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Sample records for aedc 16-foot transonic

  1. 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.

  2. 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.

  3. 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.

  4. 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.

  5. ACT Missile Model In Langley 16 Foot Transonic Tunnel

    NASA Technical Reports Server (NTRS)

    1994-01-01

    The photograph shows a 15-percent scale model of the ACT advanced missile concept in the Langley 16-Foot Transonic Tunnel. The model featured independently controlled reaction jets near the nose and the tail of the model. Aerodynamic control was provided by four fins that were located near the tail.

  6. 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.

  7. 16-foot transonic tunnel test section flowfield survey

    NASA Technical Reports Server (NTRS)

    Yetter, J. A.; Abeyounis, W. K.

    1994-01-01

    A flow survey has been made of the test section of the NASA Langley Research Center 16-Foot Transonic Tunnel at subsonic and supersonic speeds. The survey was performed using five five-hole pyramid-head probes mounted at 14 inch intervals on a survey rake. Probes were calibrated at freestream Mach numbers from 0.50 to 0.95 and from 1.18 to 1.23. Flowfield surveys were made at Mach numbers from 0.50 to 0.90 and at Mach 1.20. The surveys were made at tunnel stations 130.6, 133.6, and 136.0. By rotating the survey rake through 180 degrees, a cylindrical volume of the test section 4.7 feet in diameter and 5.4 feet long centered about the tunnel centerline was surveyed. Survey results showing the measured test section upflow and sideflow characteristics and local Mach number distributions are presented. The report documents the survey probe calibration techniques used, summarizes the procedural problems encountered during testing, and identifies the data discrepancies observed during the post-test data analysis.

  8. Computations for the 16-foot transonic tunnel, NASA, Langley Research Center, revision 1

    NASA Technical Reports Server (NTRS)

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

    1987-01-01

    The equations used by the 16 foot transonic tunnel in the data reduction programs are presented in eight modules. Each module consists of equations necessary to achieve a specific purpose. These modules are categorized in the following groups: tunnel parameters; jet exhaust measurements; skin friction drag; balance loads and model attitudes calculations; internal drag (or exit-flow distributions); pressure coefficients and integrated forces; thrust removal options; and turboprop options. This document is a companion document to NASA TM-83186, A User's Guide to the Langley 16 Foot Transonic Tunnel, August 1981.

  9. 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.

  10. 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.

  11. 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.

  12. Investigation of very low blockage ratio boattail models in the Langley 16-foot transonic tunnel

    NASA Technical Reports Server (NTRS)

    Reubush, D. E.

    1976-01-01

    An investigation at an angle of attack of 0 deg was conducted in a 16 foot transonic tunnel at Mach numbers from 0.4 to 1.05 to determine the limits in Mach number at which valid boattail pressure drag data may be obtained with very low blockage ratio bodies. Extreme care was exercised when examining any data taken at subsonic Mach numbers very near 1.0 and lower than the supersonic Mach number at which shock reflections miss the model. Boattail pressure coefficient distributions did not indicate any error, but when integrated boattail pressure drag data was plotted as a function of Mach number, data which were in error were identified.

  13. Comparison of interference-free numerical results with sample experimental data for the AEDC wall-interference model at transonic and subsonic flow conditions

    NASA Technical Reports Server (NTRS)

    Newman, P. A.; Allison, D. O.

    1974-01-01

    Numerical results obtained from two computer programs recently developed with NASA support and now available for use by others are compared with some sample experimental data taken on a rectangular-wing configuration in the AEDC 16-Foot Transonic Tunnel at transonic and subsonic flow conditions. This data was used in an AEDC investigation as reference data to deduce the tunnel-wall interference effects for corresponding data taken in a smaller tunnel. The comparisons were originally intended to see how well a current state-of-the-art transonic flow calculation for a simple 3-D wing agreed with data which was felt by experimentalists to be relatively interference-free. As a result of the discrepancies between the experimental data and computational results at the quoted angle of attack, it was then deduced from an approximate stress analysis that the sting had deflected appreciably. Thus, the comparisons themselves are not so meaningful, since the calculations must be repeated at the proper angle of attack. Of more importance, however, is a demonstration of the utility of currently available computational tools in the analysis and correlation of transonic experimental data.

  14. 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.

  15. Two-dimensional converging-diverging rippled nozzles at transonic speeds. [performed in the Langley 16-Foot Transonic Tunnel

    NASA Technical Reports Server (NTRS)

    Carlson, John R.; Asbury, Scott C.

    1994-01-01

    An experimental investigation was performed in the Langley 16-Foot Transonic tunnel to determine the effects of external and internal flap rippling on the aerodynamics of a nonaxisymmetric nozzle. Data were obtained at several Mach numbers from static conditions to 1.2 over a range of nozzle pressure ratios. Nozzles with chordal boattail angles of 10, 20, and 30 degrees, with and without surface rippling, were tested. No effect on discharge coefficient due to surface rippling was observed. Internal thrust losses due to surface rippling were measured and attributed to a combination of additional internal skin friction and shock losses. External nozzle drag for the baseline configurations were generally less than that for the rippled configurations at all free-stream Mach numbers tested. The difference between the baseline and rippled nozzle drag levels generally increased with increasing boat tail angle. The thrust-minus-drag level for each rippled nozzle configuration was less than the equivalent baseline configuration for each Mach number at the design nozzle pressure ratio.

  16. 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.

  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. 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.

  19. 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

  20. 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.

  1. Impingement of Boundary-Reflected Disturbances Originating at the Nose of a Body of Revolution in the Langley Research Center 16-Foot Transonic Tunnel

    NASA Technical Reports Server (NTRS)

    Re, Richard, J.; Capone, Francis J.

    1998-01-01

    An investigation has been conducted in the Langley 16-Foot Transonic Tunnel to determine boundary-reflected disturbance lengths at low supersonic Mach numbers in the octagonally shaped test section. A body of revolution that had a nose designed to produce a bow shock and flow field similar to that about the nose of a supersonic transport configuration was used. The impingement of reflected disturbances on the model was determined from static pressures measured on the surface of the model. Test variables included Mach number (0.90 to 1.25), model angle of attack (nominally -10, 0, and 10), and model roll angle.

  2. Aeropropulsive characteristics of twin nonaxisymmetric vectoring nozzles installed with forward-swept and aft-swept wings. [in the Langley 16 Foot Transonic Tunnel

    NASA Technical Reports Server (NTRS)

    Capone, F. J.

    1981-01-01

    An investigation was conducted in the Langley 16 Foot Transonic Tunnel to determine the aeropropulsive characteristics of a single expansion ramp nozzle (SERN) and a two dimensional convergent divergent nozzle (2-D C-D) installed with both an aft swept and a forward swept wing. The SERN was tested in both an upright and an inverted position. The effects of thrust vectoring at nozzle vector angles from -5 deg to 20 deg were studied. This investigation was conducted at Mach numbers from 0.40 to 1.20 and angles of attack from -2.0 deg to 16 deg. Nozzle pressure ratio was varied from 1.0 (jet off) to about 9.0. Reynolds number based on the wing mean geometric chord varied from about 3 million to 4.8 million, depending upon free stream number.

  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. Effect of simulated in-flight thrust reversing on vertical-tail loads of F-18 and F-15 airplane models. [conducted in the Langley 16-foot transonic tunnel

    NASA Technical Reports Server (NTRS)

    Bare, E. A.; Berrier, B. L.; Capone, F. J.

    1981-01-01

    Investigations were conducted in the Langley 16-Foot Transonic Tunnel to provide data on a 0.10-scale model of the prototype F-18 airplane and a 0.047-scale model of the F-15 three-surface configuration (canard, wing, and horizontal tails). Test data were obtained at static conditions and at Mach numbers from 0.6 to 1.2 over an angle-of-attack range from 2 deg to 15 deg. Nozzle pressure ratio was varied from jet off to about 8.0.

  6. 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.

  7. Results from a Sting Whip Correction Verification Test at the Langley 16-Foot Transonic Tunnel

    NASA Technical Reports Server (NTRS)

    Crawford, B. L.; Finley, T. D.

    2002-01-01

    In recent years, great strides have been made toward correcting the largest error in inertial Angle of Attack (AoA) measurements in wind tunnel models. This error source is commonly referred to as 'sting whip' and is caused by aerodynamically induced forces imparting dynamics on sting-mounted models. These aerodynamic forces cause the model to whip through an arc section in the pitch and/or yaw planes, thus generating a centrifugal acceleration and creating a bias error in the AoA measurement. It has been shown that, under certain conditions, this induced AoA error can be greater than one third of a degree. An error of this magnitude far exceeds the target AoA goal of 0.01 deg established at NASA Langley Research Center (LaRC) and elsewhere. New sting whip correction techniques being developed at LaRC are able to measure and reduce this sting whip error by an order of magnitude. With this increase of accuracy, the 0.01 deg AoA target is achievable under all but the most severe conditions.

  8. Calculation of transonic aileron buzz

    NASA Technical Reports Server (NTRS)

    Steger, J. L.; Bailey, H. E.

    1979-01-01

    An implicit finite-difference computer code that uses a two-layer algebraic eddy viscosity model and exact geometric specification of the airfoil has been used to simulate transonic aileron buzz. The calculated results, which were performed on both the Illiac IV parallel computer processor and the Control Data 7600 computer, are in essential agreement with the original expository wind-tunnel data taken in the Ames 16-Foot Wind Tunnel just after World War II. These results and a description of the pertinent numerical techniques are included.

  9. An investigation of the aerodynamic characteristics of a 0.00548 scale model (model no. 486) of the space shuttle 146-inch diameter solid rocket booster at angels of attack from 113 deg to 180 deg in the AEDC PWT 4-foot transonic wind tunnel (SA16F)

    NASA Technical Reports Server (NTRS)

    Ramsey, P. E.

    1976-01-01

    An experimental investigation (SA16F) was conducted in the AEDC PWT 4T to determine the entry static stability of a 0.00548 scale space shuttle solid rocket booster (SRB). The primary objective was to improve the definition of the aerodynamic characteristics in the angle of attack range beyond 90 deg in the vicinity of the entry trim point. The SRB scale model consisted of the reentry configuration with all major protuberances. A simulated heat shield around the engine nozzle was also included. Data were obtained for a 60 deg side mounted sting and a straight nose mounted sting. The angle of attack range for the side mounted sting was 113 deg to 147 deg and for the nose mounted sting 152 deg to 187 deg. The Mach number range consisted of 0.4 to 1.2 at roll angles of 0 and 90 deg. The resulting 6-component aerodynamic force data was presented as the variation of coefficients with angle of attack for each Mach number and roll angle.

  10. AEDC sensor T and E methodology

    NASA Astrophysics Data System (ADS)

    1995-07-01

    For space surveillance systems it is extremely costly, and, in some cases, impractical to provide effective field testing to evaluate operational issues. Even with sufficient systems on orbit, there are significant problems in providing the threat targets. Not only is the launch of a target missile very expensive, but it is also impractical to launch eight or ten missiles simultaneously. System performance is measured at only one set of parameters, and the possibility of acquiring zero data is real. Operational testing of space assets in ground test facilities has been limited as well by the lack of adequate target and threat simulations. AEDC's new direct write scene generation (DWSG) capabilities have overcome this limitation. These capabilities, along with recently upgraded state-of-the-art sensor test facilities and closed-loop DWSG developments, can provide ground-based testing at all levels of system acquisition. This report describes the proposed new approach to surveillance and seeker testing for the next century, which includes the marriage of the traditional DT&E with a new capability to perform early operational assessments and operational T&E prior to deployment. This T&E methodology takes full advantage of existing DoD facility investments, provides testing at lower costs, enables better investigation of the sensor design envelope with actual hardware, allows concurrent DT&E and OT&E, and reduces flight risk for a program.

  11. Enthalpy probe measurements in AEDC`s arc-heated test facilities

    SciTech Connect

    Carver, D.B.; Fryer, J.M.

    1995-08-01

    The Arnold Engineering Development Center (AEDC) has recently demonstrated a new probe for measurement of enthalpy within the free jet of its arc-heated test units. The probe, developed under contract by SPARTA, Inc., is a steady-state device that uses the conservation of mass principle to define total enthalpy of the flow that enters the probe. Dual sonic orifices are used to meter the flow. A heat exchanger between the orifices reduces the energy so that the total temperature can be measured at the entrance to the second orifice. Plenum pressures are measured at the entrance to each orifice. Equating the flow rates through each orifice makes it possible to solve for the unknown - the total enthalpy entering the probe. Bonded foil technology is the innovation that permitted a steady-state probe to be developed; the probe`s intricate cooling passages provided protection from the harsh thermal environment. The probe has been successfully demonstrated at a stagnation pressure of 15 atm and total enthalpy of 2,700 Btu/lbm. The paper includes design concepts, operating principles, and a presentation of test results.

  12. Analysis and test of a 16-foot radial rib reflector developmental model

    NASA Technical Reports Server (NTRS)

    Birchenough, Shawn A.

    1989-01-01

    Analytical and experimental modal tests were performed to determine the vibrational characteristics of a 16-foot diameter radial rib reflector model. Single rib analyses and experimental tests provided preliminary information relating to the reflector. A finite element model predicted mode shapes and frequencies of the reflector. The analyses correlated well with the experimental tests, verifying the modeling method used. The results indicate that five related, characteristic mode shapes form a group. The frequencies of the modes are determined by the relative phase of the radial ribs.

  13. Performance characteristics of an isolated coannular plug nozzle at transonic speeds

    NASA Technical Reports Server (NTRS)

    Mercer, C. E.; Burley, J. R., II

    1985-01-01

    The Langley 16-Foot Transonic Tunnel was used to evaluate the performance characteristics of a coannular plug nozzle at static conditions (Mach number of 0) and at Mach numbers from 0.65 to 1.20. Jet total pressure ratio was varied from 1.0 (jet off) to 10.0. Thirty-seven configurations generated by the combination of three geometric variables - plug angle, shroud boattail length (fixed exit radius), and shroud extension length - were tested.

  14. Orbital Debris Assesment Tesing in the AEDC Range G

    NASA Technical Reports Server (NTRS)

    Polk, Marshall; Woods, David; Roebuck, Brian; Opiela, John; Sheaffer, Patti; Liou, J.-C.

    2015-01-01

    The space environment presents many hazards for satellites and spacecraft. One of the major hazards is hypervelocity impacts from uncontrolled man-made space debris. Arnold Engineering Development Complex (AEDC), The National Aeronautics and Space Administration (NASA), The United States Air Force Space and Missile Systems Center (SMC), the University of Florida, and The Aerospace Corporation configured a large ballistic range to perform a series of hypervelocity destructive impact tests in order to better understand the effects of space collisions. The test utilized AEDC's Range G light gas launcher, which is capable of firing projectiles up to 7 km/s. A non-functional full-scale representation of a modern satellite called the DebriSat was destroyed in the enclosed range enviroment. Several modifications to the range facility were made to ensure quality data was obtained from the impact events. The facility modifcations were intended to provide a high impact energy to target mass ratio (>200 J/g), a non-damaging method of debris collection, and an instrumentation suite capable of providing information on the physics of the entire imapct event.

  15. Testing of focal plane arrays at the AEDC

    NASA Astrophysics Data System (ADS)

    Nicholson, Randy A.; Mead, Kimberly D.; Smith, Robert W.

    1992-07-01

    A facility was developed at the Arnold Engineering Development Center (AEDC) to provide complete radiometric characterization of focal plane arrays (FPAs). The highly versatile facility provides the capability to test single detectors, detector arrays, and hybrid FPAs. The primary component of the AEDC test facility is the Focal Plane Characterization Chamber (FPCC). The FPCC provides a cryogenic, low-background environment for the test focal plane. Focal plane testing in the FPCC includes flood source testing, during which the array is uniformly irradiated with IR radiation, and spot source testing, during which the target radiation is focused onto a single pixel or group of pixels. During flood source testing, performance parameters such as power consumption, responsivity, noise equivalent input, dynamic range, radiometric stability, recovery time, and array uniformity can be assessed. Crosstalk is evaluated during spot source testing. Spectral response testing is performed in a spectral response test station using a three-grating monochromator. Because the chamber can accommodate several types of testing in a single test installation, a high throughput rate and good economy of operation are possible.

  16. Forward-swept wing configuration designed for high maneuverability by use of a transonic computational method

    NASA Technical Reports Server (NTRS)

    Mann, M. J.; Mercer, C. E.

    1986-01-01

    A transonic computational analysis method and a transonic design procedure have been used to design the wing and the canard of a forward-swept-wing fighter configuration for good transonic maneuver performance. A model of this configuration was tested in the Langley 16-Foot Transonic Tunnel. Oil-flow photographs were obtained to examine the wind flow patterns at Mach numbers from 0.60 to 0.90. The transonic theory gave a reasonably good estimate of the wing pressure distributions at transonic maneuver conditions. Comparison of the forward-swept-wing configuration with an equivalent aft-swept-wing-configuration showed that, at a Mach number of 0.90 and a lift coefficient of 0.9, the two configurations have the same trimmed drag. The forward-swept wing configuration was also found to have trimmed drag levels at transonic maneuver conditions which are comparable to those of the HiMAT (highly maneuverable aircraft technology) configuration and the X-29 forward-swept-wing research configuration. The configuration of this study was also tested with a forebody strake.

  17. Uncertainty analysis of the AEDC 7V chamber

    NASA Astrophysics Data System (ADS)

    Crider, Dustin; Lowry, Heard; Nicholson, Randy; Mead, Kimberly

    2005-05-01

    For over 30 years, the Space Systems Test Facility and space chambers at the Arnold Engineering Development Center (AEDC) have been used to perform space sensor characterization, calibration, and mission simulation testing of space-based, interceptor, and airborne sensors. In partnership with the Missile Defense Agency (MDA), capability upgrades are continuously pursued to keep pace with evolving sensor technologies. Upgrades to sensor test facilities require rigorous facility characterization and calibration activities that are part of AEDC's annual activities to comply with Major Range Test Facility Base processes to ensure quality metrology and test data. This paper discusses the ongoing effort to characterize and quantify Aerospace Chamber 7V measurement uncertainties. The 7V Chamber is a state-of-the-art cryogenic/vacuum facility providing calibration and high-fidelity mission simulation for infrared seekers and sensors against a low-infrared background. One of its key features is the high fidelity of the radiometric calibration process. Calibration of the radiometric sources used is traceable to the National Institute of Standards and Technology and provides relative uncertainties on the order of two to three percent, based on measurement data acquired during many test periods. Three types of sources of measurement error and top-level uncertainties have been analyzed; these include radiometric calibration, target position, and spectral output. The approach used and presented is to quantify uncertainties of each component in the optical system and then build uncertainty diagrams and easily updated databases to detail the uncertainty for each optical system. The formalism, equations, and corresponding analyses are provided to help describe how the specific quantities are derived and currently used. This paper presents the uncertainty methodology used and current results.

  18. 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.

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

    NASA Technical Reports Server (NTRS)

    Usry, J. W.; Wallace, J. W.

    1971-01-01

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

  20. 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.

  1. Inlet flow field investigation. Part 1: Transonic flow field survey

    NASA Technical Reports Server (NTRS)

    Yetter, J. A.; Salemann, V.; Sussman, M. B.

    1984-01-01

    A wind tunnel investigation was conducted to determine the local inlet flow field characteristics of an advanced tactical supersonic cruise airplane. A data base for the development and validation of analytical codes directed at the analysis of inlet flow fields for advanced supersonic airplanes was established. Testing was conducted at the NASA-Langley 16-foot Transonic Tunnel at freestream Mach numbers of 0.6 to 1.20 and angles of attack from 0.0 to 10.0 degrees. Inlet flow field surveys were made at locations representative of wing (upper and lower surface) and forebody mounted inlet concepts. Results are presented in the form of local inlet flow field angle of attack, sideflow angle, and Mach number contours. Wing surface pressure distributions supplement the flow field data.

  2. 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

    In 1955, work was started on the conversion of a subsonic wind tunnel to a 16-foot transonic tunnel with Freon-12 or air as the test medium. The new facility, designated the Transonic Dynamics Tunnel (TDT), became fully operational in 1960. A description is presented of aeroelastic testing and research performed in the TDT since 1960. It is pointed out that wind-tunnel tests of aeroelastic models require specialized experimental techniques seldom found in other types of wind-tunnel studies. Attention is given to model mount systems, launch vehicle models, aircraft models, aircraft buffet, gust response, stability derivative measurements, and subcritical testing techniques. Aspects of vehicle development testing are considered along with aeroelastic 'fixes', aeroelastic 'surprises', approaches for controlling aeroelastic effects, and unsteady pressure measurements.

  3. Pitot pressure measurements in flow fields behind circular-arc nozzles with exhaust jets at subsonic free-stream Mach numbers. [langley 16 foot transonic tunnel

    NASA Technical Reports Server (NTRS)

    Mason, M. L.; Putnam, L. E.

