a consequence of the blowing.

  4. Aerodynamic Measurements on a Large Splitter Plate for the NASA Langley Transonic Dynamics Tunnel

    NASA Technical Reports Server (NTRS)

    Schuster, David M.

    2001-01-01

    Tests conducted in the NASA Langley Research Center Transonic Dynamics Tunnel (TDT) assess the aerodynamic characteristics of a splitter plate used to test some semispan models in this facility. Aerodynamic data are analyzed to determine the effect of the splitter plate on the operating characteristics of the TDT, as well as to define the range of conditions over which the plate can be reasonably used to obtain aerodynamic data. Static pressures measurements on the splitter plate surface and the equipment fairing between the wind tunnel wall and the splitter plate are evaluated to determine the flow quality around the apparatus over a range of operating conditions. Boundary layer rake data acquired near the plate surface define the viscous characteristics of the flow over the plate. Data were acquired over a range of subsonic, transonic and supersonic conditions at dynamic pressures typical for models tested on this apparatus. Data from this investigation should be used as a guide for the design of TDT models and tests using the splitter plate, as well as to guide future splitter plate design for this facility.

  5. Numerical investigation of the near wake of generic space launcher systems at transonic and supersonic flows

    NASA Astrophysics Data System (ADS)

    Statnikov, V.; Glatzer, C.; Meiß, J.-H.; Meinke, M.; Schröder, W.

    2013-06-01

    Numerical simulations of the near wake of generic rocket configurations are performed at transonic and supersonic freestream conditions to improve the understanding of the highly intricate near wake structures. The Reynolds number in both flow regimes is 106 based on the main body diameter, i. e., specific freestream conditions of ESA's Ariane launcher trajectory. The geometry matches models used in experiments in the framework of the German Transregional Collaborative Research Center TRR40. Both axisymmetric wind tunnel models possess cylindrical sting supports, representing a nozzle to allow investigations of a less disturbed wake flow. A zonal approach consisting of a Reynolds averaged Navier-Stokes (RANS) and a large-eddy simulation (LES) is applied. It is shown that the highly unsteady transonic wake flow at Ma∞ = 0.7 is characterized by the expanding separated shear layer, while the Mach 6.0 wake is defined by a shock, expansion waves, and a recompression region. In both cases, an instantaneous view on the base characteristics reveals complex azimuthal flow structures even for axisymmetric geometries. The flow regimes are discussed by comparing the aerodynamic characteristics, such as the size of the recirculation region and the turbulent kinetic energy.

  6. Performance Test of Laser Velocimeter System for the Langley 16-foot Transonic Tunnel

    NASA Technical Reports Server (NTRS)

    Meyers, J. F.; Hunter, W. W., Jr.; Reubush, D. E.; Nichols, C. E., Jr.; Hepner, T. E.; Lee, J. W.

    1985-01-01

    An investigation in the Langley 16-Foot Transonic Tunnel has been conducted in which a laser velocimeter was used to measure free-stream velocities from Mach 0.1 to 1.0 and the flow velocities along the stagnating streamline of a hemisphere-cylinder model at Mach 0.8 and 1.0. The flow velocity was also measured at Mach 1.0 along the line 0.533 model diameters below the model. These tests determined the performance characteristics of the dedicated two-component laser velocimeter at flow velocities up to Mach 1.0 and the effects of the wind tunnel environment on the particle-generating system and on the resulting size of the generated particles. To determine these characteristics, the measured particle velocities along the stagnating streamline at the two Mach numbers were compared with the theoretically predicted gas and particle velocities calculated using a transonic potential flow method. Through this comparison the mean detectable particle size (2.1 micron) along with the standard deviation of the detectable particles (0.76 micron) was determined; thus the performance characteristics of the laser velocimeter were established.

  7. 5. VIEW LOOKING NORTH AT 8FOOT TRANSONIC PRESSURE TUNNEL PLENUM ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    5. VIEW LOOKING NORTH AT 8-FOOT TRANSONIC PRESSURE TUNNEL PLENUM FLOOR AREA. NOTE SCHLIEREN OPTICAL SYSTEM ON STRUCTURE AT RIGHT CENTER. - NASA Langley Research Center, 8-Foot Transonic Pressure Tunnel, 640 Thornell Avenue, Hampton, Hampton, VA

  8. 1. VIEW LOOKING SOUTHEAST AT EXTERIOR OF 8FOOT TRANSONIC PRESSURE ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    1. VIEW LOOKING SOUTHEAST AT EXTERIOR OF 8-FOOT TRANSONIC PRESSURE TUNNEL. NOTE EXPANSION RINGS. - NASA Langley Research Center, 8-Foot Transonic Pressure Tunnel, 640 Thornell Avenue, Hampton, Hampton, VA

  9. 2. VIEW LOOKING EASTNORTHEAST AT EXTERIOR OF 8FOOT TRANSONIC PRESSURE ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    2. VIEW LOOKING EAST-NORTHEAST AT EXTERIOR OF 8-FOOT TRANSONIC PRESSURE TUNNEL (BUILDING 640). NOTE NACA LOGO OVER DOORWAY. - NASA Langley Research Center, 8-Foot Transonic Pressure Tunnel, 640 Thornell Avenue, Hampton, Hampton, VA

  10. A survey of cryogenic wind tunnels

    NASA Technical Reports Server (NTRS)

    Kilgore, R. A.; Dress, D. A.

    1985-01-01

    Following the development of the cryogenic wind tunnel at the NASA Langley Research Center in 1972, a large number of cryogenic wind-tunnel projects have been undertaken at various research establishments around the world. The purpose of this paper is to describe some of the more significant of these projects. Described in this paper are cryogenic wind-tunnel projects in China (CARDC), England (College of Aeronautics at Cranfield, RAE-Bedford, and University of Southampton), 'Europe' (Pilot European Transonic Windtunnel at NAL-Amsterdam, and the European Transonic Windtunnel proposed for DFVLR-Koeln), France (ONERA-CERT), Germany (DFVLR-Koeln, and DFVLR-Goettingen), Japan (NAL, University of Tsukuba, and National Defense Academy), Sweden (Rollab), and the United States (Douglas Aircraft Co., University of Illinois at Urbana-Champaign, and NASA-Langley).

  11. Aerodynamic load distributions at transonic speeds for a close-coupled wing-canard configuration: Tabulated pressure data

    NASA Technical Reports Server (NTRS)

    Washburn, K. E.; Gloss, B. B.

    1978-01-01

    Wind tunnel studies are reported on both the canard and wing surfaces of a model that is geometrically identical to one used in several force and moment tests to provide insight into the various aerodynamic interference effects. In addition to detailed pressures measurements, the pressures were integrated to illustrate the effects of Mach number, canard location, and canard-wing interference on various aerodynamic parameters. Transonic pressure tunnel Mach numbers ranged from 0.70 to 1.20 for data taken from 0 deg to approximately 16 deg angle-of-attack at 0 deg sideslip.

  12. Test-to-Test Repeatability of Results From a Subsonic Wing-Body Configuration in the National Transonic Facility

    NASA Technical Reports Server (NTRS)

    Mineck, Raymond E.; Pendergraft, Odis C., Jr.

    2000-01-01

    Results from three wind tunnel tests in the National Transonic Facility of a model of an advanced-technology, subsonic-transport wing-body configuration have been analyzed to assess the test-to-test repeatability of several aerodynamic parameters. The scatter, as measured by the prediction interval, in the longitudinal force and moment coefficients increases as the Mach number increases. Residual errors with and without the ESP tubes installed suggest a bias leading to lower drag with the tubes installed. Residual errors as well as average values of the longitudinal force and moment coefficients show that there are small bias errors between the different tests.

  13. Adaptation of two-dimensional transonic analysis code, TSFOIL, for use in modeling adaptive wall test sections

    NASA Technical Reports Server (NTRS)

    Webb, J.

    1984-01-01

    The effects of the elastic deformation of the wind tunnel wall held to a streamline-like shape were simulated numerically. The wall itself is simulated by a finite element model and is allowed to deform under the pressure loading developing in the wind tunnel with an airfoil model present. A modified version of the transonic analysis program, TSFOIL, is then used to determine the resulting flow field with the effects of the deformed wall included, the shape of the wall and the flow field. Once a result from a particular operating condition was obtained, the pressure distribution on the airfoil in the wind tunnel model is compared with solutions generated by TSFOIL in the free air mode.

  14. Unsteady transonic flow calculations for realistic aircraft configurations

    NASA Technical Reports Server (NTRS)

    Batina, John T.; Seidel, David A.; Bland, Samuel R.; Bennett, Robert M.

    1987-01-01

    A transonic unsteady aerodynamic and aeroelasticity code has been developed for application to realistic aircraft configurations. The new code is called CAP-TSD which is an acronym for Computational Aeroelasticity Program - Transonic Small Disturbance. The CAP-TSD code uses a time-accurate approximate factorization (AF) algorithm for solution of the unsteady transonic small-disturbance equation. The AF algorithm is very efficient for solution of steady and unsteady transonic flow problems. It can provide accurate solutions in only several hundred time steps yielding a significant computational cost savings when compared to alternative methods. The new code can treat complete aircraft geometries with multiple lifting surfaces and bodies including canard, wing, tail, control surfaces, launchers, pylons, fuselage, stores, and nacelles. Applications are presented for a series of five configurations of increasing complexity to demonstrate the wide range of geometrical applicability of CAP-TSD. These results are in good agreement with available experimental steady and unsteady pressure data. Calculations for the General Dynamics one-ninth scale F-16C aircraft model are presented to demonstrate application to a realistic configuration. Unsteady results for the entire F-16C aircraft undergoing a rigid pitching motion illustrated the capability required to perform transonic unsteady aerodynamic and aeroelastic analyses for such configurations.

  15. Transonic rotor tip design using numerical optimization

    NASA Technical Reports Server (NTRS)

    Tauber, Michael E.; Langhi, Ronald G.

    1985-01-01

    The aerodynamic design procedure for a new blade tip suitable for operation at transonic speeds is illustrated. For the first time, 3 dimensional numerical optimization was applied to rotor tip design, using the recent derivative of the ROT22 code, program R22OPT. Program R22OPT utilized an efficient quasi-Newton optimization algorithm. Multiple design objectives were specified. The delocalization of the shock wave was to be eliminated in forward flight for an advance ratio of 0.41 and a tip Mach number of 0.92 at psi = 90 deg. Simultaneously, it was sought to reduce torque requirements while maintaining effective restoring pitching moments. Only the outer 10 percent of the blade span was modified; the blade area was not to be reduced by more than 3 percent. The goal was to combine the advantages of both sweptback and sweptforward blade tips. A planform that featured inboard sweepback was combined with a sweptforward tip and a taper ratio of 0.5. Initially, the ROT22 code was used to find by trial and error a planform geometry which met the design goals. This configuration had an inboard section with a leading edge sweep of 20 deg and a tip section swept forward at 25 deg; in addition, the airfoils were modified.

  16. Accelerated Transonic Flow past a curvature discontinuity

    NASA Astrophysics Data System (ADS)

    de Cointet, Thomas; Ruban, Anatoly

    2014-11-01

    The aim of this talk is to investigate High Reynolds number Transonic flow past a discontinuity in body curvature. Starting with the inviscid flow outside the boundary layer, our analysis will focus on the flow in a vicinity of the point of discontinuity, where a solution of the Euler equations will be sought in self-similar form. This reduces the Euler equations to an ordinary differential equation. The analysis of this equation shows that the pressure gradient on the airfoil surface develops a strong singularity, which is proportional to (x0 - x) - 1 / 3 as the discontinuity point x0 is approached. We then study the response of the boundary layer to this extremely favourable pressure gradient. We show that the boundary layer splits into two parts, the main body of the boundary layer that becomes inviscid on approach to the singularity, and a thin viscous sublayer situated near the wall. The analysis of the behaviour of the solution in the viscous sublayer shows that Prandtl's hierarchical concept breaks down in a small region surrounding the singular point, where the viscous-inviscid interaction model should be used. In the final part of this talk we present a full formulation of the viscous-inviscid interaction problem and discuss numerical results.

  17. Subsonic-transonic stall flutter study

    NASA Technical Reports Server (NTRS)

    Stardter, H.

    1979-01-01

    The objective of the Subsonic/Transonic Stall Flutter Program was to obtain detailed measurements of both the steady and unsteady flow field surrounding a rotor and the mechanical state of the rotor while it was operating in both steady and flutter modes to provide a basis for future analysis and for development of theories describing the flutter phenomenon. The program revealed that while all blades flutter at the same frequency, they do not flutter at the same amplitude, and their interblade phase angles are not equal. Such a pattern represents the superposition of a number of rotating nodal diameter patterns, each characterized by a different amplitude and different phase indexing, but each rotating at a speed that results in the same flutter frequency as seen in the rotor system. Review of the steady pressure contours indicated that flutter may alter the blade passage pressure distribution. The unsteady pressure amplitude contour maps reveal regions of high unsteady pressure amplitudes near the leading edge, lower amplitudes near the trailing.

  18. Flow Control in a Transonic Diffuser

    NASA Astrophysics Data System (ADS)

    Gartner, Jeremy; Amitay, Michael

    2014-11-01

    In some airplanes such as fighter jets and UAV, short inlet ducts replace the more conventional ducts due to their shorter length. However, these ducts are associated with low length-to-diameter ratio and low aspect ratio and, thus, experience massive separation and the presence of secondary flow structures. These flow phenomena are undesirable as they lead to pressure losses and distortion at the Aerodynamic Interface Plane (AIP), where the engine face is located. It causes the engine to perform with a lower efficiency as it would with a straight duct diffuser. Different flow control techniques were studied on the short inlet duct, with the goal to reattach the flow and minimize the distortions at the AIP. Due to the complex interaction between the separation and the secondary flow structures, the necessity to understand the flow mechanisms, and how to control them at a more fundamental level, a new transonic diffuser with an upper ramp and a straight floor was designed and built. The objective of this project is to explore the effectiveness of different flow control techniques in a high subsonic (up to Mach 0.8) diffuser, so that the quasi two-dimensional separation and the formation of secondary flow structure can be isolated using a canonical flow field. Supported by Northrop Grumman.

  19. Test results at transonic speeds on a contoured over-the-wing propfan model

    NASA Technical Reports Server (NTRS)

    Levin, Alan D.; Smeltzer, Donald B.; Smith, Ronald C.

    1986-01-01

    A semispan wing/body model with a powered highly loaded propeller has been tested to provide data on the propulsion installation drag of advanced propfan-powered aircraft. The model had a supercritical wing with a contoured over-the-wing nacelle. It was tested in the Ames Research Center's (ARC) 14-foot Transonic Wind Tunnel at a total pressure of 1 atm. The test was conducted at angles of attack from -0.5 to 4 deg at Mach numbers ranging from 0.6 to 0.8. The test objectives were to determine propeller performance, exhaust jet effects, propeller slipstream interference drag, and total powerplant installation drag. Test results indicated a total powerplant installation drag of 82 counts (0.0082) at a Mach number of 0.8 and a lift coefficient of 0.5, which is approximately 29 percent of a typical airplane cruise drag.

  20. Influences of Models on the Unsteady Pressure Characteristics of the NASA National Transonic Facility

    NASA Technical Reports Server (NTRS)

    Jones, Gregory; Balakrishna, Sundareswara; DeMoss, Joshua; Goodliff, Scott; Bailey, Matthew

    2015-01-01

    Pressure fluctuations have been measured over the course of several tests in the National Transonic Facility to study unsteady phenomenon both with and without the influence of a model. Broadband spectral analysis will be used to characterize the length scales of the tunnel. Special attention will be given to the large-scale, low frequency data that influences the Mach number and force and moment variability. This paper will also discuss the significance of the vorticity and sound fields that can be related to the Common Research Model and will also highlight the comparisons to an empty tunnel configuration. The effectiveness of vortex generators placed at the interface of the test section and wind tunnel diffuser showed promise in reducing the empty tunnel unsteadiness, however, the vortex generators were ineffective in the presence of a model.

  1. Description of recent changes in the Langley 6- by 28-inch transonic tunnel

    NASA Technical Reports Server (NTRS)

    Sewall, W. G.

    1981-01-01

    Calibrations were obtained in the Langley 6 by 28-inch transonic tunnel with newly installed controllable reentry flaps and test section floor and ceiling. Using available theory, the top and bottom slotted walls were redesigned for minimum wind tunnel interference errors of blockage and stream-line curvature. To minimize Mach number gradients along the tunnel axis downstream of the model, controllable flaps were installed to regulate the flow reentering the test section through the slotted walls. The flap setting is independent of stagnation pressure and varies only with Mach number. The freestream Mach number is determined from the pressrue measured at a station 66.04 cm upstream of the model station. The model has no significant influence on the vertical Mach number distribution at this station. This method of Mach number determination appears to be more accurate than one using the plenum pressure.

  2. Form drag, skin friction, and vortex shedding frequencies for subsonic and transonic crossflows on circular cylinder

    NASA Technical Reports Server (NTRS)

    Murthy, V. S.; Rose, W. C.

    1977-01-01

    A series of wind-tunnel tests covering a range of Mach numbers and Reynolds numbers in subsonic and transonic flows was conducted on a circular cylinder placed normal to the flow. Form drag coefficients were determined from surface-pressure measurements and displayed as a function of Mach number to show the drag rise phenomenon. Buried wire gages arranged on the model surface were used to measure skin-friction distributions and vortex-shedding frequencies at different flow conditions. It was found that detectable periodic shedding ceases above M = 0.9. The measured skin-friction distributions indicate the positions of mean separation points clearly; these values are documented for the different flow conditions.

  3. Past, Present, and Future Capabilities of the Transonic Dynamics Tunnel from an Aeroelasticity Perspective

    NASA Technical Reports Server (NTRS)

    Cole, Stanley R.; Garcia, Jerry L.

    2000-01-01

    The NASA Langley Transonic Dynamics Tunnel (TDT) has provided a unique capability for aeroelastic testing for forty years. The facility has a rich history of significant contributions to the design of many United States commercial transports, military aircraft, launch vehicles, and spacecraft. The facility has many features that contribute to its uniqueness for aeroelasticity testing, perhaps the most important feature being the use of a heavy gas test medium to achieve higher test densities. Higher test medium densities substantially improve model-building requirements and therefore simplify the fabrication process for building aeroelastically scaled wind tunnel models. Aeroelastic scaling for the heavy gas results in lower model structural frequencies. Lower model frequencies tend to a make aeroelastic testing safer. This paper will describe major developments in the testing capabilities at the TDT throughout its history, the current status of the facility, and planned additions and improvements to its capabilities in the near future.

  4. Test Activities in the Langley Transonic Dynamics Tunnel and a Summary of Recent Facility Improvements

    NASA Technical Reports Server (NTRS)

    Cole, Stanley R.; Johnson, R. Keith; Piatak, David J.; Florance, Jennifer P.; Rivera, Jose A., Jr.

    2003-01-01

    The Langley Transonic Dynamics Tunnel (TDT) has provided a unique capability for aeroelastic testing for over forty years. The facility has a rich history of significant contributions to the design of many United States commercial transports, military aircraft, launch vehicles, and spacecraft. The facility has many features that contribute to its uniqueness for aeroelasticity testing, perhaps the most important feature being the use of a heavy gas test medium to achieve higher test densities compared to testing in air. Higher test medium densities substantially improve model-building requirements and therefore simplify the fabrication process for building aeroelastically scaled wind tunnel models. This paper describes TDT capabilities that make it particularly suited for aeroelasticity testing. The paper also discusses the nature of recent test activities in the TDT, including summaries of several specific tests. Finally, the paper documents recent facility improvement projects and the continuous statistical quality assessment effort for the TDT.

  5. Real-time data acquisition system for the NASA Langley transonic dynamics tunnel

    NASA Technical Reports Server (NTRS)

    Cole, P. H.

    1979-01-01

    The hardware configuration of the Transonic Dynamics Wind Tunnel Data Acquisition System (DAS) which consists of an analog front end that can process up to 260 channels of data is presented. The DAS also has a multi-channel analog-to-digital subsystem that can process up to 50,000 samples of data per sec, and a digital computer with standard and nonstandard devices, with graphics capability. The software configuration of the DAS and complex hardware/software interfaces are described, which can provide automatic amplifier gain and offset adjustment for each data channel. Finally, a summary of specific DAS applications is given including the real-time processing of dynamic deflection data, unsteady pressure measurements, and flutter and buffet data.

  6. Wall Boundary Layer Measurements for the NASA Langley Transonic Dynamics Tunnel

    NASA Technical Reports Server (NTRS)

    Wieseman, Carol D.; Bennett, Robert M.

    2007-01-01

    Measurements of the boundary layer parameters in the NASA Langley Transonic Dynamics tunnel were conducted during extensive calibration activities following the facility conversion from a Freon-12 heavy-gas test medium to R-134a. Boundary-layer rakes were mounted on the wind-tunnel walls, ceiling, and floor. Measurements were made over the range of tunnel operation envelope in both heavy gas and air and without a model in the test section at three tunnel stations. Configuration variables included open and closed east sidewall wall slots, for air and R134a test media, reentry flap settings, and stagnation pressures over the full range of tunnel operation. The boundary layer thickness varied considerably for the six rakes. The thickness for the east wall was considerably larger that the other rakes and was also larger than previously reported. There generally was some reduction in thickness at supersonic Mach numbers, but the effect of stagnation pressure, and test medium were not extensive.

  7. Design and construction of models for the National Transonic Facility, part 1

    NASA Technical Reports Server (NTRS)

    Young, C. P., Jr.

    1985-01-01

    The design and construction of models for the National Transonic Facility (NTF) has resulted in significant technology developments in many areas. This lecture covers the development of design criteria and major research and development work that has contributed to the successful design and fabrication models for testing at full scale Reynolds number the NTF. Emphasis is placed on the materials aspect of the design and fabrication proces, including metallic materials, mechanical properties characterization, new steel alloy development, fracture toughness enhancement, and identification of fillers and solders suitable for use in cryogenic models. Quantitative data are provided which will be of value to the potential user of NTF or for application to the design and fabrication of model systems for other cryogenic wind tunnels.

  8. Application of FLEET Velocimetry in the NASA Langley 0.3-meter Transonic Cryogenic Tunnel

    NASA Technical Reports Server (NTRS)

    Burns, Ross A.; Danehy, Paul M.; Halls, Benjamin R.; Jiang, Naibo

    2015-01-01

    Femtosecond laser electronic excitation and tagging (FLEET) velocimetry is demonstrated in a large-scale transonic cryogenic wind tunnel. Test conditions include total pressures, total temperatures, and Mach numbers ranging from 15 to 58 psia, 200 to 295 K, and 0.2 to 0.75, respectively. Freestream velocity measurements exhibit accuracies within 1 percent and precisions better than 1 m/s. The measured velocities adhere closely to isentropic flow theory over the domain of temperatures and pressures that were tested. Additional velocity measurements are made within the tunnel boundary layer; virtual trajectories traced out by the FLEET signal are indicative of the characteristic turbulent behavior in this region of the flow, where the unsteadiness increases demonstrably as the wall is approached. Mean velocities taken within the boundary layer are in agreement with theoretical velocity profiles, though the fluctuating velocities exhibit a greater deviation from theoretical predictions.

  9. A wake traverse technique for use in a 2 dimensional transonic flexible walled test section

    NASA Technical Reports Server (NTRS)

    Wolf, S. W. D.

    1982-01-01

    Reported two dimensional validation data from the Transonic Self-Streamlining Wind Tunnel (TSWT) concerns model lift. The models tested provided data on their pressure distributions. This information was numerically integrated over the model surface to determine lift, pressure drag and pitching moment. However, the pressure drag is only a small component of the total drag at nominal angles of attack and cannot be used to assess the quality of flow simulation. An intrusive technique for obtaining information on the total drag of a model in TSWT is described. The technique adopted is the wake traverse method. The associated tunnel hardware and control and data reduction software are outlined and some experimental results are presented for discussion.

  10. A PC-based simulation of the National Transonic Facitity's safety microprocessor

    NASA Technical Reports Server (NTRS)

    Thibodeaux, J. J.; Kilgore, W. A.; Balakrishna, S.

    1993-01-01

    A brief study was undertaken to demonstrate the feasibility of using a state-of-the-art off-the-shelf high speed personal computer for simulating a microprocessor presently used for wind tunnel safety purposes at Langley Research Center's National Transonic Facility (NTF). Currently, there is no active display of tunnel alarm/alert safety information provided to the tunnel operators, but rather such information is periodically recorded on a process monitoring computer printout. This does not provide on-line situational information nor permit rapid identification of safety operational violations which are able to halt tunnel operations. It was therefore decided to simulate the existing algorithms and briefly evaluate a real-time display which could provide both position and trouble shooting information.

  11. Aerodynamic effect of coolant ejection in the rear part of transonic rotor blades

    NASA Astrophysics Data System (ADS)

    Kost, F. H.; Holmes, A. T.

    1985-09-01

    An investigation of transonic turbine blades designed by Rolls-Royce/Bristol concerning the aerodynamic penalties of coolant flow for two alternative cooling configurations is discussed. Rolls-Royce designed a blade with a thick trailing edge where the coolant is ejected through slots in the trailing edge and a second blade with a thin trailing edge where coolant is ejected through a row of holes on the pressure side and a row of holes on the suction side. Tests were performed in a plane cascade wind tunnel. The results indicate the sensitivity of the blade performance to cooling configuration and coolant flow rate. By combining measured data from blade surface and wake traverses it was possible to separate the various loss mechanisms. Therefore, the separate losses due to the momentum of the coolant, change of base pressure, and change of blade friction could be determined quantitatively as a function of coolant flow rate.

  12. Electrical noise reduction techniques contributing to improved data quality at the National Transonic Facility

    NASA Technical Reports Server (NTRS)

    Mcphee, J. R.

    1988-01-01

    In initial use, the high-speed digital data acquisition systems at Langley Research Center's National Transonic Facility produced data containing unacceptably high noise levels. Described is a process whereby the contributing noise sources were identified and eliminated. The effects of 60 Hz power, system grounding, EMI/RFI, and other problems are discussed and the corrective action taken is outlined. The overall effort resulted in an improvement of greater than 5:1 in system performance. Although the report describes a system specifically used for wind tunnel data acquisition, the corrective techniques employed are generally applicable to large scale high-speed data systems where signal resolution in the low microvolts range is important.

  13. Transonic Navier-Stokes solutions of three-dimensional afterbody flows

    NASA Technical Reports Server (NTRS)

    Compton, William B., III; Thomas, James L.; Abeyounis, William K.; Mason, Mary L.