    1979-01-01

    The flow field behind a circular arc nozzle with exhaust jet was studied at subsonic free stream Mach numbers. A conical probe was used to measure the pitot pressure in the jet and free stream regions. Pressure data were recorded for two nozzle configurations at nozzle pressure ratios of 2.0, 2.9, and 5.0. At each set of test conditions, the probe was traversed from the jet center line into the free stream region at seven data acquisition stations. The survey began at the nozzle exit and extended downstream at intervals. The pitot pressure data may be applied to the evaluation of computational flow field models, as illustrated by a comparison of the flow field data with results of inviscid jet plume theory.

  4. The AEDC aerospace chamber 7V: An advanced test capability for infrared surveillance and seeker sensors

    NASA Technical Reports Server (NTRS)

    Simpson, W. R.

    1994-01-01

    An advanced sensor test capability is now operational at the Air Force Arnold Engineering Development Center (AEDC) for calibration and performance characterization of infrared sensors. This facility, known as the 7V, is part of a broad range of test capabilities under development at AEDC to provide complete ground test support to the sensor community for large-aperture surveillance sensors and kinetic kill interceptors. The 7V is a state-of-the-art cryo/vacuum facility providing calibration and mission simulation against space backgrounds. Key features of the facility include high-fidelity scene simulation with precision track accuracy and in-situ target monitoring, diffraction limited optical system, NIST traceable broadband and spectral radiometric calibration, outstanding jitter control, environmental systems for 20 K, high-vacuum, low-background simulation, and an advanced data acquisition system.

  5. Transonic airfoil design code

    NASA Technical Reports Server (NTRS)

    Bauer, F.; Garabedian, P.; Korn, D.

    1980-01-01

    Program aids in design of shockless airfoils, assists development of fuel-conserving, supercritical wings. Algorithm calculates approximate airfoil shape given prescribed pressure distribution. This allows design of families of transonic airfoils for use in aircraft wings or turbine and compressor blades. Program is written in FORTRAN IV for batch execution on CDC-6000.

  6. 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.

  7. 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.

  8. Projection technologies for imaging sensor calibration, characterization, and HWIL testing at AEDC

    NASA Astrophysics Data System (ADS)

    Lowry, H. S.; Breeden, M. F.; Crider, D. H.; Steely, S. L.; Nicholson, R. A.; Labello, J. M.

    2010-04-01

    The characterization, calibration, and mission simulation testing of imaging sensors require continual involvement in the development and evaluation of radiometric projection technologies. Arnold Engineering Development Center (AEDC) uses these technologies to perform hardware-in-the-loop (HWIL) testing with high-fidelity complex scene projection technologies that involve sophisticated radiometric source calibration systems to validate sensor mission performance. Testing with the National Institute of Standards and Technology (NIST) Ballistic Missile Defense Organization (BMDO) transfer radiometer (BXR) and Missile Defense Agency (MDA) transfer radiometer (MDXR) offers improved radiometric and temporal fidelity in this cold-background environment. The development of hardware and test methodologies to accommodate wide field of view (WFOV), polarimetric, and multi/hyperspectral imaging systems is being pursued to support a variety of program needs such as space situational awareness (SSA). Test techniques for the acquisition of data needed for scene generation models (solar/lunar exclusion, radiation effects, etc.) are also needed and are being sought. The extension of HWIL testing to the 7V Chamber requires the upgrade of the current satellite emulation scene generation system. This paper provides an overview of pertinent technologies being investigated and implemented at AEDC.

  9. Analysis of heat-transfer measurements from 2 AEDC wind tunnels on the Shuttle external tank

    NASA Technical Reports Server (NTRS)

    Nutt, K. W.

    1984-01-01

    Previous aerodynamic heating tests have been conducted in the AEDC/VKF Supersonic Wind Tunnel (A) to aid in defining the design thermal environment for the space shuttle external tank. The quality of these data has been under discussion because of the effects of low tunnel enthalpy and slow model injection rates. Recently the AEDC/VKF Hypersonic Wind Tunnel (C) has been modified to provide a Mach 4 capability that has significantly higher tunnel enthalpy with more rapid model injection rates. Tests were conducted in Tunnel C at Mach 4 to obtain data on the external tank for comparison with Tunnel A results. Data were obtained on a 0.0175 scale model of the Space Shuttle Integrated Vehicle at Re/ft = 4 x 10 to the 6th power with the tunnel stagnation temperature varying from 740 to 1440 R. Model attitude varied from an angle of attack of -5 to 5 deg and an angle of sideslip of -3 to 3 deg. One set of data was obtained in Tunnel C at Re/ft = 6.9 x 10 to the 6th for comparison with flight data. Data comparisons between the two tunnels for numerous regions on the external tank are given.

  10. Airfoil modification effects on subsonic and transonic pressure distributions and performance for the EA-6B airplane

    NASA Technical Reports Server (NTRS)

    Allison, Dennis O.; Sewall, William G.

    1995-01-01

    Longitudinal characteristics and wing-section pressure distributions are compared for the EA-6B airplane with and without airfoil modifications. The airfoil modifications were designed to increase low-speed maximum lift for maneuvering, while having a minimal effect on transonic performance. Section contour changes were confined to the leading-edge slat and trailing-edge flap regions of the wing. Experimental data are analyzed from tests in the Langley 16-Foot Transonic Tunnel on the baseline and two modified wing-fuselage configurations with the slats and flaps in their retracted positions. Wing modification effects on subsonic and transonic performance are seen in wing-section pressure distributions of the various configurations at similar lift coefficients. The modified-wing configurations produced maximum lift coefficients which exceeded those of the baseline configuration at low-speed Mach numbers (0.300 and 0.400). This benefit was related to the behavior of the wing upper surface leading-edge suction peak and the behavior of the trailing-edge pressure. At transonic Mach numbers (0.725 to 0.900), the wing modifications produced a somewhat stronger nose-down pitching moment, a slightly higher drag at low-lift levels, and a lower drag at higher lift levels.

  11. Transonic swirling nozzle flow

    NASA Technical Reports Server (NTRS)

    Keith, Theo G., Jr.; Pawlas, Gary E.

    1991-01-01

    A numerical model of viscous transonic swirling flow in axisymmetric nozzles is developed. MacCormack's implicit Gauss-Seidel method is applied to the thin-layer Navier-Stokes equations in transformed coordinates. Numerical results are compared with experimental data to validate the method. The effect of swirl and viscosity on nozzle performance are demonstrated by examining wall pressures, Mach contours, and integral parameters.

  12. 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.

  13. Effect of afterbody geometry on aerodynamic characteristics of isolated nonaxisymmetric afterbodies at transonic Mach numbers

    NASA Technical Reports Server (NTRS)

    Bangert, Linda S.; Carson, George T., Jr.

    1992-01-01

    A parametric study was conducted in the Langley 16-Foot Transonic Tunnel on an isolated nonaxisymmetic fuselage model that simulates a twin-engine fighter. The effects of aft-end closure distribution (top/bottom) nozzle-flap boattail angle versus nozzle-sidewall boattail angle) and afterbody and nozzle corner treatment (sharp or radius) were investigated. Four different closure distributions with three different corner radii were tested. Tests were conducted over a range of Mach numbers from 0.40 to 1.25 and over a range of angles of attack from -3 to 9 degrees. Solid plume simulators were used to simulate the jet exhaust. For a given closure distribution in the range of Mach numbers tested, the sharp-corner nozzles generally had the highest drag, and the 2-in. corner-radius nozzles generally had the lowest drag. The effect of closure distribution on afterbody drag was highly dependent on configuration and flight condition.

  14. Measurement of flow fields in a large transonic wind tunnel using a laser velocimeter

    NASA Technical Reports Server (NTRS)

    Putnam, L. E.; Meyers, J. F.

    1975-01-01

    An investigation to determine the feasibility of using a laser velocimeter for measuring the mean flow velocities about airplane models in the Langley 16-foot transonic tunnel has recently been completed. The laser velocimeter was a two-component fringe-type used in the back scatter mode. The tunnel airflow was seeded with oil droplets to provide scattering sources for the laser velocimeter. Measurements of the tunnel free-stream velocity were in good agreement with the tunnel calibration, and measurements of the velocity along the stagnating streamline of a hemisphere model, when adjusted for particle lag, were in good agreement with theoretical predictions. The study showed that the laser, optics, and electronics system operated satisfactorily, but that further development is required to reduce scattering particle size and substance accumulation in the tunnel.

  15. 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.

  16. Transonic Flow Past Cone Cylinders

    NASA Technical Reports Server (NTRS)

    Solomon, George E

    1955-01-01

    Experimental results are presented for transonic flow post cone-cylinder, axially symmetric bodies. The drag coefficient and surface Mach number are studied as the free-stream Mach number is varied and, wherever possible, the experimental results are compared with theoretical predictions. Interferometric results for several typical flow configurations are shown and an example of shock-free supersonic-to-subsonic compression is experimentally demonstrated. The theoretical problem of transonic flow past finite cones is discussed briefly and an approximate solution of the axially symmetric transonic equations, valid for a semi-infinite cone, is presented.

  17. Transonic conical flow

    NASA Technical Reports Server (NTRS)

    Agopian, K. G.

    1974-01-01

    The problem of inviscid, steady transonic conical flow, formulated in terms of the small disturbance theory, is studied. The small disturbance equation and similarity rules are presented, and a boundary value problem is formulated for the case of a supersonic freestream Mach number. The equation for the perturbation potential is solved numerically using an elliptic finite difference system. The difference equations are solved with a point relaxation algorithm that is also capable of capturing the shock wave during the iteration procedure by using the boundary conditions at the shock. Numerical calculations, for shock location, pressure distribution and drag coefficient, are presented for a family of nonlifting conical wings. The theory of slender wings is also presented and analytical results for pressure and drag coefficients are obtained.

  18. Effects of varying podded nacelle-nozzle installations on transonic aeropropulsive characteristics of a supersonic fighter aircraft

    NASA Technical Reports Server (NTRS)

    Capone, F. J.; Reubush, D. E.

    1983-01-01

    The aeropropulsive characteristics of an advanced twin engine fighter designed for supersonic cruise was investigated in the 16 foot Transonic Tunnel. The performance characteristics of advanced nonaxisymmetric nozzles installed in various nacelle locations, the effects of thrust induced forces on overall aircraft aerodynamics, the trim characteristics, and the thrust reverser performance were evaluated. The major model variables included nozzle power setting; nozzle duct aspect ratio; forward, mid, and aft nacelle axial locations; inboard and outboard underwing nacelle locations; and underwing and overwing nacelle locations. Thrust vectoring exhaust nozzle configurations included a wedge nozzle, a two dimensional convergent divergent nozzle, and a single expansion ramp nozzle, each with deflection angles up to 30 deg. In addition to the nonaxisymmetric nozzles, an axisymmetric nozzle installation was also tested. The use of a canard for trim was also assessed.

  19. Transonic aerodynamic characteristics of a supersonic cruise aircraft research model with the engines suspended above the wing

    NASA Technical Reports Server (NTRS)

    Mercer, C. E.; Carson, G. T., Jr.

    1979-01-01

    The influence of upper-surface nacelle exhaust flow on the aerodynamic characteristics of a supersonic cruise aircraft research configuration was investigated in a 16 foot transonic tunnel over a range of Mach numbers from 0.60 to 1.20. The arrow-wing transport configuration with engines suspended over the wing was tested at angles of attack from -4 deg to 6 deg and jet total pressure ratios from 1 to approximately 13. Wing-tip leading edge flap deflections of -10 deg to 10 deg were tested with the wing-body configuration. Various nacelle locations (chordwise, spanwise, and vertical) were tested over the ranges of Mach numbers, angles of attack, and jet total-pressure ratios. The results show that reflecting the wing-tip leading edge flap from 0 deg to -10 deg increased the maximum lift-drag ratio by 1.0 at subsonic speeds. Jet exhaust interference effects were negligible.

  20. Continued Development of a Global Heat Transfer Measurement System at AEDC Hypervelocity Wind Tunnel 9

    NASA Technical Reports Server (NTRS)

    Kurits, Inna; Lewis, M. J.; Hamner, M. P.; Norris, Joseph D.

    2007-01-01

    Heat transfer rates are an extremely important consideration in the design of hypersonic vehicles such as atmospheric reentry vehicles. This paper describes the development of a data reduction methodology to evaluate global heat transfer rates using surface temperature-time histories measured with the temperature sensitive paint (TSP) system at AEDC Hypervelocity Wind Tunnel 9. As a part of this development effort, a scale model of the NASA Crew Exploration Vehicle (CEV) was painted with TSP and multiple sequences of high resolution images were acquired during a five run test program. Heat transfer calculation from TSP data in Tunnel 9 is challenging due to relatively long run times, high Reynolds number environment and the desire to utilize typical stainless steel wind tunnel models used for force and moment testing. An approach to reduce TSP data into convective heat flux was developed, taking into consideration the conditions listed above. Surface temperatures from high quality quantitative global temperature maps acquired with the TSP system were then used as an input into the algorithm. Preliminary comparison of the heat flux calculated using the TSP surface temperature data with the value calculated using the standard thermocouple data is reported.

  1. Liquid Rocket Engine Testing - Historical Lecture: Simulated Altitude Testing at AEDC

    NASA Technical Reports Server (NTRS)

    Dougherty, N. S.

    2010-01-01

    The span of history covered is from 1958 to the present. The outline of this lecture draws from historical examples of liquid propulsion testing done at AEDC primarily for NASA's Marshall Space Flight Center (NASA/MSFC) in the Saturn/Apollo Program and for USAF Space and Missile Systems dual-use customers. NASA has made dual use of Air Force launch vehicles, Test Ranges and Tracking Systems, and liquid rocket altitude test chambers / facilities. Examples are drawn from the Apollo/ Saturn vehicles and the testing of their liquid propulsion systems. Other examples are given to extend to the family of the current ELVs and Evolved ELVs (EELVs), in this case, primarily to their Upper Stages. The outline begins with tests of the XLR 99 Engine for the X-15 aircraft, tests for vehicle / engine induced environments during flight in the atmosphere and in Space, and vehicle staging at high altitude. The discussion is from the author's perspective and background in developmental testing.

  2. Transonic airframe propulsion integration

    NASA Technical Reports Server (NTRS)

    Coltrin, Robert E.; Sanders, Bobby W.; Bencze, Daniel P.

    1992-01-01

    This chart shows the time line for HSR propulsion/airframe integration program. HSR Phase 1 efforts are underway in both propulsion and aerodynamics. The propulsion efforts focus on cycles, inlets combustors and nozzles that will be required to reduce nitrogen oxide (NOX) at cruise and noise at takeoff and landing to acceptable levels. The aerodynamic efforts concentrate on concepts that will reduce sonic booms and increase the lift/drag (L/D) ratio for the aircraft. The Phase 2 critical propulsion component technology program will focus on large scale demonstrators of the inlet, fan, combustor, and nozzle. The hardware developed here will feed into the propulsion system program which will demonstrate overall system technology readiness, particularly in the takeoff and supersonic cruise speed ranges. The Phase 2 aerodynamic performance and vehicle integration program will provide a validated data base for advanced airframe/control/integration concepts over the full HSR speed range. The results of this program will also feed into the propulsion system demonstration program, particularly in the critical transonic arena.

  3. Jump conditions in transonic equilibria

    SciTech Connect

    Guazzotto, L.; Betti, R.; Jardin, S. C.

    2013-04-15

    In the present paper, the numerical calculation of transonic equilibria, first introduced with the FLOW code in Guazzotto et al.[Phys. Plasmas 11, 604 (2004)], is critically reviewed. In particular, the necessity and effect of imposing explicit jump conditions at the transonic discontinuity are investigated. It is found that 'standard' (low-{beta}, large aspect ratio) transonic equilibria satisfy the correct jump condition with very good approximation even if the jump condition is not explicitly imposed. On the other hand, it is also found that high-{beta}, low aspect ratio equilibria require the correct jump condition to be explicitly imposed. Various numerical approaches are described to modify FLOW to include the jump condition. It is proved that the new methods converge to the correct solution even in extreme cases of very large {beta}, while they agree with the results obtained with the old implementation of FLOW in lower-{beta} equilibria.

  4. 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.

  5. Role Of High Speed Photography In The Testing Capabilities Of The Arnold Engineering Development Center (AEDC) Range And Track Facilities

    NASA Astrophysics Data System (ADS)

    Hendrix, Roy E.; Dugger, Paul H.

    1983-03-01

    Since the onset of user testing in the AEDC aeroballistic ranges in 1961, concentrated efforts in such areas as model launching techniques, test environment simulation, and specialized instrumentation have been made to enhance the usefulness of these test facilities. A wide selection of specialized instrumentation has been developed over the years to provide, among other features, panoramic photographic coverage of test models during flight. Pulsed ruby lasers, xenon flash lamps, visible-light spark sources, and flash X-ray systems are employed as short-duration radiation sources in various front-light and back-light photographic systems. Visible-light and near infrared image intensifier diodes are used to achieve high-speed shuttering in photographic pyrometry systems that measure surface temperatures of test models in flight. Turbine-driven framing cameras are used to provide multiframe photography of such high-speed phenomena as impact debris formation and model encounter with erosive fields. As a result, the capabilities of these ballistic range test units have increased significantly in regard to the types of tests that can be accommodated and to the quality and quantity of data that can be provided. Presently, five major range and companion track facilities are active in conducting hypervelocity testing in AEDC's von K6rman Gas Dynamics Facility (VKF): Ranges G, K, and S-1 and Tracks G and K. The following types of tests are conducted in these test units: ablation/erosion, transpiration-cooled nosetip (TCNT), nosetip transition, heat transfer, aerodynamic, cannon projectile, rocket contrail, reentry physics, and hypervelocity impact. The parallel achievements in high-speed photography and testing capabilities are discussed, and the significant role of photographic systems in the development of the overall testing capabilities of the AEDC range and track facilities is illustrated in numerous examples of photographic results.

  6. 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.

  7. Development and installation of a continuous water-monitoring system for the AEDC. Final report, Mar 92-May 92

    SciTech Connect

    Przybyciel, M.; Behm, J.; Sampey, T.

    1992-08-01

    A system to sample and analyze water from Rowland Creek at AEDC for hydrocarbon contaminants has been developed under a Small Business Innovation Research (SBIR) program contract. The online continuous water monitoring system involves the combination of two gas chromatographs (GC). The first instrument combines a gas/liquid sparger and a GC. The sparger uses an inert gas, helium, to remove and concentrate volatile organic chemicals which are then sequentially analyzed using a GC. The second gas chromatographic instrument involves the analysis of nonvolatile water samples through direct injection of water.

  8. A Transonic Wind-Tunnel Investigation of the Longitudinal Aerodynamic Characteristics of a Model of the Lockheed XF-104 Airplane

    NASA Technical Reports Server (NTRS)

    Hieser, Gerald; Reid, Charles F.

    1954-01-01

    The transonic longitudinal aerodynamic characteristics of a 0.0858-scale model of the Lockheed XF-104 airplane have been obtained from tests at the Langley 16-foot transonic tunnel. The results of the investigation provide some general information applicable to the transonic properties of thin, low-aspect-ratio, unswept wing configurations utilizing a high horizontal tail . The model employs a horizontal tail mounted at the top of the vertical tail and a wing with an aspect ratio of 2.5, a taper ratio of 0.385, and 3.4-percent-thick airfoil sections. The lift, drag, and static longitudinal pitching moment were measured at Mach numbers from 0.80 t o 1.09 and angles of attack from -2.5 deg to 22.5 deg. Some of the dynamic longitudinal stability properties of the airplane have been predicted from the test results. In addition, some visual flow studies on the wing surfaces obtained at Mach numbers of 0.80 and 1.00 are included. Results of the investigation show that the transonic rise in drag coefficient at zero lift is about 0.030. At high angles of attack, the model becomes longitudinally unstable at Mach numbers from 0.80 t o 0.90, whereas a reduction in static stability is experienced when very high angles of attack are reached at Mach numbers above 0.90. Longitudinal dynamic stability calculations show that the longitudinal control is good at angles of attack below the unstable break in the static pitching-moment curves, but a typical corrective control applied after the occurrence of neutral stability has little effect in averting pitch-up.

  9. Wind-Tunnel Measurements of Effect of Dive-Recovery Flaps at Transonic Speeds on Models of a Seaplane and a Transport

    NASA Technical Reports Server (NTRS)

    Heath, Atwood R., Jr.; Ward, Robert J.

    1959-01-01

    The effects of wing-lower-surface dive-recovery flaps on the aero- dynamic characteristics of a transonic seaplane model and a transonic transport model having 40 deg swept wings have been investigated in the Langley 16-foot transonic tunnel. The seaplane model had a wing with an aspect ratio of 5.26, a taper ratio of 0.333, and NACA 63A series airfoil sections streamwise. The transport model had a wing with an aspect ratio of 8, a taper ratio of 0.3, and NACA 65A series airfoil sections perpendicular to the quarter-chord line. The effects of flap deflection, flap longitudinal location, and flap sweep were generally investigated for both horizontal-tail-on and horizontal-tail-off configurations. Model force and moment measurements were made for model angles of attack from -5 deg to 14 deg in the Mach number range from 0.70 to 1.075 at Reynolds numbers of 2.95 x 10(exp 6) to 4.35 x 10(exp 6). With proper longitudinal location, wing-lower-surface dive-recovery flaps produced lift and pitching-moment increments that increased with flap deflection. For the transport model a flap located aft on the wing proved to be more effective than one located more forward., both flaps having the same span and approximately the same deflection. For the seaplane model a high horizontal tail provided added effectiveness for the deflected-flap configuration.