    1989-01-01

    The performance of a three-dimensional Navier-Stokes solution technique in predicting the transonic flow past a nonaxisymmetric nozzle was investigated. The investigation was conducted at free-stream Mach numbers ranging from 0.60 to 0.94 and an angle of attack of 0 degrees. The numerical solution procedure employs the three-dimensional, unsteady, Reynolds-averaged Navier-Stokes equations written in strong conservation form, a thin layer assumption, and the Baldwin-Lomax turbulence model. The equations are solved by using the finite-volume principle in conjunction with an approximately factored upwind-biased numerical algorithm. In the numerical procedure, the jet exhaust is represented by a solid sting. Wind-tunnel data with the jet exhaust simulated by high pressure air were also obtained to compare with the numerical calculations.

  14. Computational Analysis of the Transonic Dynamics Tunnel Using FUN3D

    NASA Technical Reports Server (NTRS)

    Chwalowski, Pawel; Quon, Eliot; Brynildsen, Scott E.

    2016-01-01

    This paper presents results from an exploratory two-year effort of applying Computational Fluid Dynamics (CFD) to analyze the empty-tunnel flow in the NASA Langley Research Center Transonic Dynamics Tunnel (TDT). The TDT is a continuous-flow, closed circuit, 16- x 16-foot slotted-test-section wind tunnel, with capabilities to use air or heavy gas as a working fluid. In this study, experimental data acquired in the empty tunnel using the R-134a test medium was used to calibrate the computational data. The experimental calibration data includes wall pressures, boundary-layer profiles, and the tunnel centerline Mach number profiles. Subsonic and supersonic flow regimes were considered, focusing on Mach 0.5, 0.7 and Mach 1.1 in the TDT test section. This study discusses the computational domain, boundary conditions, and initial conditions selected and the resulting steady-state analyses using NASA's FUN3D CFD software.

  15. Aeroelasticity matters: Some reflections on two decades of testing in the NASA Langley transonic dynamics tunnel

    NASA Technical Reports Server (NTRS)

    Reed, W. H., III

    1981-01-01

    Testing of wind-tunnel aeroelastic models is a well established, widely used means of studying flutter trends, validating theory and investigating flutter margins of safety of new vehicle designs. The Langley Transonic Dynamics Tunnel was designed specifically for work on dynamics and aeroelastic problems of aircraft and space vehicles. A cross section of aeroelastic research and testing in the facility since it became operational more than two decades ago is presented. Examples selected from a large store of experience illustrate the nature and purpose of some major areas of work performed in the tunnel. These areas include: specialized experimental techniques; development testing of new aircraft and launch vehicle designs; evaluation of proposed "fixes" to solve aeroelastic problems uncovered during development testing; study of unexpected aeroelastic phenomena (i.e., "surprises"); control of aeroelastic effects by active and passive means; and, finally, fundamental research involving measurement of unsteady pressures on oscillating wings and control surface.

  16. Development of a Free-to-Roll Transonic Test Capability (Invited)

    NASA Astrophysics Data System (ADS)

    Capone, F. J.; Owens, D. B.; Hall, R. M.

    2003-01-01

    As part of the NASA/Navy Abrupt Wing Stall Program, a relatively low-cost, rapid-access wind-tunnel free-to-roll rig was developed. This rig combines the use of conventional models and test apparatuses to evaluate both transonic performance and wing-drop/rock tendencies in a single tunnel entry. A description of the test hardware as well as a description of the experimental procedures is given. The free-to-roll test rig has been used successfully to assess the static and dynamic characteristics of three different configurations--two configurations that exhibit uncommanded lateral motions, (pre-production F/A-18E and AV-8B), and one that did not (F/A-18C).

  17. Method to predict external store carriage characteristics at transonic speeds

    NASA Technical Reports Server (NTRS)

    Rosen, Bruce S.

    1988-01-01

    Development of a computational method for prediction of external store carriage characteristics at transonic speeds is described. The geometric flexibility required for treatment of pylon-mounted stores is achieved by computing finite difference solutions on a five-level embedded grid arrangement. A completely automated grid generation procedure facilitates applications. Store modeling capability consists of bodies of revolution with multiple fore and aft fins. A body-conforming grid improves the accuracy of the computed store body flow field. A nonlinear relaxation scheme developed specifically for modified transonic small disturbance flow equations enhances the method's numerical stability and accuracy. As a result, treatment of lower aspect ratio, more highly swept and tapered wings is possible. A limited supersonic freestream capability is also provided. Pressure, load distribution, and force/moment correlations show good agreement with experimental data for several test cases. A detailed computer program description for the Transonic Store Carriage Loads Prediction (TSCLP) Code is included.

  18. Finite elements and finite differences for transonic flow calculations

    NASA Technical Reports Server (NTRS)

    Hafez, M. M.; Murman, E. M.; Wellford, L. C.

    1978-01-01

    The paper reviews the chief finite difference and finite element techniques used for numerical solution of nonlinear mixed elliptic-hyperbolic equations governing transonic flow. The forms of the governing equations for unsteady two-dimensional transonic flow considered are the Euler equation, the full potential equation in both conservative and nonconservative form, the transonic small-disturbance equation in both conservative and nonconservative form, and the hodograph equations for the small-disturbance case and the full-potential case. Finite difference methods considered include time-dependent methods, relaxation methods, semidirect methods, and hybrid methods. Finite element methods include finite element Lax-Wendroff schemes, implicit Galerkin method, mixed variational principles, dual iterative procedures, optimal control methods and least squares.

  19. Overview of Transonic to Hypersonic Stage Separation Tool Development for Multi-Stage-to-Orbit Concepts

    NASA Technical Reports Server (NTRS)

    Murphy, Kelly J.; Bunning, Pieter G.; Pamadi, Bandu N.; Scallion, William I.; Jones, Kenneth M.

    2004-01-01

    An overview of research efforts at NASA in support of the stage separation and ascent aerothermodynamics research program is presented. The objective of this work is to develop a synergistic suite of experimental, computational, and engineering tools and methods to apply to vehicle separation across the transonic to hypersonic speed regimes. Proximity testing of a generic bimese wing-body configuration is on-going in the transonic (Mach numbers 0.6, 1.05, and 1.1), supersonic (Mach numbers 2.3, 3.0, and 4.5) and hypersonic (Mach numbers 6 and 10) speed regimes in four wind tunnel facilities at the NASA Langley Research Center. An overset grid, Navier-Stokes flow solver has been enhanced and demonstrated on a matrix of proximity cases and on a dynamic separation simulation of the bimese configuration. Steady-state predictions with this solver were in excellent agreement with wind tunnel data at Mach 3 as were predictions via a Cartesian-grid Euler solver. Experimental and computational data have been used to evaluate multi-body enhancements to the widely-used Aerodynamic Preliminary Analysis System, an engineering methodology, and to develop a new software package, SepSim, for the simulation and visualization of vehicle motions in a stage separation scenario. Web-based software will be used for archiving information generated from this research program into a database accessible to the user community. Thus, a framework has been established to study stage separation problems using coordinated experimental, computational, and engineering tools.

  20. Recent Enhancements to the National Transonic Facility (Mixed Mode Operations)

    NASA Technical Reports Server (NTRS)

    Kilgore, W. Allen; Chan, David; Balakrishna, S.; Wahls, Richard A.

    2006-01-01

    The U.S. National Transonic Facility continues to make enhancements to provide quality data in a safe, efficient and cost effective method for aerodynamic ground testing. Recent enhancements discussed in this paper include the development of a Mixed-mode of operations that combine Air-mode operations with Nitrogen-mode operations. This implementation and operational results of this new Mixed-mode expands the ambient temperature transonic region of testing beyond the Air-mode limitations at a significantly reduced cost over Nitrogen Mode operation.

  1. On transonic flow models for optimized design and experiment

    NASA Astrophysics Data System (ADS)

    Stodůlka, Jiří; Sobieczky, Helmut

    2014-03-01

    In the paper the near sonic flow theory for flows with small perturbations to sonic parallel flow is developed. This theory stands on the basis of potential flow of a compressible fluid and enables to receive an exact solution of the flow parameters past transonic cusped airfoils and their geometrical description. Generated airfoil shapes are tested using CFD ANSYS Fluent code to validate the results. Obtained numerical results from all-round commercial code show good accordance with the theory and confirm their value for future work in transonic design.

  2. Slender body theory and Space Shuttle transonic aerodynamics

    NASA Technical Reports Server (NTRS)

    Malmuth, N. D.; Wu, C. C.; Cole, J. D.

    1985-01-01

    A computational implementation of transonic slender body theory and the equivalence rule has been utilized to study transonic flow field around the Space Shuttle Orbiter. The far field is described by a nonlinear axisymmetric Karman-Guderley model and the near field by a cross flow Laplace equation boundary value problem. The latter is treated using a source panel method. Preliminary comparisons with experiments give encouraging indications that the model can be useful for quick turnaround estimates. Areas of refinement to obtain more accurate predictions are discussed.

  3. Numerical studies of unsteady transonic flow over oscillating airfoil

    NASA Technical Reports Server (NTRS)

    Chyu, W. J.; Davis, S. S.

    1985-01-01

    A finite difference solution to the Navier-Stokes equations combined with a time varying grid generation technique was used to compute unsteady transonic flow over an oscillating airfoil. These computations were compared with experimental data (obtained at Ames Research Center) which form part of the AGARD standard configuration for aeroelastic analysis. A variety of approximations to the full Navier-Stokes equations was used to determine the effect of frequency, shock wave motion, flow separation, and airfoil geometry on unsteady pressures and overall air loads. Good agreement is shown between experiment and theory with the limiting factor being the lack of a reliable turbulence model for high Reynolds number, unsteady transonic flows.

  4. Numerical studies of transverse curvature effects on transonic flow stability

    NASA Technical Reports Server (NTRS)

    Macaraeg, M. G.; Daudpota, Q. I.

    1992-01-01

    A numerical study of transverse curvature effects on compressible flow temporal stability for transonic to low supersonic Mach numbers is presented for axisymmetric modes. The mean flows studied include a similar boundary-layer profile and a nonsimilar axisymmetric boundary-layer solution. The effect of neglecting curvature in the mean flow produces only small quantitative changes in the disturbance growth rate. For transonic Mach numbers (1-1.4) and aerodynamically relevant Reynolds numbers (5000-10,000 based on displacement thickness), the maximum growth rate is found to increase with curvature - the maximum occurring at a nondimensional radius (based on displacement thickness) between 30 and 100.

  5. Implicit, nonswitching, vector-oriented algorithm for steady transonic flow

    NASA Technical Reports Server (NTRS)

    Lottati, I.

    1983-01-01

    A rapid computation of a sequence of transonic flow solutions has to be performed in many areas of aerodynamic technology. The employment of low-cost vector array processors makes the conduction of such calculations economically feasible. However, for a full utilization of the new hardware, the developed algorithms must take advantage of the special characteristics of the vector array processor. The present investigation has the objective to develop an efficient algorithm for solving transonic flow problems governed by mixed partial differential equations on an array processor.

  6. Semi-span model testing in the National Transonic Facility

    NASA Technical Reports Server (NTRS)

    Chokani, Ndaona; Milholen, William E., II

    1993-01-01

    A semi-span testing technique has been proposed for the NASA Langley Research Center's National Transonic Facility (NTF). Semi-span testing has several advantages including (1) larger model size, giving increased Reynolds number capability; (2) improved model fidelity, allowing ease of flap and slat positioning which ultimately improves data quality; and (3) reduced construction costs compared with a full-span model. In addition, the increased model size inherently allows for increased model strength, reducing aeroelastic effects at the high dynamic pressure levels necessary to simulate flight Reynolds numbers. The Energy Efficient Transport (EET) full-span model has been modified to become the EET semi-span model. The full-span EET model was tested extensively at both NASA LRC and NASA Ames Research Center. The available full-span data will be useful in validating the semi-span test strategy in the NTF. In spite of the advantages discussed above, the use of a semi-span model does introduce additional challenges which must be addressed in the testing procedure. To minimize the influence of the sidewall boundary layer on the flow over the semi-span model, the model must be off-set from the sidewall. The objective is to remove the semi-span model from the sidewall boundary layer by use of a stand-off geometry. When this is done however, the symmetry along the centerline of the full-span model is lost when the semi-span model is mounted on the wind tunnel sidewall. In addition, the large semi-span model will impose a significant pressure loading on the sidewall boundary layer, which may cause separation. Even under flow conditions where the sidewall boundary layer remains attached, the sidewall boundary layer may adversely effect the flow over the semi-span model. Also, the increased model size and sidewall mounting requires a modified wall correction strategy. With these issues in mind, the semi-span model has been well instrumented with surface pressure taps to

  7. Parker winds revisited: an extension to disc winds

    NASA Astrophysics Data System (ADS)

    Waters, Timothy R.; Proga, Daniel

    2012-11-01

    A simple 1D dynamical model of thermally driven disc winds is proposed, based on the results of recent, 2.5D axisymmetric simulations. Our formulation of the disc wind problem is in the spirit of the original Parker and Bondi problems, namely we assume an elementary flow configuration consisting of an outflow following pre-defined trajectories in the presence of a central gravitating point mass. Viscosity and heat conduction are neglected. We consider two different streamline geometries, both comprised of straight lines in the (x, z)-plane: (i) streamlines that converge to a geometric point located at (x, z) = (0, -d) and (ii) streamlines that emerge at a constant inclination angle from the disc mid-plane (the x-axis, as we consider geometrically thin accretion discs). The former geometry is commonly used in kinematic models to compute synthetic spectra, while the latter, which exhibits self-similarity, is likely unused for this purpose, although it easily can be with existing kinematic models. We make the case that it should be, i.e. geometry (ii) leads to transonic wind solutions with substantially different properties owing to its lack of streamline divergence. Both geometries can be used to complement recent efforts to estimate photoevaporative mass-loss rates from protoplanetary discs. Pertinent to understanding our disc wind results, which are also applicable to X-ray binaries and active galactic nuclei, is a focused discussion on lesser known properties of classic Parker wind solutions. We find that the parameter space corresponding to decelerating Parker wind solutions is made larger due to rotation and leads instead to disc wind solutions that always accelerate after the bulk velocity is slowed to a minimum value. Surprisingly, Keplerian rotation may allow for two different transonic wind solutions for the same physical conditions.

  8. Cryogenic Wind Tunnel Models. Design and Fabrication

    NASA Technical Reports Server (NTRS)

    Young, C. P., Jr. (Compiler); Gloss, B. B. (Compiler)

    1983-01-01

    The principal motivating factor was the National Transonic Facility (NTF). Since the NTF can achieve significantly higher Reynolds numbers at transonic speeds than other wind tunnels in the world, and will therefore occupy a unique position among ground test facilities, every effort is being made to ensure that model design and fabrication technology exists to allow researchers to take advantage of this high Reynolds number capability. Since a great deal of experience in designing and fabricating cryogenic wind tunnel models does not exist, and since the experience that does exist is scattered over a number of organizations, there is a need to bring existing experience in these areas together and share it among all interested parties. Representatives from government, the airframe industry, and universities are included.

  9. The cryogenic wind tunnel for high Reynolds number testing

    NASA Technical Reports Server (NTRS)

    Kilgore, R. A.; Dress, D. A.

    1982-01-01

    The development of cryogenic wind tunnels is reviewed with reference to the theory and advantages of cryogenic tunnels, problems common to wind tunnels and their solution, and application of cryogenic wind tunnels to high Reynolds number testing. It is shown that cryogenic wind tunnels can achieve full-scale Reynolds number with reasonable tunnel size, dynamic pressure, and drive power; the use of such tunnels also makes it possible to separate the effects of Reynolds number, Mach number, and aeroelasticity. Application of the cryogenic tunnel concept is illustrated by three examples, namely an atmospheric low-speed cryogenic tunnel, a 0.3-meter transonic cryogenic tunnel, and the National Transonic Facility now nearing completion.

  10. Basis Function Approximation of Transonic Aerodynamic Influence Coefficient Matrix

    NASA Technical Reports Server (NTRS)

    Li, Wesley W.; Pak, Chan-gi

    2011-01-01

    A technique for approximating the modal aerodynamic influence coefficients matrices by using basis functions has been developed and validated. An application of the resulting approximated modal aerodynamic influence coefficients matrix for a flutter analysis in transonic speed regime has been demonstrated. This methodology can be applied to the unsteady subsonic, transonic, and supersonic aerodynamics. The method requires the unsteady aerodynamics in frequency-domain. The flutter solution can be found by the classic methods, such as rational function approximation, k, p-k, p, root-locus et cetera. The unsteady aeroelastic analysis for design optimization using unsteady transonic aerodynamic approximation is being demonstrated using the ZAERO flutter solver (ZONA Technology Incorporated, Scottsdale, Arizona). The technique presented has been shown to offer consistent flutter speed prediction on an aerostructures test wing 2 configuration with negligible loss in precision in transonic speed regime. These results may have practical significance in the analysis of aircraft aeroelastic calculation and could lead to a more efficient design optimization cycle.

  11. Basis Function Approximation of Transonic Aerodynamic Influence Coefficient Matrix

    NASA Technical Reports Server (NTRS)

    Li, Wesley Waisang; Pak, Chan-Gi

    2010-01-01

    A technique for approximating the modal aerodynamic influence coefficients [AIC] matrices by using basis functions has been developed and validated. An application of the resulting approximated modal AIC matrix for a flutter analysis in transonic speed regime has been demonstrated. This methodology can be applied to the unsteady subsonic, transonic and supersonic aerodynamics. The method requires the unsteady aerodynamics in frequency-domain. The flutter solution can be found by the classic methods, such as rational function approximation, k, p-k, p, root-locus et cetera. The unsteady aeroelastic analysis for design optimization using unsteady transonic aerodynamic approximation is being demonstrated using the ZAERO(TradeMark) flutter solver (ZONA Technology Incorporated, Scottsdale, Arizona). The technique presented has been shown to offer consistent flutter speed prediction on an aerostructures test wing [ATW] 2 configuration with negligible loss in precision in transonic speed regime. These results may have practical significance in the analysis of aircraft aeroelastic calculation and could lead to a more efficient design optimization cycle

  12. Unsteady transonic algorithm improvements for realistic aircraft applications

    NASA Technical Reports Server (NTRS)

    Batina, John T.

    1987-01-01

    Improvements to a time-accurate approximate factorization (AF) algorithm were implemented for steady and unsteady transonic analysis of realistic aircraft configurations. These algorithm improvements were made to the CAP-TSD (Computational Aeroelasticity Program - Transonic Small Disturbance) code developed at the Langley Research Center. The code permits the aeroelastic analysis of complete aircraft in the flutter critical transonic speed range. The AF algorithm of the CAP-TSD code solves the unsteady transonic small-disturbance equation. The algorithm improvements include: an Engquist-Osher (E-O) type-dependent switch to more accurately and efficiently treat regions of supersonic flow; extension of the E-O switch for second-order spatial accuracy in these regions; nonreflecting far field boundary conditions for more accurate unsteady applications; and several modifications which accelerate convergence to steady-state. Calculations are presented for several configurations including the General Dynamics one-ninth scale F-16C aircraft model to evaluate the algorithm modifications. The modifications have significantly improved the stability of the AF algorithm and hence the reliability of the CAP-TSD code in general.

  13. Transonic unsteady aerodynamic and aeroelastic calculations about airfoils and wings

    NASA Technical Reports Server (NTRS)

    Goorjian, Peter M.; Guruswamy, Guru P.

    1988-01-01

    Recent advances in the numerical simulation of unsteady transonic flow around airfoils and wings are surveyed, with an emphasis on the treatment of aeroelastic effects. The fundamental physical principles involved are discussed, and the numerical implementation of the methods is considered. Typical results are presented in extensive graphs and diagrams and briefly characterized, with reference to experimental data.

  14. Wind tunnel testing of low-drag airfoils

    NASA Technical Reports Server (NTRS)

    Harvey, W. Donald; Mcghee, R. J.; Harris, C. D.

    1986-01-01

    Results are presented for the measured performance recently obtained on several airfoil concepts designed to achieve low drag by maintaining extensive regions of laminar flow without compromising high-lift performance. The wind tunnel results extend from subsonic to transonic speeds and include boundary-layer control through shaping and suction. The research was conducted in the NASA Langley 8-Ft Transonic Pressure Tunnel (TPT) and Low Turbulence Pressure Tunnel (LTPT) which have been developed for testing such low-drag airfoils. Emphasis is placed on identifying some of the major factors influencing the anticipated performance of low-drag airfoils.

  15. Measurement of recovery temperature on an airfoil in the Langley 0.3-m transonic cryogenic tunnel

    NASA Technical Reports Server (NTRS)

    Johnson, C. B.; Adcock, J. B.

    1981-01-01

    Experimental measurements of recovery temperature were made on an airfoil in the Langley 0.3-m Transonic Cryogenic Tunnel at Mach numbers of 0.60 and 0.84 over a Reynolds number per meter range from about 15,000,000 to about 335,000,000. The measured recovery temperatures were considerably below those associated with ideal-gas ambient temperature wind tunnels. This difference was accentuated as the stagnation pressure increased and the total temperature decreased. A boundary-layer code modified for use with cryogenic nitrogen adequately predicted the measured adiabatic wall temperature at all conditions. A quantitative, on-line assessment of the nonadiabatic condition of a model can be made during the operation of a cryogenic wind tunnel by using a correlation for the adiabatic wall temperature which is only a function of total temperature, total pressure, and local Mach number on the model.

  16. Wind-Tunnel/Flight Correlation, 1981

    NASA Technical Reports Server (NTRS)

    Mckinney, L. W. (Editor); Baals, D. D. (Editor)

    1982-01-01

    Wind-tunnel/flight correlation activities are reviewed to assure maximum effectiveness of the early experimental programs of the National Transonic Facility (NTF). Topics included a status report of the NTF, the role of tunnel-to-tunnel correlation, a review of past flight correlation research and the resulting data base, the correlation potential of future flight vehicles, and an assessment of the role of computational fluid dynamics.

  17. Models for cryogenic wind tunnels

    NASA Technical Reports Server (NTRS)

    Lawing, Pierce L.

    1989-01-01

    Model requirements, types of model construction methods, and research in new ways to build models are discussed. The 0.3-m Transonic Cryogenic Tunnel was in operation for 16 years and many 2-D airfoil pressure models were tested. In addition there were airfoil models dedicated to transition detection techniques and other specialized research. There were also a number of small 3-D models tested. A chronological development in model building technique is described which led to the construction of many successful models. The difficulties of construction are illustrated by discussing several unsuccessful model fabrication attempts. The National Transonic Facility, a newer and much larger tunnel, was used to test a variety of models including a submarine, transport and fighter configurations, and the Shuttle Orbiter. A new method of building pressure models was developed and is described. The method is centered on the concept of bonding together plates with pressure channels etched into the bond planes, which provides high density pressure instrumentation with minimum demand on parent model material. With care in the choice of materials and technique, vacuum brazing can be used to produce strong bonds without blocking pressure channels and with no bonding voids between channels. Using multiple plates, a 5 percent wing with 96 orifices was constructed and tested in a transonic cryogenic wind tunnel. Samples of test data are presented and future applications of the technology are suggested.

  18. Transonic Unsteady Aerodynamics of the F/A-18E at Conditions Promoting Abrupt Wing Stall

    NASA Technical Reports Server (NTRS)

    Schuster, David M.; Byrd, James E.

    2003-01-01

    A transonic wind tunnel test of an 8% F/A-18E model was conducted in the NASA Langley Research Center (LaRC) 16-Foot Transonic Tunnel (16-Ft TT) to investigate the Abrupt Wing Stall (AWS) characteristics of this aircraft. During this test, both steady and unsteady measurements of balance loads, wing surface pressures, wing root bending moments, and outer wing accelerations were performed. The test was conducted with a wide range of model configurations and test conditions in an attempt to reproduce behavior indicative of the AWS phenomenon experienced on full-scale aircraft during flight tests. This paper focuses on the analysis of the unsteady data acquired during this test. Though the test apparatus was designed to be effectively rigid. model motions due to sting and balance flexibility were observed during the testing, particularly when the model was operating in the AWS flight regime. Correlation between observed aerodynamic frequencies and model structural frequencies are analyzed and presented. Significant shock motion and separated flow is observed as the aircraft pitches through the AWS region. A shock tracking strategy has been formulated to observe this phenomenon. Using this technique, the range of shock motion is readily determined as the aircraft encounters AWS conditions. Spectral analysis of the shock motion shows the frequencies at which the shock oscillates in the AWS region, and probability density function analysis of the shock location shows the propensity of the shock to take on a bi-stable and even tri-stable character in the AWS flight regime.

  19. Aerodynamic analysis for aircraft with nacelles, pylons, and winglets at transonic speeds

    NASA Technical Reports Server (NTRS)

    Boppe, Charles W.

    1987-01-01

    A computational method has been developed to provide an analysis for complex realistic aircraft configurations at transonic speeds. Wing-fuselage configurations with various combinations of pods, pylons, nacelles, and winglets can be analyzed along with simpler shapes such as airfoils, isolated wings, and isolated bodies. The flexibility required for the treatment of such diverse geometries is obtained by using a multiple nested grid approach in the finite-difference relaxation scheme. Aircraft components (and their grid systems) can be added or removed as required. As a result, the computational method can be used in the same manner as a wind tunnel to study high-speed aerodynamic interference effects. The multiple grid approach also provides high boundary point density/cost ratio. High resolution pressure distributions can be obtained. Computed results are correlated with wind tunnel and flight data using four different transport configurations. Experimental/computational component interference effects are included for cases where data are available. The computer code used for these comparisons is described in the appendices.

  20. Recent Improvements in Semi-Span Testing at the National Transonic Facility (Invited)

    NASA Technical Reports Server (NTRS)

    Gatlin, Gregory M.; Tomek, William G.; Payne, Francis M.; Griffiths, Robert C.

    2006-01-01

    Three wind tunnel investigations of a commercial transport, high-lift, semi-span configuration have recently been conducted in the National Transonic Facility at the NASA Langley Research Center. Throughout the course of these investigations multiple improvements have been developed in the facility semi-span test capability. The primary purpose of the investigations was to assess Reynolds number scale effects on a modern commercial transport configuration up to full-scale flight test conditions (Reynolds numbers on the order of 27 million). The tests included longitudinal aerodynamic studies at subsonic takeoff and landing conditions across a range of Reynolds numbers from that available in conventional wind tunnels up to flight conditions. The purpose of this paper is to discuss lessons learned and improvements incorporated into the semi-span testing process. Topics addressed include enhanced thermal stabilization and moisture reduction procedures, assessments and improvements in model sealing techniques, compensation of model reference dimensions due to test temperature, significantly improved semi-span model access capability, and assessments of data repeatability.

  1. A method for data base management and analysis for wind tunnel data

    NASA Technical Reports Server (NTRS)

    Biser, Aileen O.

    1987-01-01

    To respond to the need for improved data base management and analysis capabilities for wind-tunnel data at the Langley 16-Foot Transonic Tunnel, research was conducted into current methods of managing wind-tunnel data and a method was developed as a solution to this need. This paper describes the development of the data base management and analysis method for wind-tunnel data. The design and implementation of the software system are discussed and examples of its use are shown.

  2. Transonic Flows of Bethe-Zel'dovich-Thompson Fluids

    NASA Astrophysics Data System (ADS)

    Cramer, Mark; Andreyev, Aleksandr

    2013-11-01

    We examine steady transonic flows of Bethe-Zel'dovich-Thompson (BZT) fluids over thin turbine blades or airfoils. BZT fluids are ordinary fluids having a region of negative fundamental derivative over a finite range of pressures and temperatures in the single phase regime. We present the transonic small disturbance equation, shock jump conditions, and shock existence conditions capable of capturing the qualitative behavior of BZT fluids. The flux function is seen to be quartic in the pressure or density perturbation rather than the quadratic (convex) flux function of the perfect gas theory. We show how this nonconvex flux function can be used to predict and explain the complex flows possible. Numerical solutions using a successive line relaxation (SLR) scheme are presented. New results of interest include shock-splitting, collisions between expansion and compression shocks, two compressive bow shocks in supersonic flows, and the observation of as many as three normal stern shocks following an oblique trailing edge shock.

  3. Analysis of transonic flow about lifting wing-body configurations

    NASA Technical Reports Server (NTRS)

    Barnwell, R. W.

    1975-01-01

    An analytical solution was obtained for the perturbation velocity potential for transonic flow about lifting wing-body configurations with order-one span-length ratios and small reduced-span-length ratios and equivalent-thickness-length ratios. The analysis is performed with the method of matched asymptotic expansions. The angles of attack which are considered are small but are large enough to insure that the effects of lift in the region far from the configuration are either dominant or comparable with the effects of thickness. The modification to the equivalence rule which accounts for these lift effects is determined. An analysis of transonic flow about lifting wings with large aspect ratios is also presented.

  4. Validation of Blockage Interference Corrections in the National Transonic Facility

    NASA Technical Reports Server (NTRS)

    Walker, Eric L.

    2007-01-01

    A validation test has recently been constructed for wall interference methods as applied to the National Transonic Facility (NTF). The goal of this study was to begin to address the uncertainty of wall-induced-blockage interference corrections, which will make it possible to address the overall quality of data generated by the facility. The validation test itself is not specific to any particular modeling. For this present effort, the Transonic Wall Interference Correction System (TWICS) as implemented at the NTF is the mathematical model being tested. TWICS uses linear, potential boundary conditions that must first be calibrated. These boundary conditions include three different classical, linear. homogeneous forms that have been historically used to approximate the physical behavior of longitudinally slotted test section walls. Results of the application of the calibrated wall boundary conditions are discussed in the context of the validation test.

  5. Assessment of the National Transonic Facility for Laminar Flow Testing

    NASA Technical Reports Server (NTRS)

    Crouch, Jeffrey D.; Sutanto, Mary I.; Witkowski, David P.; Watkins, A. Neal; Rivers, Melissa B.; Campbell, Richard L.

    2010-01-01

    A transonic wing, designed to accentuate key transition physics, is tested at cryogenic conditions at the National Transonic Facility at NASA Langley. The collaborative test between Boeing and NASA is aimed at assessing the facility for high-Reynolds number testing of configurations with significant regions of laminar flow. The test shows a unit Reynolds number upper limit of 26 M/ft for achieving natural transition. At higher Reynolds numbers turbulent wedges emanating from the leading edge bypass the natural transition process and destroy the laminar flow. At lower Reynolds numbers, the transition location is well correlated with the Tollmien-Schlichting-wave N-factor. The low-Reynolds number results suggest that the flow quality is acceptable for laminar flow testing if the loss of laminar flow due to bypass transition can be avoided.

  6. Transonic stability and control of aircraft using CFD methods

    NASA Technical Reports Server (NTRS)

    Vinh, Lam-Son; Edwards, John W.; Seidel, David A.; Batina, John T.

    1988-01-01

    Implementation of a capability to calculate longitudinal short-period response in the CAP-TSD (Computational Aeroelasticity Program - Transonic Small Disturbance) finite-difference code is described. The code, developed recently at the NASA Langley Research Center, is capable of solving steady and unsteady flows about complete aircraft configurations and is used primarily for aeroelastic calculations in the critical transonic speed range. The longitudinal short-period equations of motion in state-space form have been coupled to the time-accurate lift and moment calculated by the program. Transient responses to an elevator pulse for free-flying aircraft demonstrate the new capability. A trim routine is also added to the code to obtain trim automatically during steady-state flow field convergence. Stability and control derivatives are estimated from the calculated transient response by a maximum likelihood estimation program. Results for a fighter configuration and a general aviation configuration are presented to assess the capability.

  7. Studies in a transonic rotor aerodynamics and noise facility

    NASA Technical Reports Server (NTRS)

    Wright, S. E.; Lee, D. J.; Crosby, W.

    1984-01-01

    The design, construction and testing of a transonic rotor aerodynamics and noise facility was undertaken, using a rotating arm blade element support technique. This approach provides a research capability intermediate between that of a stationary element in a moving flow and that of a complete rotating blade system, and permits the acoustic properties of blade tip elements to be studied in isolation. This approach is an inexpensive means of obtaining data at high subsonic and transonic tip speeds on the effect of variations in tip geometry. The facility may be suitable for research on broad band noise and discrete noise in addition to high-speed noise. Initial tests were conducted over the Mach number range 0.3 to 0.93 and confirmed the adequacy of the acoustic treatment used in the facility to avoid reflection from the enclosure.

  8. Laser velocimetry applied to transonic and supersonic aerodynamics

    NASA Technical Reports Server (NTRS)

    Johnson, D. A.; Bachalo, W. D.; Moddaress, D.

    1976-01-01

    Measurements obtained with laser velocimetry in a Mach 2.9 separated turbulent boundary layer and in the transonic flow past a two-dimensional airfoil section are presented and compared to data realized by conventional techniques. Agreement in mean velocities was realized where the pressure measurements could be considered reliable; however, in regions of instantaneous reverse velocities, the laser results were found to be consistent with the physics of the flow whereas the pressure data were not. Streamwise turbulence intensities are also presented. In the transonic airfoil study, velocity measurements obtained immediately outside the upper surface boundary layer of a 6-inch chord NACA 64A010 airfoil are compared to edge velocities inferred from surface pressure measurements. For free-stream Mach numbers of 0.6 and 0.8, the agreement in results was very good. "Dual scatter" optical arrangements in conjunction with a single particle, counter-type signal processor were employed in these investigations.

  9. A parametric study of transonic blade-vortex interaction noise

    NASA Technical Reports Server (NTRS)

    Lyrintzis, A. S.

    1991-01-01

    Several parameters of transonic blade-vortex interactions (BVI) are being studied and some ideas for noise reduction are introduced and tested using numerical simulation. The model used is the two-dimensional high frequency transonic small disturbance equation with regions of distributed vorticity (VTRAN2 code). The far-field noise signals are obtained by using the Kirchhoff method with extends the numerical 2-D near-field aerodynamic results to the linear acoustic 3-D far-field. The BVI noise mechanisms are explained and the effects of vortex type and strength, and angle of attack are studied. Particularly, airfoil shape modifications which lead to noise reduction are investigated. The results presented are expected to be helpful for better understanding of the nature of the BVI noise and better blade design.

  10. Convergence acceleration and shock fitting for transonic aerodynamics computations

    NASA Technical Reports Server (NTRS)

    Hafez, M. M.; Cheng, H. K.

    1975-01-01

    Two problems in computational fluid dynamics are studied in the context of transonic small-disturbance theory - namely, (1) how to speed up the convergence for currently available iterative procedures, and (2) how a shock-fitting method may be adapted to existing relaxation procedures with minimal alterations in computer programming and storage requirements. The paper contributes to a clarification of error analyses for sequence transformations based on the power method (including also the nonlinear transforms of Aitken, Shanks, and Wilkinson), and to developing a cyclic iterative procedure applying the transformations. Examples testing the procedure for a model Dirichlet problem and for a transonic airfoil problem show that savings in computer time by a factor of three to five are generally possible, depending on accuracy requirements and the particular iterative procedure used.-

  11. Unsteady transonic potential flow over a flexible fuselage

    NASA Technical Reports Server (NTRS)

    Gibbons, Michael D.

    1993-01-01

    A flexible fuselage capability has been developed and implemented within version 1.2 of the CAP-TSD code. The capability required adding time dependent terms to the fuselage surface boundary conditions and the fuselage surface pressure coefficient. The new capability will allow modeling the effect of a flexible fuselage on the aeroelastic stability of complex configurations. To assess the flexible fuselage capability several steady and unsteady calculations have been performed for slender fuselages with circular cross-sections. Steady surface pressures are compared with experiment at transonic flight conditions. Unsteady cross-sectional lift is compared with other analytical results at a low subsonic speed and a transonic case has been computed. The comparisons demonstrate the accuracy of the flexible fuselage modifications.

  12. Videogrammetric Model Deformation Measurement Technique for Wind Tunnel Applications

    NASA Technical Reports Server (NTRS)

    Barrows, Danny A.

    2006-01-01

    Videogrammetric measurement technique developments at NASA Langley were driven largely by the need to quantify model deformation at the National Transonic Facility (NTF). This paper summarizes recent wind tunnel applications and issues at the NTF and other NASA Langley facilities including the Transonic Dynamics Tunnel, 31-Inch Mach 10 Tunnel, 8-Ft high Temperature Tunnel, and the 20-Ft Vertical Spin Tunnel. In addition, several adaptations of wind tunnel techniques to non-wind tunnel applications are summarized. These applications include wing deformation measurements on vehicles in flight, determining aerodynamic loads based on optical elastic deformation measurements, measurements on ultra-lightweight and inflatable space structures, and the use of an object-to-image plane scaling technique to support NASA s Space Exploration program.

  13. Cryogenic Balance Technology at the National Transonic Facility

    NASA Technical Reports Server (NTRS)

    Parker, P. A.

    2001-01-01

    This paper provides an overview of force measurement at the National Transonic Facility (NTF). The NTF has unique force measurement requirements that dictate an integration of all aspects of balance design, production, and calibration. An overview of current force measurement capabilities is provided along with new balance development efforts. Research activities in the areas of thermal compensation and balance calibration are presented. Also, areas of future research are detailed.

  14. Some examples of unsteady transonic flows over airfoils

    NASA Technical Reports Server (NTRS)

    Ballhaus, W. F.; Magnus, R.; Yoshihara, H.

    1975-01-01

    A finite difference flutter analysis is presented for the NACA 64A-410 airfoil at M equals 0.72, where the incidence is abruptly changed from 2 to 4 degrees. The effect of gust loads is studied, and the unsteady flow adjusting process is displayed. The semi-implicit procedure of Ballhaus and Lomax (1974) is used to solve the small disturbance transonic potential equation. The physical aspects of the results, rather than the numerical details, are emphasized.

  15. Calculation of transonic flows using an extended integral equation method

    NASA Technical Reports Server (NTRS)

    Nixon, D.

    1976-01-01

    An extended integral equation method for transonic flows is developed. In the extended integral equation method velocities in the flow field are calculated in addition to values on the aerofoil surface, in contrast with the less accurate 'standard' integral equation method in which only surface velocities are calculated. The results obtained for aerofoils in subcritical flow and in supercritical flow when shock waves are present compare satisfactorily with the results of recent finite difference methods.

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

    NASA Technical Reports Server (NTRS)

    Flores, J.; Holst, T. L.; Kaynak, Unver; Gundy, K.; Thomas, S. D.

    1986-01-01

    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.

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

    NASA Technical Reports Server (NTRS)

    Flores, J.; Gundy, K.; Gundy, K.; Gundy, K.; Gundy, K.; Gundy, K.

    1986-01-01

    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.

  18. Transonic Free-To-Roll Analysis of the F/A-18E and F-35 Configurations

    NASA Technical Reports Server (NTRS)

    Owens, D. Bruce; McConnell, Jeffrey K.; Brandon, Jay M.; Hall, Robert M.

    2004-01-01

    The free-to-roll technique is used as a tool for predicting areas of uncommanded lateral motions. Recently, the NASA/Navy/Air Force Abrupt Wing Stall Program extended the use of this technique to the transonic speed regime. Using this technique, this paper evaluates various wing configurations on the pre-production F/A-18E aircraft and the Joint Strike Fighter (F-35) aircraft. The configurations investigated include leading and trailing edge flap deflections, fences, leading edge flap gap seals, and vortex generators. These tests were conducted in the NASA Langley 16-Foot Transonic Tunnel. The analysis used a modification of a figure-of-merit developed during the Abrupt Wing Stall Program to discern configuration effects. The results showed how the figure-of-merit can be used to schedule wing flap deflections to avoid areas of uncommanded lateral motion. The analysis also used both static and dynamic wind tunnel data to provide insight into the uncommanded lateral behavior. The dynamic data was extracted from the time history data using parameter identification techniques. In general, modifications to the pre-production F/A-18E resulted in shifts in angle-of-attack where uncommanded lateral activity occurred. Sealing the gap between the inboard and outboard leading-edge flaps on the Navy version of the F-35 eliminated uncommanded lateral activity or delayed the activity to a higher angle-of-attack.

  19. High Reynolds number tests of a Douglas DLBA 032 airfoil in the Langley 0.3-meter transonic cryogenic tunnel

    NASA Technical Reports Server (NTRS)

    Johnson, Charles B.; Dress, David A.; Hill, Acquilla S.; Wilcox, Peter A.; Bui, Minh H.

    1986-01-01

    A wind-tunnel investigation of a Douglas advanced-technology airfoil was conducted in the Langley 0.3-Meter Transonic Cryogenic Tunnel (0.3-m TCT). The temperature was varied from 227 K (409 R) to 100 K (180 R) at pressures ranging from about 159 kPa (1.57 atm) to about 514 kPa (5.07 atm). Mach number was varied from 0.50 to 0.78. These variables provided a Reynolds number range (based on airfoil chord) from 6.0 to 30.0 x 10 to the 6th power. This investigation was specifically designed to: (1) test a Douglas airfoil from moderately low to flight-equivalent Reynolds numbers, and (2) evaluate sidewall-boundary-layer effects on transonic airfoil performance characteristics by a systematic variation of Mach number, Reynolds number, and sidewall-boundary-layer removal. Data are included which demonstrate the effects of fixing transition, Mach number, Reynolds number, and sidewall-boundary-layer removal on the aerodynamic characteristics of the airfoil. Also included are remarks on model design and model structural integrity.

  20. Transonic single-mode flutter and buffet of a low aspect ratio wing having a subsonic airfoil shape

    NASA Technical Reports Server (NTRS)

    Erickson, L. L.

    1974-01-01

    Transonic flutter and buffet results obtained from wind-tunnel tests of a low aspect ratio semispan wing model are presented. The tests were conducted to investigate potential transonic aeroelastic problems of vehicles having subsonic airfoil sections. The model employed NACA 00XX-64 airfoil sections in the streamwise direction and had a 14 deg leading edge sweep angle. Aspect ratio, and average thickness were 4.0, 0.35, and 8 percent, respectively. The model was tested at Mach numbers from 0.6 to 0.95 at angles of attack from 0 deg to 15 deg. Two zero lift flutter conditions were found that involved essentially single normal mode vibrations. With boundary layer trips on the model, flutter occurred in a narrow Mach number range centered at about Mach 0.90. The frequency and motion of this flutter were like that of the first normal mode vibration. With the trips removed flutter occurred at a slightly high Mach number but in a mode strongly resembling that of the second normal mode.

  1. Transonic flow past a wedge profile with detached bow wave

    NASA Technical Reports Server (NTRS)

    Vincenti, Walter G; Wagoner, Cleo B

    1952-01-01

    A theoretical study has been made of the aerodynamic characteristics at zero angle of attack of a thin, doubly symmetrical double-wedge profile in the range of supersonic flight speed in which the bow wave is detached. The analysis utilizes the equations of the transonic small-disturbance theory and involves no assumptions beyond those implicit in this theory. The mixed flow about the front half of the profile is calculated by relaxation solution of boundary conditions along the shock polar and sonic line. The purely subsonic flow about the rear of the profile is found by means of the method of characteristics specialized to the transonic small-disturbance theory. Complete calculations were made for four values of the transonic similarity parameter. These were found sufficient to bridge the gap between the previous results of Guderley and Yoshihara at a Mach number of 1 and the results which are readily obtained when the bow wave is attached and the flow is completely supersonic.

  2. Fourier time spectral method for subsonic and transonic flows

    NASA Astrophysics Data System (ADS)

    Zhan, Lei; Liu, Feng; Papamoschou, Dimitri

    2016-06-01

    The time accuracy of the exponentially accurate Fourier time spectral method (TSM) is examined and compared with a conventional 2nd-order backward difference formula (BDF) method for periodic unsteady flows. In particular, detailed error analysis based on numerical computations is performed on the accuracy of resolving the local pressure coefficient and global integrated force coefficients for smooth subsonic and non-smooth transonic flows with moving shock waves on a pitching airfoil. For smooth subsonic flows, the Fourier TSM method offers a significant accuracy advantage over the BDF method for the prediction of both the local pressure coefficient and integrated force coefficients. For transonic flows where the motion of the discontinuous shock wave contributes significant higher-order harmonic contents to the local pressure fluctuations, a sufficient number of modes must be included before the Fourier TSM provides an advantage over the BDF method. The Fourier TSM, however, still offers better accuracy than the BDF method for integrated force coefficients even for transonic flows. A problem of non-symmetric solutions for symmetric periodic flows due to the use of odd numbers of intervals is uncovered and analyzed. A frequency-searching method is proposed for problems where the frequency is not known a priori. The method is tested on the vortex shedding problem of the flow over a circular cylinder.

  3. Effects of increased leading-edge thickness on performance of a transonic rotor blade. [in single stage transonic compressor

    NASA Technical Reports Server (NTRS)

    Reid, L.; Urasek, D. C.

    1972-01-01

    A single-stage transonic compressor was tested with two rotor blade leading-edge configurations to investigate the effect of increased leading-edge thickness on the performance of a transonic blade row. The original rotor blade configuration was modified by cutting back the leading edge sufficiently to double the blade leading-edge thickness and thus the blade gap blockage in the tip region. At design speed this modification resulted in a decrease in rotor overall peak efficiency of four points. The major portion of this decrement in rotor overall peak efficienty was attributed to the flow conditions in the outer 30 percent of the blade span. At 70 and 90 percent of design speed, the modification had very little effect on rotor overall performance.

  4. Transonic Experimental Observations of Abrupt Wing Stall on an F/A-18E Model (Invited)

    NASA Technical Reports Server (NTRS)

    McMillin, S. Naomi; Hall, Robert M.; Lamar, John E.

    2003-01-01

    A transonic wind tunnel test of an 8% F/A-18E model was conducted in the NASA Langley Research Center (LaRC) 16 ft Transonic Tunnel (16-ft TT) to investigate on-surface flow physics during stall. The technical approach employed focused on correlating static (or time-averaged) and unsteady wind-tunnel test data to the unsteady wing-stall events using force, moment, pressure, and pressure-sensitive-paint measurements. This paper focuses on data obtained on the pre-production configuration of the F/A-18E aircraft at Mach number of 0.90. The flow unsteadiness occurring on the wing as the wing went through the stall process was captured using the time history of balance and pressure measurements and by calculating the root mean square (RMS) for a number of instrument signals. The second step was to gather global perspectives on the pressures influencing the wing stall process. The abrupt wing stall experienced by the 8% F/A-18E Model was observed to be an unsteady event triggered by the rapid advancement of separation, which had migrated forward from the trailing edge, to the leading-edge flap hingeline over a very small increment in angle of attack. The angle of attack at which this stall occurred varied, from run to run, over an 1 deg increment. The abrupt wing stall was observed, using pressure-sensitive-paint, to occur simultaneously on both wing panels or asymmetrically. The pressure-sensitive paint data and wingroot bending moment data were essential in providing insight to the flow structures occurring over the wing and the possible asymmetry of those flow structures. A repeatability analysis conducted on eight runs of static data provided a quick and inexpensive examination of the unsteady aerodynamic characteristics of abrupt wing stall. The results of the repeatability analysis agreed extremely well with data obtained using unsteady measurement techniques. This approach could be used to identify test conditions for more complex unsteady data measurements using

  5. The cryogenic wind tunnel

    NASA Technical Reports Server (NTRS)

    Kilgore, R. A.

    1976-01-01

    Based on theoretical studies and experience with a low speed cryogenic tunnel and with a 1/3-meter transonic cryogenic tunnel, the cryogenic wind tunnel concept was shown to offer many advantages with respect to the attainment of full scale Reynolds number at reasonable levels of dynamic pressure in a ground based facility. The unique modes of operation available in a pressurized cryogenic tunnel make possible for the first time the separation of Mach number, Reynolds number, and aeroelastic effects. By reducing the drive-power requirements to a level where a conventional fan drive system may be used, the cryogenic concept makes possible a tunnel with high productivity and run times sufficiently long to allow for all types of tests at reduced capital costs and, for equal amounts of testing, reduced total energy consumption in comparison with other tunnel concepts.

  6. LaRC design analysis report for National Transonic Facility for 9% nickel tunnel shell. Volume 1: Finite difference analysis of cone/cylinder junction

    NASA Technical Reports Server (NTRS)

    Ramsey, J. W., Jr.; Taylor, J. T.; Wilson, J. F.; Gray, C. E., Jr.; Leatherman, A. D.; Rooker, J. R.; Allred, J. W.

    1976-01-01

    The results of extensive computer (finite element, finite difference and numerical integration), thermal, fatigue, and special analyses of critical portions of a large pressurized, cryogenic wind tunnel (National Transonic Facility) are presented. The computer models, loading and boundary conditions are described. Graphic capability was used to display model geometry, section properties, and stress results. A stress criteria is presented for evaluation of the results of the analyses. Thermal analyses were performed for major critical and typical areas. Fatigue analyses of the entire tunnel circuit are presented.

  7. LaRC design analysis report for National Transonic Facility for 304 stainless steel tunnel shell. Volume 1S: Finite difference analysis of cone/cylinder junction

    NASA Technical Reports Server (NTRS)

    Ramsey, J. W., Jr.; Taylor, J. T.; Wilson, J. F.; Gray, C. E., Jr.; Leatherman, A. D.; Rooker, J. R.; Allred, J. W.

    1976-01-01

    The results of extensive computer (finite element, finite difference and numerical integration), thermal, fatigue, and special analyses of critical portions of a large pressurized, cryogenic wind tunnel (National Transonic Facility) are presented. The computer models, loading and boundary conditions are described. Graphic capability was used to display model geometry, section properties, and stress results. A stress criteria is presented for evaluation of the results of the analyses. Thermal analyses were performed for major critical and typical areas. Fatigue analyses of the entire tunnel circuit are presented.

  8. A fan pressure ratio correlation in terms of Mach number and Reynolds number for the Langley 0.3 meter transonic cryogenic tunnel

    NASA Technical Reports Server (NTRS)

    Lawing, P. L.; Adcock, J. B.; Ladson, C. L.

    1980-01-01

    Calibration data for the two dimensional test section of the Langley 0.3 Meter Transonic Cryogenic Tunnel were used to develop a Mach number-Reynolds number correlation for the fan pressure ratio in terms of test section conditions. Well established engineering relationships combined to form an equation which is functionally analogous to the correlation. A geometric loss coefficient which is independent of Reynolds number or Mach number was determined. Present and anticipated uses of this concept include improvement of tunnel control schemes, comparison of efficiencies for operationally similar wind tunnels, prediction of tunnel test conditions and associated energy usage, and determination of Reynolds number scaling laws for similar fluid flow systems.

  9. Operating Characteristics of the Multiple Critical Venturi System and Secondary Calibration Nozzles Used for Weight-Flow Measurements in the Langley 16-Foot Transonic Tunnel

    NASA Technical Reports Server (NTRS)

    Berrier, B. L.; Leavitt, L. D.; Bangert, L. S.

    1985-01-01

    An investigation has been conducted in the Langley 16 Foot Transonic Tunnel to determine the weight flow measurement characteristics of a multiple critical Venturi system and the nozzle discharge coefficient characteristics of a series of convergent calibration nozzles. The effects on model discharge coefficient of nozzle throat area, model choke plate open area, nozzle pressure ratio, jet total temperature, and number and combination of operating Venturis were investigated. Tests were conducted at static conditions (tunnel wind off) at nozzle pressure ratios from 1.3 to 7.0.

  10. Selected computations of transonic cavity flows

    NASA Technical Reports Server (NTRS)

    Atwood, Christopher A.

    1993-01-01

    An efficient diagonal scheme implemented in an overset mesh framework has permitted the analysis of geometrically complex cavity flows via the Reynolds averaged Navier-Stokes equations. Use of rapid hyperbolic and algebraic grid methods has allowed simple specification of critical turbulent regions with an algebraic turbulence model. Comparisons between numerical and experimental results are made in two dimensions for the following problems: a backward-facing step; a resonating cavity; and two quieted cavity configurations. In three-dimensions the flow about three early concepts of the stratospheric Observatory For Infrared Astronomy (SOFIA) are compared to wind-tunnel data. Shedding frequencies of resolved shear layer structures are compared against experiment for the quieted cavities. The results demonstrate the progress of computational assessment of configuration safety and performance.

  11. Control of the NASA Langley 16-Foot Transonic Tunnel with the Self-Organizing Feature Map

    NASA Technical Reports Server (NTRS)

    Motter, Mark A.

    1998-01-01

    A predictive, multiple model control strategy is developed based on an ensemble of local linear models of the nonlinear system dynamics for a transonic wind tunnel. The local linear models are estimated directly from the weights of a Self Organizing Feature Map (SOFM). Local linear modeling of nonlinear autonomous systems with the SOFM is extended to a control framework where the modeled system is nonautonomous, driven by an exogenous input. This extension to a control framework is based on the consideration of a finite number of subregions in the control space. Multiple self organizing feature maps collectively model the global response of the wind tunnel to a finite set of representative prototype controls. These prototype controls partition the control space and incorporate experimental knowledge gained from decades of operation. Each SOFM models the combination of the tunnel with one of the representative controls, over the entire range of operation. The SOFM based linear models are used to predict the tunnel response to a larger family of control sequences which are clustered on the representative prototypes. The control sequence which corresponds to the prediction that best satisfies the requirements on the system output is applied as the external driving signal. Each SOFM provides a codebook representation of the tunnel dynamics corresponding to a prototype control. Different dynamic regimes are organized into topological neighborhoods where the adjacent entries in the codebook represent the minimization of a similarity metric which is the essence of the self organizing feature of the map. Thus, the SOFM is additionally employed to identify the local dynamical regime, and consequently implements a switching scheme than selects the best available model for the applied control. Experimental results of controlling the wind tunnel, with the proposed method, during operational runs where strict research requirements on the control of the Mach number were met, are

  12. Flight-determined derivatives and dynamic characteristics for the HL-10 lifting body vehicle at subsonic and transonic Mach numbers

    NASA Technical Reports Server (NTRS)

    Strutz, L. W.