  10. 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.

  11. 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.

  12. 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.

  13. 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.

  14. Development of HWIL Testing Capabilities for Satellite Target Emulation at AEDC

    NASA Astrophysics Data System (ADS)

    Lowry, H.; Crider, D.; Burns, J.; Thompson, R.; Goldsmith, G., II; Sholes, W.

    Programs involved in Space Situational Awareness (SSA) need the capability to test satellite sensors in a Hardware-in-the-Loop (HWIL) environment. Testing in a ground system avoids the significant cost of on-orbit test targets and the resulting issues such as debris mitigation, and in-space testing implications. The space sensor test facilities at AEDC consist of cryo-vacuum chambers that have been developed to project simulated targets to air-borne, space-borne, and ballistic platforms. The 7V chamber performs calibration and characterization of surveillance and seeker systems, as well as some mission simulation. The 10V chamber is being upgraded to provide real-time target simulation during the detection, acquisition, discrimination, and terminal phases of a seeker mission. The objective of the Satellite Emulation project is to upgrade this existing capability to support the ability to discern and track other satellites and orbital debris in a HWIL capability. It would provide a baseline for realistic testing of satellite surveillance sensors, which would be operated in a controlled environment. Many sensor functions could be tested, including scene recognition and maneuvering control software, using real interceptor hardware and software. Statistically significant and repeatable datasets produced by the satellite emulation system can be acquired during such test and saved for further analysis. In addition, the robustness of the discrimination and tracking algorithms can be investigated by a parametric analysis using slightly different scenarios; this will be used to determine critical points where a sensor system might fail. The radiometric characteristics of satellites are expected to be similar to the targets and decoys that make up a typical interceptor mission scenario, since they are near ambient temperature. Their spectral reflectivity, emissivity, and shape must also be considered, but the projection systems employed in the 7V and 10V chambers should be

  15. Effect of Afterbody-Ejector Configurations on the Performance at Transonic Speeds of a Pylon-Supported Nacelle Model having a Hot-Jet Exhaust

    NASA Technical Reports Server (NTRS)

    Swihart, John M.; Mercer, Charles E.; Norton, Harry T., Jr.

    1959-01-01

    An investigation of several afterbody-ejector configurations on a pylon-supported nacelle model has been completed in the Langley 16-foot transonic tunnel at Mach numbers from 0.80 to 1.05. The propulsive performance of two nacelle afterbodies with low boattailing and long ejector spacing was compared with a configuration corresponding to a turbojet-engine installation having a highly boattailed afterbody with a short ejector. The jet exhaust was simulated with a hydrogen peroxide turbojet simulator. The angle of attack was maintained at 0 deg, and the average Reynolds number based on body length was 20 x 10(exp 6). The results of the investigation indicated that the configuration with a conical afterbody with smooth transition to a 15 deg boattail angle had large beneficial jet effects on afterbody pressure-drag coefficient and had the best thrust-minus-drag performance of the afterbody-ejector configurations investigated.

  16. Fuselage and nozzle pressure distributions of a 1/12-scale F-15 propulsion model at transonic speeds. Effect of fuselage modifications and nozzle variables

    NASA Technical Reports Server (NTRS)

    Pendergraft, O. C., Jr.; Carson, G. T., Jr.

    1984-01-01

    Static pressure coefficient distributions on the forebody, afterbody, and nozzles of a 1/12 scale F-15 propulsion model was determined in the 16 foot transonic tunnel for Mach numbers from 0.60 to 1.20, angles of attack from -2 deg to 7 deg and ratio of jet total pressure to free stream static pressure from 1 up to about 7, depending on Mach number. The effects of nozzle geometry and horizontal tail deflection on the pressure distributions were investigated. Boundary layer total pressure profiles were determined at two locations ahead of the nozzles on the top nacelle surface. Reynolds number varied from about 1.0 x 10 to the 7th power per meter, depending on Mach number.

  17. A Parametric Investigation of Nozzle Planform and Internal/External Geometry at Transonic Speeds

    NASA Technical Reports Server (NTRS)

    Cler, Daniel L.

    1995-01-01

    An experimental investigation of multidisciplinary (scarfed trailing edge) nozzle divergent flap geometry was conducted at transonic speeds in the NASA Langley 16-Foot Transonic Tunnel. The geometric parameters investigated include nozzle planform, nozzle contouring location (internal and/or external), and nozzle area ratio (area ratio 1.2 and 2.0). Data were acquired over a range of Mach Numbers from 0.6 to 1.2, angle-of-attack from 0.0 degrees to 9.6 degrees and nozzle pressure ratios from 1.0 to 20.0. Results showed that increasing the rate of change internal divergence angle across the width of the nozzle or increasing internal contouring will decrease static, aeropropulsive and thrust removed drag performance regardless of the speed regime. Also, increasing the rate of change in boattail angle across the width of the nozzle or increasing external contouring will provide the lowest thrust removed drag. Scarfing of the nozzle trailing edges reduces the aeropropulsive performance for the most part and adversely affects the nozzle plume shape at higher nozzle pressure ratios thus increasing the thrust removed drag. The effects of contouring were primary in nature and the effects of planform were secondary in nature. Larger losses occur supersonically than subsonically when scarfing of nozzle trailing edges occurs. The single sawtooth nozzle almost always provided lower thrust removed drag than the double sawtooth nozzles regardless the speed regime. If internal contouring is required, the double sawtooth nozzle planform provides better static and aeropropulsive performance than the single sawtooth nozzle and if no internal contouring is required the single sawtooth provides the highest static and aeropropulsive performance.

  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. An experimental investigation of propfan installations on an upswept supercritical wing at transonic Mach numbers

    NASA Technical Reports Server (NTRS)

    Bartlett, G. R.

    1985-01-01

    An investigation has been conducted in the Langley 16 Foot Transonic Tunnel to determine propfan installation and slipstream interference effects on an unswept supercritical wing. This data can be used for verification of existing and developing theoretical codes as well as giving an understanding of the flow interactions associated with propeller/nacelle/wing integration. The investigation was conducted over a Mach number range of 0.5 to 0.8 and at angles of attack from 0 deg to 3 deg. The propeller was powered by an air turbine simulator and the exhaust from the air turbine was used to simulate the exhaust from the propfan nacelle. Reynolds number based on wing chord varied from 3 to 4 million. Results indicate that the propfan causes an increase in the wing lift coefficient. It was found that most of the propeller induced swirl is recovered by the wing. The propeller slipstream also causes a large favorable leading edge suction peak on the upwash side and a smaller unfavorable decrease on the downwash side.

  20. Transonic Investigation of Two-Dimensional Nozzles Designed for Supersonic Cruise

    NASA Technical Reports Server (NTRS)

    Capone, Francis J.; Deere, Karen A.

    2001-01-01

    An experimental and computational investigation has been conducted to determine the off-design uninstalled drag characteristics a of a two-dimensional convergent-divergent nozzle designed for a supersonic cruise civil transport. The main objective of this investigation was to determine the effects of varying nozzle external flap curvature and sidewall boattail angle and curvature on nozzle drag The experimental investigation was conducted in the Langley 16-Foot Transonic Tunnel at Mach numbers from 0.80 to 1.20 at nozzle pressure ratios up to nine. Three-dimensional simulations of nozzle performance were obtained with the computational fluid dynamics code PAB using turbulence closure and nonlinear Reynolds stress modeling. The results of this investigation indicate that excellent correlation between experimental and predicted results was obtained for the nozzle with a moderate amount of boattail curvature. The nozzle with an external flap having a sharp shoulder (no curvature) had the lowest nozzle pressure drag. At a Mach number of 1.2, sidewall pressure drag doubled as sidewall boattail angle was increased from 4 to 8 deg. Reducing the height of the sidewall caused large decreases in both the sidewall and flap pressure drags.

  1. Transonic Investigation of Two-Dimensional Nozzles Designed for Supersonic Cruise

    NASA Technical Reports Server (NTRS)

    Capone, Francis J.; Deere, Karen A.

    2015-01-01

    An experimental and computational investigation has been conducted to determine the off-design uninstalled drag characteristics of a two-dimensional convergent-divergent nozzle designed for a supersonic cruise civil transport. The overall objectives were to: (1) determine the effects of nozzle external flap curvature and sidewall boattail variations on boattail drag; (2) develop an experimental data base for 2D nozzles with long divergent flaps and small boattail angles and (3) provide data for correlating computational fluid dynamic predictions of nozzle boattail drag. The experimental investigation was conducted in the Langley 16-Foot Transonic Tunnel at Mach numbers from 0.80 to 1.20 at nozzle pressure ratios up to 9. Three-dimensional simulations of nozzle performance were obtained with the computational fluid dynamics code PAB3D using turbulence closure and nonlinear Reynolds stress modeling. The results of this investigation indicate that excellent correlation between experimental and predicted results was obtained for the nozzle with a moderate amount of boattail curvature. The nozzle with an external flap having a sharp shoulder (no curvature) had the lowest nozzle pressure drag. At a Mach number of 1.2, sidewall pressure drag doubled as sidewall boattail angle was increased from 4deg to 8deg. Reducing the height of the sidewall caused large decreases in both the sidewall and flap pressure drags. Summary

  2. National Transonic Facility Characterization Status

    NASA Technical Reports Server (NTRS)

    Bobbitt, C., Jr.; Everhart, J.; Foster, J.; Hill, J.; McHatton, R.; Tomek, W.

    2000-01-01

    This paper describes the current status of the characterization of the National Transonic Facility. The background and strategy for the tunnel characterization, as well as the current status of the four main areas of the characterization (tunnel calibration, flow quality characterization, data quality assurance, and support of the implementation of wall interference corrections) are presented. The target accuracy requirements for tunnel characterization measurements are given, followed by a comparison of the measured tunnel flow quality to these requirements based on current available information. The paper concludes with a summary of which requirements are being met, what areas need improvement, and what additional information is required in follow-on characterization studies.

  3. 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.

  4. Numerical computation of aeroelastically corrected transonic loads

    NASA Technical Reports Server (NTRS)

    Chipman, R.; Waters, C.; Mackenzie, D.

    1979-01-01

    A numerical scheme is presented for the computation of transonic aerodynamic loads on flexible wings. The method consists of iteratively applying the loads computed by a 3D transonic aerodynamics code to a structural model to obtain elastic twist, and then recomputing the loads. Because this iteration is performed concurrently with the iterations performed in computing the aerodynamics, flexible loads are obtained in roughly the same amount of computing time as required to obtain rigid loads. Applications of this method to a flexible supercritical transonic transport wing are presented and compared with model test data.

  5. Transonic interactions of unsteady vortical flows

    NASA Technical Reports Server (NTRS)

    Mccroskey, W. J.; Srinivasan, G. R.

    1984-01-01

    Unsteady interactions of strong concentrated vortices, distributed gusts, and sharp-edged gusts with stationary airfoils were analyzed in two-dimensional transonic flow. A simple and efficient method for introducing such vortical disturbances was implemented in numerical codes that range from inviscid transonic small disturbance to thin-layer Navier Stokes. The numerical results demonstrate the large distortions in the overall flow field and in the surface air loads that are produced by various vortical interactions. The results of the different codes are in excellent qualitative agreement, but, as might expected, the transonic small-disturbance calculations are deficient in the important region near the leading edge.

  6. Simulation Of Unsteady, Inviscid, Rotational, Transonic Flow

    NASA Technical Reports Server (NTRS)

    Damodaran, Murali

    1992-01-01

    Report describes numerical simulation of two-dimensional, unsteady, inviscid rotational, transonic flow about rigid airfoil in such motions as pitching or plunging oscillations. Study demonstrates potential utility of computation in analyses of aeroelasticity of airfoils.

  7. Computed Flows In A Transonic Turbine

    NASA Technical Reports Server (NTRS)

    Rangwalla, A. A.; Madavan, N. K.; Johnson, P. D.

    1993-01-01

    Report presents computational study of flow in first stage of three alternative versions of proposed transonic turbine. Study demonstrates application of computational fluid dynamics to predict performance and analyze effects of changes in designs of these advanced machines.

  8. Recent advances in transonic computational aeroelasticity

    NASA Technical Reports Server (NTRS)

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

    1988-01-01

    A transonic unsteady aerodynamic and aeroelasticity code called CAP-TSD was developed for application to realistic aircraft configurations. The code permits the calculation of steady and unsteady flows about complete aircraft configurations for aeroelastic analysis in the flutter critical transonic speed range. The CAP-TSD code uses a time accurate approximate factorization algorithm for solution of the unsteady transonic small disturbance potential equation. An overview is given of the CAP-TSD code development effort and results are presented which demonstrate various capabilities of the code. Calculations are presented for several configurations including the General Dynamics 1/9 scale F-16 aircraft model and the ONERA M6 wing. Calculations are also presented from a flutter analysis of a 45 deg sweptback wing which agrees well with the experimental data. Descriptions are presented of the CAP-TSD code and algorithm details along with results and comparisons which demonstrate these recent developments in transonic computational aeroelasticity.

  9. 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.

  10. Transonic airfoil design using Cartesian coordinates

    NASA Technical Reports Server (NTRS)

    Carlson, L. A.

    1976-01-01

    A numerical technique for designing transonic airfoils having a prescribed pressure distribution (the inverse problem) is presented. The method employs the basic features of Jameson's iterative solution for the full potential equation, except that inverse boundary conditions and Cartesian coordinates are used. The method is a direct-inverse approach that controls trailing-edge closure. Examples show the application of the method to design aft-cambered and other airfoils specifically for transonic flight.

  11. A model for transonic plasma flow

    SciTech Connect

    Guazzotto, Luca; Hameiri, Eliezer

    2014-02-15

    A linear, two-dimensional model of a transonic plasma flow in equilibrium is constructed and given an explicit solution in the form of a complex Laplace integral. The solution indicates that the transonic state can be solved as an elliptic boundary value problem, as is done in the numerical code FLOW [Guazzotto et al., Phys. Plasmas 11, 604 (2004)]. Moreover, the presence of a hyperbolic region does not necessarily imply the presence of a discontinuity or any other singularity of the solution.

  12. 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.

  13. 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.

  14. Longitudinal Aerodynamic Characteristics of a Wing-Body-Tail Model Having a Highly Tapered, Cambered 45 degree Swept Wing of Aspect Ratio 4 at Transonic Speeds

    NASA Technical Reports Server (NTRS)

    West, F. E., Jr.

    1959-01-01

    The longitudinal aerodynamic characteristics of a wing-body-horizontal-tail configuration designed for efficient performance at transonic speeds has been investigated at Mach numbers from 0.80 to 1.03 in the Langley 16-foot transonic tunnel. The effect of adding an outboard leading-edge chord-extension to the highly tapered 45 deg. swept wing was also obtained. The average Reynolds number for this investigation was 6.7 x 10(exp 6) based on the wing mean aerodynamic chord. The relatively low tail placement as well as the addition of a chord-extension achieved some alleviation of the pitchup tendencies of the wing-fuselage configuration. The maximum trimmed lift-drag ratio was 16.5 up to a Mach number of 0.9, with the moment center located at the quarter-chord point of the mean aerodynamic chord. For the untrimmed case, the maximum lift-drag ratio was approximately 19.5 up to a Mach number of 0.9.

  15. 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.

  16. An experimental investigation of an advanced turboprop installation on a swept wing at subsonic and transonic speeds

    NASA Technical Reports Server (NTRS)

    Carlson, John R.; Pendergraft, Odis C., Jr.

    1987-01-01

    An investigation was conducted in the Langley 16-Foot Transonic Tunnel to determine the effects of a turboprop-nacelle installation on the pressure distributions over a swept, supercritical wing. The tests were conducted at Mach numbers from 0.20 to 0.80, at angles of attack from 0 to 5 degrees, nacelle nozzle pressure ratios from 1.0 to 1.6, and at propeller tip speeds from 700 to 800 ft/sec. The results of this study indicate that the turboprop nacelle interference, with and without power, on a swept wing is greater on the inboard wing panel than on the outboard wing panel. The over-the-wing nacelle installation with the propeller upwash on the inboard panel had flow separation problems at a Mach number of 0.80. No severe flow separation problems appear to exist for either propeller rotation direction for the under-the-wing nacelle installation. The local flow disturbances caused by the under-the-wing nacelle installation were in general less severe than for the over-the-wing nacelle installation.

  17. Experience with transonic unsteady aerodynamic calculations

    NASA Technical Reports Server (NTRS)

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

    1984-01-01

    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.

  18. Turbulence and modeling in transonic flow

    NASA Technical Reports Server (NTRS)

    Rubesin, Morris W.; Viegas, John R.

    1989-01-01

    A review is made of the performance of a variety of turbulence models in the evaluation of a particular well documented transonic flow. This is done to supplement a previous attempt to calibrate and verify transonic airfoil codes by including many more turbulence models than used in the earlier work and applying the calculations to an experiment that did not suffer from uncertainties in angle of attack and was free of wind tunnel interference. It is found from this work, as well as in the earlier study, that the Johnson-King turbulence model is superior for transonic flows over simple aerodynamic surfaces, including moderate separation. It is also shown that some field equation models with wall function boundary conditions can be competitive with it.

  19. Transonic wing analysis using advanced computational methods

    NASA Technical Reports Server (NTRS)

    Henne, P. A.; Hicks, R. M.

    1978-01-01

    This paper discusses the application of three-dimensional computational transonic flow methods to several different types of transport wing designs. The purpose of these applications is to evaluate the basic accuracy and limitations associated with such numerical methods. The use of such computational methods for practical engineering problems can only be justified after favorable evaluations are completed. The paper summarizes a study of both the small-disturbance and the full potential technique for computing three-dimensional transonic flows. Computed three-dimensional results are compared to both experimental measurements and theoretical results. Comparisons are made not only of pressure distributions but also of lift and drag forces. Transonic drag rise characteristics are compared. Three-dimensional pressure distributions and aerodynamic forces, computed from the full potential solution, compare reasonably well with experimental results for a wide range of configurations and flow conditions.

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  4. Some iterative schemes for transonic potential flows

    NASA Technical Reports Server (NTRS)

    Wong, Y. S.; Hafez, M. M.

    1985-01-01

    The minimal residual (MR) method for the numerical solution of transonic potential flows is closely related to the conjugate gradient method, which has found widespread use in the solution of large sparse, symmetric, and positive-definite linear equations. The primary advantage of the MR method is its applicability to both symmetric and nonsymmetric matrices.

  5. 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.

  6. 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.

  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. 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%.

  9. Viscous Transonic Airfoil Workshop compendium of results

    NASA Technical Reports Server (NTRS)

    Holst, Terry L.

    1987-01-01

    Results from the Viscous Transonic Airfoil Workshop held at the AIAA 25th Aerospace Sciences Meeting at Reno, NV in January 1987, are compared with each other and with experimental data. Test cases used in this workshop include attached and separated transonic flows for three different airfoils: the NACA 0012 airfoil, the RAE 2822 airfoil, and the Jones airfoil. A total of 23 sets of numerical results from 15 different author groups are included. The numerical methods used vary widely and include: 16 Navier-Stokes methods, 2 Euler/boundary-layer methods, and 5 full-potential/boundary-layer methods. The results indicate a high degree of sophistication among the numerical methods with generally good agreement between the various computed and experimental results for attached or moderately-separated cases. The agreement for cases with larger separation is only fair and suggests additional work is required in this area.

  10. 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.

  11. 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.

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  16. 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

  17. 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.

  18. 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.

  19. 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.

  20. The STD/MHD codes - Comparison of analyses with experiments at AEDC/HPDE, Reynolds Metal Co., and Hercules, Inc. [for MHD generator flows

    NASA Technical Reports Server (NTRS)

    Vetter, A. A.; Maxwell, C. D.; Swean, T. F., Jr.; Demetriades, S. T.; Oliver, D. A.; Bangerter, C. D.

    1981-01-01

    Data from sufficiently well-instrumented, short-duration experiments at AEDC/HPDE, Reynolds Metal Co., and Hercules, Inc., are compared to analyses with multidimensional and time-dependent simulations with the STD/MHD computer codes. These analyses reveal detailed features of major transient events, severe loss mechanisms, and anomalous MHD behavior. In particular, these analyses predicted higher-than-design voltage drops, Hall voltage overshoots, and asymmetric voltage drops before the experimental data were available. The predictions obtained with these analyses are in excellent agreement with the experimental data and the failure predictions are consistent with the experiments. The design of large, high-interaction or advanced MHD experiments will require application of sophisticated, detailed and comprehensive computational procedures in order to account for the critical mechanisms which led to the observed behavior in these experiments.

  1. Pressure distributions obtained on a 0.10-scale model of the space shuttle Orbiter's forebody in the AEDC 16T propulsion wind tunnel

    NASA Technical Reports Server (NTRS)

    Siemers, P. M., III; Henry, M. W.