    1972-01-01

    The HL-10 lifting body stability and control derivatives were determined by using an analog-matching technique and compared with derivatives obtained from wind-tunnel results. The flight derivatives were determined as a function of angle of attack for a subsonic configuration at Mach 0.7 and for a transonic configuration at Mach 0.7, 0.9, and 1.2. At an angle of attack of 14 deg, data were obtained for a Mach number range from 0.6 to 1.4. The flight and wind-tunnel derivatives were in general agreement, with the possible exception of the longitudinal and lateral damping derivatives. Some differences were noted between the vehicle dynamic response characteristics calculated from flight-determined derivatives and those predicted by the wind-tunnel results. However, the only difference the pilots noted between the response of the vehicle in flight and the response of a simulator programed with wind-tunnel-predicted data was that the damping generally was higher in the flight vehicle.

  13. 4. VIEW LOOKING NORTHNORTHEAST AT TEST SECTION OF 8FOOT TRANSONIC ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    4. VIEW LOOKING NORTH-NORTHEAST AT TEST SECTION OF 8-FOOT TRANSONIC PRESSURE TUNNEL SHOWING ACCESS PORT TO TEST SECTION (RIGHT) AND PLENUM SURROUNDING AREA. - NASA Langley Research Center, 8-Foot Transonic Pressure Tunnel, 640 Thornell Avenue, Hampton, Hampton, VA

  14. Design features and operational characteristics of the Langley pilot transonic cryogenic tunnel

    NASA Technical Reports Server (NTRS)

    Kilgore, R. A.

    1974-01-01

    A fan-driven transonic cryogenic tunnel was designed, and its purging, cooldown, and warmup times were determined satisfactory. Cooling with liquid nitrogen is at the power levels required for transonic testing. Good temperature distributions are obtained by using a simple nitrogen injection system.

  15. A Microsoft Project-Based Planning, Tracking, and Management Tool for the National Transonic Facility's Model Changeover Process

    NASA Technical Reports Server (NTRS)

    Vairo, Daniel M.

    1998-01-01

    The removal and installation of sting-mounted wind tunnel models in the National Transonic Facility (NTF) is a multi-task process having a large impact on the annual throughput of the facility. Approximately ten model removal and installation cycles occur annually at the NTF with each cycle requiring slightly over five days to complete. The various tasks of the model changeover process were modeled in Microsoft Project as a template to provide a planning, tracking, and management tool. The template can also be used as a tool to evaluate improvements to this process. This document describes the development of the template and provides step-by-step instructions on its use and as a planning and tracking tool. A secondary role of this document is to provide an overview of the model changeover process and briefly describe the tasks associated with it.

  16. Uniqueness of transonic shock solutions in a duct for steady potential flow

    NASA Astrophysics Data System (ADS)

    Chen, Gui-Qiang; Yuan, Hairong

    We study the uniqueness of solutions with a transonic shock in a duct in a class of transonic shock solutions, which are not necessarily small perturbations of the background solution, for steady potential flow. We prove that, for given uniform supersonic upstream flow in a straight duct, there exists a unique uniform pressure at the exit of the duct such that a transonic shock solution exists in the duct, which is unique modulo translation. For any other given uniform pressure at the exit, there exists no transonic shock solution in the duct. This is equivalent to establishing a uniqueness theorem for a free boundary problem of a partial differential equation of second order in a bounded or unbounded duct. The proof is based on the maximum/comparison principle and a judicious choice of special transonic shock solutions as a comparison solution.

  17. On the application of transonic similarity rules to wings of finite span

    NASA Technical Reports Server (NTRS)

    Spreiter, John R

    1953-01-01

    The transonic aerodynamic characteristics of wings of finite span are discussed from the point of view of a unified small perturbation theory for subsonic, transonic, and supersonic flows about thin wings. This approach avoids certain ambiguities which appear if one studies transonic flows by means of equations derived under the more restrictive assumption that the local velocities are everywhere close to sonic velocity. The relation between the two methods of analysis of transonic flow is examined, the similarity rules and known solutions of transonic flow theory are reviewed, and the asymptotic behavior of the lift, drag, and pitching-moment characteristics of wings of large and small aspect ratio is discussed. It is shown that certain methods of data presentation are advantageous for the effective display of these characteristics.

  18. Parker winds revisited: An extension to disk winds

    NASA Astrophysics Data System (ADS)

    Waters, Timothy R.

    A simple 1D dynamical model of thermally driven disk winds is proposed, based on the results of recent, 2.5D axi-symmetric simulations. Our formulation of the disk wind problem is in the spirit of the original Parker (1958) and Bondi (1952) problems, namely we assume an elementary flow configuration consisting of an outflow following pre-defined trajectories in the presence of a central gravitating point mass. Viscosity and heat conduction are neglected. We consider two different streamline geometries, both comprised of straight lines in the (x,z)-plane: (i) streamlines that converge to a geometric point located at (x,z)=(0,-- d) and (ii) streamlines that emerge at a constant inclination angle from the disk midplane (the x-axis, as we consider geometrically thin accretion discs). The former geometry is commonly used in kinematic models to compute synthetic spectra, while the latter, which exhibits self-similarity, is likely unused for this purpose, although it easily can be with existing kinematic models. We make the case that it should be, i.e. that geometry (ii) leads to transonic wind solutions with substantially different properties owing to its lack of streamline divergence. Pertinent to understanding our disk wind results, which are applicable to X-ray binaries, active galactic nuclei, and circumstellar discs, is a focused discussion on lesser known properties of Parker wind solutions. Parker winds are of wide applicability and have recently been used to predict photoevaporative mass loss rates from protoplanetary discs, but not without shortcomings, as we address. In addition, the analytical solutions of Parker winds are ideal for assessing and validating the accuracy of hydrodynamical simulations. Geometry (i) contains the spherically symmetric Parker wind solution as a special case, while one instance of geometry (ii) has been used as a testbed problem for hydrodynamic simulations performed in cylindrical coordinates. We present a parameter survey of our

  19. Contributions of the Transonic Dynamics Tunnel to the Testing of Active Control of Aeroelastic Response

    NASA Technical Reports Server (NTRS)

    Perry, Boyd, III; Noll, Thomas E.; Scott, Robert C.

    2000-01-01

    By the 1960s, researchers began to investigate the feasibility of using active controls technology (ACT) for increasing the capabilities of military and commercial aircraft. Since then many researchers, too numerous to mention, have investigated and demonstrated the usefulness of ACT for favorably modifying the aeroelastic response characteristics of flight vehicles. As a result, ACT entered the limelight as a viable tool for answering some very difficult design questions and had the potential for obtaining structural weight reductions optimizing maneuvering performance, and satisfying the multimission requirements being imposed on future military and commercial aircraft designs. Over the past 40 years, the NASA Langley Research Center (LaRC) has played a major role in developing ACT in part by its participation in many wind-tunnel programs conducted in the Transonic Dynamics Tunnel (TDT). These programs were conducted for the purposes of: (1) establishing concept feasibility; (2) demonstrating proof of concept; and (3) providing data for validating new modeling, analysis, and design methods. This paper provides an overview of the ACT investigations conducted in the TDT. For each program discussed herein, the objectives of the effort, the testing techniques, the test results, any, signIficant findings, and the lessons learned with respect to ACT testing are presented.

  20. Implementation of the WICS Wall Interference Correction System at the National Transonic Facility

    NASA Technical Reports Server (NTRS)

    Iyer, Venkit; Martin, Lockheed; Everhart, Joel L.; Bir, Pamela J.; Ulbrich, Norbert

    2000-01-01

    The Wall Interference Correction System (WICS) is operational at the National Transonic Facility (NTF) of NASA Langley Research Center (NASA LaRC) for semispan and full span tests in the solid wall (slots covered) configuration, The method is based on the wall pressure signature method for computing corrections to the measured parameters. It is an adaptation of the WICS code operational at the 12 ft pressure wind tunnel (12ft PWT) of NASA Ames Research Center (NASA ARC). This paper discusses the details of implementation of WICS at the NTF including, tunnel calibration, code modifications for tunnel and support geometry, changes made for the NTF wall orifices layout, details of interfacing with the tunnel data processing system, and post-processing of results. Example results of applying WICS to a semispan test and a full span test are presented. Comparison with classical correction results and an analysis of uncertainty in the corrections are also given. As a special application of the code, the Mach number calibration data from a centerline pipe test was computed by WICS. Finally, future work for expanding the applicability of the code including online implementation is discussed.

  1. Active load control during rolling maneuvers. [performed in the Langley Transonic Dynamics Tunnel

    NASA Technical Reports Server (NTRS)

    Woods-Vedeler, Jessica A.; Pototzky, Anthony S.; Hoadley, Sherwood T.

    1994-01-01

    A rolling maneuver load alleviation (RMLA) system has been demonstrated on the active flexible wing (AFW) wind tunnel model in the Langley Transonic Dynamics Tunnel (TDT). The objective was to develop a systematic approach for designing active control laws to alleviate wing loads during rolling maneuvers. Two RMLA control laws were developed that utilized outboard control-surface pairs (leading and trailing edge) to counteract the loads and that used inboard trailing-edge control-surface pairs to maintain roll performance. Rolling maneuver load tests were performed in the TDT at several dynamic pressures that included two below and one 11 percent above open-loop flutter dynamic pressure. The RMLA system was operated simultaneously with an active flutter suppression system above open-loop flutter dynamic pressure. At all dynamic pressures for which baseline results were obtained, torsion-moment loads were reduced for both RMLA control laws. Results for bending-moment load reductions were mixed; however, design equations developed in this study provided conservative estimates of load reduction in all cases.

  2. The New Heavy Gas Testing Capability in the NASA Langley Transonic Dynamics Tunnel

    NASA Technical Reports Server (NTRS)

    Cole, Stanley R.; Rivera, Jose A., Jr.

    1997-01-01

    The NASA Langley Transonic Dynamics Tunnel (TDT) has provided a unique capability for aeroelastic testing for over thirty-five years. The facility has a rich history of significant contributions to the design of many United States commercial transports and military aircraft. The facility has many features which contribute to its uniqueness for aeroelasticity testing; however, perhaps the most important facility capability is the use of a heavy gas test medium to achieve higher test densities. Higher test medium densities substantially improve model building requirements and therefore simplify the fabrication process for building aeroelastically scaled wind-tunnel models. The heavy gas also provides other testing benefits, including reduction in the power requirements to operate the facility during testing. Unfortunately, the use of the original heavy gas has been curtailed due to environmental concerns. A new gas, referred to as R-134a, has been identified as a suitable replacement for the former TDT heavy gas. The TDT is currently undergoing a facility upgrade to allow testing in R-134a heavy gas. This replacement gas will result in an operational test envelope, model scaling advantages, and general testing capabilities similar to those available with the former TDT heavy gas. As such, the TDT is expected to remain a viable facility for aeroelasticity research and aircraft dynamic clearance testing well into the 21st century. This paper describes the anticipated advantages and facility calibration plans for the new heavy gas and briefly reviews several past test programs that exemplify the possible benefits of heavy gas testing.

  3. Subsonic, transonic, and supersonic stability and control characteristics of the -147B space shuttle orbiter

    NASA Technical Reports Server (NTRS)

    Mennell, R. C.

    1973-01-01

    Experimental aerodynamic investigations were conducted on 0.015 scale representations of two Space Shuttle Orbiter configurations in a trisonic wind tunnel from June 20, 1973 to June 30, 1973. The primary test objective was to define subsonic, transonic, and supersonic stability and control characteristics of the -147B Orbiter. Six-component aerodynamic force and moment data for the -147B Orbiter were recorded over an angle of attack range of -2 deg to 30 deg at Mach numbers of 0.6, 0.9, 1.2, 2.0, and 3.0. Reynolds numbers of 5.0, 7.0, 8.0, and 9.0 x 100000 6/ft were tested at Mach numbers less than 2.0 while testing at Mach 2.0 and 3.0 was conducted at a Reynolds number of 11.0 x 100000/ft. Eleven deflections of 0 deg, +15 deg, -20, deg and -40 deg; body flap deflections of 0 deg, +13.75 deg and -14.25 deg; and rudder flare angles of 24.92 deg and 54.92 deg were tested on the -147B Orbiter over the entire Mach number range. Testing of the -139B Orbiter was for data verification and configuration comparison purposes only.

  4. Qualitative evaluation of a flush air data system at transonic speeds and high angles of attack

    NASA Technical Reports Server (NTRS)

    Larson, Terry J.; Whitmore, Stephen A.; Ehernberger, L. J.; Johnson, J. Blair; Siemers, Paul M., III

    1987-01-01

    Flight tests were performed on an F-14 aircraft to evaluate the use of flush pressure orifices on the nose section for obtaining air data at transonic speeds over a large range of flow angles. This program was part of a flight test and wind tunnel program to assess the accuracies of such systems for general use on aircraft. It also provided data to validate algorithms developed for the shuttle entry air data system designed at NASA Langley. Data were obtained for Mach numbers between 0.60 and 1.60, for angles of attack up to 26.0 deg, and for sideslip angles up to 11.0 deg. With careful calibration, a flush air data system with all flush orifices can provide accurate air data information over a large range of flow angles. Several orificies on the nose cap were found to be suitable for determination of stagnation pressure. Other orifices on the nose section aft of the nose cap were shown to be suitable for determination of static pressure. Pairs of orifices on the nose cap provided the most sensitive measurements for determining angles of attack and sideslip, although orifices located farther aft on the nose section could also be used.

  5. Implementation of the WICS Wall Interference Correction System at the National Transonic Facility

    NASA Technical Reports Server (NTRS)

    Iyer, Venkit; Everhart, Joel L.; Bir, Pamela J.; Ulbrich, Norbert

    2000-01-01

    The Wall Interference Correction System (WICS) is operational at the National Transonic Facility (NTF) of NASA Langley Research Center (NASA LaRC) for semispan and full span tests in the solid wall (slots covered) configuration. The method is based on the wall pressure signature method for computing corrections to the measured parameters. It is an adaptation of the WICS code operational at the 12 ft pressure wind tunnel (12ft PWT) of NASA Ames Research Center (NASA ARC). This paper discusses the details of implementation of WICS at the NTF including tunnel calibration, code modifications for tunnel and support geometry, changes made for the NTF wall orifices layout, details of interfacing with the tunnel data processing system, and post-processing of results. Example results of applying WICS to a semispan test and a full span test are presented. Comparison with classical correction results and an analysis of uncertainty in the corrections are also given. As a special application of the code, the Mach number calibration data from a centerline pipe test was computed by WICS. Finally, future work for expanding the applicability of the code including online implementation is discussed.

  6. A study of multi-body aerodynamic interference at transonic Mach numbers

    NASA Technical Reports Server (NTRS)

    Cottrell, Charles J.; Martinez, Agusto; Chapman, Gary T.

    1987-01-01

    A wind tunnel experiment involving single, double, and triple combinations of mutually interfering generic, unfinned aircraft stores has been conducted. Each combination of stores was tested at Mach numbers from 0.60 to 1.20 and at angles of attack from 0 to 25 deg for the single store and from 0 to 6 deg for the double and triple store configurations. Extensive axial and circumferential pressure and flow visualization data at each store location were obtained. Euler solutions for each configuration at 0 deg incidence have been generated and compared with experimental data. This comparison indicates an Euler flow solver can yield accurate predictions of the location and magnitude of multibody interference provided an appropriate grid is used and the viscous effects associated with these configurations remain small. The data indicate multibody interference in the transonic region increases as the freestream Mach number approaches 1 from either direction, and subsides as the Mach number moves away from sonic conditions. This interference is characterized by a large, localized reduction in pressure on the inboard surfaces of the bodies which results in forces that draw the configuration closer together.

  7. Numerical simulation of the tip vortex off a low-aspect-ratio wing at transonic speed

    NASA Technical Reports Server (NTRS)

    Mansour, N. N.

    1984-01-01

    The viscous transonic flow around a low aspect ratio wing was computed by an implicit, three dimensional, thin-layer Navier-Stokes solver. The grid around the geometry of interest is obtained numerically as a solution to a Dirichlet problem for the cube. A low aspect ratio wing with large sweep, twist, taper, and camber is the chosen geometry. The topology chosen to wrap the mesh around the wing with good tip resolution is a C-O type mesh. The flow around the wing was computed for a free stream Mach number of 0.82 at an angle of attack of 5 deg. At this Mach number, an oblique shock forms on the upper surface of the wing, and a tip vortex and three dimensional flow separation off the wind surface are observed. Particle path lines indicate that the three dimensional flow separation on the wing surface is part of the roots of the tip vortex formation. The lifting of the tip vortex before the wing trailing edge is observed by following the trajectory of particles release around the wing tip.

  8. Methodology of Blade Unsteady Pressure Measurement in the NASA Transonic Flutter Cascade

    NASA Technical Reports Server (NTRS)

    Lepicovsky, J.; McFarland, E. R.; Capece, V. R.; Jett, T. A.; Senyitko, R. G.

    2002-01-01

    In this report the methodology adopted to measure unsteady pressures on blade surfaces in the NASA Transonic Flutter Cascade under conditions of simulated blade flutter is described. The previous work done in this cascade reported that the oscillating cascade produced waves, which for some interblade phase angles reflected off the wind tunnel walls back into the cascade, interfered with the cascade unsteady aerodynamics, and contaminated the acquired data. To alleviate the problems with data contamination due to the back wall interference, a method of influence coefficients was selected for the future unsteady work in this cascade. In this approach only one blade in the cascade is oscillated at a time. The majority of the report is concerned with the experimental technique used and the experimental data generated in the facility. The report presents a list of all test conditions for the small amplitude of blade oscillations, and shows examples of some of the results achieved. The report does not discuss data analysis procedures like ensemble averaging, frequency analysis, and unsteady blade loading diagrams reconstructed using the influence coefficient method. Finally, the report presents the lessons learned from this phase of the experimental effort, and suggests the improvements and directions of the experimental work for tests to be carried out for large oscillation amplitudes.

  9. Survey of Primary Flow Measurement Parameters at the NASA Langley Transonic Dynamics Tunnel

    NASA Technical Reports Server (NTRS)

    Piatak, David J.

    2003-01-01

    An assessment of the methods and locations used to measure the primary flow conditions in the NASA Langley Transonic Dynamics Tunnel was conducted during calibration activities following the facility conversion from a Freon-12 heavy-gas test medium to R-134a. A survey of stagnation pressure, plenum static pressure, and stagnation temperature was undertaken at many pertinent locations in the settling chamber, plenum, and contraction section of the wind tunnel and these measurements were compared to those of the existing primary flow measurement systems. Local flow velocities were measured in the settling chamber using a pitot probe. Results illustrate that small discrepancies exist between measured primary tunnel flow conditions and the survey measurements. These discrepancies in tunnel stagnation pressure, plenum pressure, and stagnation temperature were found to be approximately +/- 1-3 psf and 2-3 degrees Fahrenheit. The propagation of known instrument errors in measured primary flow conditions and its impact on tunnel Mach number, dynamic pressure, flow velocity, and Reynolds number have been investigated analytically and shown to require careful attention when considering the uncertainty in measured test section conditions.

  10. Measurements of Flow Turbulence in the NASA Langley Transonic Dynamics Tunnel

    NASA Technical Reports Server (NTRS)

    Wiesman, Carol D.; Sleeper, Robert K.

    2005-01-01

    An assessment of the flow turbulence in the NASA Langley Transonic Dynamics Tunnel (TDT) was conducted during calibration activities following the facility conversion from a Freon-12 heavy-gas test medium to an R134a heavy-gas test medium. Total pressure, static pressure, and acoustic pressure levels were measured at several locations on a stingmounted rake. The test measured wall static pressures at several locations although this paper presents only those from one location. The test used two data acquisition systems, one sampling at 1000 Hz and the second sampling at 125 000 Hz, for acquiring time-domain data. This paper presents standard deviations and power spectral densities of the turbulence points throughout the wind tunnel envelope in air and R134a. The objective of this paper is to present the turbulence characteristics for the test section. No attempt is made to assess the causes of the turbulence. The present paper looks at turbulence in terms of pressure fluctuations. Reference 1 looked at tunnel turbulence in terms of velocity fluctuations.

  11. A study of transonic aerodynamic analysis methods for use with a hypersonic aircraft synthesis code

    NASA Technical Reports Server (NTRS)

    Sandlin, Doral R.; Davis, Paul Christopher

    1992-01-01

    A means of performing routine transonic lift, drag, and moment analyses on hypersonic all-body and wing-body configurations were studied. The analysis method is to be used in conjunction with the Hypersonic Vehicle Optimization Code (HAVOC). A review of existing techniques is presented, after which three methods, chosen to represent a spectrum of capabilities, are tested and the results are compared with experimental data. The three methods consist of a wave drag code, a full potential code, and a Navier-Stokes code. The wave drag code, representing the empirical approach, has very fast CPU times, but very limited and sporadic results. The full potential code provides results which compare favorably to the wind tunnel data, but with a dramatic increase in computational time. Even more extreme is the Navier-Stokes code, which provides the most favorable and complete results, but with a very large turnaround time. The full potential code, TRANAIR, is used for additional analyses, because of the superior results it can provide over empirical and semi-empirical methods, and because of its automated grid generation. TRANAIR analyses include an all body hypersonic cruise configuration and an oblique flying wing supersonic transport.

  12. Experimental Investigation of the DLR-F6 Transport Configuration in the National Transonic Facility

    NASA Technical Reports Server (NTRS)

    Gatlin, Gregory M.; Rivers, Melissa B.; Goodliff, Scott L.; Rudnik, Ralf; Sitzmann, Martin

    2008-01-01

    An experimental aerodynamic investigation of the DLR (German Aerospace Center) F6 generic transport configuration has been conducted in the NASA NTF (National Transonic Facility) for CFD validation within the framework of the AIAA Drag Prediction Workshop. Force and moment, surface pressure, model deformation, and surface flow visualization data have been obtained at Reynolds numbers of both 3 million and 5 million. Flow-through nacelles and a side-of-body fairing were also investigated on this wing-body configuration. Reynolds number effects on trailing edge separation have been assessed, and the effectiveness of the side-of-body fairing in eliminating a known region of separated flow has been determined. Data obtained at a Reynolds number of 3 million are presented together for comparison with data from a previous wind tunnel investigation in the ONERA S2MA facility. New surface flow visualization capabilities have also been successfully explored and demonstrated in the NTF for the high pressure and moderately low temperature conditions required in this investigation. Images detailing wing surface flow characteristics are presented.

  13. Adjoint-based airfoil shape optimization in transonic flow

    NASA Astrophysics Data System (ADS)

    Gramanzini, Joe-Ray

    The primary focus of this work is efficient aerodynamic shape optimization in transonic flow. Adjoint-based optimization techniques are employed on airfoil sections and evaluated in terms of computational accuracy as well as efficiency. This study examines two test cases proposed by the AIAA Aerodynamic Design Optimization Discussion Group. The first is a two-dimensional, transonic, inviscid, non-lifting optimization of a Modified-NACA 0012 airfoil. The second is a two-dimensional, transonic, viscous optimization problem using a RAE 2822 airfoil. The FUN3D CFD code of NASA Langley Research Center is used as the ow solver for the gradient-based optimization cases. Two shape parameterization techniques are employed to study their effect and the number of design variables on the final optimized shape: Multidisciplinary Aerodynamic-Structural Shape Optimization Using Deformation (MASSOUD) and the BandAids free-form deformation technique. For the two airfoil cases, angle of attack is treated as a global design variable. The thickness and camber distributions are the local design variables for MASSOUD, and selected airfoil surface grid points are the local design variables for BandAids. Using the MASSOUD technique, a drag reduction of 72.14% is achieved for the NACA 0012 case, reducing the total number of drag counts from 473.91 to 130.59. Employing the BandAids technique yields a 78.67% drag reduction, from 473.91 to 99.98. The RAE 2822 case exhibited a drag reduction from 217.79 to 132.79 counts, a 39.05% decrease using BandAids.

  14. The National Transonic Facility - Status and operational planning

    NASA Technical Reports Server (NTRS)

    Howell, R. R.

    1980-01-01

    The National Transonic Facility is scheduled to be completed in 1982. Several technical concerns remain, including thermal stress constraints which may limit the rate at which temperatures can be changed, seal performance in the cryogenic environment, and the tunnel process controls which could affect data acquisition rate. The current design affords the capability of dealing with all these concerns if they become problems. The outstanding instrument need is a real-time model surface deformation measurement system. Finally, an examination of the occupancy cost and the cost of liquid nitrogen for high Reynolds number tests indicates that operating costs should not be an inhibiting factor in the use of the facility.

  15. Transonic cascade flow prediction using the Navier-Stokes equations

    NASA Technical Reports Server (NTRS)

    Arnone, A.; Stecco, S. S.

    1991-01-01

    This paper presents results which summarize the work carried out during the last three years to improve the efficiency and accuracy of numerical predictions in turbomachinery flow calculations. A new kind of nonperiodic c-type grid is presented and a Runge-Kutta scheme with accelerating strategies is used as a flow solver. The code capability is presented by testing four different blades at different exit Mach numbers in transonic regimes. Comparison with experiments shows the very good reliability of the numerical prediction. In particular, the loss coefficient seems to be correctly predicted by using the well-known Baldwin-Lomax turbulence model.

  16. Airfoil shape and thickness effects on transonic airloads and flutter

    NASA Technical Reports Server (NTRS)

    Bland, S. R.; Edwards, J. W.

    1983-01-01

    A transient pulse technique is used to obtain harmonic forces from a time-marching solution of the complete unsteady transonic small perturbation potential equation. The unsteady pressures and forces acting on a model of the NACA 64A010 conventional airfoil and the MBB A-3 supercritical airfoil over a range of Mach numbers are examined in detail. Flutter calculations at constant angle of attack show a similar flutter behavior for both airfoils, except for a boundary shift in Mach number associated with corresponding Mach number shift in the unsteady aerodynamic forces. Differences in the static aeroelastic twist behavior for the two airfoils are significant.

  17. Airfoil shape and thickness effects on transonic airloads and flutter

    NASA Technical Reports Server (NTRS)

    Bland, S. R.; Edwards, J. W.

    1983-01-01

    A transient pulse technique is used to obtain harmonic forces from a time-marching solution of the complete unsteady transonic small perturbation potential evaluation. The unsteady pressures and forces acting on a model of the NACA 64A010 conventional airfoil and the MBB A-3 supercritical airfoil over a range of Mach numbers are examined in detail. Flutter calculations at constant angle of attack show a similar flutter behavior for both airfoils, except for a boundary shift in Mach number associated with a corresponding Mach number shift in the unsteady aerodynamic forces. Differences in the static aeroelastic twist behavior for the two airfoils are significant.

  18. Shock wave-turbulent boundary layer interactions in transonic flow

    NASA Technical Reports Server (NTRS)

    Adamson, T. C., Jr.; Messiter, A. F.

    1976-01-01

    The method of matched asymptotic expansions is used in analyzing the structure of the interaction region formed when a shock wave impinges on a turbulent flat plate boundary layer in transonic flow. Solutions in outer regions, governed by inviscid flow equations, lead to relations for the wall pressure distribution. Solutions in the inner regions, governed by equations in which Reynolds and/or viscous stresses are included, lead to a relation for the wall shear stress. Solutions for the wall pressure distribution are reviewed for both oblique and normal incoming shock waves. Solutions for the wall shear stress are discussed.

  19. Development of a nonlinear unsteady transonic flow theory

    NASA Technical Reports Server (NTRS)

    Stahara, S. S.; Spreiter, J. R.

    1973-01-01

    A nonlinear, unsteady, small-disturbance theory capable of predicting inviscid transonic flows about aerodynamic configurations undergoing both rigid body and elastic oscillations was developed. The theory is based on the concept of dividing the flow into steady and unsteady components and then solving, by method of local linearization, the coupled differential equation for unsteady surface pressure distribution. The equations, valid at all frequencies, were derived for two-dimensional flows, numerical results, were obtained for two classses of airfoils and two types of oscillatory motions.

  20. Calculations of transonic boattail flow at small angle of attack

    NASA Technical Reports Server (NTRS)

    Nakayama, A.; Chow, W. L.

    1979-01-01

    A transonic flow past a boattailed afterbody under a small angle of attack was examined. It is known that the viscous effect offers significant modifications of the pressure distribution on the afterbody. Thus, the formulation for the inviscid flow was based on the consideration of a flow past a nonaxisymmetric body. The full three dimensional potential equation was solved through numerical relaxation, and quasi-axisymmetric boundary layer calculations were performed to estimate the displacement effect. It was observed again that the viscous effects were not negligible. The trend of the final results agreed well with the experimental data.

  1. Viscous transonic flow computation over Space Shuttle configuration

    NASA Technical Reports Server (NTRS)

    Fujii, K.; Kutler, P.

    1984-01-01

    A thin-layer Navier-Stokes code capable of predicting steady-state viscous flows is applied to the transonic flow over a Space Shuttle configuration. The code is written in the generalized coordinate system, and the grid-generation code of Fujii (1983) is used for the discretization of the flow field. The flow-field computation is done using the CRAY 1S computer at NASA Ames. The computed result is physically reasonable, even though no experimental data is available for the comparison purpose.

  2. Transonic Turbulent Flow Predictions With Two-Equation Turbulence Models

    NASA Technical Reports Server (NTRS)

    Liou, William W.; Shih, Tsan-Hsing

    1996-01-01

    Solutions of the Favre-averaged Navier-Stokes equations for two well-documented transonic turbulent flows are compared in detail with existing experimental data. While the boundary layer in the first case remains attached, a region of extensive flow separation has been observed in the second case. Two recently developed k-epsilon, two-equation, eddy-viscosity models are used to model the turbulence field. These models satisfy the realizability constraints of the Reynolds stresses. Comparisons with the measurements are made for the wall pressure distribution, the mean streamwise velocity profiles, and turbulent quantities. Reasonably good agreement is obtained with the experimental data.

  3. Aeroelastic Tailoring of Transport Wings Including Transonic Flutter Constraints

    NASA Technical Reports Server (NTRS)

    Stanford, Bret K.; Wieseman, Carol D.; Jutte, Christine V.

    2015-01-01

    Several minimum-mass optimization problems are solved to evaluate the effectiveness of a variety of novel tailoring schemes for subsonic transport wings. Aeroelastic stress and panel buckling constraints are imposed across several trimmed static maneuver loads, in addition to a transonic flutter margin constraint, captured with aerodynamic influence coefficient-based tools. Tailoring with metallic thickness variations, functionally graded materials, balanced or unbalanced composite laminates, curvilinear tow steering, and distributed trailing edge control effectors are all found to provide reductions in structural wing mass with varying degrees of success. The question as to whether this wing mass reduction will offset the increased manufacturing cost is left unresolved for each case.

  4. Rotor wake characteristics of a transonic axial flow fan

    NASA Technical Reports Server (NTRS)

    Hathaway, M. D.; Gertz, J.; Epstein, A.; Strazisar, A. J.

    1985-01-01

    State of the art turbomachinery flow analysis codes are not capable of predicting the viscous flow features within turbomachinery blade wakes. Until efficient 3D viscous flow analysis codes become a reality there is therefore a need for models which can describe the generation and transport of blade wakes and the mixing process within the wake. To address the need for experimental data to support the development of such models, high response pressure measurements and laser anemometer velocity measurements were obtained in the wake of a transonic axial flow fan rotor.

  5. Transonics and fighter aircraft: Challenges and opportunities for CFD

    NASA Technical Reports Server (NTRS)

    Miranda, Luis R.

    1989-01-01

    The application of computational fluid dynamics (CFD) to fighter aircraft design and development is discussed. Methodology requirements for the aerodynamic design of fighter aircraft are briefly reviewed. The state-of-the-art of computational methods for transonic flows in the light of these requirements is assessed and the techniques found most adequate for the subject application are identified. Highlights from some proof-of-feasibility Euler and Navier-Stokes computations about a complete fighter aircraft configuration are presented. Finally, critical issues and opportunities for design application of CFD are discussed.

  6. Calculation of unsteady transonic aerodynamics for oscillating wings with thickness

    NASA Technical Reports Server (NTRS)

    Ruo, S. Y.; Theisen, J. G.

    1975-01-01

    An analytical approach is presented to account for some of the nonlinear characteristics of the transonic flow equation for finite thickness wings undergoing harmonic oscillation at sonic flight speed in an inviscid, shock-free fluid. The thickness effect is accounted for in the analysis through use of the steady local Mach number distribution over the wing at its mean position by employing the local linearization concept and a coordinate transformation. Computed results are compared with that of the linearized theory and experiments. Based on the local linearization concept, an alternate formulation avoiding the limitations of the coordinate transformation method is presented.

  7. Calculation of unsteady transonic flows using the integral equation method

    NASA Technical Reports Server (NTRS)

    Nixon, D.

    1978-01-01

    The basic integral equations for a harmonically oscillating airfoil in a transonic flow with shock waves are derived; the reduced frequency is assumed to be small. The problems associated with shock wave motion are treated using a strained coordinate system. The integral equation is linear and consists of both line integrals and surface integrals over the flow field which are evaluated by quadrature. This leads to a set of linear algebraic equations that can be solved directly. The shock motion is obtained explicitly by enforcing the condition that the flow is continuous except at a shock wave. Results obtained for both lifting and nonlifting oscillatory flows agree satisfactorily with other accurate results.

  8. Subsonic/transonic stall flutter investigation of a rotating rig

    NASA Technical Reports Server (NTRS)

    Jutras, R. R.; Fost, R. B.; Chi, R. M.; Beacher, B. F.

    1981-01-01

    Stall flutter is investigated by obtaining detailed quantitative steady and aerodynamic and aeromechanical measurements in a typical fan rotor. The experimental investigation is made with a 31.3 percent scale model of the Quiet Engine Program Fan C rotor system. Both subsonic/transonic (torsional mode) flutter and supersonic (flexural) flutter are investigated. Extensive steady and unsteady data on the blade deformations and aerodynamic properties surrounding the rotor are acquired while operating in both the steady and flutter modes. Analysis of this data shows that while there may be more than one traveling wave present during flutter, they are all forward traveling waves.

  9. Highlights of experience with a flexible walled test section in the NASA Langley 0.3-meter transonic cryogenic tunnel

    NASA Technical Reports Server (NTRS)

    Wolf, Stephen W. D.; Ray, Edward J.

    1988-01-01

    The unique combination of adaptive wall technology with a contonuous flow cryogenic wind tunnel is described. This powerful combination allows wind tunnel users to carry out 2-D tests at flight Reynolds numbers with wall interference essentially eliminated. Validation testing was conducted to support this claim using well tested symmetrical and cambered airfoils at transonic speeds and high Reynolds numbers. The test section hardware has four solid walls, with the floor and ceiling flexible. The method of adapting/shaping the floor and ceiling to eliminate top and bottom wall interference at its source is outlined. Data comparisons for different size models tested and others in several sophisticated 2-D wind tunnels are made. In addition, the effects of Reynolds number, testing at high lift with associated large flexible wall movements, the uniqueness of the adapted wall shapes, and the effects of sidewall boundary layer control are examined. The 0.3-m TCT is now the most advanced 2-D research facility anywhere.

  10. Flow visualization in the Langley 0.3-meter Transonic Cryogenic Tunnel and preliminary plans for the National Transonic Facility

    NASA Technical Reports Server (NTRS)

    Rhodes, D. B.; Jones, S. B.

    1982-01-01

    Design problems associated with the integration of flow visualization in cryogenic facilities are discussed. The possible effects from the cryogenic environment (i.e., window distortion due to thermal contraction both in the mounts and in the window material itself and turbulence in the flow due to injected LN2) are examined. The flow visualization techniques studied are schlieren, shadowgraph, moire deflectometry, and holographic interferometry. The test beds for this work are a Langley in-house cryogenic test chamber and the 0.3-Meter Transonic Cryogenic Tunnel.

  11. Evaluation of a flexible wall testing technique to minimize wall interferences in the NASA Langley 0.3-m Transonic Cryogenic Tunnel

    NASA Technical Reports Server (NTRS)

    Wolf, Stephen W. D.

    1988-01-01

    Free air simulations in conventional transonic wind tunnels require improvement which adaptive wall testing techniques can provide primarily by minimizing wall interferences. In addition, these techniques offer other substantial advantages such as increased Reynolds number capability and a reduction in tunnel drive power. The combination of an adaptive wall test section with a continuous flow cryogenic wind tunnel is unique. The test section has four solid walls with two flexible walls mounted between rigid sidewalls. This modification of an existing major facility stresses the practicalities of the testing technique. These practicalities were evaluated in terms of flexible wall test section design and operation. Increased hardware and operating complexity of the new test section is offset by a significant improvement in real-time data accuracy in two-dimensional testing. Validation testing has expanded the experience with flexible walled test sections into the realms of flight Reynolds numbers and high lift. Data accuracy has been assessed with regard to test section geometry and operating tolerances. The successful evaluation of the testing technique in two-dimensional testing has proved that a production type operation is possible, with suitable control system and test section design. The 0.3-m Transonic Cryogenic Tunnel with an adaptive wall test section currently represents the most advanced two-dimensional facility anywhere.

  12. Possible safety hazards associated with the operation of the 0.3-m transonic cryogenic tunnel at the NASA Langley Research Center

    NASA Technical Reports Server (NTRS)

    Webster, T. J.

    1982-01-01

    The 0.3 m Transonic Cryogenic Tunnel (TCT) at the NASA Langley Research Center was built in 1973 as a facility intended to be used for no more than 60 hours in order to verify the validity of the cryogenic wind tunnel concept at transonic speeds. The role of the 0.3 m TCT has gradually changed until now, after over 3000 hours of operation, it is classified as a major NASA research facility and, under the administration of the Experimental Techniques Branch, it is used extensively for the testing of airfoils at high Reynolds numbers and for the development of various technologies related to the efficient operation and use of cryogenic wind tunnels. The purpose of this report is to document the results of a recent safety analysis of the 0.3 m TCT facility. This analysis was made as part of an on going program with the Experimental Techniques Branch designed to ensure that the existing equipment and current operating procedures of the 0.3 m TCT facility are acceptable in terms of today's standards of safety for cryogenic systems.

  13. Results for the hybrid laminar flow control experiment conducted in the NASA Langley 8-foot transonic pressure tunnel on a 7-foot chord model

    NASA Technical Reports Server (NTRS)

    Bobbitt, Percy J.; Ferris, James C.; Harvey, William D.; Goradia, Suresh H.

    1992-01-01

    A description is given of the development of, and results from, the hybrid laminar flow control (HLFC) experiment conducted in the NASA LaRC 8 ft Transonic Pressure Tunnel on a 7 ft chord, 23 deg swept model. The methods/codes used to obtain the contours of the HLFC model surface and to define the suction requirements are outlined followed by a discussion of the model construction, suction system, instrumentation, and some example results from the wind tunnel tests. Included in the latter are the effects of Mach number, suction level, and the extent of suction. An assessment is also given of the effect of the wind tunnel environment on the suction requirements. The data show that, at or near the design Mach number, large extents of laminar flow can be achieved with suction mass flows over the first 25 percent, or less, of the chord. Top surface drag coefficients with suction extending from the near leading edge to 20 percent of the chord were approximately 40 percent lower than those obtained with no suction. The results indicate that HLFC can be designed for transonic speeds with lift and drag coefficients approaching those of LFC designs but with much smaller extents and levels of suction.

  14. Computations of Internal and External Axisymmetric Nozzle Aerodynamics at Transonic Speeds

    NASA Technical Reports Server (NTRS)

    Dalbello, Teryn; Georgiadis, Nicholas; Yoder, Dennis; Keith, Theo

    2003-01-01

    Computational Fluid Dynamics (CFD) analyses of axisymmetric circular-arc boattail nozzles have been completed in support of NASA's Next Generation Launch Technology Program to investigate the effects of high-speed nozzle geometries on the nozzle internal flow and the surrounding boattail regions. These computations span the very difficult transonic flight regime, with shock-induced separations and strong adverse pressure gradients. External afterbody and internal nozzle pressure distributions computed with the Wind code are compared with experimental data. A range of turbulence models were examined in Wind, including an Explicit Algebraic Stress model (EASM). Computations on two nozzle geometries have been completed at freestream Mach numbers ranging from 0.6 to 0.9, driven by nozzle pressure ratios (NPR) ranging from 2.9 to 5. Results obtained on converging-only geometry indicate reasonable agreement to experimental data, with the EASM and Shear Stress Transport (SST) turbulence models providing the best agreement. Calculations completed on a converging-diverging geometry involving large-scale internal flow separation did not converge to a true steady-state solution when run with variable timestepping (steady-state). Calculations obtained using constant timestepping (time-accurate) indicate less variations in flow properties compared with steady-state solutions. This failure to converge to a steady-state solution was found to be the result of difficulties in using variable time-stepping with large-scale separations present in the flow. Nevertheless, time-averaged boattail surface pressure coefficient and internal nozzle pressures show fairly good agreement with experimental data. The SST turbulence model demonstrates the best over-all agreement with experimental data.

  15. Flow Visualization and Laser Velocimetry for Wind Tunnels

    NASA Technical Reports Server (NTRS)

    Hunter, W. W., Jr. (Editor); Foughner, J. T., Jr. (Editor)

    1982-01-01

    The need for flow visualization and laser velocimetry were discussed. The purpose was threefold: (1) provide a state-of-the-art overview; (2) provide a forum for industry, universities, and government agencies to address problems in developing useful and productive flow visualization and laser velocimetry measurement techniques; and (3) provide discussion of recent developments and applications of flow visualization and laser velocimetry measurement techniques and instrumentation systems for wind tunnels including the 0.3-Meter Transonic Cryogenic Tunnel.

  16. Viscous three-dimensional calculations of transonic fan performance

    NASA Technical Reports Server (NTRS)

    Chima, Rodrick V.

    1992-01-01

    A 3-D flow analysis code was used to compute the design speed operating line of a transonic fan rotor, and the results were compared with experimental data. The code is an explicit finite difference code with an algebraic turbulence model. The transonic fan, called Rotor 67, was tested experimentally at NASA Lewis conventional aerodynamic probes and with user anemometry and was included as one of the AGARD test cases for the computation of internal flows. The experimental data are described. Maps of total pressure ratio and adiabatic efficiency vs mass flow were computed and are compared with the experimental maps, with good agreement. Detailed comparisons between calculations and experiment are made at two operating points, one near peak efficiency and the other near stall. Blade-to-blade contour plots are used to show the shock structure. Comparisons of circumferentially integrated flow quantities downstream of the rotor show spanwise distributions of several aerodynamic parameters. Calculated Mach number distributions are compared with laser anemometer data within the blade row and the wake to quantify the accuracy of the calculations. Particle traces are used to show the nature of secondary flow.

  17. Transonic Drag Prediction Using an Unstructured Multigrid Solver

    NASA Technical Reports Server (NTRS)

    Mavriplis, D. J.; Levy, David W.

    2001-01-01

    This paper summarizes the results obtained with the NSU-3D unstructured multigrid solver for the AIAA Drag Prediction Workshop held in Anaheim, CA, June 2001. The test case for the workshop consists of a wing-body configuration at transonic flow conditions. Flow analyses for a complete test matrix of lift coefficient values and Mach numbers at a constant Reynolds number are performed, thus producing a set of drag polars and drag rise curves which are compared with experimental data. Results were obtained independently by both authors using an identical baseline grid and different refined grids. Most cases were run in parallel on commodity cluster-type machines while the largest cases were run on an SGI Origin machine using 128 processors. The objective of this paper is to study the accuracy of the subject unstructured grid solver for predicting drag in the transonic cruise regime, to assess the efficiency of the method in terms of convergence, cpu time, and memory, and to determine the effects of grid resolution on this predictive ability and its computational efficiency. A good predictive ability is demonstrated over a wide range of conditions, although accuracy was found to degrade for cases at higher Mach numbers and lift values where increasing amounts of flow separation occur. The ability to rapidly compute large numbers of cases at varying flow conditions using an unstructured solver on inexpensive clusters of commodity computers is also demonstrated.

  18. Flow Disturbance Measurements in the National Transonic Facility

    NASA Technical Reports Server (NTRS)

    King, Rudolph A.; Andino, Marlyn Y.; Melton, Latunia; Eppink, Jenna; Kegerise, Michael A.

    2013-01-01

    Recent flow measurements have been acquired in the National Transonic Facility to assess the test-section unsteady flow environment. The primary purpose of the test is to determine the feasibility of the facility to conduct laminar-flow-control testing and boundary-layer transition-sensitive testing at flight-relevant operating conditions throughout the transonic Mach number range. The facility can operate in two modes, warm and cryogenic test conditions for testing full and semispan-scaled models. Data were acquired for Mach and unit Reynolds numbers ranging from 0.2 less than or equal to M less than or equal to 0.95 and 3.3 × 10(exp 6) less than Re/m less than 220×10(exp 6) collectively at air and cryogenic conditions. Measurements were made in the test section using a survey rake that was populated with 19 probes. Roll polar data at selected conditions were obtained to look at the uniformity of the flow disturbance field in the test section. Data acquired included mean total temperatures, mean and fluctuating static/total pressures, and mean and fluctuating hot-wire measurements. This paper focuses primarily on the unsteady pressure and hot-wire results. Based on the current measurements and previous data, an assessment was made that the facility may be a suitable facility for ground-based demonstrations of laminar-flow technologies at flight-relevant conditions in the cryogenic mode.

  19. Cavity Unsteady-Pressure Measurements at Subsonic and Transonic Speeds

    NASA Technical Reports Server (NTRS)

    Tracy, Maureen B.; Plentovich, E. B.

    1997-01-01

    An experimental investigation was conducted in the Langley 8-Foot Transonic Pressure Tunnel to determine the flow characteristics of rectangular cavities with varying relative dimensions at subsonic and transonic speeds. Cavities were tested with width-to-depth ratios of 1, 4, 8, and 16 for length-to-depth ratios l/h of 1 through 17.5. The maximum cavity depth was 2.4 in., and the turbulent boundary layer approaching the cavity was approximately 0.5 in. thick. Unsteady- and mean static-pressure measurements were made at free-stream Mach numbers from 0.20 to 0.95 at a unit Reynolds number per foot of approximately 3 x 10(exp 6); however, only unsteady-pressure results are presented in this paper. Results indicate that as l/h increases, cavity flows changed from resonant to nonresonant with resonant amplitudes decreasing gradually. Resonant spectra are obtained largely in cavities with mean static-pressure distributions characteristic of open and transitional flows. Resonance sometimes occurred for closed flow. Increasing cavity width or decreasing cavity depth while holding l/h fixed had the effect of increasing resonant amplitudes and sometimes induced resonance. The effects due to changes in width are more pronounced. Decreasing Mach number has the effect of broadening the resonances.

  20. Flow Disturbance Characterization Measurements in the National Transonic Facility

    NASA Technical Reports Server (NTRS)

    King, Rudolph A.; Andino, Marlyn Y.; Melton, Latunia; Eppink, Jenna; Kegerise, Michael A.; Tsoi, Andrew

    2012-01-01

    Recent flow measurements have been acquired in the National Transonic Facility (NTF) to assess the unsteady flow environment in the test section. The primary purpose of the test is to determine the feasibility of the NTF to conduct laminar-flow-control testing and boundary-layer transition sensitive testing. The NTF can operate in two modes, warm (air) and cold/cryogenic (nitrogen) test conditions for testing full and semispan scaled models. The warm-air mode enables low to moderately high Reynolds numbers through the use of high tunnel pressure, and the nitrogen mode enables high Reynolds numbers up to flight conditions, depending on aircraft type and size, utilizing high tunnel pressure and cryogenic temperatures. NASA's Environmentally Responsible Aviation (ERA) project is interested in demonstrating different laminar-flow technologies at flight-relevant operating conditions throughout the transonic Mach number range and the NTF is well suited for the initial ground-based demonstrations. Roll polar data at selected test conditions were obtained to look at the uniformity of the flow disturbance field in the test section. Data acquired from the rake probes included mean total temperatures, mean and fluctuating static/total pressures, and mean and fluctuating hot-wire measurements. . Based on the current measurements and previous data, an assessment was made that the NTF is a suitable facility for ground-based demonstrations of laminar-flow technologies at flight-relevant conditions in the cryogenic mode.

  1. Shock Excursion in Transonic Convex-Corner Flows

    NASA Astrophysics Data System (ADS)

    Chung, Kung-Ming; Chang, Po-Hsiung; Chang, Keh-Chin

    2016-06-01

    A typical feature on a transonic airfoil is associated with a quasi-normal shock on the upper surface. When a shock is strong enough, development of a separation bubble changes the flowfield significantly. High intensity of pressure fluctuations is associated with unsteady shock wave. An experimental study was conducted to investigate transonic convex-corner flows, including single- and round-convex corners. Peak pressure fluctuations and zero-crossing frequency are associated with incoming Mach number, total turning angle, and separation length. In the cases of R100 and R200, the level of zero-crossing frequency decreases with increasing turning angle as observed in the single convex-corner flows. An increase in the level of pressure fluctuations corresponds to lower zero-cross frequency. In the cases of R300, there is a roughly constant level of surface pressure fluctuations or the variation of pressure fluctuations with the turning angle is less significant. The zero-crossing frequency would be independent of turning angle in the cases of R300.

  2. Transonic analysis and design of axisymmetric bodies in nonuniform flow

    NASA Technical Reports Server (NTRS)

    Chang, Jen-Fu; Lan, C. Edward

    1987-01-01

    An inviscid nonuniform axisymmetric transonic code was developed for applications in analysis and design. Propfan slipstream effect on pressure distribution for a body with and without sting was investigated. Results show that nonuniformity causes pressure coefficient to be more negative and shock strength to be stronger and more rearward. Sting attached to a body reduced the pressure peak and moves the rear shock forward. Extent and Mach profile shapes of the nonuniformity region appeared to have little effect on the pressure distribution. Increasing nonuniformity magnitude made pressure coefficient more negative and moved the shock rearward. Design study was conducted with the CONMIN optimizer for an ellipsoid and a body with the NACA-0012 counter. For the ellipsoid, the general trend showed that to reduce the pressure drag, the front portion of the body should be thinner and the contour of the rear portion should be flatter than the ellipsoid. For the design of a body with a sharp trailing edge in transonic flow with an initial shape given by the NACA-0012 contour, the pressure drag was reduced by decreasing the nose radius and increasing the thickness in the aft portion. Drag reduction percentages are given.

  3. Quasi-normal acoustic oscillations in the transonic Bondi flow

    NASA Astrophysics Data System (ADS)

    Chaverra, Eliana; Sarbach, Olivier

    2016-01-01

    We analyze the dynamics of nonspherical acoustic perturbations of the transonic Bondi flow, describing the steady radial accretion of a polytropic perfect fluid into a gravity center. The propagation of such perturbations can be described by a wave equation on the curved effective background geometry determined by the acoustic metric introduced by Unruh in the context of experimental black hole evaporation. We show that for the transonic Bondi flow, Unruh's acoustic metric describes an analogue black hole and that the acoustic perturbations undergo quasi-normal oscillations. The associated quasi-normal frequencies are computed and they are proven to scale like the surface gravity of the acoustic black hole. This provides an explanation for results given in an earlier work, where it was shown that the acoustic perturbations of a relativistic fluid accreted by a nonrotating black hole possess quasi-normal modes, and where it was found empirically that the associated frequencies scaled like the surface gravity of the analogue black hole in the limit where the radius of the sonic horizon is much larger than the Schwarzschild radius.

  4. Global convergence of inexact Newton methods for transonic flow

    NASA Technical Reports Server (NTRS)

    Young, David P.; Melvin, Robin G.; Bieterman, Michael B.; Johnson, Forrester T.; Samant, Satish S.

    1990-01-01

    In computational fluid dynamics, nonlinear differential equations are essential to represent important effects such as shock waves in transonic flow. Discretized versions of these nonlinear equations are solved using iterative methods. In this paper an inexact Newton method using the GMRES algorithm of Saad and Schultz is examined in the context of the full potential equation of aerodynamics. In this setting, reliable and efficient convergence of Newton methods is difficult to achieve. A poor initial solution guess often leads to divergence or very slow convergence. This paper examines several possible solutions to these problems, including a standard local damping strategy for Newton's method and two continuation methods, one of which utilizes interpolation from a coarse grid solution to obtain the initial guess on a finer grid. It is shown that the continuation methods can be used to augment the local damping strategy to achieve convergence for difficult transonic flow problems. These include simple wings with shock waves as well as problems involving engine power effects. These latter cases are modeled using the assumption that each exhaust plume is isentropic but has a different total pressure and/or temperature than the freestream.