    1986-01-01

    Pressure distribution test data obtained on a 0.10-scale model of the forward fuselage of the Space Shuttle Orbiter are presented without analysis. The tests were completed in the AEDC 16T Propulsion Wind Tunnel. The 0.10-scale model was tested at angles of attack from -2 deg to 18 deg and angles of side slip from -6 to 6 deg at Mach numbers from 0.25 to 1/5 deg. The tests were conducted in support of the development of the Shuttle Entry Air Data System (SEADS). In addition to modeling the 20 SEADS orifices, the wind-tunnel model was also instrumented with orifices to match Development Flight Instrumentation (DFI) port locations that existed on the Space Shuttle Orbiter Columbia (OV-102) during the Orbiter Flight Test program. This DFI simulation has provided a means of comparisons between reentry flight pressure data and wind-tunnel and computational data.

  2. 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.

  3. 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.

  4. Lateral Stability and Control Measurements of a 0.0858-Scale Model of the Lockheed XF-104 Airplane at Transonic Speeds

    NASA Technical Reports Server (NTRS)

    Arabian, Donald D.; Schmeer, James W.

    1955-01-01

    An investigation of the lateral stability and control effectiveness of a 0.0858-scale model of the Lockheed XF-104 airplane has been conducted in the Langley 16-foot transonic tunnel. The model has a low aspect ratio, 3.4-percent-thick wing with negative dihedral. The horizontal tail is located on top of the vertical tail. The investigation was made through a Mach number range of 0.80 to 1.06 at sideslip angles of -5 deg. to 5 deg. and angles of attack from 0 deg. to 16 deg. The control effectiveness of the aileron, rudder, and yaw damper were determined through the Mach number and angle-of-attack range. The results of the investigation indicated that the directional stability derivative was stable and that positive effective dihedral existed throughout the lift-coefficient range and Mach number range tested. The total aileron effectiveness, which in general produced favorable yaw with rolling moment, remained fairly constant for lift coefficients up to about 0.8 for the Mach number range tested. Yawing-moment effectiveness of the rudder changed little through the Mach number range. However, the yaw damper effectiveness decreased about 30 percent at the intermediate test Mach numbers.

  5. Effects of afterbody boattail design and empennage arrangement on aeropropulsive characteristics of a twin-engine fighter model at transonic speeds

    NASA Technical Reports Server (NTRS)

    Bangert, Linda S.; Leavitt, Laurence D.; Reubush, David E.

    1987-01-01

    The effects of empennage arrangement and afterbody boattail design of nonaxisymmetric nozzles on the aeropropulsive characteristics of a twin-engine fighter-type model have been determined in an investigation conducted in the Langley 16-Foot Transonic Tunnel. Three nonaxisymmetric and one twin axisymmetric convergent-divergent nozzle configurations were tested with three different tail arrangements: a two-tail V-shaped arrangement; a staggered, conventional three-tail arrangement; and a four-tail arrangement similar to that on the F-18. Two of the nonaxisymmetric nozzles were also vectorable. Tests were conducted at Mach numbers from 0.60 to 1.20 over an angle-of-attack range from -3 deg to 9 deg. Nozzle pressure ratio was varied from 1 (jet off) to approximately 12, depending on Mach number. Results indicate that at design nozzle pressure ratio, the medium aspect ratio nozzle (with equal boattail angles on the nozzle sidewalls and upper and lower flaps) had the lowest zero angle of attack drag of the nonaxisymmetric nozzles for all tail configurations at subsonic Mach numbers. The drag levels of the twin axisymmetric nozzles were competitive with those of the medium-aspect-ratio nozzle at subsonic Mach number.

  6. Aeropropulsive characteristics of twin single-expansion-ramp vectoring nozzles installed with forward-swept wings and canards. [transonic tunnel tests

    NASA Technical Reports Server (NTRS)

    Mason, M. L.; Capone, F. J.

    1983-01-01

    The Langley 16 foot transonic tunnel was used to determine the aeropropulsive characteristics of twin single-expansion-ramp vectoring nozzles installed in a wing-body configuration with forward-swept wings. The configuration was tested with and without fixed canards. The test conditions included free-stream Mach numbers of 0.60, 0.90, and 1.20. The model angle of attack ranged from -2 deg to 14 deg; the nozzle pressure ratio ranged from 1.0 (jet off) to 9.0. The Reynolds number based on the wing mean aerodynamic chord varied from 3.0 x 10 to the 6th power to 4.8 x 10 to the 6th power, depending on Mach number. Aerodynamic characteristics were analyzed to determine the effects of thrust vectoring and the canard effects on the wing-afterbody-nozzle and the wing-afterbody portions of the model. Thrust vectoring had no effect on the angle of attack for the onset of flow separation on the wing but resulted in reduced drag at angle-of-attack values above that required for wing flow separation. The canard was found to have little effect on the thrust-induced lift resulting from vectoring, since canard effects occurred primarily on the wing.

  7. Transonic potential flow in hyperbolic nozzles

    NASA Technical Reports Server (NTRS)

    Park, M.; Caughey, D. A.

    1986-01-01

    The full potential equation for the classical problem of transonic flow through a hyperbolic nozzle (with or without a shock wave) is solved in conservation form using the finite volume method of Jameson and Caughey (1977). Either a firstor a second-order numerical viscosity is added in the direction of the flow, explicitly, in conservation form. A multigrid alternating direction implicit method is used to solve the difference equations, and the results obtained are compared with analytical and numerical results from previous researches.

  8. 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.

  9. Computational methods for unsteady transonic flows

    NASA Technical Reports Server (NTRS)

    Edwards, John W.; Thomas, James L.

    1987-01-01

    Computational methods for unsteady transonic flows are surveyed with emphasis upon applications to aeroelastic analysis and flutter prediction. Computational difficulty is discussed with respect to type of unsteady flow; attached, mixed (attached/separated) and separated. Significant early computations of shock motions, aileron buzz and periodic oscillations are discussed. The maturation of computational methods towards the capability of treating complete vehicles with reasonable computational resources is noted and a survey of recent comparisons with experimental results is compiled. The importance of mixed attached and separated flow modeling for aeroelastic analysis is discussed and recent calculations of periodic aerodynamic oscillations for an 18 percent thick circular arc airfoil are given.

  10. Computational methods for unsteady transonic flows

    NASA Technical Reports Server (NTRS)

    Edwards, John W.; Thomas, J. L.

    1987-01-01

    Computational methods for unsteady transonic flows are surveyed with emphasis on prediction. Computational difficulty is discussed with respect to type of unsteady flow; attached, mixed (attached/separated) and separated. Significant early computations of shock motions, aileron buzz and periodic oscillations are discussed. The maturation of computational methods towards the capability of treating complete vehicles with reasonable computational resources is noted and a survey of recent comparisons with experimental results is compiled. The importance of mixed attached and separated flow modeling for aeroelastic analysis is discussed, and recent calculations of periodic aerodynamic oscillations for an 18 percent thick circular arc airfoil are given.

  11. 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.

  12. 0.3 Meter Transonic Cryogenic Tunnel

    NASA Technical Reports Server (NTRS)

    1985-01-01

    Full Description: The Langley 0.3-Meter Transonic Cryogenic tunnel (0.3-m TCT) is used for testing two-dimensional airfoil sections and other models at high Reynolds numbers. The tunnel can operate continuously over a range of Mach numbers from about 0.1 to above 1.2, with a stagnation pressure from 14.7 to 88.0 psia (1 to 6 atmospheres) and a stagnation temperature from -320F to 130F (78 K to 328 K). This results in a maximum Reynolds number capability in excess of 100 x 106 per foot. The adaptive walls, floor, and ceiling in the 13-in. by 13-in. (33-cm by 33-cm) test section can be moved to the free-stream streamline shape, eliminating or reducing the wall effects on the model. The combination of flight Reynolds numbers capability and minimal wall interference makes the 0.3-m TCT a powerful tool for aeronautical research at transonic speeds. The Mach number, pressure, temperature, and adaptive wall shape are automatically controlled. The test section has computer-controlled angle of attack and traversing wake survey-probe systems. The facility has been modified to also use alternate test media--a heavy gas (sulfur hexafluoride, SF6), or air, both with a newly installed heat exchanger.

  13. Geometrical acoustics and transonic helicopter sound

    NASA Technical Reports Server (NTRS)

    Isom, Morris; Purcell, Timothy W.; Strawn, Roger C.

    1987-01-01

    A new method is presented for predicting the impulsive noise generated by a transonic rotor blade. The method is a combined approach involving computational fluid dynamics and geometrical acoustics. A full-potential finite-difference method is used to obtain the pressure field close to the blade. A Kirchhoff integral formulation is then used to extend these finite-difference results into the far field. This Kirchhoff formula is based on geometrical acoustics approximations. It requires initial data across a plane at the sonic radius in a blade-fixed coordinate system. This data is provided by the finite-difference solution. Acoustic pressure predictions show good agreement with hover experimental data for cases with hover tip Mach numbers of 0.88 through 0.96. The cases above 0.92 tip Mach number are dominated by non-linear transonic effects seen as strong shocks on and off the blade tip. This paper gives the first successful predictions of far-field acoustic pressures for high-speed impulsive noise over a range of Mach numbers after delocalization.

  14. High-transonic-speed transport aircraft study

    NASA Technical Reports Server (NTRS)

    Kulfan, R. M.

    1974-01-01

    An initial design study of high-transonic-speed transport aircraft has been completed. Five different design concepts were developed. These included fixed swept wing, variable-sweep wing, delta wing, double-fuselage yawed-wing, and single-fuselage yawed-wing aircraft. The boomless supersonic design objectives of range = 5560 km (3000 nmi), payload = 18,143 kg (40,000 lb), Mach = 1.2, and FAR Part 36 aircraft noise levels were achieved by the single-fuselage yawed-wing configuration with a gross weight of 211,828 kg (467,000 lb). A noise level of 15 EPNdB below FAR Part 36 requirements was obtained with a gross weight increase to 226,796 kg (500,000 lb). The off-design subsonic range capability for this configuration exceeded the Mach 1.2 design range by more than 20%. Although wing aeroelastic divergence was a primary design consideration for the yawed-wing concepts, the graphite-epoxy wings of this study were designed by critical gust and maneuver loads rather than by divergence requirements. The transonic nacelle drag is shown to be very sensitive to the nacelle installation. A six-degree-of-freedom dynamic stability analysis indicated that the control coordination and stability augmentation system would require more development than for a symmetrical airplane but is entirely feasible. A three-plane development plan is recommended to establish the full potential of the yawed-wing concept.

  15. Numerical calculations of two dimensional, unsteady transonic flows with circulation

    NASA Technical Reports Server (NTRS)

    Beam, R. M.; Warming, R. F.

    1974-01-01

    The feasibility of obtaining two-dimensional, unsteady transonic aerodynamic data by numerically integrating the Euler equations is investigated. An explicit, third-order-accurate, noncentered, finite-difference scheme is used to compute unsteady flows about airfoils. Solutions for lifting and nonlifting airfoils are presented and compared with subsonic linear theory. The applicability and efficiency of the numerical indicial function method are outlined. Numerically computed subsonic and transonic oscillatory aerodynamic coefficients are presented and compared with those obtained from subsonic linear theory and transonic wind-tunnel data.

  16. Computational, unsteady transonic aerodynamics and aeroelasticity about airfoils and wings

    NASA Technical Reports Server (NTRS)

    Goorjian, Peter M.; Guruswamy, Guru P.

    1987-01-01

    Research in the area of computational, unsteady transonic flows about airfoils and wings, including aeroelastic effects is reviewed. In the last decade, there have been extensive developments in computational methods in response to the need for computer codes with which to study fundamental aerodynamic and aeroelastic problems in the critical transonic regime. For example, large commercial aircraft cruise most effectively in the transonic flight regime and computational fluid dynamics (CDF) provides a new tool, which can be used in combination with test facilities to reduce the costs, time, and risks of aircraft development.

  17. 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.

  18. Wing analysis using a transonic potential flow computational method

    NASA Technical Reports Server (NTRS)

    Henne, P. A.; Hicks, R. M.

    1978-01-01

    The ability of the method to compute wing transonic performance was determined by comparing computed results with both experimental data and results computed by other theoretical procedures. Both pressure distributions and aerodynamic forces were evaluated. Comparisons indicated that the method is a significant improvement in transonic wing analysis capability. In particular, the computational method generally calculated the correct development of three-dimensional pressure distributions from subcritical to transonic conditions. Complicated, multiple shocked flows observed experimentally were reproduced computationally. The ability to identify the effects of design modifications was demonstrated both in terms of pressure distributions and shock drag characteristics.

  19. Analysis of viscous transonic flow over airfoil sections

    NASA Technical Reports Server (NTRS)

    Huff, Dennis L.; Wu, Jiunn-Chi; Sankar, L. N.

    1987-01-01

    A full Navier-Stokes solver has been used to model transonic flow over three airfoil sections. The method uses a two-dimensional, implicit, conservative finite difference scheme for solving the compressible Navier-Stokes equations. Results are presented as prescribed for the Viscous Transonic Airfoil Workshop to be held at the AIAA 25th Aerospace Sciences Meeting. The NACA 0012, RAE 2822 and Jones airfoils have been investigated for both attached and separated transonic flows. Predictions for pressure distributions, loads, skin friction coefficients, boundary layer displacement thickness and velocity profiles are included and compared with experimental data when possible. Overall, the results are in good agreement with experimental data.

  20. 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).

  1. Techniques for correcting approximate finite difference solutions. [considering transonic flow

    NASA Technical Reports Server (NTRS)

    Nixon, D.

    1978-01-01

    A method of correcting finite-difference solutions for the effect of truncation error or the use of an approximate basic equation is presented. Applications to transonic flow problems are described and examples are given.

  2. Unsteady transonic flow calculations for interfering lifting surface configurations

    NASA Technical Reports Server (NTRS)

    Batina, J. T.

    1985-01-01

    Unsteady transonic flow calculations are presented for aerodynamically interfering lifting surface configurations. Calculations are performed by extending the XTRAN3S (Version 1.5) unsteady transonic small-disturbance code to allow the treatment of an additional lifting surface. The research was conducted as a first-step toward developing the capability to treat a complete flight vehicle. Grid generation procedures for swept tapered interfering lifting surface applications of XTRAN3S are described. Transonic calculations are presented for wing-tail and canard-wing configurations for several values of mean angle of attack. The effects of aerodynamic interference on transonic steady pressure distributions and steady and oscillatory spanwise lift distributions are demonstrated. Results due to wing, tail, or canard pitching motions are presented and discussed in detail.

  3. Design optimization of axisymmetric bodies in nonuniform transonic flow

    NASA Technical Reports Server (NTRS)

    Lan, C. Edward

    1989-01-01

    An inviscid transonic code capable of designing an axisymmetric body in a uniform or nonuniform flow was developed. The design was achieved by direct optimiation by coupling an analysis code with an optimizer. Design examples were provided for axisymmetric bodies with fineness ratios of 8.33 and 5 at different Mach numbers. It was shown that by reducing the nose radius and increasing the afterbody thickness of initial shapes obtained from symmetric NACA four-digit airfoil contours, wave drag could be reduced by 29 percent for a body of fineness ratio 8.33 in a nonuniform transonic flow of M = 0.98 to 0.995. The reduction was 41 percent for a body of fineness ratio 5 in a uniform transonic flow of M = 0.925 and 65 percent for the same body but in a nonuniform transonic flow of M = 0.90 to 0.95.

  4. Active Suppression of the Transonic Flutter Using Sliding Mode Control

    NASA Astrophysics Data System (ADS)

    Degaki, Takanori; Suzuki, Shinji

    This paper describes two-dimensional active flutter suppression to cope with the transonic dip using the sliding mode control. The airfoil model has plunge and pitch degrees of freedom with leading and trailing edge control surfaces. The aerodynamic forces acting on the airfoil, lift and pitching moment, are calculated by solving Euler's equations using computational fluid dynamics. At a specific altitude, flutter occurs between Mach number of 0.7 and 0.88, which corresponds to the transonic dip. The sliding mode control makes the airfoil to be stable all through the Mach number including the transonic dip. The sliding mode controller gives wider flutter margin than a linear quadratic regulator. These characteristics indicate that the sliding mode control is useful for active flutter suppression in the transonic flight.

  5. 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

  6. 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

  7. 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

  8. 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

  9. 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

  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. 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

  12. Transonic Unsteady Aerodynamics and Aeroelasticity 1987, part 1

    NASA Technical Reports Server (NTRS)

    Bland, Samuel R. (Compiler)

    1989-01-01

    Computational fluid dynamics methods have been widely accepted for transonic aeroelastic analysis. Previously, calculations with the TSD methods were used for 2-D airfoils, but now the TSD methods are applied to the aeroelastic analysis of the complete aircraft. The Symposium papers are grouped into five subject areas, two of which are covered in this part: (1) Transonic Small Disturbance (TSD) theory for complete aircraft configurations; and (2) Full potential and Euler equation methods.

  13. 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.

  14. 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.

  15. Mach number effects on transonic aeroelastic forces and flutter characteristics

    NASA Technical Reports Server (NTRS)

    Mohr, Ross W.; Batina, John T.; Yang, Henry T. Y.

    1988-01-01

    Transonic aeroelastic stability analysis and flutter calculations are presented for a generic transport-type wing based on the use of the CAP-TSD (Computational Aeroelasticity Program - Transonic Small Disturbance) finite-difference code. The CAP-TSD code was recently developed for transonic unsteady aerodynamic and aeroelastic analysis of complete aircraft configurations. A binary aeroelastic system consisting of simple bending and torsion modes was used to study aeroelastic behavior at transonic speeds. Generalized aerodynamic forces are presented for a wide range of Mach number and reduced frequency. Aeroelastic characteristics are presented for variations in freestream Mach number, mass ratio, and bending-torsion frequency ratio. Flutter boundaries are presented which have two transonic dips in flutter speed. The first dip is the usual transonic dip involving a bending-dominated flutter mode. The second dip is characterized by a single degree-of-freedom torsion oscillation. These aeroelastic results are physically interpreted and shown to be related to the steady state shock location and changes in generalized aerodynamic forces due to freestream Mach number.

  16. Results of tests using a 0.0125-scale model (70-QT) of the space shuttle vehicle orbiter in the AEDC VKF tunnel B (IA22), volume 2

    NASA Technical Reports Server (NTRS)

    Daileda, J. J.; Marroquin, J.

    1977-01-01

    Tabulated data of an experimental investigation are presented which was conducted in the AEDC/VKF Tunnel B to obtain interaction effects of RCS thruster jet plumes on SSV aerodynamics during staging to simulate RTLS abort. Interaction effects of the orbiter RCS thruster jet plumes on the orbiter and ET aerodynamics were investigated. RCS thruster jet plumes were simulated using both air and a 15 percent argon 85 percent helium gas mixture. The ET angle of attack range was -40 to +25 deg at sideslip angles of 0, 3, and 6 degrees. Orbiter angle of attack was varied from -15 to +10 degrees at sideslip angles of 0 and 3 deg. External tank full scale separation distances simulated were 0 to 1400 in. axially; 0 to 54 in. laterally; and a range of -100 to 1000 in. vertically. Data were also obtained on the ET in the interference-free flow field. Quiescent (no tunnel flow) thruster plume interaction data were obtained on the orbiter and orbiter-ET combination. Tests were conducted at Mach number 6 and a Reynolds number of 0.86 million per foot.

  17. 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.

  18. 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.

  19. Analysis of three-dimensional transonic compressors

    NASA Technical Reports Server (NTRS)

    Bourgeade, A.

    1984-01-01

    A method for computing the three-dimensional transonic flow around the blades of a compressor or of a propeller is given. The method is based on the use of the velocity potential, on the hypothesis that the flow is inviscid, irrotational and isentropic. The equation of the potential is solved in a transformed space such that the surface of the blade is mapped into a plane where the periodicity is implicit. This equation is in a nonconservative form and is solved with the help of a finite difference method using artificial time. A computer code is provided and some sample results are given in order to demonstrate the influence of three-dimensional effects and the blade's rotation.

  20. Transonic Blunt Body Aerodynamic Coefficients Computation

    NASA Astrophysics Data System (ADS)

    Sancho, Jorge; Vargas, M.; Gonzalez, Ezequiel; Rodriguez, Manuel

    2011-05-01

    In the framework of EXPERT (European Experimental Re-entry Test-bed) accurate transonic aerodynamic coefficients are of paramount importance for the correct trajectory assessment and parachute deployment. A combined CFD (Computational Fluid Dynamics) modelling and experimental campaign strategy was selected to obtain accurate coefficients. A preliminary set of coefficients were obtained by CFD Euler inviscid computation. Then experimental campaign was performed at DNW facilities at NLR. A profound review of the CFD modelling was done lighten up by WTT results, aimed to obtain reliable values of the coefficients in the future (specially the pitching moment). Study includes different turbulence modelling and mesh sensitivity analysis. Comparison with the WTT results is explored, and lessons learnt are collected.

  1. Design of a transonically profiled wing

    NASA Technical Reports Server (NTRS)

    Kiekebusch, B.