  5. Transonic Flow Around Swept Wings: Revisiting Von Karman's Similarity Rule

    NASA Astrophysics Data System (ADS)

    Kirkman, Jeffrey J.

    Modern aircraft are expected to fly faster and more efficiently than their predecessors. To improve aerodynamic efficiency, designers must carefully consider and handle shock wave formation. Presently, many designers utilize computationally heavy optimization methods to design wings. While these methods may work, they do not provide insight. This thesis aims to better understand fundamental methods that govern wing design. In order to further understand the flow in the transonic regime, this work revisits the Transonic Similarity Rule. This rule postulates an equivalent incompressible geometry to any high speed geometry in flight and postulates a "stretching" analogy. This thesis utilizes panel methods and Computational Fluid Dynamics (CFD) to show that the "stretching" analogy is incorrect, but instead the flow is transformed by a nonlinear "scaling" of the flow velocity. This work also presents data to show the discrepancies between many famous authors in deriving the accurate Critical Pressure Coefficient (Cp*) equation for both swept and unswept wing sections. The final work of the thesis aims to identify the correct predictive methods for the Critical Pressure Coefficient.

  6. Eulerian-Lagrangian Simulations of Transonic Flutter Instabilities

    NASA Technical Reports Server (NTRS)

    Bendiksen, Oddvar O.

    1994-01-01

    This paper presents an overview of recent applications of Eulerian-Lagrangian computational schemes in simulating transonic flutter instabilities. This approach, the fluid-structure system is treated as a single continuum dynamics problem, by switching from an Eulerian to a Lagrangian formulation at the fluid-structure boundary. This computational approach effectively eliminates the phase integration errors associated with previous methods, where the fluid and structure are integrated sequentially using different schemes. The formulation is based on Hamilton's Principle in mixed coordinates, and both finite volume and finite element discretization schemes are considered. Results from numerical simulations of transonic flutter instabilities are presented for isolated wings, thin panels, and turbomachinery blades. The results suggest that the method is capable of reproducing the energy exchange between the fluid and the structure with significantly less error than existing methods. Localized flutter modes and panel flutter modes involving traveling waves can also be simulated effectively with no a priori knowledge of the type of instability involved.

  7. Laser velocimetry applied to transonic and supersonic aerodynamics

    NASA Technical Reports Server (NTRS)

    Johnson, D. A.; Bachalo, W. D.; Moddaress, D.

    1976-01-01

    As a further demonstration of the capabilities of laser velocity in compressible aerodynamics, measurements obtained in a Mach 2.9 separated turbulent boundary layer and in the transonic flow past a two-dimensional airfoil section are presented and compared to data realized by conventional techniques. In the separated-flow study, the comparisons were made against pitot-static pressure data. Agreement in mean velocities was realized where the pressure measurements could be considered reliable; however, in regions of instantaneous reverse velocities, the laser results were found to be consistent with the physics of the flow whereas the pressure data were not. The laser data obtained in regions of extremely high turbulence suggest that velocity biasing does not occur if the particle occurrence rate is low relative to the turbulent fluctuation rate. Streamwise turbulence intensities are also presented. In the transonic airfoil study, velocity measurements obtained immediately outside the upper surface boundary layer of a 6-inch chord MACA 64A010 airfoil are compared to edge velocities inferred from surface pressure measurements. For free-stream Mach numbers of 0.6 and 0.8, the agreement in results was very good. Dual scatter optical arrangements in conjunction with a single particle, counter-type signal processor were employed in these investigations. Half-micron-diameter polystyrene spheres and naturally occurring condensed oil vapor acted as light scatterers in the two respective flows. Bragg-cell frequency shifting was utilized in the separated flow study.

  8. Transonic blade-vortex interactions noise: A parametric study

    NASA Technical Reports Server (NTRS)

    Lyrintzis, A. S.; Xue, Y.

    1990-01-01

    Transonic Blade-Vortex Interactions (BVI) are simulated numerically and the noise mechanisms are investigated. The 2-D high frequency transonic small disturbance equation is solved numerically (VTRAN2 code). An Alternating Direction Implicit (ADI) scheme with monotone switches is used; viscous effects are included on the boundary and the vortex is simulated by the cloud-in-cell method. The Kirchoff method is used for the extension of the numerical 2-D near field aerodynamic results to the linear acoustic 3-D far field. The viscous effect (shock/boundary layer interaction) on BVI is investigated. The different types of shock motion are identified and compared. Two important disturbances with different directivity exist in the pressure signal and are believed to be related to the fluctuating lift and drag forces. Noise directivity for different cases is shown. The maximum radiation occurs at an angle between 60 and 90 deg below the horizontal for an airfoil fixed coordinate system and depends on the details of the airfoil shape. Different airfoil shapes are studied and classified according to the BVI noise produced.

  9. Transonic swept wings studied by the lifting-line theory

    NASA Technical Reports Server (NTRS)

    Cheng, H. K.; Meng, S. Y.; Chow, R.; Smith, R. C.

    1981-01-01

    Transonic swept wings are analyzed as a lifting-line problem under a small-disturbance approximation. Basic concepts and principal results of the asymptotic theory are discussed. The study focuses on straight oblique wings and V-shaped swept wings, of which the local centerline curvature can be equated to zero. The three-dimensional (3-D) perturbation of the nonlinear component flow admits a similarity flow structure but requires that all wing sections are generated from a single airfoil profile; the reduced 2-D problems in this case are solved only once for all span stations. Examples of solutions involving high subcritical and slightly supercritical component flows are demonstrated and compared with surface pressure data from 3-D computer codes based on the full-potential equation (FLO 22). Except in the neighborhood of leading edges, where the small-disturbance assumption breaks down, and in the vicinities of wing tips and the symmetry plane, where neither the theory nor the 3-D codes may claim full validity, reasonable agreement is consistently found. The explicit results from the upwash analysis, along with the similarity flow structure, provides a rational approach to the control of 3-D effects in transonic aerodynamic design studies.

  10. Two-Fluid Equilibrium for Transonic Poloidal Flows

    NASA Astrophysics Data System (ADS)

    Guazzotto, Luca; Betti, Riccardo

    2012-03-01

    Much analytical and numerical work has been done in the past on ideal MHD equilibrium in the presence of macroscopic flow. In recent years, several authors have worked on equilibrium formulations for a two-fluid system, in which inertial ions and massless electrons are treated as distinct fluids. In this work, we present our approach to the formulation of the two-fluid equilibrium problem. Particular attention is given to the relation between the two-fluid equations and the equilibrium equations for the single-fluid ideal MHD system. Our purpose is to reconsider the results of one-fluid calculation with the more accurate two-fluid model, referring in particular to the so-called transonic discontinuities, which occur when the poloidal velocity spans a range crossing the poloidal sound speed (i.e., the sound speed reduced by a factor Bp/B). It is expected that the one-fluid discontinuity will be resolved into a sharp gradient region by the two-fluid model. Also, contrary to the ideal MHD case, in the two-fluid model the equations governing the equilibrium are elliptic in the whole range of interest for transonic equilibria. The numerical solution of the two-fluid system of equations is going to be based on a code built on the structure of the existing ideal-MHD code FLOW.

  11. Viscous three-dimensional calculations of transonic fan performance

    NASA Technical Reports Server (NTRS)

    Chima, Rodrick V.

    1991-01-01

    A 3-D flow analysis code was used to compute the design speed operating line of a transonic fan rotor, and the results were compared with experimental data. The code is an explicit finite difference code with an algebraic turbulence model. The transonic fan, called rotor 67, was tested experimentally at NASA-Lewis with conventional aerodynamic probes and with user anemometry and was included as one of the AGARD test cases for the computation of internal flows. The experimental data are described. Maps of total pressure ratio and adiabatic efficiency versus mass flow were computed and are compared with the experimental maps, with good agreement. Detailed comparisons between calculations and experiment are made at two operating points, one near peak efficiency and the other near stall. Blade-to-blade contour plots are used to show the shock structure. Comparisons of circumferentially integrated flow quantities downstream of the rotor show spanwise distributions of several aerodynamic parameters. Calculated Mach number distributions are compared with laser anemometer data within the blade row and the wake to quantify the accuracy of the calculations. Particle traces are used to show the nature of secondary flow.

  12. The design of models for cryogenic wind tunnels. [mechanical properties and loads

    NASA Technical Reports Server (NTRS)

    Gillespie, V. P.

    1977-01-01

    Factors to be considered in the design and fabrication of models for cryogenic wind tunnels include high model loads imposed by the high operating pressures, the mechanical and thermodynamic properties of materials in low temperature environments, and the combination of aerodynamic loads with the thermal environment. Candidate materials are being investigated to establish criteria for cryogenic wind tunnel models and their installation. Data acquired from these tests will be provided to users of the National Transonic Facility.

  13. Cryogenic wind tunnels for high Reynolds number testing

    NASA Technical Reports Server (NTRS)

    Lawing, P. L.; Kilgore, R. A.; Mcguire, P. D.

    1986-01-01

    A compilation of lectures presented at various Universities over a span of several years is discussed. A central theme of these lectures has been to present the research facility in terms of the service it provides to, and its potential effect on, the entire community, rather than just the research community. This theme is preserved in this paper which deals with the cryogenic transonic wind tunnels at Langley Research Center. Transonic aerodynamics is a focus both because of its crucial role in determining the success of aeronautical systems and because cryogenic wind tunnels are especially applicable to the transonics problem. The paper also provides historical perspective and technical background for cryogenic tunnels, culminating in a brief review of cryogenic wind tunnel projects around the world. An appendix is included to provide up to date information on testing techniques that have been developed for the cryogenic tunnels at Langley Research Center. In order to be as inclusive and as current as possible, the appendix is less formal than the main body of the paper. It is anticipated that this paper will be of particular value to the technical layman who is inquisitive as to the value of, and need for, cryogneic tunnels.

  14. Numerical studies of porous airfoils in transonic flow. Ph.D. Thesis. Final Report, 1 Jun. 1985 - 31 Aug. 1986

    NASA Technical Reports Server (NTRS)

    Chow, C. Y.

    1986-01-01

    A numerical tool is constructed to examine the effects of a porous surface on transonic airfoil performance and to help understand the flow structure of passive shockwave/boundary layer interactions. The porous region is located near the shock with a cavity underneath it. This study is composed of two parts. Solved in the first part, with an inviscid-flow approach, is the transonic full-potential equation associated with transpiration boundary conditions which are obtained from porosity modeling. The numerical results indicate that a porous airfoil has a wave drag lower than that of a solid airfoil. The observed lambda-shock structure in the wind-tunnel testing can be predicted. Furthermore, the lift could be increased with an appropriate porosity distribution. In the second part of this work, the modified version of either an interactive boundary layer (IBL) algorithm or a thin-layer Navier-Stokes (TLNS) algorithm is used to study the outer flow, while a stream-function formulation is used to model the inner flow in the shallow cavity. The coupling procedure at the porous surface is based on Darcy's law and the assumption of a constant total pressure in the cavity. In addition, a modified Baldwin-Lomax turbulence model is used to describe the transpired turbulent boundary layer in the TLNS approach, while the Cebeci turbulence model is used in the IBL approach. According to the present analysis, a porous surface can reduce the wave drag appreciably, but can also increase the viscous losses. As has been observed experimentally, the numerical results indicate that the total drag is reduced at higher Mach numbers and increased at lower Mach numbers when the angles of attack are small. Furthermore, the streamline pattern of passive shock/boundary layer interaction are revealed.

  15. Unsteady transonic small-disturbance theory including entropy and vorticity effects

    NASA Technical Reports Server (NTRS)

    Batina, John T.

    1988-01-01

    Modifications to unsteady transonic small disturbance theory to include entropy and vorticity effects are presented. The modifications were implemented in the CAP-TSD (Computational Aeroelasticity Program - Transonic Small Disturbance) code. The code permits the aeroelastic analysis of complete aircraft configurations in the flutter critical transonic speed range. Entropy and vorticity effects were incorporated within the solution procedure to more accurately analyze flows with strong shock waves. The modified code includes these effects while retaining the relative simplicity and cost efficiency of the TSD formulation. Detailed descriptions are presented of the entropy and vorticity modifications along with calculated results and comparisons which assess the modified theory. These results are in good agreement with parallel Euler calculations and with experimental data. Therefore, the present method now provides the aeroelastician with an affordable capability to analyze relatively difficult transonic flows without having to solve the computationally more expensive Euler equations.

  16. Numerical computation of transonic flow governed by the full-potential equation

    NASA Technical Reports Server (NTRS)

    Holst, T. L.

    1983-01-01

    Numerical solution techniques for solving transonic flow fields governed by the full potential equation are discussed. In a general sense relaxation schemes suitable for the numerical solution of elliptic partial differential equations are presented and discussed with emphasis on transonic flow applications. The presentation can be divided into two general categories: An introductory treatment of the basic concepts associated with the numerical solution of elliptic partial differential equations and a more advanced treatment of current procedures used to solve the full potential equation for transonic flow fields. The introductory material is presented for completeness and includes a brief introduction (Chapter 1), governing equations (Chapter 2), classical relaxation schemes (Chapter 3), and early concepts regarding transonic full potential equation algorithms (Chapter 4).

  17. Unsteady transonic flow past airfoils in rigid-body motion. [UFLO5

    SciTech Connect

    Chang, I C

    1981-03-01

    With the aim of developing a fast and accurate computer code for predicting the aerodynamic forces needed for a flutter analysis, some basic concepts in computational transonics are reviewed. The unsteady transonic flow past airfoils in rigid body motion is adequately described by the potential flow equation as long as the boundary layer remains attached. The two dimensional unsteady transonic potential flow equation in quasilinear form with first order radiation boundary conditions is solved by an alternating direction implicit scheme in an airfoil attached sheared parabolic coordinate system. Numerical experiments show that the scheme is very stable and is able to resolve the higher nonlinear transonic effects for filter analysis within the context of an inviscid theory.

  18. Unsteady transonic small-disturbance theory including entropy and vorticity effects

    NASA Technical Reports Server (NTRS)

    Batina, John T.

    1988-01-01

    Modifications to unsteady transonic small-disturbance theory to include entropy and vorticity effects are presented. The modifications have been implemented in the CAP-TSD (Computational Aeroelasticity Program - Transonic Small Disturbance) code developed recently at the NASA Langley Research Center. The code permits the aeroelastic analysis of complete aircraft configurations in the flutter critical transonic speed range. Entropy and vorticity effects have been incorporated within the solution procedure to more accurately analyze flows with strong shock waves. The modified code includes these effects while retaining the relative simplicity and cost efficiency of the TSD formulation. The paper presents detailed descriptions of the entropy and vorticity modifications along with calculated results and comparisons which assess the modified theory. These results are in good agreement with parallel Euler calculations and with experimental data. Therefore, the present method now provides the aeroelastician with an affordable capability to analyze relatively difficult transonic flows without having to solve the computationally more expensive Euler equations.

  19. Unsteady transonic flow calculations for two-dimensional canard-wing configurations with aeroelastic applications

    NASA Technical Reports Server (NTRS)

    Batina, J. T.

    1985-01-01

    Unsteady transonic flow calculations for aerodynamically interfering airfoil configurations are performed as a first-step toward solving the three-dimensional canard-wing interaction problem. These calculations are performed by extending the XTRAN2L two-dimensional unsteady transonic small-disturbance code to include an additional airfoil. Unsteady transonic forces due to plunge and pitch motions of a two-dimensional canard and wing are presented. Results for a variety of canard-wing separation distances reveal the effects of aerodynamic interference on unsteady transonic airloads. Aeroelastic analyses employing these unsteady airloads demonstrate the effects of aerodynamic interference on aeroelastic stability and flutter. For the configurations studied, increases in wing flutter speed result with the inclusion of the aerodynamically interfering canard.

  20. Unsteady transonic flow calculations for two-dimensional canard-wing configurations with aeroelastic applications

    NASA Technical Reports Server (NTRS)

    Batina, J. T.

    1985-01-01

    Unsteady transonic flow calculations for aerodynamically interfering airfoil configurations are performed as a first step toward solving the three dimensional canard wing interaction problem. These calculations are performed by extending the XTRAN2L two dimensional unsteady transonic small disturbance code to include an additional airfoil. Unsteady transonic forces due to plunge and pitch motions of a two dimensional canard and wing are presented. Results for a variety of canard wing separation distances reveal the effects of aerodynamic interference on unsteady transonic airloads. Aeroelastic analyses employing these unsteady airloads demonstrate the effects of aerodynamic interference on aeroelastic stability and flutter. For the configurations studied, increases in wing flutter speed result with the inclusion of the aerodynamically interfering canard.

  1. The effect of Reynolds number on transonic compressor blade rotor section

    NASA Astrophysics Data System (ADS)

    Beheshti Amiri, H.; Shahrabi Farahani, A.; Khazaei, H.

    2016-10-01

    In this paper, the effect of Reynolds number on transonic compressor blade rotor section is investigated. After passing through the first transonic compressor stages , the flow becomes remarkably compressed. In the present work, it is intended to numerically investigate the effects of the inflow Reynolds number on the unique incidence, flow losses, deviation angle, and shock position, at three different important points of "Minimum Loss" and "Choked Flow" in started conditions and "Stall Operation" in un-started conditions.

  2. Numerical study on the correlation of transonic single-degree-of-freedom flutter and buffet

    NASA Astrophysics Data System (ADS)

    Gao, ChuanQiang; Zhang, WeiWei; Liu, YiLang; Ye, ZhengYin; Jiang, YueWen

    2015-08-01

    Transonic single-degree-of-freedom (SDOF) flutter and transonic buffet are the typical and complex aeroelastic phenomena in the transonic flow. In this study, transonic aeroelastic issues of an elastic airfoil are investigated using Unsteady Reynolds-Averaged Navier-Stokes (URANS) equations. The airfoil is free to vibrate in SDOF of pitching. It is found that, the coupling system may be unstable and SDOF self-excited pitching oscillations occur in pre-buffet flow condition, where the free-stream angle of attack (AOA) is lower than the buffet onset of a stationary airfoil. In the theory of classical aeroelasticity, this unstable phenomenon is defined as flutter. However, this transonic SDOF flutter is closely related to transonic buffet (unstable aerodynamic models) due to the following reasons. Firstly, the SDOF flutter occurs only when the free-stream AOA of the spring suspended airfoil is slightly lower than that of buffet onset, and the ratio of the structural characteristic frequency to the buffet frequency is within a limited range. Secondly, the response characteristics show a high correlation between the SDOF flutter and buffet. A similar "lock-in" phenomenon exists, when the coupling frequency follows the structural characteristic frequency. Finally, there is no sudden change of the response characteristics in the vicinity of buffet onset, that is, the curve of response amplitude with the free-stream AOA is nearly smooth. Therefore, transonic SDOF flutter is often interwoven with transonic buffet and shows some complex characteristics of response, which is different from the traditional flutter.

  3. Transonic small disturbances equation applied to the solution of two-dimensional nonsteady flows

    NASA Technical Reports Server (NTRS)

    Couston, M.; Angelini, J. J.; Mulak, P.

    1980-01-01

    Transonic nonsteady flows are of large practical interest. Aeroelastic instability prediction, control figured vehicle techniques or rotary wings in forward flight are some examples justifying the effort undertaken to improve knowledge of these problems is described. The numerical solution of these problems under the potential flow hypothesis is described. The use of an alternating direction implicit scheme allows the efficient resolution of the two dimensional transonic small perturbations equation.

  4. Numerical Modeling of Flow Control in a Boundary-Layer-Ingesting Offset Inlet Diffuser at Transonic Mach Numbers

    NASA Technical Reports Server (NTRS)

    Allan Brian G.; Owens, Lewis, R.

    2006-01-01

    This paper will investigate the validation of a NASA developed, Reynolds-averaged Navier-Stokes (RANS) flow solver, OVERFLOW, for a boundary-layer-ingesting (BLI) offset (S-shaped) inlet in transonic flow with passive and active flow control devices as well as the baseline case. Numerical simulations are compared to wind tunnel results of a BLI inlet conducted at the NASA Langley 0.3-Meter Transonic Cryogenic Tunnel. Comparisons of inlet flow distortion, pressure recovery, and inlet wall pressures are performed. The numerical simulations are compared to the BLI inlet data at a freestream Mach number of 0.85 and a Reynolds number of approximately 2 million based on the length of the fan-face diameter. The numerical simulations with and without wind tunnel walls are performed, quantifying effects of the tunnel walls on the BLI inlet flow measurements. The wind tunnel test evaluated several different combinations of jet locations and mass flow rates as well as a vortex generator (VG) vane case. The numerical simulations will be performed on a single jet configuration for varying actuator mass flow rates at a fix inlet mass flow condition. Validation of the numerical simulations for the VG vane case will also be performed for varying inlet mass flow rates. Overall, the numerical simulations were able to predict the baseline circumferential flow distortion, DPCPavg, very well for comparisons made within the designed operating range of the BLI inlet. However the CFD simulations did predict a total pressure recovery that was 0.01 lower than the experiment. Numerical simulations of the baseline inlet flow also showed good agreement with the experimental inlet centerline surface pressures. The vane case showed that the CFD predicted the correct trends in the circumferential distortion for varying inlet mass flow but had a distortion level that was nearly twice as large as the experiment. Comparison to circumferential distortion measurements for a 15 deg clocked 40 probe

  5. Development of a multigrid transonic potential flow code for cascades

    NASA Technical Reports Server (NTRS)

    Steinhoff, John

    1992-01-01

    Finite-volume methods for discretizing transonic potential flow equations have proven to be very flexible and accurate for both two and three dimensional problems. Since they only use local properties of the mapping, they allow decoupling of the grid generation from the rest of the problem. A very effective method for solving the discretized equations and converging to a solution is the multigrid-ADI technique. It has been successfully applied to airfoil problems where O type, C type and slit mappings have been used. Convergence rates for these cases are more than an order of magnitude faster than with relaxation techniques. In this report, we describe a method to extend the above methods, with the C type mappings, to airfoil cascade problems.

  6. Transonic Flow Field Analysis for Wing-Fuselage Configurations

    NASA Technical Reports Server (NTRS)

    Boppe, C. W.

    1980-01-01

    A computational method for simulating the aerodynamics of wing-fuselage configurations at transonic speeds is developed. The finite difference scheme is characterized by a multiple embedded mesh system coupled with a modified or extended small disturbance flow equation. This approach permits a high degree of computational resolution in addition to coordinate system flexibility for treating complex realistic aircraft shapes. To augment the analysis method and permit applications to a wide range of practical engineering design problems, an arbitrary fuselage geometry modeling system is incorporated as well as methodology for computing wing viscous effects. Configuration drag is broken down into its friction, wave, and lift induced components. Typical computed results for isolated bodies, isolated wings, and wing-body combinations are presented. The results are correlated with experimental data. A computer code which employs this methodology is described.

  7. Experimental and numerical study on condensation in transonic steam flow

    NASA Astrophysics Data System (ADS)

    Majkut, Mirosław; Dykas, Sławomir; Strozik, Michał; Smołka, Krystian

    2015-09-01

    The present paper describes an experimental and numerical study of steam condensing flow in a linear cascade of turbine stator blades. The experimental research was performed on the facility of a small scale steam power plant located at Silesian University of Technology in Gliwice, Poland. The test rig of the facility allows us to perform the tests of steam transonic flows for the conditions corresponding to these which prevail in the low-pressure (LP) condensing steam turbine stages. The experimental data of steam condensing flow through the blade-to- blade stator channel were compared with numerical results obtained using the in-house CFD numerical code TraCoFlow. Obtained results confirmed a good quality of the performed experiment and numerical calculations.

  8. The design and development of transonic multistage compressors

    NASA Technical Reports Server (NTRS)

    Ball, C. L.; Steinke, R. J.; Newman, F. A.

    1988-01-01

    The development of the transonic multistage compressor is reviewed. Changing trends in design and performance parameters are noted. These changes are related to advances in compressor aerodynamics, computational fluid mechanics and other enabling technologies. The parameters normally given to the designer and those that need to be established during the design process are identified. Criteria and procedures used in the selection of these parameters are presented. The selection of tip speed, aerodynamic loading, flowpath geometry, incidence and deviation angles, blade/vane geometry, blade/vane solidity, stage reaction, aerodynamic blockage, inlet flow per unit annulus area, stage/overall velocity ratio, and aerodynamic losses are considered. Trends in these parameters both spanwise and axially through the machine are highlighted. The effects of flow mixing and methods for accounting for the mixing in the design process are discussed.

  9. Unsteady transonic aerodynamic and aeroelastic calculations about airfoils and wings

    NASA Technical Reports Server (NTRS)

    Goorjian, P. M.; Guruswamy, G. P.

    1985-01-01

    The development and application of transonic small disturbance codes for computing two dimensional flows, using the code ATRAN2, and for computing three dimensional flows, using the code ATRAN3S, are described. Calculated and experimental results are compared for unsteady flows about airfoils and wings, including several of the cases from the AGARD Standard Aeroelastic Configurations. In two dimensions, the results include AGARD priority cases for the NACA 54A006, NACA 64A010, NACA 0012, and MBB-A3 airfoils. In three dimensions, the results include flow about the F-5 wing, a typical wing, and the AGARD rectangular wings. Viscous corrections are included in some calculations, including those for the AGARD rectangular wing. For several cases, the aerodynamic and aeroelastic calculations are compared with experimental results.

  10. Viscous effect on airfoils for unsteady transonic flows

    NASA Technical Reports Server (NTRS)

    Lee, S. C.

    1982-01-01

    The viscous effect on aerodynamic performance of an arbitrary airfoil executing low frequency maneuvers during transonic flight was investigated. The small disturbance code, LTRAN2, was modified by using a conventional integral method, BLAYER, for the boundary layer and an empirical relation, viscous wedge, for simulating the suddenly thickened boundary layer behind the shock. Before the shock, only the boundary layer displacement thickness was evaluated. After the shock, the empirical wedge thickness was superimposed on the boundary layer thickness along the surface as well as in the wake region. The pressure coefficients were calculated for both steady and unsteady states. The viscous solution takes fewer iterations to obtain the converged steady state solution. Comparisons made with experimental data and the inviscid solution show that the viscous solution agrees better with the experimental data with about the same (or slightly less) amount of computational time.

  11. Accurate solutions for transonic viscous flow over finite wings

    NASA Technical Reports Server (NTRS)

    Vatsa, V. N.

    1986-01-01

    An explicit multistage Runge-Kutta type time-stepping scheme is used for solving the three-dimensional, compressible, thin-layer Navier-Stokes equations. A finite-volume formulation is employed to facilitate treatment of complex grid topologies encountered in three-dimensional calculations. Convergence to steady state is expedited through usage of acceleration techniques. Further numerical efficiency is achieved through vectorization of the computer code. The accuracy of the overall scheme is evaluated by comparing the computed solutions with the experimental data for a finite wing under different test conditions in the transonic regime. A grid refinement study ir conducted to estimate the grid requirements for adequate resolution of salient features of such flows.

  12. Improved method for transonic airfoil design-by-optimization

    NASA Technical Reports Server (NTRS)

    Kennelly, R. A., Jr.

    1983-01-01

    An improved method for use of optimization techniques in transonic airfoil design is demonstrated. FLO6QNM incorporates a modified quasi-Newton optimization package, and is shown to be more reliable and efficient than the method developed previously at NASA-Ames, which used the COPES/CONMIN optimization program. The design codes are compared on a series of test cases with known solutions, and the effects of problem scaling, proximity of initial point to solution, and objective function precision are studied. In contrast to the older method, well-converged solutions are shown to be attainable in the context of engineering design using computational fluid dynamics tools, a new result. The improvements are due to better performance by the optimization routine and to the use of problem-adaptive finite difference step sizes for gradient evaluation.

  13. Recent applications of the transonic wing analysis computer code, TWING

    NASA Technical Reports Server (NTRS)

    Subramanian, N. R.; Holst, T. L.; Thomas, S. D.

    1982-01-01

    An evaluation of the transonic-wing-analysis computer code TWING is given. TWING utilizes a fully implicit approximate factorization iteration scheme to solve the full potential equation in conservative form. A numerical elliptic-solver grid-generation scheme is used to generate the required finite-difference mesh. Several wing configurations were analyzed, and the limits of applicability of this code was evaluated. Comparisons of computed results were made with available experimental data. Results indicate that the code is robust, accurate (when significant viscous effects are not present), and efficient. TWING generally produces solutions an order of magnitude faster than other conservative full potential codes using successive-line overrelaxation. The present method is applicable to a wide range of isolated wing configurations including high-aspect-ratio transport wings and low-aspect-ratio, high-sweep, fighter configurations.

  14. Transonic flow about a thick circular-arc airfoil

    NASA Technical Reports Server (NTRS)

    Mcdevitt, J. B.; Levy, L. L., Jr.; Deiwert, G. S.

    1975-01-01

    An experimental and theoretical study of transonic flow over a thick airfoil, prompted by a need for adequately documented experiments that could provide rigorous verification of viscous flow simulation computer codes, is reported. Special attention is given to the shock-induced separation phenomenon in the turbulent regime. Measurements presented include surface pressures, streamline and flow separation patterns, and shadowgraphs. For a limited range of free-stream Mach numbers the airfoil flow field is found to be unsteady. Dynamic pressure measurements and high-speed shadowgraph movies were taken to investigate this phenomenon. Comparisons of experimentally determined and numerically simulated steady flows using a new viscous-turbulent code are also included. The comparisons show the importance of including an accurate turbulence model. When the shock-boundary layer interaction is weak the turbulence model employed appears adequate, but when the interaction is strong, and extensive regions of separation are present, the model is inadequate and needs further development.

  15. Unsteady transonic flow control around an airfoil in a channel

    NASA Astrophysics Data System (ADS)

    Hamid, Md. Abdul; Hasan, A. B. M. Toufique; Ali, Mohammad; Mitsutake, Yuichi; Setoguchi, Toshiaki; Yu, Shen

    2016-04-01

    Transonic internal flow around an airfoil is associated with self-excited unsteady shock wave oscillation. This unsteady phenomenon generates buffet, high speed impulsive noise, non-synchronous vibration, high cycle fatigue failure and so on. Present study investigates the effectiveness of perforated cavity to control this unsteady flow field. The cavity has been incorporated on the airfoil surface. The degree of perforation of the cavity is kept constant as 30%. However, the number of openings (perforation) at the cavity upper wall has been varied. Results showed that this passive control reduces the strength of shock wave compared to that of baseline airfoil. As a result, the intensity of shock wave/boundary layer interaction and the root mean square (RMS) of pressure oscillation around the airfoil have been reduced with the control method.

  16. Improved computational treatment of transonic flow about swept wings

    NASA Technical Reports Server (NTRS)

    Ballhaus, W. F.; Bailey, F. R.; Frick, J.

    1976-01-01

    Relaxation solutions to classical three-dimensional small-disturbance (CSD) theory for transonic flow about lifting swept wings are reported. For such wings, the CSD theory was found to be a poor approximation to the full potential equation in regions of the flow field that are essentially two-dimensional in a plane normal to the sweep direction. The effect of this deficiency on the capture of embedded shock waves in terms of (1) the conditions under which shock waves can exist and (2) the relations they must satisfy when they do exist is emphasized. A modified small-disturbance (MSD) equation, derived by retaining two previously neglected terms, was proposed and shown to be a consistent approximation to the full potential equation over a wider range of sweep angles. The effect of these extra terms is demonstrated by comparing CSD, MSD, and experimental wing surface pressures.

  17. Measurements of Aerodynamic Damping in the MIT Transonic Rotor

    NASA Technical Reports Server (NTRS)

    Crawley, E. F.

    1981-01-01

    A method was developed and demonstrated for the direct measurement of aerodynamic forcing and aerodynamic damping of a transonic compressor. The method is based on the inverse solution of the structural dynamic equations of motion of the blade disk system in order to determine the forces acting on the system. The disturbing and damping forces acting on a given blade are determined if the equations of motion are expressed in individual blade coordinates. If the structural dynamic equations are transformed to multiblade coordinates, the damping can be measured for blade disk modes, and related to a reduced frequency and interblade phase angle. In order to measure the aerodynamic damping in this way, the free response to a known excitation is studied.

  18. Transonic buffet behavior of Northrop F-5A aircraft

    NASA Technical Reports Server (NTRS)

    Hwang, C.; Pi, W. S.

    1975-01-01

    Flight tests were performed on an extensively instrumented F-5A aircraft to investigate the dynamic buffet pressure distribution on the wing surfaces and the responses during a series of transonic maneuvers called the windup turns. The maneuvers to maximum lift were performed at three Mach number-altitude combinations with a constant 'q' of approximately 14,360 N/sq m (300 psf). The fluctuating buffet pressure data at 24 stations on the right wing of the F-5A were acquired by miniaturized semiconductor type pressure transducers mounted on the wing. A new transducer mounting and wiring technique was applied where the interference with the natural flow condition was minimized. The data acquired in this manner were found adequate to trace the shock origin, the movement of the shock front and the development of the separated flow (shock induced or leading edge induced) on the wing surface.

  19. Video model deformation system for the National Transonic Facility

    NASA Technical Reports Server (NTRS)

    Burner, A. W.; Snow, W. L.; Goad, W. K.

    1983-01-01

    A photogrammetric closed circuit television system to measure model deformation at the National Transonic Facility is described. The photogrammetric approach was chosen because of its inherent rapid data recording of the entire object field. Video cameras are used to acquire data instead of film cameras due to the inaccessibility of cameras which must be housed within the cryogenic, high pressure plenum of this facility. A rudimentary theory section is followed by a description of the video-based system and control measures required to protect cameras from the hostile environment. Preliminary results obtained with the same camera placement as planned for NTF are presented and plans for facility testing with a specially designed test wing are discussed.

  20. A well posed boundary value problem in transonic gas dynamics

    NASA Technical Reports Server (NTRS)

    Sanz, J. M.

    1978-01-01

    A boundary value problem for the Tricomi equation was studied in connection with transonic gas dynamics. The transformed equation delta u plus 1/3Y u sub Y equals 0 in canonical coordinates was considered in the complex domain of two independent complex variables. A boundary value problem was then set by prescribing the real part of the solution on the boundary of the real unit circle. The Dirichlet problem in the upper unit semicircle with vanishing values of the solution at Y = 0 was solved explicitly in terms of the hypergeometric function for the more general Euler-Poisson-Darboux equation. An explicit representation of the solution was also given for a mixed Dirichlet and Neumann problem for the same equation and domain.

  1. Navier-Stokes analysis of transonic cascade flow

    NASA Technical Reports Server (NTRS)

    Arnone, A.; Liou, M.-S.; Povinelli, L.

    1990-01-01

    A new kind of C-type grid is proposed, this grid is non-periodic on the wake and allows minimum skewness for cascades with high turning and large camber. Reynolds-averaged Navier-Stokes equations are solved on this type of grid using a finite volume discretization and a full multigrid method which uses Runge-Kutta stepping as the driving scheme. The Baldwin-Lomax eddy-viscosity model is used for turbulence closure. A detailed numerical study is proposed for a highly loaded transonic blade. A grid independence analysis is presented in terms of pressure distribution, exit flow angles, and loss coefficient. Comparison with experiments clearly demonstrates the capability of the proposed procedure.

  2. An initial investigation into methods of computing transonic aerodynamic sensitivity coefficients

    NASA Technical Reports Server (NTRS)

    Carlson, Leland A.

    1994-01-01

    The primary accomplishments of the project are as follows: (1) Using the transonic small perturbation equation as a flowfield model, the project demonstrated that the quasi-analytical method could be used to obtain aerodynamic sensitivity coefficients for airfoils at subsonic, transonic, and supersonic conditions for design variables such as Mach number, airfoil thickness, maximum camber, angle of attack, and location of maximum camber. It was established that the quasi-analytical approach was an accurate method for obtaining aerodynamic sensitivity derivatives for airfoils at transonic conditions and usually more efficient than the finite difference approach. (2) The usage of symbolic manipulation software to determine the appropriate expressions and computer coding associated with the quasi-analytical method for sensitivity derivatives was investigated. Using the three dimensional fully conservative full potential flowfield model, it was determined that symbolic manipulation along with a chain rule approach was extremely useful in developing a combined flowfield and quasi-analytical sensitivity derivative code capable of considering a large number of realistic design variables. (3) Using the three dimensional fully conservative full potential flowfield model, the quasi-analytical method was applied to swept wings (i.e. three dimensional) at transonic flow conditions. (4) The incremental iterative technique has been applied to the three dimensional transonic nonlinear small perturbation flowfield formulation, an equivalent plate deflection model, and the associated aerodynamic and structural discipline sensitivity equations; and coupled aeroelastic results for an aspect ratio three wing in transonic flow have been obtained.

  3. SWIFT Code Assessment for Two Similar Transonic Compressors

    NASA Technical Reports Server (NTRS)

    Chima, Rodrick V.

    2009-01-01

    One goal of the NASA Fundamental Aeronautics Program is the assessment of computational fluid dynamic (CFD) codes used for the design and analysis of many aerospace systems. This paper describes the assessment of the SWIFT turbomachinery analysis code for two similar transonic compressors, NASA rotor 37 and stage 35. The two rotors have identical blade profiles on the front, transonic half of the blade but rotor 37 has more camber aft of the shock. Thus the two rotors have the same shock structure and choking flow but rotor 37 produces a higher pressure ratio. The two compressors and experimental data are described here briefly. Rotor 37 was also used for test cases organized by ASME, IGTI, and AGARD in 1994-1998. Most of the participating codes over predicted pressure and temperature ratios, and failed to predict certain features of the downstream flowfield. Since then the AUSM+ upwind scheme and the k- turbulence model have been added to SWIFT. In this work the new capabilities were assessed for the two compressors. Comparisons were made with overall performance maps and spanwise profiles of several aerodynamic parameters. The results for rotor 37 were in much better agreement with the experimental data than the original blind test case results although there were still some discrepancies. The results for stage 35 were in very good agreement with the data. The results for rotor 37 were very sensitive to turbulence model parameters but the results for stage 35 were not. Comparison of the rotor solutions showed that the main difference between the two rotors was not blade camber as expected, but shock/boundary layer interaction on the casing.

  4. On laminar separation at a corner point in transonic flow

    NASA Astrophysics Data System (ADS)

    Ruban, A. I.; Turkyilmaz, I.

    2000-11-01

    The separation of the laminar boundary layer from a convex corner on a rigid body contour in transonic flow is studied based on the asymptotic analysis of the Navier Stokes equations at large values of the Reynolds number. It is shown that the flow in a small vicinity of the separation point is governed, as usual, by strong interaction between the boundary layer and the inviscid part of the flow. Outside the interaction region the Kármán Guderley equation describing transonic inviscid flow admits a self-similar solution with the pressure on the body surface being proportional to the cubic root of the distance from the separation point. Analysis of the boundary layer driven by this pressure shows that as the interaction region is approached the boundary layer splits into two parts: the near-wall viscous sublayer and the main body of the boundary layer where the flow is locally inviscid. It is interesting that contrary to what happens in subsonic and supersonic flows, the displacement effect of the boundary layer is primarily due to the inviscid part. The contribution of the viscous sublayer proves to be negligible to the leading order. Consequently, the flow in the interaction region is governed by the inviscid inviscid interaction. To describe this flow one needs to solve the Kármán Guderley equation for the potential flow region outside the boundary layer; the solution in the main part of the boundary layer was found in an analytical form, thanks to which the interaction between the boundary layer and external flow can be expressed via the corresponding boundary condition for the Kármán Guderley equation. Formulation of the interaction problem involves one similarity parameter which in essence is the Kármán Guderley parameter suitably modified for the flow at hand. The solution of the interaction problem has been constructed numerically.

  5. Unsteady design-point flow phenomena in transonic compressors

    NASA Technical Reports Server (NTRS)

    Gertz, J. B.; Epstein, A. H.

    1986-01-01

    High-frequency response probes which had previously been used exclusively in the MIT Blowndown Facility were successfully employed in two conventional steady state axial flow compressor facilities to investigate the unsteady flowfields of highly loaded transonic compressors at design point operation. Laser anemometry measurements taken simultaneously with the high response data were also analyzed. The time averaged high response data of static and total pressure agreed quite well with the conventional steady state instrumentation except for flow angle which showed a large spread in values at all radii regardless of the type of instrumentation used. In addition, the time resolved measurements confirmed earlier test results obtained in the MIT Blowdown Facility for the same compressor. The results of these tests have further revealed that the flowfields of highly loaded transonic compressors are heavily influenced by unsteady flow phenomena. The high response measurements exhibited large variations in the blade to blade flow and in the blade passage flow. The observed unsteadiness in the blade wakes is explained in terms of the rotor blades' shed vorticity in periodic vortex streets. The wakes were modeled as two-dimensional vortex streets with finite size cores. The model fit the data quite well as it was able to reproduce the average wake shape and bi-modal probability density distributions seen in the laser anemometry data. The presence of vortex streets in the blade wakes also explains the large blade to blade fluctuations seen by the high response probes which is simply due to the intermittent sampling of the vortex street as it is swept past a stationary probe.

  6. Advancing Test Capabilities at NASA Wind Tunnels

    NASA Technical Reports Server (NTRS)

    Bell, James

    2015-01-01

    NASA maintains twelve major wind tunnels at three field centers capable of providing flows at 0.1 M 10 and unit Reynolds numbers up to 45106m. The maintenance and enhancement of these facilities is handled through a unified management structure under NASAs Aeronautics and Evaluation and Test Capability (AETC) project. The AETC facilities are; the 11x11 transonic and 9x7 supersonic wind tunnels at NASA Ames; the 10x10 and 8x6 supersonic wind tunnels, 9x15 low speed tunnel, Icing Research Tunnel, and Propulsion Simulator Laboratory, all at NASA Glenn; and the National Transonic Facility, Transonic Dynamics Tunnel, LAL aerothermodynamics laboratory, 8 High Temperature Tunnel, and 14x22 low speed tunnel, all at NASA Langley. This presentation describes the primary AETC facilities and their current capabilities, as well as improvements which are planned over the next five years. These improvements fall into three categories. The first are operations and maintenance improvements designed to increase the efficiency and reliability of the wind tunnels. These include new (possibly composite) fan blades at several facilities, new temperature control systems, and new and much more capable facility data systems. The second category of improvements are facility capability advancements. These include significant improvements to optical access in wind tunnel test sections at Ames, improvements to test section acoustics at Glenn and Langley, the development of a Supercooled Large Droplet capability for icing research, and the development of an icing capability for large engine testing. The final category of improvements consists of test technology enhancements which provide value across multiple facilities. These include projects to increase balance accuracy, provide NIST-traceable calibration characterization for wind tunnels, and to advance optical instruments for Computational Fluid Dynamics (CFD) validation. Taken as a whole, these individual projects provide significant

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

    NASA Technical Reports Server (NTRS)

    Johnson, W. G., Jr.; Hill, A. S.; Ray, E. J.; Rozendaal, R. A.; Butler, T. W.

    1982-01-01

    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.

  8. Transonic galactic outflows in a dark matter halo with a central black hole and its application to the Sombrero galaxy

    NASA Astrophysics Data System (ADS)

    Igarashi, Asuka; Mori, Masao; Nitta, Shin-ya

    2014-10-01

    We have classified possible transonic solutions of galactic outflows in the gravitational potential of the dark matter halo (DMH) and supermassive black hole (SMBH) under the assumptions of isothermal, spherically symmetric and steady state. It is clarified that the gravity of SMBH adds a new branch of transonic solutions with the transonic point in very close proximity to the centre in addition to the outer transonic point generated by the gravity of DMH. Because these two transonic solutions have substantially different mass fluxes and starting points, these solutions may have different influences on the evolution of galaxies and the release of metals into intergalactic space. We have applied our model to the Sombrero galaxy and obtained a new type of galactic outflow: a slowly accelerated transonic outflow through the transonic point at very distant region (≃126 kpc). In this galaxy, previous works reported that although the trace of the galactic outflow is observed by X-ray, the gas density distribution is consistent with the hydrostatic state. We have clarified that the slowly accelerating outflow has a gas density profile quite similar to that of the hydrostatic solution in the widely spread subsonic region. Thus, the slowly accelerating transonic solution cannot be distinguished from the hydrostatic solution in the observed region (≤25 kpc) even if slow transonic flow exists. Our model provides a new perspective of galactic outflows and is applicable even to quiescent galaxies with inactive star formation.

  9. High Reynolds number transonic tests on a NACA 0012 airfoil in the Langley 0.3-meter transonic cryogenic tunnel

    NASA Technical Reports Server (NTRS)

    Ladson, Charles L.; Hill, S. Acquilla

    1987-01-01

    Tests were conducted in the two-dimensional test section of the Langley 0.3-m Transonic Cryogenic Tunnel on a NACA 0012 airfoil to obtain aerodynamic data as a part of the Advanced Technology Airfoil Test (ATAT) program. The test program covered a Mach number range of 0.30 to 0.82 and a Reynolds number range of 3.0 to 45.0 x 10 to the 6th. The stagnation pressure was varied between 1.2 and 6.0 atmospheres and the stagnation temperature was varied between 300 K and 90 K to obtain these test conditions. Plots of the spanwise variation of drag coefficient as a function of normal force coefficient and the variation of the basic aerodynamic characteristics with angle of attack are shown. The data are presented uncorrected for wall interference effects and without analysis.

  10. A Factorial Data Rate and Dwell Time Experiment in the National Transonic Facility

    NASA Technical Reports Server (NTRS)

    DeLoach, R.

    2000-01-01

    This report is an introductory tutorial on the application of formal experiment design methods to wind tunnel testing, for the benefit of aeronautical engineers with little formal experiment design training. It also describes the results of a Study to determine whether increases in the sample rate and dwell time of the National Transonic Facility data system Would result in significant changes in force and moment data. Increases in sample rate from 10 samples per second to 50 samples per second were examined, as were changes in dwell time from one second per data point to two seconds. These changes were examined for a representative aircraft model in a range of tunnel operating conditions defined by angles of attack from 0 deg to 3.8 degrees, total pressure from 15.0 psi to 24.1 psi, and Mach numbers from 0.52 to 0.82. No statistically significant effect was associated with the change in sample rate. The change in dwell time from one second to two seconds affected axial force measurements, and to a lesser degree normal force measurements. This dwell effect comprises a "rectification error" caused by incomplete cancellation of the positive and negative elements of certain low frequency dynamic components that are not rejected by the one-Hz low-pass filters of the data system. These low frequency effects may be due to tunnel circuit phenomena and other sources. The magnitude of the dwell effect depends on dynamic pressure, with angle of attack and Mach number influencing the strength of this dependence. An analysis is presented which suggests that the magnitude of the rectification error depends on the ratio of measurement dwell time to the period of the low-frequency dynamics, as well as the amplitude of the dynamics The essential conclusion of this analysis is that extending the dwell time (or, equivalently, replicating short-dwell data points) reduces the rectification error.

  11. Further buffeting tests in a cryogenic wind tunnel

    NASA Technical Reports Server (NTRS)

    Mabey, D. G.; Boyden, R. P.; Johnson, W. G., Jr.

    1992-01-01

    Further measurements of buffeting, using wing-root strain gauges, were made in the NASA Langley 0.3 m Cryogenic Wind Tunnel to refine techniques which will be used in larger cryogenic facilities such as the United States National Transonic Facility (NTF) and European Transonic Wind Tunnel (ETW). The questions addressed included the relative importance of variations in frequency parameter and Reynolds number, the choice of model material (considering both stiffness and damping) and the effects of static aeroelastic distortion. The main series of tests was made on half models of slender 65 deg delta wings with a sharp leading edge. The three delta wings had the same planform but widely different bending stiffness and frequencies (obtained by varying both the material and the thickness of the wings). It was known that the flow on this configuration would be insensitive to variations in Reynold number. Additional tests were made on one unswept half-wing of aspect ratio 1.5 with an NPL 9510 aerofoil section, known to be sensitive to variations in Reynolds number at transonic speeds. For brevity the test Mach numbers were restricted to M = 0.21 and 0.35 for the delta wings and to M = 0.30 for the unswept wing.

  12. Critical point of the solar wind by radio sounding data

    NASA Astrophysics Data System (ADS)

    Lotova, N. A.; Oraevsky, V. N.; Pisarenko, Ya. V.; Vladimirskii, K. V.

    1995-06-01

    Results of the close-to-Sun plasmas sounding at the transonic region of the solar wind, where the sub-to supersonic flow transition proceeds (at 10 to 40 solar radii from the Sun), are presented. Natural sources of two types were used, water vapour maser sources at 1.35 cm and guasars at 2.9 m wavelength. scattering observations cover the period of 1986 to 1993, Russian Academy of Sciences telescopes RT-22 and DCR-1000 were used, IPS index and scattering angle being the immediate results of observations. Extensive studies of the scintillation index and scattering angle radial profiles reveal a remarkable structural detail, 'transonic region forrunner'-narrow region of diminished scattering close to the internal border of the extended transonic region with its characteristic enhanced scattering. Comparisons of the scattering and plasma velocity profiles let it possible to determine the critical point positions by the comparatively simple scattering observations. This new possibility widely improves the process of the basic data accumulation in the fundamental problem of the solar wind acceleration mechanism.

  13. Flutter tests (IS4) of the 0.0125-scale shuttle reflection plane model 30-OTS in the Langley Research Center 26-inch transonic blowdown tunnel test no. 547

    NASA Technical Reports Server (NTRS)

    Kotch, M. A.

    1974-01-01

    A series of slab wing flutter models with rigid orbiter fuselage, external tank, and SRB models of the space shuttle were tested, in a reflection plane arrangement, in the NASA Langley Research Center's 26-inch Transonic Blowdown Tunnel. Model flutter boundaries were obtained for both a wing-alone configuration and a wing-with-orbiter, tank and SRB configuration. Additional test points were taken of the wing-with-orbiter configuration, as a correlation with the wing-alone condition. A description of the wind tunnel models and test procedures utilized in the experiment are provided.

  14. The Method of Complex Characteristics for Design of Transonic Compressors.

    NASA Astrophysics Data System (ADS)

    Bledsoe, Margaret Randolph

    We calculate shockless transonic flows past two -dimensional cascades of airfoils characterized by a prescribed speed distribution. The approach is to find solutions of the partial differential equation (c('2)-u('2)) (PHI)(,xx) - 2uv (PHI)(,xy) + (c('2)-v('2)) (PHI)(,yy) = 0 by the method of complex characteristics. Here (PHI) is the velocity potential, so (DEL)(PHI) = (u,v), and c is the local speed of sound. Our method consists in noting that the coefficients of the equation are analytic, so that we can use analytic continuation, conformal mapping, and a spectral method in the hodograph plane to determine the flow. After complex extension we obtain canonical equations for (PHI) and for the stream function (psi) as well as an explicit map from the hodograph plane to complex characteristic coordinates. In the subsonic case, a new coordinate system is defined in which the flow region corresponds to the interior of an ellipse. We construct special solutions of the flow equations in these coordinates by solving characteristic initial value problems in the ellipse with initial data defined by the complete system of Chebyshev polynomials. The condition (psi) = 0 on the boundary of the ellipse is used to determine the series representation of (PHI) and (psi). The map from the ellipse to the complex flow coordinates is found from data specifying the speed q as a function of the arc length s. The transonic problem for shockless flow becomes well posed after appropriate modifications of this procedure. The nonlinearity of the problem is handled by an iterative method that determines the boundary value problem in the ellipse and the map function in sequence. We have implemented this method as a computer code to design two-dimensional cascades of shockless compressor airfoils with gap-to-chord ratios as low as .5 and supersonic zones on both the upper and lower surfaces. The method may be extended to solve more general boundary value problems for second order partial

  15. Navier-Stokes simulations of unsteady transonic flow phenomena

    NASA Technical Reports Server (NTRS)

    Atwood, C. A.

    1992-01-01

    Numerical simulations of two classes of unsteady flows are obtained via the Navier-Stokes equations: a blast-wave/target interaction problem class and a transonic cavity flow problem class. The method developed for the viscous blast-wave/target interaction problem assumes a laminar, perfect gas implemented in a structured finite-volume framework. The approximately factored implicit scheme uses Newton subiterations to obtain the spatially and temporally second-order accurate time history of the blast-waves with stationary targets. The inviscid flux is evaluated using either of two upwind techniques, while the full viscous terms are computed by central differencing. Comparisons of unsteady numerical, analytical, and experimental results are made in two- and three-dimensions for Couette flows, a starting shock-tunnel, and a shock-tube blockage study. The results show accurate wave speed resolution and nonoscillatory discontinuity capturing of the predominantly inviscid flows. Viscous effects were increasingly significant at large post-interaction times. While the blast-wave/target interaction problem benefits from high-resolution methods applied to the Euler terms, the transonic cavity flow problem requires the use of an efficient scheme implemented in a geometrically flexible overset mesh environment. Hence, the Reynolds averaged Navier-Stokes equations implemented in a diagonal form are applied to the cavity flow class of problems. Comparisons between numerical and experimental results are made in two-dimensions for free shear layers and both rectangular and quieted cavities, and in three-dimensions for Stratospheric Observatory For Infrared Astronomy (SOFIA) geometries. The acoustic behavior of the rectangular and three-dimensional cavity flows compare well with experiment in terms of frequency, magnitude, and quieting trends. However, there is a more rapid decrease in computed acoustic energy with frequency than observed experimentally owing to numerical

  16. Self-sustained shock oscillations on airfoils at transonic speeds

    NASA Astrophysics Data System (ADS)

    Lee, B. H. K.