    1978-01-01

    The application of well known design concepts with the combined use of thick transonic profiles to aircraft wing design was investigated. Optimization in terms of weight and operational costs was emphasized. It is shown that the usual design criteria and concepts are too restricted and do not sufficiently represent the physical processes over the wing. Suggestions are made for improving this situation, and a design example given. Compared with a wing design according to previously used criteria, the new design is found to be superior in the most important functions. It is concluded that an isobar concept adjusted to the planform in conjunction with an 'organically' designed wing will lead to the weight optimum solutions of wing profiles.

  2. 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.

  3. Demonstration of PIV in a Transonic Compressor

    NASA Technical Reports Server (NTRS)

    Wernet, Mark P.

    1997-01-01

    Particle Imaging Velocimetry (PIV) is a powerful measurement technique which can be used as an alternative or complementary approach to Laser Doppler Velocimetry (LDV) in a wide range of research applications. PIV data are measured simultaneously at multiple points in space, which enables the investigation of the non-stationary spatial structures typically encountered in turbomachinery. Many of the same issues encountered in the application of LDV techniques to rotating machinery apply in the application of PIV. Preliminary results from the successful application of the standard 2-D PIV technique to a transonic axial compressor are presented. The lessons learned from the application of the 2-D PIV technique will serve as the basis for applying 3-component PIV techniques to turbomachinery.

  4. 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.

  5. Nonclassical aileron buzz in transonic flow

    NASA Technical Reports Server (NTRS)

    Bendiksen, Oddvar O.

    1993-01-01

    A computational study of inviscid, transonic aileron and trailing-edge buzz instabilities is presented. A mixed Eulerian-Lagrangian formulation is used to model the fluid-structure system and to obtain a system of space-discretized equations that is time-marched to simulate the aeroelastic behavior of the wing-aileron system. Results obtained suggest that shock-induced separation may not be an essential driving force behind all buzz phenomena. Several examples are shown where the shock motion interacts with the aileron motion to extract energy from the flow. If the trailing-edge region is sufficiently flexible and the shocks are at the trailing edge, a trailing-edge buzz instability appears possible.

  6. 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.

  7. 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.

  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. 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

  11. Results of an investigation of hypersonic viscous interaction effects of the space shuttle orbiter using a 0.010 scale model (51-0) in the AEDC-VKF tunnel F (OA160)

    NASA Technical Reports Server (NTRS)

    Elder, D. J.

    1975-01-01

    An experimental aerodynamic investigation was conducted in the AEDC-VKF Hypervelocity Wind Tunnel (Tunnel F) at a nomial Mach number of 19 to determine hypersonic viscous interaction effects on the space shuttle orbiter. The tests were conducted at an angle of attack of 30 degrees over a free-stream Reynolds number (based on fuselage length) variation from 0.1 to 0.4 million. Viscous interaction parameter was varied from 0.02 to 0.06. Six component static stability force and moment data were measured by an internally compensated internal strain gage balance. Resulting data are presented.

  12. Results of tests of a Rockwell International space shuttle orbiter (-139 configuration) 0.0175-scale model (no. 29-0) in AEDC tunnel F to determine hypersonic heating effects (OH11)

    NASA Technical Reports Server (NTRS)

    Quan, M.

    1975-01-01

    Results from wind tunnel tests to determine hypersonic aerodynamic heating rates on a NASA/Rockwell Space Shuttle Orbiter are reported. The tests were to determine Mach number effects, if any, and to obtain overall heating rate data at high Mach numbers from 10.5 to 16. The model used was a 0.0175-scale model built to Rockwell Orbiter lines VL70-000139. The model identity number is 29-0. These tests, designated OH11, were conducted in the AEDC Tunnel F.

  13. Comparison of computational results of a few representative three-dimensional transonic potential flow analysis programs

    NASA Technical Reports Server (NTRS)

    Tanaka, K.; Hirose, H.

    1986-01-01

    The development of transonic aerodynamic computation methods and specific examples, as well as examples of three-dimensional transonic computation in design, are discussed. The case of the transonic transport and the case of the small transport are analyzed. Requirements for programs of the future are itemized.

  14. 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.

  15. Transonic airfoil design for helicopter rotor applications

    NASA Technical Reports Server (NTRS)

    Hassan, Ahmed A.; Jackson, B.

    1989-01-01

    Despite the fact that the flow over a rotor blade is strongly influenced by locally three-dimensional and unsteady effects, practical experience has always demonstrated that substantial improvements in the aerodynamic performance can be gained by improving the steady two-dimensional charateristics of the airfoil(s) employed. The two phenomena known to have great impact on the overall rotor performance are: (1) retreating blade stall with the associated large pressure drag, and (2) compressibility effects on the advancing blade leading to shock formation and the associated wave drag and boundary-layer separation losses. It was concluded that: optimization routines are a powerful tool for finding solutions to multiple design point problems; the optimization process must be guided by the judicious choice of geometric and aerodynamic constraints; optimization routines should be appropriately coupled to viscous, not inviscid, transonic flow solvers; hybrid design procedures in conjunction with optimization routines represent the most efficient approach for rotor airfroil design; unsteady effects resulting in the delay of lift and moment stall should be modeled using simple empirical relations; and inflight optimization of aerodynamic loads (e.g., use of variable rate blowing, flaps, etc.) can satisfy any number of requirements at design and off-design conditions.

  16. 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.

  17. Applications of a transonic wing design method

    NASA Technical Reports Server (NTRS)

    Campbell, Richard L.; Smith, Leigh A.

    1989-01-01

    A method for designing wings and airfoils at transonic speeds using a predictor/corrector approach was developed. The procedure iterates between an aerodynamic code, which predicts the flow about a given geometry, and the design module, which compares the calculated and target pressure distributions and modifies the geometry using an algorithm that relates differences in pressure to a change in surface curvature. The modular nature of the design method makes it relatively simple to couple it to any analysis method. The iterative approach allows the design process and aerodynamic analysis to converge in parallel, significantly reducing the time required to reach a final design. Viscous and static aeroelastic effects can also be accounted for during the design or as a post-design correction. Results from several pilot design codes indicated that the method accurately reproduced pressure distributions as well as the coordinates of a given airfoil or wing by modifying an initial contour. The codes were applied to supercritical as well as conventional airfoils, forward- and aft-swept transport wings, and moderate-to-highly swept fighter wings. The design method was found to be robust and efficient, even for cases having fairly strong shocks.

  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. Computation of Transonic Flows Using Potential Methods

    NASA Technical Reports Server (NTRS)

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

    1997-01-01

    The proposed paper will describe the state of the art associated with numerical solution of the full or exact velocity potential equation for solving transonic, external-aerodynamic flows. The presentation will begin with a review of the literature emphasizing research activities of the past decade. Next, the various forms of the full or exact velocity potential equation, the equation's corresponding mathematical characteristics, and the derivation assumptions will be presented and described in detail. Impact of the derivation assumptions on simulation accuracy, especially with respect to shock wave capture, will be presented and discussed relative to the more complete Euler or Navier-Stokes formulations. The technical presentation will continue with a description of recently developed full potential numerical approach characteristics. This description will include governing equation nondimensionalization, physical-to-computational-domain mapping procedures, a limited description of grid generation requirements, the spatial discretization scheme, numerical implementation of boundary conditions, and the iteration scheme. The next portion of the presentation will present and discuss numerical results for several two- and three-dimensional aerodynamic applications. Included in the results section will be a discussion and demonstration of a typical grid refinement analysis for determining spatial convergence of the numerical solution and level of solution accuracy. Computer timings for a variety of full potential applications will be compared and contrasted with similar results for the Euler equation formulation. Finally. the presentation will end with concluding remarks and recommendations for future work.

  20. Nonlinear Green's function method for unsteady transonic flows

    NASA Technical Reports Server (NTRS)

    Tseng, K.; Morino, L.

    1982-01-01

    Advantages to employing Green's function in describing unsteady three-dimensional transonic flows are explored. The development of the function for application to linear subsonic and supersonic unsteady aerodynamics is reviewed. It is shown that unique solutions are possible for external flows, with all functional expressions being defined in Prandtl-Glauert space. The development of methods of using the Green's function for transonic flows is traced, noting the necessity of including the effects of significant nonlinear terms. The steady-state problem is considered to demonstrate the shock-capturing ability of the method and the usefulness of the function in the incompressible, subsonic, transonic, and supersonic areas of potential unsteady three-dimensional flows around complex configurations. Computational time is asserted to be an order of magnitude less than with finite difference methods.

  1. 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.

  2. Transonic flow control by means of local energy deposition

    NASA Astrophysics Data System (ADS)

    Aul'Chenko, S. M.; Zamuraev, V. P.; Kalinina, A. P.

    2011-11-01

    Experimental data for the feasibility of transonic flow control by means of energy deposition are generalized. Energy supplied to the immediate vicinity of a body in stream before a compression shock is found to result in the nonlinear interaction of introduced disturbances with the shock and the surface in zones extended along the surface. A new, explosive gasdynamic mechanism behind the shift of the compression shock is discovered. It is shown that the nonlinear character of the interaction may considerably decrease the wave resistance of, e.g., transonic airfoils. It is found that energy supply from without stabilizes a transonic flow about an airfoil—the effect similar to the Khristianovich stabilization effect. The dependence of the energy deposition optimal frequency on the energy source parameters and Mach number of the incoming flow at which the resistance drops to the greatest extent is obtained. The influence of the real thermodynamic properties and viscosity of air is studied.

  3. Solution of steady and unsteady transonic-vortex flows using Euler and full-potential equations

    NASA Technical Reports Server (NTRS)

    Kandil, Osama A.; Chuang, Andrew H.; Hu, Hong

    1989-01-01

    Two methods are presented for inviscid transonic flows: unsteady Euler equations in a rotating frame of reference for transonic-vortex flows and integral solution of full-potential equation with and without embedded Euler domains for transonic airfoil flows. The computational results covered: steady and unsteady conical vortex flows; 3-D steady transonic vortex flow; and transonic airfoil flows. The results are in good agreement with other computational results and experimental data. The rotating frame of reference solution is potentially efficient as compared with the space fixed reference formulation with dynamic gridding. The integral equation solution with embedded Euler domain is computationally efficient and as accurate as the Euler equations.

  4. Aerodynamic results of a separation effects test on a 0.01-scale model (52-OTS) of integrated SSV in the AEDC/VKF 40-by-40 inch supersonic wind tunnel A, volume 1

    NASA Technical Reports Server (NTRS)

    Campbell, J. H., II

    1975-01-01

    Experimental aerodynamic investigations were conducted, during the period July 18-19, 1974, in the AEDC/VKF Tunnel A facility on a 0.01-scale model (52-OTS) of the integrated space shuttle vehicle, including only one SRB. The purpose of the investigation was to obtain data for close-in proximity (SRB to orbiter/tank) effects with the orbiter/tank combination at relatively high alpha and beta attitudes, and with the SRB separation motors off. The AEDC Captive Trajectory System (CTS), which supported the SRB, was used in conjunction with the tunnel primary sector (supporting the orbiter/tank) to obtain grid type separation effects data. The one symmetrical SRB model was used interchangeably to obtain both right-hand and left-hand SRB data. Free-stream data were also obtained for the orbiter/tank and for the SRB. This data was used to provide baselines for proximity effects. The entire investigation was conducted at a free-stream Mach number of 4.5 with unit Reynolds number ranging from 4.0 to 6.5 million per foot.

  5. 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.

  6. 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.

  7. 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.

  8. 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.

  9. 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.

  10. 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.

  11. 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.

  12. Issac, Jason Cherian ses in transonic flow

    NASA Technical Reports Server (NTRS)

    Issac, Jason Cherion; Kapania, Rakesh K.

    1993-01-01

    Flutter analysis of a two degree of freedom airfoil in compressible flow is performed using a state-space representation of the unsteady aerodynamic behavior. Indicial response functions are used to represent the normal force and moment response of the airfoil. The structural equations of motion of the airfoil with bending and torsional degrees of freedom are coupled to the unsteady air loads and the aeroelastic system so modelled is solved as an eigenvalue problem to determine the stability. The aeroelastic equations are also directly integrated with respect to time and the time-domain results compared with the results from the eigenanalysis. A good agreement is obtained. The derivatives of the flutter speed obtained from the eigenanalysis are calculated with respect to the mass and stiffness parameters by both analytical and finite-difference methods for various transonic Mach numbers. The experience gained from the two degree of freedom model is applied to study the sensitivity of the flutter response of a wing with respect to various shape parameters. The parameters being considered are as follows: (1) aspect ratio; (2) surface area of the wing; (3) taper ratio; and (4) sweep. The wing deflections are represented by Chebyshev polynomials. The compressible aerodynamic state-space model used for the airfoil section is extended to represent the unsteady aerodynamic forces on a generally laminated tapered skewed wing. The aeroelastic equations are solved as an eigenvalue problem to determine the flutter speed of the wing. The derivatives of the flutter speed with respect to the shape parameters are calculated by both analytical and finite difference methods.

  13. Computed Aeroelastic Motions Of Wings In Transonic Flows

    NASA Technical Reports Server (NTRS)

    Guruswamy, Guru P.; Obayashi, Shigeru

    1995-01-01

    Report describes computational simulations of aeroelastic motions of delta and swept wings in transonic flows. Study directed toward understanding aerodynamic behavior and enhancing maneuverability of fighter airplanes equipped with such wings. Also has implications for gas pumps and turbines, in which flows near tips of vanes and blades reach supersonic speeds.

  14. 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

  15. 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.

  16. 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.

  17. Intermittent Flow Regimes in a Transonic Fan Airfoil Cascade

    NASA Technical Reports Server (NTRS)

    Lepicovsky, J.; McFarland, E. R.; Chima, R. V.; Capece, V. R.; Hayden, J.

    2002-01-01

    A study was conducted in the NASA Glenn Research Center linear cascade on the intermittent flow on the suction surface of an airfoil section from the tip region of a modern low aspect ratio fan blade. Experimental results revealed that, at a large incidence angle, a range of transonic inlet Mach numbers exist where the leading-edge shock-wave pattern was unstable. Flush mounted high frequency response pressure transducers indicated large local jumps in the pressure in the leading edge area, which generates large intermittent loading on the blade leading edge. These measurements suggest that for an inlet Mach number between 0.9 and 1.0 the flow is bi-stable, randomly switching between subsonic and supersonic flows. Hence, it appears that the change in overall flow conditions in the transonic region is based on the frequency of switching between two stable flow states rather than on the continuous increase of the flow velocity. To date, this flow behavior has only been observed in a linear transonic cascade. Further research is necessary to confirm this phenomenon occurs in actual transonic fans and is not the byproduct of an endwall restricted linear cascade.

  18. Transonic Shock Problem for the Euler System in a Nozzle

    NASA Astrophysics Data System (ADS)

    Xin, Zhouping; Yan, Wei; Yin, Huicheng

    2009-10-01

    In this paper, we study the well-posedness problem on transonic shocks for steady ideal compressible flows through a two-dimensional slowly varying nozzle with an appropriately given pressure at the exit of the nozzle. This is motivated by the following transonic phenomena in a de Laval nozzle. Given an appropriately large receiver pressure P r , if the upstream flow remains supersonic behind the throat of the nozzle, then at a certain place in the diverging part of the nozzle, a shock front intervenes and the flow is compressed and slowed down to subsonic speed, and the position and the strength of the shock front are automatically adjusted so that the end pressure at exit becomes P r , as clearly stated by Courant and Friedrichs [Supersonic flow and shock waves, Interscience Publishers, New York, 1948 (see section 143 and 147)]. The transonic shock front is a free boundary dividing two regions of C 2,α flow in the nozzle. The full Euler system is hyperbolic upstream where the flow is supersonic, and coupled hyperbolic-elliptic in the downstream region Ω+ of the nozzle where the flow is subsonic. Based on Bernoulli’s law, we can reformulate the problem by decomposing the 3 × 3 Euler system into a weakly coupled second order elliptic equation for the density ρ with mixed boundary conditions, a 2 × 2 first order system on u 2 with a value given at a point, and an algebraic equation on ( ρ, u 1, u 2) along a streamline. In terms of this reformulation, we can show the uniqueness of such a transonic shock solution if it exists and the shock front goes through a fixed point. Furthermore, we prove that there is no such transonic shock solution for a class of nozzles with some large pressure given at the exit.

  19. A statistical approach to the experimental evaluation of transonic turbine airfoils in a linear cascade

    SciTech Connect

    Shelton, M.L.; Gregory, B.A. ); Doughty, R.L.; Kiss, T.; Moses, H.L. . Mechanical Engineering Dept.)

    1993-07-01

    In aircraft engine design (and in other applications), small improvements in turbine efficiency may be significant. Since analytical tools for predicting transonic turbine losses are still being developed, experimental efforts are required to evaluate various designs, calibrate design methods, and validate CFD analysis tools. However, these experimental efforts must be very accurate to measure the performance differences to the levels required by the highly competitive aircraft engine market. Due to the sensitivity of transonic and supersonic flow fields, it is often difficult to obtain the desired level of accuracy. In this paper, a statistical approach is applied to the experimental evaluation of transonic turbine airfoils in the VPI and SU transonic cascade facility in order to quantify the differences between three different transonic turbine airfoils. This study determines whether the measured performance differences between the three different airfoils are statistically significant. This study also assesses the degree of confidence in the transonic cascade testing process at VPI and SU.

  20. Report of the panel on dynamics and aeroelasticity. [transonic tunnel capabilities

    NASA Technical Reports Server (NTRS)

    Houbolt, J.

    1977-01-01

    Model scaling for flutter analysis is reviewed. Characteristics of the Langley Transonic Dynamics Tunnel (TDT) are described and several features are recommended for inclusion in the National Transonic Facility. Problem areas suggested for the NTF include: Reynolds number effects on control surface unsteady aerodynamics; effects of Reynolds number on buffet onset and loads; transonic unsteady aerodynamics; and Reynolds number effects on flutter characteristics of wing planforms and airfoils.

  1. Aerodynamic results of a separation effects test conducted in the AEDC 40 by 40 inch tunnel A facility on the Rockwell International launch configuration 3 (model-OTS) integrated vehicle (IA13), volume 1

    NASA Technical Reports Server (NTRS)

    Campbell, J. H., II

    1975-01-01

    Experimental aerodynamic investigations were conducted from July 5 through July 17, 1973, on a 0.01 scale model. The AEDC captive trajectory system was utilized in conjunction with the tunnel primary sector to obtain grid-type data for external tank abort from the orbiter, and for nominal separation of one solid rocket booster from the orbiter-tank combination. Booster separation was investigated with and without separation motors plume simulation. The plumes were generated by eight M sub j = 2.15 nozzles using a 1500 psia cold air supply. Free stream data were obtained for all models (orbiter, tank, orbiter-tank, and right-hand booster) to provide baselines for evaluation of proximity effects.

  2. Results of investigations (OA77 and OA78) on an 0.015-scale 140A/B configuration space shuttle vehicle orbiter model 49-0 in the AEDC VKF B and C wind tunnels, revision A

    NASA Technical Reports Server (NTRS)

    Gillins, R. L.

    1975-01-01

    Aerodynamic data obtained from wind tunnel tests of an 0.015-scale 140A/B configuration SSV Orbiter model in the AEDC VKF B and C wind tunnels are presented. Tests were conducted at Mach numbers of 6 and 8 in the B tunnel and at a Mach number of 10 to in the C tunnel to verify hypersonic stability and control characteristics, determine control surface effectiveness, and investigate Reynolds number effects of the 140A/B configuration. Force data were obtained for various control surface settings and Reynolds numbers in the angle-of-attack range of 15 deg to 45 deg and at angles of sideslip of -5 deg to +10 deg. Data were obtained for a few configurations at angles of attack from -27 deg to 45 deg. Control surface variables included elevon, rudder, speedbrake and bodyflap deflections. The effects of an alternate wing leading edge shape were investigated to determine its hypersonic stability and control characteristics.

  3. Implicit approximate-factorization schemes for the efficient solution of steady transonic flow problems

    NASA Technical Reports Server (NTRS)

    Ballhaus, W. F.; Jameson, A.; Albert, J.

    1977-01-01

    Implicit approximate-factorization algorithms (AF) are developed for the solution of steady-state transonic flow problems. The performance of the AF solution method is evaluated relative to that of the standard solution method for transonic flow problems, successive line over-relaxation (SLOR). Both methods are applied to the solution of the nonlinear, two-dimensional transonic small-disturbance equation. Results indicate that the AF method requires substantially less computer time than SLOR to solve the nonlinear finite-difference matrix equation for a transonic flow field. This increase in computational efficiency is achieved with no appreciable increase in computer storage or coding complexity.

  4. 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.

  5. A linearized Euler analysis of unsteady transonic flows in turbomachinery

    SciTech Connect

    Hall, K.C.; Clark, W.S.; Lorence, C.B. . Dept. of Mechanical Engineering and Materials Science)

    1994-07-01

    A computational method for efficiently predicting unsteady transonic flows in two- and three-dimensional cascades is presented. The unsteady flow is modeled using a linearized Euler analysis whereby the unsteady flow field is decomposed into a nonlinear mean flow plus a linear harmonically varying unsteady flow. The equations that govern the perturbation flow, the linearized Euler equations, are linear variable coefficient equations. For transonic flows containing shocks, shock capturing is used to model the shock impulse (the unsteady load due to the harmonic motion of the shock). A conservative Lax-Wendroff scheme is used to obtain a set of linearized finite volume equations that describe the harmonic small disturbance behavior of the flow. Conditions under which such a discretization will correctly predict the shock impulse are investigated. Computational results are presented that demonstrate the accuracy and efficiency of the present method as well as the essential role of unsteady shock impulse loads on the flutter stability of fans.

  6. 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.

  7. A finite-difference method for transonic airfoil design.

    NASA Technical Reports Server (NTRS)

    Steger, J. L.; Klineberg, J. M.

    1972-01-01

    This paper describes an inverse method for designing transonic airfoil sections or for modifying existing profiles. Mixed finite-difference procedures are applied to the equations of transonic small disturbance theory to determine the airfoil shape corresponding to a given surface pressure distribution. The equations are solved for the velocity components in the physical domain and flows with embedded shock waves can be calculated. To facilitate airfoil design, the method allows alternating between inverse and direct calculations to obtain a profile shape that satisfies given geometric constraints. Examples are shown of the application of the technique to improve the performance of several lifting airfoil sections. The extension of the method to three dimensions for designing supercritical wings is also indicated.

  8. 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.

  9. Three-dimensional shock structure in a transonic flutter cascade

    NASA Technical Reports Server (NTRS)

    Boldman, D. R.; Buggele, A. E.; Decker, A. J.

    1982-01-01

    Rapid double-pulse holography was employed to obtain detailed, two-dimensional images of the shock forming during simulated flutter in a transonic flowfield. The experiment comprised a linear cascade of airfoils externally oscillated in torsion and viewed tangentially at the shock surface. Three biconvex airfoils were subjected to harmonic pitching motion about the midchord axis at a frequency of 0.53 while immersed in a Mach 0.81 flow. Failure to produce observable shocks led to use of choked flow with a Mach number near one, of which 50 holograms were taken. The images revealed a narrow shock surface with a spanwise variation in the shock properties. The method is concluded to be useful for examining transonic flowfield shocks in the presence of airfoil flutter.

  10. 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.

  11. The transonic multi-foil Augmentor-Wing

    NASA Technical Reports Server (NTRS)

    Farbridge, J. E.; Smith, R. C.

    1977-01-01

    The paper describes the development of a transonic blown multi-foil Augmentor-Wing airfoil section that has a thickness/chord (t/c) value of 0.18. In comparison with an unblown single-foil supercritical section of the same overall t/c the new multi-foil section is characterized by an increased drag rise Mach number, increased buffet boundaries, and a reduction in 'effective' drag due to blowing. Potential advantages of the Augmentor-Wing are considered and the testing of three high-speed models in a trisonic pressurized wind tunnel (possessing a two-dimensional transonic insert) is discussed. The data indicate that a very thick wing is feasible since separations toward the rear of the main foil can be controlled both by shroud location and augmentor blowing.

  12. Experimental transonic flutter characteristics of supersonic cruise configurations

    NASA Technical Reports Server (NTRS)

    Durham, Michael H.; Cole, Stanley R.; Cazier, F. W., Jr.; Keller, Donald F.; Parker, Ellen C.; Wilkie, W. Keats

    1990-01-01

    The flutter characteristics of a generic arrow-wing supersonic transport configuration are studied. The wing configuration has a 3 percent biconvex airfoil and a leading-edge sweep of 73 deg out to a cranked tip with a 60 deg leading-edge sweep. The ground vibration tests and flutter test procedure are described. The effects of flutter on engine nacelles, fuel loading, wing-mounted vertical fin, wing angle-of-attack, and wing tip mass and stiffness distributions are analyzed. The data reveal that engine nacelles reduce the transonic flutter dynamic pressure by 25-30 percent; fuel loadings decrease dynamic pressures by 25 percent; 4-6 deg wing angles-of-attack cause steep transonic boundaries; and 5-10 percent changes in flutter dynamic pressures are the result of the wing-mounted vertical fin and wing-tip mass and stiffness distributions.

  13. Geared-elevator flutter study. [transonic flutter characteristics of empennage

    NASA Technical Reports Server (NTRS)

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

    1976-01-01

    The paper describes an experimental and analytical study of the transonic flutter characteristics of an empennage flutter model having an all-movable horizontal tail with a geared elevator. Two configurations were flutter tested: one with a geared elevator and one with a locked elevator with the model cantilever-mounted on a sting in the wind tunnel. The geared-elevator configuration fluttered experimentally at about 20% higher dynamic pressures than the locked-elevator configuration. The experimental flutter boundary was nearly flat at transonic speeds for both configurations. It was found that an analysis which treated the elevator as a discrete surface predicted flutter dynamic pressure levels better than analyses which treated the stabilizer and elevator as a warped surface. Warped-surface methods, however, predicted more closely the experimental flutter frequencies and Mach number trends.

  14. Recent developments in finite element analysis for transonic airfoils

    NASA Technical Reports Server (NTRS)

    Hafez, M. M.; Murman, E. M.

    1979-01-01

    The prediction of aerodynamic forces in the transonic regime generally requires a flow field calculation to solve the governing non-linear mixed elliptic-hyperbolic partial differential equations. Finite difference techniques were developed to the point that design and analysis application are routine, and continual improvements are being made by various research groups. The principal limitation in extending finite difference methods to complex three-dimensional geometries is the construction of a suitable mesh system. Finite element techniques are attractive since their application to other problems have permitted irregular mesh elements to be employed. The purpose of this paper is to review the recent developments in the application of finite element methods to transonic flow problems and to report some recent results.

  15. Transonic wind-tunnel tests of a lifting parachute model

    NASA Technical Reports Server (NTRS)

    Foughner, J. T., Jr.; Reed, J. F.; Wynne, E. C.

    1976-01-01

    Wind-tunnel tests have been made in the Langley transonic dynamics tunnel on a 0.25-scale model of Sandia Laboratories' 3.96-meter (13-foot), slanted ribbon design, lifting parachute. The lifting parachute is the first stage of a proposed two-stage payload delivery system. The lifting parachute model was attached to a forebody representing the payload. The forebody was designed and installed in the test section in a manner which allowed rotational freedom about the pitch and yaw axes. Values of parachute axial force coefficient, rolling moment coefficient, and payload trim angles in pitch and yaw are presented through the transonic speed range. Data are presented for the parachute in both the reefed and full open conditions. Time history records of lifting parachute deployment and disreefing tests are included.

  16. A tomographic technique for aerodynamics at transonic speeds

    NASA Technical Reports Server (NTRS)

    Lee, G.

    1985-01-01

    Computer aided tomography (CAT) provides a means of noninvasively measuring the air density distribution around an aerodynamic model. This technique is global in that a large portion of the flow field can be measured. A test of the applicability of CAT to transonic velocities was studied. A hemispherical-nose cylinder afterbody model was tested at a Mach number of 0.8 with a new laser holographic interferometer at the 2- by 2-Foot Transonic Wind Tunnel. Holograms of the flow field were taken and were reconstructed into interferograms. The fringe distribution (a measure of the local densities) was digitized for subsequent data reduction. A computer program based on the Fourier-transform technique was developed to convert the fringe distribution into three-dimensional densities around the model. Theoretical aerodynamic densities were calculated for evaluating and assessing the accuracy of the data obtained from the tomographic method.

  17. 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.

  18. Shockless design and analysis of transonic blade shapes

    NASA Technical Reports Server (NTRS)

    Dulikravich, D. S.; Sobieczky, H.

    1981-01-01

    A fast computer program was developed to eliminate the shocks by slightly altering portions of the contour of a given airfoil in the cascade. The program can be used in two basic modes: (1) An analysis for steady, transonic, potential flow through a given planar cascade of airfoils and (2) a design for converting a given cascade into a shockless transonic cascade. The design mode can automatically be followed by the analysis mode, which confirms that the flow field is shock free. The program generates its own multilevel boundary conforming computational grids and solves a full potential equation in a fully conservative form. The shockless design is performed by implementing Sobieczky's fictitious-gas elliptic continuation concept.

  19. Prediction of unsteady transonic flow around missile configurations

    NASA Technical Reports Server (NTRS)

    Nixon, D.; Reisenthel, P. H.; Torres, T. O.; Klopfer, G. H.

    1990-01-01

    This paper describes the preliminary development of a method for predicting the unsteady transonic flow around missiles at transonic and supersonic speeds, with the final goal of developing a computer code for use in aeroelastic calculations or during maneuvers. The basic equations derived for this method are an extension of those derived by Klopfer and Nixon (1989) for steady flow and are a subset of the Euler equations. In this approach, the five Euler equations are reduced to an equation similar to the three-dimensional unsteady potential equation, and a two-dimensional Poisson equation. In addition, one of the equations in this method is almost identical to the potential equation for which there are well tested computer codes, allowing the development of a prediction method based in part on proved technology.

  20. 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.

  1. 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.-

  2. 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.

  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. 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.

  5. Design of transonic airfoil sections using a similarity theory

    NASA Technical Reports Server (NTRS)

    Nixon, D.

    1978-01-01

    A study of the available methods for transonic airfoil and wing design indicates that the most powerful technique is the numerical optimization procedure. However, the computer time for this method is relatively large because of the amount of computation required in the searches during optimization. The optimization method requires that base and calibration solutions be computed to determine a minimum drag direction. The design space is then computationally searched in this direction; it is these searches that dominate the computation time. A recent similarity theory allows certain transonic flows to be calculated rapidly from the base and calibration solutions. In this paper the application of the similarity theory to design problems is examined with the object of at least partially eliminating the costly searches of the design optimization method. An example of an airfoil design is presented.

  6. 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.

  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. A computational design method for transonic turbomachinery cascades

    NASA Technical Reports Server (NTRS)

    Sobieczky, H.; Dulikravich, D. S.

    1982-01-01

    This paper describes a systematical computational procedure to find configuration changes necessary to modify the resulting flow past turbomachinery cascades, channels and nozzles, to be shock-free at prescribed transonic operating conditions. The method is based on a finite area transonic analysis technique and the fictitious gas approach. This design scheme has two major areas of application. First, it can be used for design of supercritical cascades, with applications mainly in compressor blade design. Second, it provides subsonic inlet shapes including sonic surfaces with suitable initial data for the design of supersonic (accelerated) exits, like nozzles and turbine cascade shapes. This fast, accurate and economical method with a proven potential for applications to three-dimensional flows is illustrated by some design examples.

  9. Numerical studies of unsteady transonic flow over an oscillating airfoil

    NASA Technical Reports Server (NTRS)

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

    1984-01-01

    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.

  10. 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.

  11. National Transonic Facility Fan Blade prepreg material characterization tests

    NASA Technical Reports Server (NTRS)

    Klich, P. J.; Richards, W. H.; Ahl, E. L., Jr.

    1981-01-01

    The test program for the basic prepreg materials used in process development work and planned fabrication of the national transonic facility fan blade is presented. The basic prepreg materials and the design laminate are characterized at 89 K, room temperature, and 366 K. Characterization tests, test equipment, and test data are discussed. Material tests results in the warp direction are given for tensile, compressive, fatigue (tension-tension), interlaminar shear and thermal expansion.

  12. 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.

  13. Preliminary calibration and test results from the National Transonic Facility

    NASA Technical Reports Server (NTRS)

    Mckinney, Linwood W.; Fuller, Dennis E.

    1986-01-01

    The National Transonic Facility (NTF) was operated to design condition of 120 million Reynolds number at a Mach number of 1.0. All systems were checked out except plenum isolation valves; modifications are being made to heaters on the actuators. Initial steady-state calibration indicates excellent steady flow characteristics. The first test of the Pathfinder 1 model indicated significant Reynolds number effects. Some effect of temperature on instrumentation were obtained. The cause of these effects is being evaluated.

  14. Spectral multigrid methods with applications to transonic potential flow

    NASA Technical Reports Server (NTRS)

    Streett, C. L.; Zang, T. A.; Hussaini, M. Y.

    1983-01-01

    Spectral multigrid methods are demonstrated to be a competitive technique for solving the transonic potential flow equation. The spectral discretization, the relaxation scheme, and the multigrid techniques are described in detail. Significant departures from current approaches are first illustrated on several linear problems. The principal applications and examples, however, are for compressible potential flow. These examples include the relatively challenging case of supercritical flow over a lifting airfoil.

  15. Spectral multigrid methods with applications to transonic potential flow

    NASA Technical Reports Server (NTRS)

    Streett, C. L.; Zang, T. A.; Hussaini, M. Y.

    1985-01-01

    Spectral multigrid methods are demonstrated to be a competitive technique for solving the transonic potential flow equation. The spectral discretization, the relaxation scheme, and the multigrid techniques are described in detail. Significant departures from current approaches are first illustrated on several linear problems. The principal applications and examples, however, are for compressible potential flow. These examples include the relatively challenging case of supercritical flow over a lifting airfoil.

  16. 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.

  17. 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.

  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. Unsteady transonic flow calculations for wing-fuselage configurations

    NASA Technical Reports Server (NTRS)

    Batina, J. T.

    1986-01-01

    Unsteady transonic flow calculations are presented for wing-fuselage configurations. Calculations are performed by extending the XTRAN3S unsteady transonic small-disturbance code to allow the treatment of a fuselage. Details of the XTRAN3S fuselage modeling are discussed in the context of the small-disturbance equation. Transonic calculations are presented for three wing-fuselage configurations with leading edge sweep angles ranging from 0 deg to 46.76 deg. Simple bending and torsion modal oscillations of the wing are calculated. Sectional lift and moment coefficients for the wing-alone and wing-fuselage cases are compared and the effects of fuselage aerodynamic interference on the unsteady wing loading are revealed. Tabulated generalized aerodynamic forces used in flutter analyses, indicate small changes in the real in-phase component and as much as a 30% change in the imaginary component when the fuselage is included in the calculation. These changes result in a 2 to 5% increase in total magnitude and a several degree increase in phase.

  20. 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.

  1. 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.

  2. 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.

  3. Fourier time spectral method for subsonic and transonic flows

    NASA Astrophysics Data System (ADS)

    Zhan, Lei; Liu, Feng; Papamoschou, Dimitri

    2016-01-01

    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.

  4. Transonic flows of dense gases over finite wings

    NASA Astrophysics Data System (ADS)

    Cinnella, P.

    2008-04-01

    Transonic inviscid flows of dense gases of the Bethe-Zel'dovich-Thompson (BZT) type over finite wings are numerically investigated. BZT gases are fluids of the retrograde type (i.e., that superheat when expanded), which exhibit a region of negative values of the fundamental derivative of gas dynamics Γ. As a consequence, they display, in the transonic and supersonic regime, nonclassical gas dynamic behaviors, such as rarefaction shock waves and mixed shock/fan waves. The peculiar properties of BZT fluids have received increased interest in recent years because of their possible application in energy-conversion cycles. The present research aims at providing insight about the transonic aerodynamics of BZT fluids past finite wings, roughly representative of isolated turbine blades with infinite tip leakage. This represents an important step toward the design of advanced turbine blades by using organic working fluids. An investigation of the flow patterns and aerodynamic performance for several choices of the upstream thermodynamic conditions is provided, and the advantages of using BZT working fluids instead of classical ones are discussed.

  5. A numerical study of flutter in a transonic fan

    SciTech Connect

    Isomura, K.; Giles, M.B.

    1998-07-01

    The bending mode flutter of a modern transonic fan has been studied using a quasi-three-dimensional viscous unsteady CFD code. The type of flutter in this research is that of a highly loaded blade with a tip relative Mach number just above unity, commonly referred to as transonic stall flutter. This type of flutter is often encountered in modern wide chord fans without a part span shroud. The CFD simulation uses an upwinding scheme with Roe`s third-order flux differencing, and Johnson and King`s turbulence model with the later modification due to Johnson and Coakley. A dynamic transition point model is developed using the e{double_prime} method and Schubauer and Klebanoff`s experimental data. The calculations of the flow in this fan reveal that the source of the flutter of 1H1 transonic fan is an oscillation of the passage shock, rather than a stall. As the blade loading increases, the passage shock moves forward. Just before the passage shock unstarts, the stability of the passage shock decreases, and a small blade vibration causes the shock to oscillate with a large amplitude between unstarted and started positions. The dominant component of the blade excitation force is due to the foot of the oscillating passage shock on the blade pressure surface.

  6. 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.

  7. 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.

  8. 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

  9. 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.

  10. 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.

  11. Prediction of transonic flutter for a supercritical wing by modified strip analysis and comparison with experiment

    NASA Technical Reports Server (NTRS)

    Yates, E. C., Jr.; Wynne, E. C.; Farmer, M. G.; Desmarais, R. N.

    1981-01-01

    Use of a supercritical airfoil can adversely affect wing flutter speeds in the transonic range. As adequate theories for three dimensional unsteady transonic flow are not yet available, the modified strip analysis was used to predict the transonic flutter boundary for the supercritical wing. The steady state spanwise distributions of section lift curve slope and aerodynamic center, required as input for the flutter calculations, were obtained from pressure distributions. The calculated flutter boundary is in agreement with experiment in the subsonic range. In the transonic range, a transonic bucket is calculated which closely resembles the experimental one with regard to both shape and depth, but it occurs at about 0.04 Mach number lower than the experimental one.

  12. 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.

  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. A transonic-small-disturbance wing design methodology

    NASA Technical Reports Server (NTRS)

    Phillips, Pamela S.; Waggoner, Edgar G.; Campbell, Richard L.

    1988-01-01

    An automated transonic design code has been developed which modifies an initial airfoil or wing in order to generate a specified pressure distribution. The design method uses an iterative approach that alternates between a potential-flow analysis and a design algorithm that relates changes in surface pressure to changes in geometry. The analysis code solves an extended small-disturbance potential-flow equation and can model a fuselage, pylons, nacelles, and a winglet in addition to the wing. A two-dimensional option is available for airfoil analysis and design. Several two- and three-dimensional test cases illustrate the capabilities of the design code.

  15. Initial research program for the National Transonic Facility

    NASA Technical Reports Server (NTRS)

    Gloss, B. B.

    1984-01-01

    The construction and checkout of the National Transonic Facility (NTF) have been completed, and detailed calibration is now in progress. The initial NTF research program covers a wide range of study areas falling into three major elements: (1) the assessment of Reynolds number sensitivities for a broad range of configurations and flow phenomena; (2) validation of the ability of NTF to simulate full-scale aerodynamics; and (3) the development of test techniques for improved test simulations in existing wind tunnels. This paper, therefore, is a status report on these various elements of the initial NTF research program.

  16. A hybrid algorithm for transonic airfoil and wing design

    NASA Technical Reports Server (NTRS)

    Campbell, Richard L.; Smith, Leigh A.

    1987-01-01

    The present method for the design of transonic airfoils and wings employs a predictor/corrector approach in which an analysis code calculates the flowfield for an initial geometry, then modifies it on the basis of the difference between calculated and target pressures. This allows the design method to be straightforwardly coupled with any existing analysis code, as presently undertaken with several two- and three-dimensional potential flow codes. The results obtained indicate that the method is robust and accurate, even in the cases of airfoils with strongly supercritical flow and shocks. The design codes are noted to require computational resources typical of current pure-inverse methods.

  17. An inverse method with regularity condition for transonic airfoil design

    NASA Technical Reports Server (NTRS)

    Zhu, Ziqiang; Xia, Zhixun; Wu, Liyi

    1991-01-01

    It is known from Lighthill's exact solution of the incompressible inverse problem that in the inverse design problem, the surface pressure distribution and the free stream speed cannot both be prescribed independently. This implies the existence of a constraint on the prescribed pressure distribution. The same constraint exists at compressible speeds. Presented here is an inverse design method for transonic airfoils. In this method, the target pressure distribution contains a free parameter that is adjusted during the computation to satisfy the regularity condition. Some design results are presented in order to demonstrate the capabilities of the method.

  18. Refined numerical solution of the transonic flow past a wedge

    NASA Technical Reports Server (NTRS)

    Liang, S.-M.; Fung, K.-Y.

    1985-01-01

    A numerical procedure combining the ideas of solving a modified difference equation and of adaptive mesh refinement is introduced. The numerical solution on a fixed grid is improved by using better approximations of the truncation error computed from local subdomain grid refinements. This technique is used to obtain refined solutions of steady, inviscid, transonic flow past a wedge. The effects of truncation error on the pressure distribution, wave drag, sonic line, and shock position are investigated. By comparing the pressure drag on the wedge and wave drag due to the shocks, a supersonic-to-supersonic shock originating from the wedge shoulder is confirmed.

  19. Interaction of multiple supersonic jets with a transonic flow field

    NASA Technical Reports Server (NTRS)

    Seginer, A.; Manela, J.

    1983-01-01

    The influence of multiple high pressure, supersonic, radial or tangential jets, that are injected from the circumference of the base plane of an axisymmetric body, on its longitudinal aerodynamic coefficients in transonic flow is studied experimentally. The interaction of the jets with the body flow field increases the pressures on the forebody, thus altering its lift and static stability characteristics. It is shown that, within the range of parameters studied. This interaction has a stabilizing effect on the body. The contribution to lift and stability is significant at small angles of attack and decreases nonlinearly at higher angles when the crossflow mechanism becomes dominant.