    2001-02-01

    Self-sustained shock wave oscillations on airfoils at transonic flow conditions are associated with the phenomenon of buffeting. The physical mechanisms of the periodic shock motion are not yet fully understood even though experiments performed over fifty years ago have demonstrated the presence of oscillatory shock waves on the airfoil surfaces at high subsonic speeds. The unsteady pressure fluctuations generated by the low-frequency large-amplitude shock motions are highly undesirable from the structural integrity and aircraft maneuverability point of view. For modern supercritical wing design with thick profiles, the shock-induced fluctuations are particularly severe and methods to reduce the shock wave amplitudes to lower values or even to delay the oscillations to higher Mach numbers or incidence angles will result in expanding the buffet boundary of the airfoil. This review begins with a recapitulation of the classical work on shock-induced bubble separation and trailing edge separation of a turbulent boundary layer. The characteristics of the unsteady pressure fluctuations are used to classify the types of shock-boundary layer interaction. The various modes of shock wave motion for different flow conditions and airfoil configurations are described. The buffet boundaries obtained using the standard trailing edge pressure divergence technique and an alternative approach of measuring the divergence of normal fluctuating forces are compared to show the equivalence. The mechanisms of self-sustained shock oscillations are discussed for symmetrical circular-arc airfoils at zero incidence and for supercritical airfoils at high incidence angles with fully separated flows. The properties of disturbances in the wake are examined from linear stability analysis of two-dimensional compressible flows. The advances in high-speed computing make predictions of buffeting flows possible. Navier-Stokes solvers and approximate boundary layer-inviscid flow interaction methods are

  17. Transonic Shock Oscillations and Wing Flutter Calculated with an Interactive Boundary Layer Coupling Method

    NASA Technical Reports Server (NTRS)

    Edwards, John W.

    1996-01-01

    A viscous-inviscid interactive coupling method is used for the computation of unsteady transonic flows involving separation and reattachment. A lag-entrainment integral boundary layer method is used with the transonic small disturbance potential equation in the CAP-TSDV (Computational Aeroelasticity Program - Transonic Small Disturbance) code. Efficient and robust computations of steady and unsteady separated flows, including steady separation bubbles and self-excited shock-induced oscillations are presented. The buffet onset boundary for the NACA 0012 airfoil is accurately predicted and shown computationally to be a Hopf bifurcation. Shock-induced oscillations are also presented for the 18 percent circular arc airfoil. The oscillation onset boundaries and frequencies are accurately predicted, as is the experimentally observed hysteresis of the oscillations with Mach number. This latter stability boundary is identified as a jump phenomenon. Transonic wing flutter boundaries are also shown for a thin swept wing and for a typical business jet wing, illustrating viscous effects on flutter and the effect of separation onset on the wing response at flutter. Calculations for both wings show limit cycle oscillations at transonic speeds in the vicinity of minimum flutter speed indices.

  18. Unsteady transonic aerodynamics and aeroelastic calculations at low-supersonic freestreams

    NASA Technical Reports Server (NTRS)

    Guruswamy, Guru P.; Goorjian, Peter M.

    1988-01-01

    A computational procedure is presented to simulate transonic unsteady flows and corresponding aeroelasticity of wings at low-supersonic freestreams. The flow is modeled by using the transonic small-perturbation theory. The structural equations of motions are modeled using modal equations of motion directly coupled with aerodynamics. Supersonic freestreams are simulated by properly accounting for the boundary conditions based on pressure waves along the flow characteristics in streamwise planes. The flow equations are solved using the time-accurate, alternating-direction implicit finite-difference scheme. The coupled aeroelastic equations of motion are solved by an integration procedure based on the time-accurate, linear-acceleration method. The flow modeling is verified by comparing calculations with experiments for both steady and unsteady flows at supersonic freestreams. The unsteady computations are made for oscillating wings. Comparisons of computed results with experiments show good agreement. Aeroelastic responses are computed for a rectangular wing at Mach numbers ranging from subtransonic to upper-transonic (supersonic) freestreams. The extension of the transonic dip into the upper transonic regime is illustrated.

  19. A forward-swept wing configuration designed for high maneuverability by use of three-dimensional transonic theory

    NASA Technical Reports Server (NTRS)

    Mann, M. J.; Mercer, C. E.

    1985-01-01

    Supercritical technology has been applied to the design of a forward-swept-wing fighter configuration, and an assessment has been made of the relative performance of forward versus aft sweep. The wing and canard for this forward-swept wing configuration were designed for transonic maneuver by the use of a transonic computational analysis method and a transonic design procedure. The computational method calculates the transonic flow over a canard-wing-fuselage combination so that the strong transonic induced-flow effects of the canard on the wing are taken into account. A model of this configuration was constructed and was tested in the Langley 16-Foot Transonic Tunnel. The transonic theory gave a reasonably good estimate of the wing pressure distributions at transonic maneuver conditions. Comparison of this configuration with an equivalent aft-swept wing configuration showed that, at a Mach number of 0.9 and a lift coefficient of 0.9, the two configurations have essentially the same drag. This forward-swept wing configuration was also found to have very good maneuver performance relative to the Rockwell International HiMAT highly-maneuverable aircraft configuration.

  20. Finite element analysis of transonic flows in cascades: Importance of computational grids in improving accuracy and convergence

    NASA Technical Reports Server (NTRS)

    Ecer, A.; Akay, H. U.

    1981-01-01

    The finite element method is applied for the solution of transonic potential flows through a cascade of airfoils. Convergence characteristics of the solution scheme are discussed. Accuracy of the numerical solutions is investigated for various flow regions in the transonic flow configuration. The design of an efficient finite element computational grid is discussed for improving accuracy and convergence.

  1. F-15B transonic flight research testbed aircraft in flight

    NASA Technical Reports Server (NTRS)

    1996-01-01

    NASA's Dryden Flight Research Center, Edwards, California, is flying a modified McDonnell-Douglas F-15B aircraft as a testbed for a variety of transonic flight experiments. The two-seat aircraft, bearing NASA tail number 836, is shown during a recent flight over the high desert carrying a Drdyen-designed Flight Test Fixture (FTF) upon which aerodynamic experiments are mounted. The FTF is a heavily instrumented fin-like structure which is mounted on the F-15B's underbelly in place of the standard external fuel tank. Since being aquired by NASA in 1993, the aircraft has been modified to include video recording, telemetry and data recording capabilities. The twin-engine aircraft flew several flights recently in support of an experiment to determine the precise location of sonic shockwave development as air passes over an airfoil. The F-15B is currently being prepared for the Boundary Layer Heat Experiment, which will explore the potential drag reduction from heating the turbulent portion of the air that passes over the fuselage of a large aircraft.

  2. Calculation of tip clearance effects in a transonic compressor rotor

    NASA Technical Reports Server (NTRS)

    Chima, R. V.

    1996-01-01

    The flow through the tip clearance region of a transonic compressor rotor (NASA rotor 37) was computed and compared to aerodynamic probe and laser anemometer data. Tip clearance effects were modeled both by gridding the clearance gap and by using a simple periodicity model across the ungridded gap. The simple model was run with both the full gap height, and with half the gap height to simulate a vena-contracta effect. Comparisons between computed and measured performance maps and downstream profiles were used to validate the models and to assess the effects of gap height on the simple clearance model. Recommendations were made concerning the use of the simple clearance model. Detailed comparisons were made between the gridded clearance gap solution and the laser anemometer data near the tip at two operating points. The computer results agreed fairly well with the data but overpredicted the extent of the casing separation and underpredicted the wake decay rate. The computations were then used to describe the interaction of the tip vortex, the passage shock, and the casing boundary layer.

  3. Calculation of Tip Clearance Effects in a Transonic Compressor Rotor

    NASA Technical Reports Server (NTRS)

    Chima, R. V.

    1998-01-01

    The flow through the tip clearance region of a transonic compressor rotor (NASA rotor 37) was computed and compared to aerodynamic probe and laser anemometer data. Tip clearance effects were modeled both by gridding the clearance gap and by using a simple periodicity model across the ungridded gap. The simple model was run with both the full gap height, and with half the gap height to simulate a vena-contracta effect. Comparisons between computed and measured performance maps and downstream profiles were used to validate the models and to assess the effects of gap height on the simple clearance model. Recommendations were made concern- ing the use of the simple clearance model Detailed comparisons were made between the gridded clearance gap solution and the laser anemometer data near the tip at two operating points. The computed results agreed fairly well with the data but overpredicted the extent of the casing separation and underpredicted the wake decay rate. The computations were then used to describe the interaction of the tip vortex, the passage shock, and the casing boundary layer.

  4. Development of an Uncertainty Model for the National Transonic Facility

    NASA Technical Reports Server (NTRS)

    Walter, Joel A.; Lawrence, William R.; Elder, David W.; Treece, Michael D.

    2010-01-01

    This paper introduces an uncertainty model being developed for the National Transonic Facility (NTF). The model uses a Monte Carlo technique to propagate standard uncertainties of measured values through the NTF data reduction equations to calculate the combined uncertainties of the key aerodynamic force and moment coefficients and freestream properties. The uncertainty propagation approach to assessing data variability is compared with ongoing data quality assessment activities at the NTF, notably check standard testing using statistical process control (SPC) techniques. It is shown that the two approaches are complementary and both are necessary tools for data quality assessment and improvement activities. The SPC approach is the final arbiter of variability in a facility. Its result encompasses variation due to people, processes, test equipment, and test article. The uncertainty propagation approach is limited mainly to the data reduction process. However, it is useful because it helps to assess the causes of variability seen in the data and consequently provides a basis for improvement. For example, it is shown that Mach number random uncertainty is dominated by static pressure variation over most of the dynamic pressure range tested. However, the random uncertainty in the drag coefficient is generally dominated by axial and normal force uncertainty with much less contribution from freestream conditions.

  5. Newton like: Minimal residual methods applied to transonic flow calculations

    NASA Technical Reports Server (NTRS)

    Wong, Y. S.

    1984-01-01

    A computational technique for the solution of the full potential equation is presented. The method consists of outer and inner iterations. The outer iterate is based on a Newton like algorithm, and a preconditioned Minimal Residual method is used to seek an approximate solution of the system of linear equations arising at each inner iterate. The present iterative scheme is formulated so that the uncertainties and difficulties associated with many iterative techniques, namely the requirements of acceleration parameters and the treatment of additional boundary conditions for the intermediate variables, are eliminated. Numerical experiments based on the new method for transonic potential flows around the NACA 0012 airfoil at different Mach numbers and different angles of attack are presented, and these results are compared with those obtained by the Approximate Factorization technique. Extention to three dimensional flow calculations and application in finite element methods for fluid dynamics problems by the present method are also discussed. The Inexact Newton like method produces a smoother reduction in the residual norm, and the number of supersonic points and circulations are rapidly established as the number of iterations is increased.

  6. Endwall Heat Transfer Measurements in a Transonic Turbine Cascade

    NASA Technical Reports Server (NTRS)

    Giel, P. W.; Thurman, D. R.; VanFossen, G. J.; Hippensteele, S. A.; Boyle, R. J.

    1996-01-01

    Turbine blade endwall heat transfer measurements are given for a range of Reynolds and Mach numbers. Data were obtained for Reynolds numbers based on inlet conditions of 0.5 and 1.0 x 106, for isentropic exit Mach numbers of 1.0 and 1.3, and for freestream turbulence intensities of 0.25% and 7.0%. Tests were conducted in a linear cascade at the NASA Lewis Transonic Turbine Blade Cascade Facility. The test article was a turbine rotor with 136' of turning and an axial chord of 12.7 cm. The large scale allowed for very detailed measurements of both flow field and surface phenomena. The intent of the work is to provide benchmark quality data for computational fluid dynamics (CFD) code and model verification. The flow field in the cascade is highly three-dimensional as a result of thick boundary layers at the test section inlet. Endwall heat transfer data were obtained using a steady-state liquid crystal technique.

  7. Drag Measurement at Transonic Speeds on a Freely Falling Body

    NASA Technical Reports Server (NTRS)

    Bailey, F. J., Jr.; Mathews, Charles W.; Thompson, Jim R.

    1945-01-01

    Direct measurements have been made of the drag of a special test body and its stabilizing tail surfaces throughout free drops from high altitudes. The data obtained have been used to establish the relation between the drag coefficient and the Mach number for the body and for the tail surfaces over a range of Mach numbers from 0.85 to 1.15. For bodies of the form tested, the drag per square foot of frontal area increased abruptly from about 3 percent of atmospheric pressure at a Mach number of 0.95 to 17 percent of atmospheric pressure at a Mach number of 1.00, then linearly with Mach number to 28 percent of atmospheric pressure at a Mach number of approximately 1.15. Some doubt exists as to the applicability of the tail drag results to the estimation of wing drag at transonic speeds because of the possibility of appreciable interference effects between the vertical and the horizontal surfaces and between the body and the tail surfaces. Insofar as they are applicable, the tail drag results indicated that with symmetrical 6-percent-thick area may be expected to increase abruptly from 4 percent of atmospheric pressure at a Mach number of 0.88 to 36 percent of atmospheric pressure at a Mach number of 1.00, then linearly with Mach number to approximately 50 percent of atmospheric pressure at a Mach number of 1.15.

  8. Performance of turbulence models for transonic flows in a diffuser

    NASA Astrophysics Data System (ADS)

    Liu, Yangwei; Wu, Jianuo; Lu, Lipeng

    2016-09-01

    Eight turbulence models frequently used in aerodynamics have been employed in the detailed numerical investigations for transonic flows in the Sajben diffuser, to assess the predictive capabilities of the turbulence models for shock wave/turbulent boundary layer interactions (SWTBLI) in internal flows. The eight turbulence models include: the Spalart-Allmaras model, the standard k - 𝜀 model, the RNG k - 𝜀 model, the realizable k - 𝜀 model, the standard k - ω model, the SST k - ω model, the v2¯ - f model and the Reynolds stress model. The performance of the different turbulence models adopted has been systematically assessed by comparing the numerical results with the available experimental data. The comparisons show that the predictive performance becomes worse as the shock wave becomes stronger. The v2¯ - f model and the SST k - ω model perform much better than other models, and the SST k - ω model predicts a little better than the v2¯ - f model for pressure on walls and velocity profile, whereas the v2¯ - f model predicts a little better than the SST k - ω model for separation location, reattachment location and separation length for strong shock case.

  9. An approximate factorization solution of the Navier-Stokes equations for transonic flow using body-fitted coordinates with application to NACA 64A010 airfoils

    NASA Technical Reports Server (NTRS)

    Copper, G. K.

    1980-01-01

    The implementation of the approximate factorization algorithm and its ability to efficiently and accurately describe transonic flow about an NACA 64A010 airfoil section is examined. The approximate factorization algorithm is developed from the nondimensional, conservative, vectorized Navier-Stokes equations expressed in curvilinear coordinates. Equations of state and transport coefficient relations appropriate to atmospheric air are appended to close the system of partial differential equations. An algebraic turbulence model is also incorporated into the equation set. This algorithm was verified by investigating the flow about an NACA 64A010 airfoil at 0, 2, and 3.5 deg angle of attack for free-stream conditions of 2,000,000 Reynolds number and 0.8 Mach number. Overall results were in good qualitative agreement with wind tunnel data sets. However, while nondimensional times of six were attained, numerical difficulties prevented any case from reaching a true steady state.

  10. High Reynolds number tests of the cast 10-2/DOA 2 airfoil in the Langley 0.3-meter transonic cryogenic tunnel, phase 2

    NASA Technical Reports Server (NTRS)

    Dress, D. A.; Stanewsky, E.; Mcguire, P. D.; Ray, E. J.

    1984-01-01

    Wind tunnel tests of an advanced technology airfoil, the CAST 10-2/DOA 2, were conducted in the Langley 0.3-Meter Transonic Cryogenic Tunnel (0.3-m TCT). This was the third of a series of tests conducted in a cooperative airfoil research program between the National Aeronautics and Space Administration and the Deutsche Forschungsund Versuchsanstalt fur Luft- und Raumfahrt e. V. For these tests, temperature was varied from 270 K to 110 K at pressures from 1.5 to 5.75 atmospheres. Mach number was varied from 0.60 to 0.80, and the Reynolds number (based on airfoil chord) was varied from 2 to 20 million. The aerodynamic data for the 7.62 cm chord airfoil model used in these tests is presented without analysis. Descriptions of the 0.3-m TCT, the airfoil model, the test instrumentation, and the testing procedures are included.

  11. Navier-Stokes simulation of external/internal transonic flow on the forebody/inlet of the AV-8B Harrier II

    NASA Technical Reports Server (NTRS)

    Mysko, Stephen J.; Chyu, Wei J.; Stortz, Michael W.; Chow, Chuen-Yen

    1993-01-01

    In this work, the computation of combined external/internal transonic flow on the complex forebody/inlet configuration of the AV-8B Harrier II is performed. The actual aircraft has been measured and its surface and surrounding domain, in which the fuselage and inlet have a common wall, have been described using structured grids. The 'thin-layer' Navier-Stokes equations were used to model the flow along with the Chimera embedded multi-block technique. A fully conservative, alternating direction implicit (ADI), approximately factored, partially fluxsplit algorithm was employed to perform the computation. Comparisons to some experimental wind tunnel data yielded good agreement for flow at zero incidence and angle of attack. The aim of this paper is to provide a methodology or computational tool for the numerical solution of complex external/internal flows.

  12. High Reynolds number tests of the CAST 10-2/DOA 2 airfoil in the Langley 0.3-meter transonic cryogenic tunnel, phase 1

    NASA Technical Reports Server (NTRS)

    Dress, D. A.; Mcguire, P. D.; Stanewsky, E.; Ray, E. J.

    1983-01-01

    A wind tunnel investigation of an advanced technology airfoil, the CAST 10-2/DOA 2, was conducted in the Langley 0.3 meter Transonic Cryogenic Tunnel (0.3 m TCT). This was the first of a series of tests conducted in a cooperative National Aeronautics and Space Administration (NASA) and the Deutsche Forschungs- und Versuchsanstalt fur Luft- und Raumfahrt e. V. (DFVLR) airfoil research program. Test temperature was varied from 280 K to 100 K to pressures from slightly above 1 to 5.8 atmospheres. Mach number was varied from 0.60 to 0.80, and the Reynolds number (based on airfoil chord) was varied from 4 x 10 to the 8th power to 45 x 10 to the 6th power. This report presents the experimental aerodynamic data obtained for the airfoil and includes descriptions of the airfoil model, the 0.3 m TCT, the test instrumentation, and the testing procedures.

  13. National Wind Tunnel Complex (NWTC)

    NASA Technical Reports Server (NTRS)

    1996-01-01

    The National Wind Tunnel Complex (NWTC) Final Report summarizes the work carried out by a unique Government/Industry partnership during the period of June 1994 through May 1996. The objective of this partnership was to plan, design, build and activate 'world class' wind tunnel facilities for the development of future-generation commercial and military aircraft. The basis of this effort was a set of performance goals defined by the National Facilities Study (NFS) Task Group on Aeronautical Research and Development Facilities which established two critical measures of improved wind tunnel performance; namely, higher Reynolds number capability and greater productivity. Initial activities focused upon two high-performance tunnels (low-speed and transonic). This effort was later descoped to a single multipurpose tunnel. Beginning in June 1994, the NWTC Project Office defined specific performance requirements, planned site evaluation activities, performed a series of technical/cost trade studies, and completed preliminary engineering to support a proposed conceptual design. Due to budget uncertainties within the Federal government, the NWTC project office was directed to conduct an orderly closure following the Systems Design Review in March 1996. This report provides a top-level status of the project at that time. Additional details of all work performed have been archived and are available for future reference.

  14. Reynolds Number Effects on Leading Edge Radius Variations of a Supersonic Transport at Transonic Conditions

    NASA Technical Reports Server (NTRS)

    Rivers, S. M. B.; Wahls, R. A.; Owens, L. R.

    2001-01-01

    A computational study focused on leading-edge radius effects and associated Reynolds number sensitivity for a High Speed Civil Transport configuration at transonic conditions was conducted as part of NASA's High Speed Research Program. The primary purposes were to assess the capabilities of computational fluid dynamics to predict Reynolds number effects for a range of leading-edge radius distributions on a second-generation supersonic transport configuration, and to evaluate the potential performance benefits of each at the transonic cruise condition. Five leading-edge radius distributions are described, and the potential performance benefit including the Reynolds number sensitivity for each is presented. Computational results for two leading-edge radius distributions are compared with experimental results acquired in the National Transonic Facility over a broad Reynolds number range.

  15. Effects of airfoil shape, thickness, camber, and angle of attack on calculated transonic unsteady airloads

    NASA Technical Reports Server (NTRS)

    Batina, J. T.

    1985-01-01

    The effects of airfoil shape, thickness, camber, and mean angle of attack on transonic unsteady airloads were investigated as calculated by the transonic small-disturbance computer code XTRAN2L. Shape effects were investigated by examining the pressure distributions, shock locations, and unsteady airloads for three 10 percent thick airfoils. NACA 0010, NACA 64A010, and parabolic arc. Thickness effects were determined by studying a single airfoil shape with three different thicknesses: NACA 0008, NACA 0010, and NACA 0012. Angle of attack and camber effects were studied by including mean angle of attack or by adding a simple parabolic camber distribution to the originally symmetric airfoils. Comparisons of unsteady airloads for different airfoil configurations show similar results caused by variations in airfoil shape, thickness, camber, or mean angle of attack. Computer costs can be reduced by limiting the number of transonic unsteady aerodynamic calculations for small changes in airfoil geometry or angle of attack.

  16. Algorithm developments for the Euler equations with calculations of transonic flows

    NASA Technical Reports Server (NTRS)

    Goorjian, Peter M.

    1987-01-01

    A new algorithm has been developed for the Euler equations that uses flux vector splitting in combination with the concept of rotating the coordinate system to the local streamwise direction. Flux vector biasing is applied along the local streamwise direction and central differencing is used transverse to the flow direction. The flux vector biasing is switched from upwind for supersonic flow to downwind-biased for subsonic flow. This switching is based on the Mach number; hence the proper domain of dependence is used in the supersonic regions and the switching occurs across shock waves. The theoretical basis and the development of the formulas for flux vector splitting are presented. Then several one-dimensional calculations are presented of steady and unsteady transonic flows, which demonstrate the stability and accuracy of the algorithm. Finally results are shown for unsteady transonic flow over an airfoil. The pressure coefficient plots show sharp transonic shock profiles, and the Mach contour plots show smoothly varying contours.

  17. Subsonic and transonic low-Reynolds-number airfoils with reduced pitching moments

    NASA Technical Reports Server (NTRS)

    Van Dam, C. P.; Hicks, R.; Reuther, J.

    1990-01-01

    A subsonic and a transonic airfoil are presented for application in a high-altitude long-endurance aircraft and a very-high-altitude aircraft, respectively. The subsonic airfoil is designed for a lift coefficient c(l) = 1.4 at a chord Reynolds number Re = 700,000 and a very low Mach number. The transonic airfoil is designed for c(l) = 1.0 at Re = 500,000 and a transonic Mach number M = 0.7. Both airfoils are developed to perform as well or better than previously designed airfoils. However, the present airfoils are developed for a constrained pitching moment to reduce aircraft trim drag and to relieve, to some extent, the torsional loads in the typically high-aspect-ratio wings. The beneficial effects of a cruise flap and of boundary-layer transition control on the off-design performance characteristics are illustrated.

  18. CFD Predictions for Transonic Performance of the ERA Hybrid Wing-Body Configuration

    NASA Technical Reports Server (NTRS)

    Deere, Karen A.; Luckring, James M.; McMillin, S. Naomi; Flamm, Jeffrey D.; Roman, Dino

    2016-01-01

    A computational study was performed for a Hybrid Wing Body configuration that was focused at transonic cruise performance conditions. In the absence of experimental data, two fully independent computational fluid dynamics analyses were conducted to add confidence to the estimated transonic performance predictions. The primary analysis was performed by Boeing with the structured overset-mesh code OVERFLOW. The secondary analysis was performed by NASA Langley Research Center with the unstructured-mesh code USM3D. Both analyses were performed at full-scale flight conditions and included three configurations customary to drag buildup and interference analysis: a powered complete configuration, the configuration with the nacelle/pylon removed, and the powered nacelle in isolation. The results in this paper are focused primarily on transonic performance up to cruise and through drag rise. Comparisons between the CFD results were very good despite some minor geometric differences in the two analyses.

  19. Extension of transonic flow computational concepts in the analysis of cavitated bearings

    NASA Technical Reports Server (NTRS)

    Vijayaraghavan, D.; Keith, T. G., Jr.; Brewe, D. E.

    1990-01-01

    An analogy between the mathematical modeling of transonic potential flow and the flow in a cavitating bearing is described. Based on the similarities, characteristics of the cavitated region and jump conditions across the film reformation and rupture fronts are developed using the method of weak solutions. The mathematical analogy is extended by utilizing a few computational concepts of transonic flow to numerically model the cavitating bearing. Methods of shock fitting and shock capturing are discussed. Various procedures used in transonic flow computations are adapted to bearing cavitation applications, for example, type differencing, grid transformation, an approximate factorization technique, and Newton's iteration method. These concepts have proved to be successful and have vastly improved the efficiency of numerical modeling of cavitated bearings.

  20. Detailed flow measurements and predictions for a three-stage transonic fan

    NASA Astrophysics Data System (ADS)

    Calvert, W. J.; Stapleton, A. W.

    1994-04-01

    Detailed flow measurements were taken at DRA Pyestock on a Rolls-Royce three-stage transonic research fan using advanced laser transit velocimetry and holography techniques to supplement the fixed pressure and temperature instrumentation. The results have been compared with predictions using the DRA S1-S2 quasi-three-dimensional flow calculation system at a range of speeds. The agreement was generally encouraging, both for the overall performance and for details of the internal flow such as positions of shock waves. Taken together with the computational efficiency of the calculations and previous experience on single-stage transonic fans and core compressors, this establishes the S1-S2 system as a viable design tool for future multistage transonic fans.