  20. 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.

  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. Fast Euler solver for transonic airfoils. I - Theory. II - Applications

    NASA Technical Reports Server (NTRS)

    Dadone, Andrea; Moretti, Gino

    1988-01-01

    Equations written in terms of generalized Riemann variables are presently integrated by inverting six bidiagonal matrices and two tridiagonal matrices, using an implicit Euler solver that is based on the lambda-formulation. The solution is found on a C-grid whose boundaries are very close to the airfoil. The fast solver is then applied to the computation of several flowfields on a NACA 0012 airfoil at various Mach number and alpha values, yielding results that are primarily concerned with transonic flows. The effects of grid fineness and boundary distances are analyzed; the code is found to be robust and accurate, as well as fast.

  3. 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.

  4. 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.

  5. 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.

  6. An analysis method for two-dimensional transonic viscous flow

    NASA Technical Reports Server (NTRS)

    Bavitz, P. C.

    1975-01-01

    A method for the approximate calculation of transonic flow over airfoils, including shock waves and viscous effects, is described. Numerical solutions are obtained by use of a computer program which is discussed in the appendix. The importance of including the boundary layer in the analysis is clearly demonstrated, as well as the need to improve on existing procedures near the trailing edge. Comparisons between calculations and experimental data are presented for both conventional and supercritical airfoils, emphasis being on the surface pressure distribution, and good agreement is indicated.

  7. Finite element analysis of periodic transonic flow problems

    NASA Technical Reports Server (NTRS)

    Fix, G. J.

    1978-01-01

    Flow about an oscillating thin airfoil in a transonic stream was considered. It was assumed that the flow field can be decomposed into a mean flow plus a periodic perturbation. On the surface of the airfoil the usual Neumman conditions are imposed. Two computer programs were written, both using linear basis functions over triangles for the finite element space. The first program uses a banded Gaussian elimination solver to solve the matrix problem, while the second uses an iterative technique, namely SOR. The only results obtained are for an oscillating flat plate.

  8. 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.

  9. Aerodynamic optimum design of transonic turbine cascades using Genetic Algorithms

    NASA Astrophysics Data System (ADS)

    Li, Jun; Feng, Zhenping; Chang, Jianzhong; Shen, Zuda

    1997-06-01

    This paper presents an aerodynamic optimum design method for transonic turbine cascades based on the Genetic Algorithms coupled to the inviscid flow Euler solver and the boundary-layer calculation. The Genetic Algorithms control the evolution of a population of cascades towards an optimum design. The fitness value of each string is evaluated using the flow solver. The design procedure has been developed and the behavior of the genetic algorithms has been tested. The objective functions of the design examples are the minimum mean-square deviation between the aimed pressure and computed pressure and the minimum amount of user expertise.

  10. Evaluation of flow field approximations for transonic compressor stages

    SciTech Connect

    Dorney, D.J.; Sharma, O.P.

    1997-07-01

    The flow through gas turbine compressors is often characterized by unsteady, transonic, and viscous phenomena. Accurately predicting the behavior of these complex multi-blade-row flows with unsteady rotor-stator interacting Navier-Stokes analyses can require enormous computer resources. In this investigation, several methods for predicting the flow field, losses, and performance quantities associated with axial compressor stages are presented. The methods studied include: (1) the unsteady fully coupled blade row technique, (2) the steady coupled blade row method, (3) the steady single blade row technique, and (4) the loosely coupled blade row method. The analyses have been evaluated in terms of accuracy and efficiency.

  11. Optimum Transonic Airfoils Based on the Euler Equations

    NASA Technical Reports Server (NTRS)

    Iollo, Angelo; Salas, Manuel, D.

    1996-01-01

    We solve the problem of determining airfoils that approximate, in a least square sense, given surface pressure distributions in transonic flight regimes. The flow is modeled by means of the Euler equations and the solution procedure is an adjoint- based minimization algorithm that makes use of the inverse Theodorsen transform in order to parameterize the airfoil. Fast convergence to the optimal solution is obtained by means of the pseudo-time method. Results are obtained using three different pressure distributions for several free stream conditions. The airfoils obtained have given a trailing edge angle.

  12. 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.

  13. 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.

  14. Vectorizable multigrid algorithms for transonic-flow calculations

    NASA Technical Reports Server (NTRS)

    Melson, N. D.

    1986-01-01

    The analysis and the incorporation into a multigrid scheme of several vectorizable algorithms are discussed. von Neumann analyses of vertical-line, horizontal-line, and alternating-direction ZEBRA algorithms were performed; and the results were used to predict their multigrid damping rates. The algorithms were then successfully implemented in a transonic conservative full-potential computer program. The convergence acceleration effect of multiple grids is shown, and the convergence rates of the vectorizable algorithms are compared with those of standard successive-line overrelaxation (SLOR) algorithms.

  15. Wind-US Unstructured Flow Solutions for a Transonic Diffuser

    NASA Technical Reports Server (NTRS)

    Mohler, Stanley R., Jr.

    2005-01-01

    The Wind-US Computational Fluid Dynamics flow solver computed flow solutions for a transonic diffusing duct. The calculations used an unstructured (hexahedral) grid. The Spalart-Allmaras turbulence model was used. Static pressures along the upper and lower wall agreed well with experiment, as did velocity profiles. The effect of the smoothing input parameters on convergence and solution accuracy was investigated. The meaning and proper use of these parameters are discussed for the benefit of Wind-US users. Finally, the unstructured solver is compared to the structured solver in terms of run times and solution accuracy.

  16. 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.

  17. Unsteady transonic flow analysis for low aspect ratio, pointed wings.

    NASA Technical Reports Server (NTRS)

    Kimble, K. R.; Ruo, S. Y.; Wu, J. M.; Liu, D. Y.

    1973-01-01

    Oswatitsch and Keune's parabolic method for steady transonic flow is applied and extended to thin slender wings oscillating in the sonic flow field. The parabolic constant for the wing was determined from the equivalent body of revolution. Laplace transform methods were used to derive the asymptotic equations for pressure coefficient, and the Adams-Sears iterative procedure was employed to solve the equations. A computer program was developed to find the pressure distributions, generalized force coefficients, and stability derivatives for delta, convex, and concave wing planforms.

  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. 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.

  1. Non-isentropic unsteady transonic small disturbance theory

    NASA Technical Reports Server (NTRS)

    Fuglsang, D. F.; Williams, M. H.

    1985-01-01

    Modifications to transonic small disturbance theory (TSD) which more accurately model the Euler equations and seem to remove the problem of nonunique potential flow solutions are presented. The modifications are implemented in the two-dimensional computer code XTRAN2L, and steady and unsteady flow calculations made for the NACA 0012, NLR 7301, and NACA 64A010A airfoils. Comparisons are made with unmodified and modified TSD, Euler, and full potential theories and with experimental data. The modified theory requires only minor coding changes in existing algorithms for calculating small disturbance flows, and results in relatively small increases in computational cost.

  2. 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.

  3. Low aspect ratio transonic rotors: Part 2. Influence of location of maximum thickness on transonic compressor performance

    SciTech Connect

    Wadia, A.R. ); Law, C.H. )

    1993-04-01

    Transonic compressor rotor performance is sensitive to variations in several known design parameters. One such parameter is the chordwise location of maximum thickness. This article reports on the design and experimental evaluation of two versions of a low aspect ratio transonic rotor that had the location of the tip blade section maximum thickness moved forward in two increments from the nominal 70% to 55 and 40% chord length, respectively. The original hub characteristics were preserved and the maximum thickness location was adjusted proportionately along the span. Although designed to satisfy identical design speed requirements, the experimental results reveal significant variation in the performance of the rotors. At design speed, the rotor with its maximum thickness located at 55% chord length attains the highest peak efficiency among the three rotors but has lowest flow rollback relative to the other two versions. To focus on current ruggedization issues for transonic blading (e.g., bird and ice ingestion), detailed comparison of test data and analysis to characterize the aerodynamic flow details responsible for the measured performance differences were confined to the two rotors with the most forward location of maximum thickness. A three-dimensional viscous flow analysis was used to identify the performance-enhancing features of the higher efficiency rotor and to provide guidance in the interpretation of the experimental measurements. The computational results of the viscous analysis show that the difference in performance between the two rotors can be attributed to the higher shock losses that result from the increased leading edge wedge angle as the maximum thickness is moved closer to the leading edge.

  4. 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.

  5. 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.

  6. 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.

  7. 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.

  8. A short history of the European Transonic Wind Tunnel ETW

    NASA Astrophysics Data System (ADS)

    Green, John; Quest, Jürgen

    2011-07-01

    This paper is written as a contribution to the celebration of 50 years of Progress in Aerospace Sciences and of the centenary of the birth of its founder, Dietrich Küchemann. It reviews the evolution of the European Transonic Wind Tunnel, ETW, from early conceptual studies to its entry into service and its current capabilities and achievements. It traces the development, from the earliest days, of experimental aerodynamics and of the basic aerodynamic understanding that gave rise to the main periods of wind tunnel building before and after World War II. By about 1960, this activity appeared to have come to a natural halt. The paper gives an account of the role of Küchemann in arguing the need in 1968 for a further step in wind tunnel capability, to provide transonic testing at high Reynolds numbers. It describes his leading role in gaining acceptance of the concept, formulating the specification and promoting studies of alternative, radical design options for the co-operative European project that became ETW. The progress of ETW through design, construction, commissioning and into full operation is recorded. The paper discusses the many technical innovations that have been introduced in order to meet customer requirements in the challenging field of aerodynamic testing in a cryogenic environment and, finally, looks to the future and the further technical challenges that it holds.

  9. Time-dependent transonic flow solutions for axial turbomachinery

    NASA Technical Reports Server (NTRS)

    Erdos, J.; Alzner, E.; Kalben, P.; Mcnally, W.; Slutsky, S.

    1975-01-01

    Three-dimensional unsteady transonic flow through an axial turbomachine stage is described in terms of a pair of two-dimensional formulations pertaining to orthogonal surfaces, namely, a blade-to-blade surface and a hub-to-casing surface. The resulting systems of nonlinear, inviscid, compressible equations of motion are solved by an explicit finite-difference technique. The blade-to-blade program includes the periodic interaction between rotor and stator blade rows. Treatment of the boundary conditions and of the blade slipstream motion by a characteristic type procedure is discussed in detail. Harmonic analysis of the acoustic far field produced by the blade row interaction, including an arbitrary initial transient, is outlined. Results from the blade-to-blade program are compared with experimental measurements of the rotating pressure field at the tip of a high-speed fan. The hub-to-casing program determines circumferentially averaged flow properties on a meridional plane. Blade row interactions are neglected in this formulation, but the force distributions over the entire blade surface for both the rotor and stator are obtained. Results from the hub-to-casing program are compared with a relaxation method solution for a subsonic rotor. Results are also presented for a quiet fan stage which includes transonic flow in both the rotor and stator and a normal shock in the stator.

  10. Numerical optimization design of advanced transonic wing configurations

    NASA Technical Reports Server (NTRS)

    Cosentino, G. B.; Holst, T. L.

    1984-01-01

    A computationally efficient and versatile technique for use in the design of advanced transonic wing configurations has been developed. A reliable and fast transonic wing flow-field analysis program, TWING, has been coupled with a modified quasi-Newton method, unconstrained optimization algorithm, QNMDIF, to create a new design tool. Fully three-dimensional wing designs utilizing both specified wing pressure distributions and drag-to-lift ration minimization as design objectives are demonstrated. Because of the high computational efficiency of each of the components of the design code, in particular the vectorization of TWING and the high speed of the Cray X-MP vector computer, the computer time required for a typical wing design is reduced by approximately an order of magnitude over previous methods. In the results presented here, this computed wave drag has been used as the quantity to be optimized (minimized) with great success, yielding wing designs with nearly shock-free (zero wave drag) pressure distributions and very reasonable wing section shapes.

  11. 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.

  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. 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.

  14. 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.

  15. 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.

  16. 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.

  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. 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.

  19. 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.

  20. 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.

  1. 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.

  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. 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.

  4. 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.

  5. 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.

  6. Application of a multi-level grid method to transonic flow calculations

    NASA Technical Reports Server (NTRS)

    South, J. C., Jr.; Brandt, A.

    1976-01-01

    A multi-level grid method was studied as a possible means of accelerating convergence in relaxation calculations for transonic flows. The method employs a hierarchy of grids, ranging from very coarse to fine. The coarser grids are used to diminish the magnitude of the smooth part of the residuals. The method was applied to the solution of the transonic small disturbance equation for the velocity potential in conservation form. Nonlifting transonic flow past a parabolic arc airfoil is studied with meshes of both constant and variable step size.

  7. A theoretical basis for extending surface-paneling methods to transonic flow

    NASA Technical Reports Server (NTRS)

    Erickson, L. L.; Strande, S. M.

    1985-01-01

    The surface integral terms in Green's third identity are often used to solve the Prandtl-Glauert (linear potential-flow) equation with panel methods. This can be done, as in the PAN AIR code, for either subsonic or supersonic flow about complete aircraft. The extension to transonic flow is suggested by the volume integral terms of Green's third identity. The mathematical basis for this extension, without the use of body-fitted grids, is presented. Supercritical transonic results computed from a two-dimensional transonic PAN AIR research code demonstrate the method.

  8. Transonic Reynolds Number and Leading-Edge Bluntness Effects on a 65 deg Delta Wing

    NASA Technical Reports Server (NTRS)

    Luckring, J. M.

    2003-01-01

    A 65 deg delta wing has been tested in the National Transonic Facility (NTF) at mean aerodynamic chord Reynolds numbers from 6 million to 120 million at subsonic and transonic speeds. The configuration incorporated a systematic variation of the leading edge bluntness. The analysis for this paper is focused on the Reynolds number and bluntness effects at transonic speeds (M=0.85) from this data set. The results show significant effects of both these parameters on the onset and progression of leading-edge vortex separation.

  9. Transonic Reynolds Number and Leading-Edge Bluntness Effects on a 65 deg Delta Wing

    NASA Technical Reports Server (NTRS)

    Luckring, J. M.

    2003-01-01

    A 65 degree delta wing has been tested in the National Transonic Facility (NTF) at mean aerodynamic chord Reynolds numbers from 6 million to 120 million at subsonic and transonic speeds. The configuration incorporated a systematic variation of the leading edge bluntness. The analysis for this paper is focused on the Reynolds number and bluntness effects at transonic speeds (M = 0.85) from this data set. The results show significant effects of both these parameters on the onset and progression of leading edge vortex separation.

  10. Transonic Reynolds Number and Leading-Edge Bluntness Effects on a 65 deg Delta Wing

    NASA Technical Reports Server (NTRS)

    Luckring, J. M.

    2003-01-01

    A 65 deg delta wing has been tested in the National Transonic Facility (NTF) at mean aerodynamic chord Reynolds numbers from 6 million to 120 million at subsonic and transonic speeds. The configuration incorporated a systematic variation of the leading edge bluntness. The analysis for this paper is focused on the Reynolds number and bluntness effects at transonic speeds (M = 0.85) from this data set. The results show significant effects of both these parameters on the onset and progression of leading- edge vortex separation.

  11. CAP-TSD: A program for unsteady transonic analysis of realistic aircraft configurations

    NASA Technical Reports Server (NTRS)

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

    1989-01-01

    The development of a new transonic code to predict unsteady flows about realistic aircraft configurations are described. An approximate factorization algorithm for solution of the unsteady transonic small disturbance equation is first described. Because of the superior stability characteristics of the AF algorithm, a new transonic aeroelasticity code was developed which is described in some detail. The new code was very easy to modify to include the additional aircraft components, so in a very short period of time the code was developed to treat complete aircraft configurations. Finally, applications are presented which demonstrate many of the geometry capabilities of the new code.

  12. Artificial compressibility methods for numerical solutions of transonic full potential equation

    NASA Technical Reports Server (NTRS)

    Hafez, M.; Murman, E.; South, J.

    1979-01-01

    New methods for transonic flow computations based on the full potential equation in conservation form are presented. The idea is to modify slightly the density (due to the artificial viscosity in the supersonic region), and solve the resulting elliptic-like problem iteratively. It is shown that standard discretization techniques (central differencing) as well as some standard iterative procedures (SOR, ADI, and explicit methods) are applicable to the modified transonic mixed-type equation. Calculations of transonic flows around cylinders and airfoils are discussed with special emphasis on the explicit methods that are suitable for vector processing on the STAR 100 computer.

  13. Calculative techniques for transonic flows about certain classes of wing-body combinations, phase 2

    NASA Technical Reports Server (NTRS)

    Stahara, S. S.; Spreiter, J. R.

    1972-01-01

    Theoretical analysis and associated computer programs were developed for predicting properties of transonic flows about certain classes of wing-body combinations. The procedures used are based on the transonic equivalence rule and employ either an arbitrarily-specified solution or the local linerization method for determining the nonlifting transonic flow about the equivalent body. The class of wind planform shapes include wings having sweptback trailing edges and finite tip chord. Theoretical results are presented for surface and flow-field pressure distributions for both nonlifting and lifting situations at Mach number one.

  14. Steady, Nonrotating, Blade-to-Blade Potential Transonic Cascade Flow Analysis Code

    NASA Technical Reports Server (NTRS)

    Dulikravich, D. S.

    1983-01-01

    CAS2D computer program numerically solves artifically time-dependent form of actual full potential equation, providing steady, nonrotating, bladeto-blade potential transonic cascade flow analysis code. CAS2D written in FORTRAN IV.

  15. Techniques for correcting approximate finite difference solutions. [applied to transonic flow

    NASA Technical Reports Server (NTRS)

    Nixon, D.

    1979-01-01

    A method of correcting finite-difference solutions for the effect of truncation error or the use of an approximate basic equation is presented. Applications to transonic flow problems are described and examples given.

  16. 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.

  17. 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.

  18. 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).

  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. 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.

  2. The effect of Reynolds number on transonic compressor blade rotor section

    NASA Astrophysics Data System (ADS)

    Beheshti Amiri, H.; Shahrabi Farahani, A.; Khazaei, H.

    2015-12-01

    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.

  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. Aerodynamic design for improved manueverability by use of three-dimensional transonic theory

    NASA Technical Reports Server (NTRS)

    Mann, M. J.; Campbell, R. L.; Ferris, J. C.

    1984-01-01

    Improvements in transonic maneuver performance by the use of three-dimensional transonic theory and a transonic design procedure were examined. The FLO-27 code of Jameson and Caughey was used to design a new wing for a fighter configuration with lower drag at transonic maneuver conditions. The wing airfoil sections were altered to reduce the upper-surface shock strength by means of a design procedure which is based on the iterative application of the FLO-27 code. The plan form of the fighter configuration was fixed and had a leading edge sweep of 45 deg and an aspect ratio of 3.28. Wind-tunnel tests were conducted on this configuration at Mach numbers from 0.60 to 0.95 and angles of attack from -2 deg to 17 deg. The transonic maneuver performance of this configuration was evaluated by comparison with a wing designed by empirical methods and a wing designed primarily by two-dimensional transonic theory. The configuration designed by the use of FLO-27 had the same or lower drag than the empirical wing and, for some conditions, lower drag than the two-dimensional design. From some maneuver conditions, the drag of the two-dimensional design was somewhat lower.

  5. 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.

  6. 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.

  7. 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.

  8. 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.

  9. 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.

  10. 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.

  11. Recent rotorcraft aeroelastic testing in the Langley Transonic Dynamics Tunnel

    NASA Technical Reports Server (NTRS)

    Yeager, William T., Jr.; Mirick, Paul H.; Wilbur, Matthew L.; Singleton, Jeffrey D.; Wilkie, W. K.; Hamouda, M.-N. H.

    1991-01-01

    Wind-tunnel testing of a properly scaled aeroelastic model helicopter rotor is considered a necessary phase in the design and development of new rotor systems. For this reason, extensive testing of aeroelastically scaled model rotors is done in the Transonic Dynamics Tunnel (TDT) located at the Langley Research Center. A unique capability of this facility, which enables proper dynamic scaling, is the use of diflourodichloromethane, or Refrigerant-12 (R-12) as a test medium. The paper presents a description of the TDT and a discussion of the benefits of using R-12 as a test medium. A description of the system used to conduct model tests is provided and examples of recent rotor tests are cited to illustrate the types of aeroelastic model rotor tests conducted in the TDT.

  12. 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.

  13. Stall flutter experiment in a transonic oscillating linear cascade

    NASA Technical Reports Server (NTRS)

    Boldman, D. R.; Buggele, A. E.; Michalson, G. M.

    1981-01-01

    Two dimensional biconvex airfoils were oscillated at reduced frequencies up to 0.5 based on semi-chord and a free stream Mach number of 0.80 to simulate transonic stall flutter in rotors. Steady-state periodicity was confirmed through end-wall pressure measurements, exit flow traverses, and flow visualization. The initial flow visualization results from flutter tests indicated that the oscillating shock on the airfoils lagged the airfoil motion by as much as 80 deg. These initial data exhibited an appreciable amount of scatter; however, a linear fit of the results indicated that the greatest shock phase lag occurred at a positive interblade phase angle. Photographs of the steady-state and unsteady flow fields reveal some of the features of the lambda shock wave on the suction surface of the airfoils.

  14. Radio maps of the solar wind transonic region

    NASA Technical Reports Server (NTRS)

    Lotova, N. A.; Vladimirskii, K. V.; Lebedev, P. N.; Korelov, O. A.

    1995-01-01

    Extensive radio astronomical exploration of near-solar plasmas was carried out in 1987-1993 at radial distances nowadays inaccessible for direct spacecraft experiments. Radio wave sounding experiments were carried out using two types of natural sources-water vapor masers at 1.35 cm and quasars at 2.9 m wavelength. Russian Academy of Sciences radio telescopes RT-22 and DCR-1000 were used. The results of daily observations formed the radial dependence of the radio wave scattering from whence structural peculiarities of the solar wind flow were derived. Simultaneous observations of several sources makes it possible to realize radio maps visualizing the plasma flow stream structure. The 1987 to 1993 maps allow to analyse the transonic region structural changes in relation to the 11-year cycle phase.

  15. 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 problem. 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.

  16. Transonic airfoil analysis and design in nonuniform flow

    NASA Technical Reports Server (NTRS)

    Chang, J. F.; Lan, C. E.

    1986-01-01

    A nonuniform transonic airfoil code is developed for applications in analysis, inverse design and direct optimization involving an airfoil immersed in propfan slipstream. Problems concerning the numerical stability, convergence, divergence and solution oscillations are discussed. The code is validated by comparing with some known results in incompressible flow. A parametric investigation indicates that the airfoil lift-drag ratio can be increased by decreasing the thickness ratio. A better performance can be achieved if the airfoil is located below the slipstream center. Airfoil characteristics designed by the inverse method and a direct optimization are compared. The airfoil designed with the method of direct optimization exhibits better characteristics and achieves a gain of 22 percent in lift-drag ratio with a reduction of 4 percent in thickness.

  17. 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.

  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. Computation of viscous transonic flow about a lifting airfoil

    NASA Technical Reports Server (NTRS)

    Walitt, L.; Liu, C. Y.

    1976-01-01

    The viscous transonic flow about a stationary body in free air was numerically investigated. The geometry chosen was a symmetric NACA 64A010 airfoil at a freestream Mach number of 0.8, a Reynolds number of 4 million based on chord, and angles of attack of 0 and 2 degrees. These conditions were such that, at 2 degrees incidence unsteady periodic motion was calculated along the aft portion of the airfoil and in its wake. Although no unsteady measurements were made for the NACA 64A010 airfoil at these flow conditions, interpolated steady measurements of lift, drag, and surface static pressures compared favorably with corresponding computed time-averaged lift, drag, and surface static pressures.

  20. 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.

  1. 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.

  2. 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.

  3. Implicit calculations of transonic flows using monotone methods

    NASA Astrophysics Data System (ADS)

    Goorjian, P. M.; van Buskirk, R.

    1981-01-01

    Implicit approximate-factorization algorithms have been developed that use monotone methods for the calculation of steady and unsteady transonic flows governed by the small-disturbance-potential equation. These algorithms use the new Engquist-Osher switch in the type-dependent differencing in place of the standard Murman-Cole switch. The resulting algorithms are more stable; hence, calculations can be done more efficiently. For steady flows, the convergence rate is about 35% faster, and for unsteady flows the allowable time step is about 10 times larger. These improvements are achieved with no increase in computer storage and with only minor modifications in codes that use the Murman-Cole switch. Also an implicit algorithm has been developed for the steady full-potential equation in one-dimension, which uses monotone methods.

  4. 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.

  5. 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.

  6. A solution to water vapor in the National Transonic Facility

    NASA Technical Reports Server (NTRS)

    Gloss, Blair B.; Bruce, Robert A.

    1989-01-01

    As cryogenic wind tunnels are utilized, problems associated with the low temperature environment are being discovered and solved. Recently, water vapor contamination was discovered in the National Transonic Facility, and the source was shown to be the internal insulation which is a closed-cell polyisocyanurate foam. After an extensive study of the absorptivity characteristics of the NTF thermal insulation, the most practical solution to the problem was shown to be the maintaining of a dry environment in the circuit at all times. Utilizing a high aspect ratio transport model, it was shown that the moisture contamination effects on the supercritical wing pressure distributions were within the accuracy of setting test conditions and as such were considered negligible for this model.

  7. Euler calculations of unsteady transonic flow in cascades

    NASA Technical Reports Server (NTRS)

    Bendiksen, Oddvar O.

    1991-01-01

    In the present paper, Euler calculations of unsteady transonic flow in cascades are presented. A finite volume scheme is used to discretize the equations, which are implemented on a blade-fitted deformable mesh. The space-discretized equations are integrated forward in time using a multistage Runge-Kutta scheme. Adaptive dissipation terms of the type proposed by Jameson and Baker are added to capture shocks and to suppress nonphysical oscillations. Phase-shifted boundary conditions are used to reduce the computational domain to a single reference passage. No assumptions of small amplitudes or small flow deflections are made. Thus, the present code makes it possible to carry out aeroelastic calculations for cases where the shock strengths and oscillation amplitudes exceed the inherent limitations of potential flow codes.

  8. 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.

  9. 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.

  10. Multi-grid calculation of transonic potential flows

    NASA Technical Reports Server (NTRS)

    Caughey, D. A.; Shmilovich, A.

    1985-01-01

    The finite-volume method discussed by Jameson and Caughey (1977), and Caughey and Jameson (1979, 1980) has made it possible to calculate the transonic potential flow past any configuration for which a suitable boundary-conforming coordinate grid can be constructed. However, computations for practical three-dimensional problems have remained quite expensive in terms of the required computer time. The reason for this is primarily related to the large number of grid cells necessary for adequate resolution in these complex three-dimensional problems, taking into account the large number of iterations required to achieve even modest convergence on these fine grids. The present chapter provides a description of work directed at removing this latter difficulty by making use of the multigrid method. Attention is given to finite-volume formulation, multigrid iteration, geometrical aspects, and computed results.

  11. A well posed boundary value problem in transonic gas dynamics

    NASA Technical Reports Server (NTRS)

    Sanz, J. M.

    1978-01-01

    A new approach considered by Garabedian and Korn (1976) to solve a problem of airfoil design has led to a transonic boundary value problem. It remains to be shown that this problem is well posed. A description is presented of an investigation in which it is shown that a corresponding problem for the Tricomi equation is well posed. The solution to the boundary value problem is characterized, in a unique way, as a sum of two particular solutions. The Poisson formulas for the unit semicircle for the Euler-Poisson-Darboux equation are considered and reflection laws for solutions of the general Euler-Poisson-Darboux equation are established. It is proved that the considered boundary value problem for the case in which the involved function is periodic and continuous is well posed within the specified class of solutions.

  12. Stereoscopic particle image velocimetry in a transonic turbine stage

    NASA Astrophysics Data System (ADS)

    Lang, H.; Mørck, T.; Woisetschläger, J.

    2002-06-01

    In order to investigate the flow field in axial turbine stages, a continuously operating transonic test turbine facility for high pressure ratios was designed at Graz University of Technology, Austria. This test facility allows optical access to the rotor and to the stator trailing edge. The three-dimensional velocity distribution in the region of stator-rotor interaction was investigated by stereoscopic particle image velocimetry. To obtain the three-dimensional velocity vectors in one plane at the mid-section of the turbine blades, a calibration-based method was used. Light-sheet delivery, seeding and triggering to four pre-defined rotor-stator positions are discussed, and an insight into the rotor-stator interaction is given, including vortex shedding in the stator wake.

  13. 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.

  14. 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.

  15. 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.

  16. 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.

  17. Numerical calculation of the transonic flow past a swept wing

    NASA Technical Reports Server (NTRS)

    Jameson, A.; Caughey, D. A.

    1977-01-01

    A numerical method is presented for analyzing the transonic potential flow past a lifting, swept wing. A finite difference approximation to the full potential equation is solved in a coordinate system which is nearly conformally mapped from the physical space in planes parallel to the symmetry plane, and reduces the wing surface to a portion of one boundary of the computational grid. A coordinate invariant, rotated difference scheme is used, and the difference equations are solved by relaxation. The method is capable of treating wings of arbitrary planform and dihedral, although approximations in treating the tips and vortex sheet make its accuracy suspect for wings of small aspect ratio. Comparisons of calculated results with experimental data are shown for examples of both conventional and supercritical transport wings. Agreement is good for both types, but it was found necessary to account for the displacement effect of the boundary layer for the supercritical wing, presumably because of its greater sensitivity to changes in effective geometry.

  18. 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.

  19. 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.

  20. 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.

  1. Improved finite difference schemes for transonic potential calculations

    NASA Technical Reports Server (NTRS)

    Hafez, M.; Osher, S.; Whitlow, W., Jr.

    1984-01-01

    Engquist and Osher (1980) have introduced a finite difference scheme for solving the transonic small disturbance equation, taking into account cases in which only compression shocks are admitted. Osher et al. (1983) studied a class of schemes for the full potential equation. It is proved that these schemes satisfy a new discrete 'entropy inequality' which rules out expansion shocks. However, the conducted analysis is restricted to steady two-dimensional flows. The present investigation is concerned with the adoption of a heuristic approach. The full potential equation in conservation form is solved with the aid of a modified artificial density method, based on flux biasing. It is shown that, with the current scheme, expansion shocks are not possible.

  2. Transonic Tones and Excess Broadband Noise in Overexpanded Supersonic Jets

    NASA Technical Reports Server (NTRS)

    Zaman, Khairul B. M. Q.

    2009-01-01

    Noise characteristics of convergent-divergent (C-D) nozzles in the overexpanded regime are the focus of this paper. The flow regime is encountered during takeoff and landing of certain airplanes and also with rocket nozzles in launch-pad environment. Experimental results from laboratory-scale single nozzles are discussed. The flow often undergoes a resonance accompanied by emission of tones (referred to as transonic tones). The phenomenon is different from the well-known screech tones. Unlike screech, the frequency increases with increasing supply pressure. There is a staging behavior odd harmonic stages occur at lower pressures while the fundamental occurs in a range of relatively higher pressures. A striking feature is that tripping of the nozzle s internal boundary layer tends to suppress the resonance. However, even in the absence of tones the broadband levels are found to be high. That is, relative to a convergent case and at same pressure ratio, the C-D nozzles are found to be noisier, often by more than 10dB. This excess broadband noise (referred to as EBBN) is further explored. Its characteristics are found to be different from the well-known broadband shockassociated noise ( BBSN ). For example, while the frequency of the BBSN peak varies with observation angle no such variation is noted with EBBN. The mechanisms of the transonic tone and the EBBN are not completely understood yet. They appear to be due to unsteady shock motion inside the nozzle. The shock drives the flow downstream like a vibrating diaphragm, and resonance takes place similarly as with acoustic resonance of a conical section having one end closed and the other end open. When the boundary layer is tripped, apparently a breakdown of azimuthal coherence suppresses the resonance. However, there is still unsteady shock motion albeit with superimposed randomness. Such random motion of the internal shock and its interaction with the separated boundary layer produces the EBBN.

  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. 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.

  5. 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.

  6. Resonance Effects in the NASA Transonic Flutter Cascade Facility

    NASA Technical Reports Server (NTRS)

    Lepicovsky, J.; Capece, V. R.; Ford, C. T.

    2003-01-01

    Investigations of unsteady pressure loadings on the blades of fans operating near the stall flutter boundary are carried out under simulated conditions in the NASA Transonic Flutter Cascade facility (TFC). It has been observed that for inlet Mach numbers of about 0.8, the cascade flowfield exhibits intense low-frequency pressure oscillations. The origins of these oscillations were not clear. It was speculated that this behavior was either caused by instabilities in the blade separated flow zone or that it was a tunnel resonance phenomenon. It has now been determined that the strong low-frequency oscillations, observed in the TFC facility, are not a cascade phenomenon contributing to blade flutter, but that they are solely caused by the tunnel resonance characteristics. Most likely, the self-induced oscillations originate in the system of exit duct resonators. For sure, the self-induced oscillations can be significantly suppressed for a narrow range of inlet Mach numbers by tuning one of the resonators. A considerable amount of flutter simulation data has been acquired in this facility to date, and therefore it is of interest to know how much this tunnel self-induced flow oscillation influences the experimental data at high subsonic Mach numbers since this facility is being used to simulate flutter in transonic fans. In short, can this body of experimental data still be used reliably to verify computer codes for blade flutter and blade life predictions? To answer this question a study on resonance effects in the NASA TFC facility was carried out. The results, based on spectral and ensemble averaging analysis of the cascade data, showed that the interaction between self-induced oscillations and forced blade motion oscillations is very weak and can generally be neglected. The forced motion data acquired with the mistuned tunnel, when strong self-induced oscillations were present, can be used as reliable forced pressure fluctuations provided that they are extracted

  7. 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.

  8. 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.

  9. 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

  10. 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

  11. 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.

  12. Analytical investigation of off-design performance of a transonic turbine

    NASA Technical Reports Server (NTRS)

    Whitney, Warren J; Stewart, Warner L

    1954-01-01

    The off-design performance and a breakdown of the losses of a transonic turbine were determined by an analytical method that was previously developed for turbines of more conservative design. The analytically obtained performance map is compared with the performance map obtained from an experimental investigation of the turbine. The rotor hub conditions of incidence angle, relative Mach number, and reaction calculated from the analytical results are compared with those calculated from experimental data. The loss breakdown obtained for the transonic turbine did not differ substantially from that previously obtained from a turbine of more conservative design, except that large stator-exit shock losses were predicted for the transonic turbine at low speeds. The trends of the rotor hub incidence angle, relative Mach number, and reaction calculated from the analytical results agreed well with those calculated from the experimental data over the performance range. These trends indicate that, compared with a turbine of more conservative design, the transonic turbine operated over a much smaller range of incidence angle, a much wider range of rotor relative Mach number, and at a considerably lower level of reaction. Good over-all agreement was obtained between the analytically predicted performance and the experimental performance, except at 40-percent design speed, where in the analysis the stator reached limiting loading before the rotor choked. Since this discrepancy resulted from errors in the simplifying assumptions used in the analysis, it is regarded as a limitation in the analytical method as applied to a transonic turbine.

  13. 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.

  14. 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.

  15. 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.

  16. 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.

  17. TAS: A Transonic Aircraft/Store flow field prediction code

    NASA Technical Reports Server (NTRS)

    Thompson, D. S.

    1983-01-01

    A numerical procedure has been developed that has the capability to predict the transonic flow field around an aircraft with an arbitrarily located, separated store. The TAS code, the product of a joint General Dynamics/NASA ARC/AFWAL research and development program, will serve as the basis for a comprehensive predictive method for aircraft with arbitrary store loadings. This report described the numerical procedures employed to simulate the flow field around a configuration of this type. The validity of TAS code predictions is established by comparison with existing experimental data. In addition, future areas of development of the code are outlined. A brief description of code utilization is also given in the Appendix. The aircraft/store configuration is simulated using a mesh embedding approach. The computational domain is discretized by three meshes: (1) a planform-oriented wing/body fine mesh, (2) a cylindrical store mesh, and (3) a global Cartesian crude mesh. This embedded mesh scheme enables simulation of stores with fins of arbitrary angular orientation.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  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. Endwall heat transfer measurements in a transonic turbine cascade

    SciTech Connect

    Giel, P.W.; Thurman, D.R.; Van Fossen, G.J.; Hippensteele, S.A.; Boyle, R.J.

    1998-04-01

    Turbine blade endwall heat transfer measurements are presented for a range of Reynolds and Mach numbers. Data were obtained for Reynolds numbers based on inlet conditions of 0.5 and 1.0 {times} 10{sup 6}, for isentropic exit Mach numbers of 1.0 and 1.3, and for free-stream 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 deg 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 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.

  4. Turbulence measurements in a transonic two-passage turbine cascade

    NASA Astrophysics Data System (ADS)

    Vicharelli, Amanda; Eaton, John K.

    2006-06-01

    This paper presents detailed turbulence measurements in a two-dimensional, transonic, double passage turbine cascade. Particle image velocimetry was used to obtain mean velocity and turbulence measurements all around a single turbine blade within about 2 mm of the blade and wall surfaces. The passage walls were designed using an optimization procedure so that the blade surface pressure distribution matches that of the blade in an infinite cascade. The resulting experimental model captures much of the complexity of a real turbine stage (including high streamline curvature, strong accelerations, and shocks) in a passage with a continuous wall shape, allowing for high measurement resolution and well controlled boundary conditions for comparison to CFD. The measurements show that in the inviscid regions of the passage the absolute level of the turbulent fluctuations does not change significantly as the flow accelerates, while the local turbulence intensity drops rapidly as the flow accelerates. These results provide a benchmark data set that can be used to improve turbulence models.

  5. Noise Prediction of NASA SR2 Propeller in Transonic Conditions

    NASA Astrophysics Data System (ADS)

    Gennaro, Michele De; Caridi, Domenico; Nicola, Carlo De

    2010-09-01

    In this paper we propose a numerical approach for noise prediction of high-speed propellers for Turboprop applications. It is based on a RANS approach for aerodynamic simulation coupled with Ffowcs Williams-Hawkings (FW-H) Acoustic Analogy for propeller noise prediction. The test-case geometry adopted for this study is the 8-bladed NASA SR2 transonic cruise propeller, and simulated Sound Pressure Levels (SPL) have been compared with experimental data available from Wind Tunnel and Flight Tests for different microphone locations in a range of Mach numbers between 0.78 and 0.85 and rotational velocities between 7000 and 9000 rpm. Results show the ability of this approach to predict noise to within a few dB of experimental data. Moreover corrections are provided to be applied to acoustic numerical results in order for them to be compared with Wind Tunnel and Flight Test experimental data, as well computational grid requirements and guidelines in order to perform complete aerodynamic and aeroacoustic calculations with highly competitive computational cost.

  6. Stator performance and unsteady loading in transonic compressor stages

    SciTech Connect

    Durali, M.; Kerrebrock, J.L.

    1998-04-01

    The structure and behavior of wakes from a transonic compressor rotor and their effect on the loading and performance of the downstream stator have been investigated experimentally. The rotor was 23.25 inches in diameter with a measured tip Mach number of 1.23 and a pressure ratio of 1.66. Time and space-resolved measurements have been completed of the rotor and stator outflow, as well of the pressure distribution on the surface of the stator blades. It was found that the wakes from this rotor have large flow angle and flow Mach number variations from the mean flow, significant pressure fluctuations, and a large degree of variation from hub to tip. There was a significant total pressure defect and practically no static pressure variation associated with the stator wakes. Wakes from the rotor exist nearly undiminished in the exit flow of the stator and decay in the annular duct behind the stator. The pressure at all points along the chord over each of the stator blades` surfaces fluctuated nearly in phase in response to the rotor wakes, that is the unsteady chordwise pressure distribution is determined mainly by the change in angle of incidence to the blade and not by the local velocity fluctuations within the passage. The unsteady forces on the stator blades, induced by the rotor wakes, were as high as 25% of the steady forces, and lagged the incidence of the wakes on the leading edge by approximately 180 deg at most radii.

  7. 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.

  8. 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.

  9. 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.

  10. Semi-span model testing in the national transonic facility

    NASA Technical Reports Server (NTRS)

    Chokani, Ndaona

    1994-01-01

    The present work was motivated by an ongoing research program at NASA Langley Research Center to develop a semi-span testing capability for the National Transonic Facility (NTF). This test technique is being investigated as a means to design and optimize high-lift devices at flight Reynolds numbers in a ground test facility. Even though the freestream Mach numbers of interest are around .20, the flow around a transport wing with high lift devices deployed may contain regions of compressible flow. Thus to properly model the flow physics, a compressible flow solver may be required. However, the application of a compressible flow solver at low Mach numbers can be problematic. The objective of this phase of the project is to directly compare the performance of two widely used three-dimensional compressible Navier-Stokes solvers at low Mach numbers to both experimental data and to results obtained from an incompressible Navier-Stokes solver. The geometries of interest are two isolated wings with different leading edge sweep angles. The compressible Navier-Stokes solvers chosen, TLNS3D-MB and CFL3D, which were developed at NASA Langley Research Center (LaRC), represent the current state-of-the-art in compressible 3-D Navier-Stokes solvers. The incompressible Navier-Stokes solver, INS3D-UP, developed recently at NASA Ames Research Center (ARC), represents the current state-of-the-art in incompressible Navier-Stokes solvers.

  11. 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.

  12. Calculation of unsteady transonic flows with mild separation by viscous-inviscid interaction

    NASA Technical Reports Server (NTRS)

    Howlett, James T.

    1992-01-01

    This paper presents a method for calculating viscous effects in two- and three-dimensional unsteady transonic flow fields. An integral boundary-layer method for turbulent viscous flow is coupled with the transonic small-disturbance potential equation in a quasi-steady manner. The viscous effects are modeled with Green's lag-entrainment equations for attached flow and an inverse boundary-layer method for flows that involve mild separation. The boundary-layer method is used stripwise to approximate three-dimensional effects. Applications are given for two-dimensional airfoils, aileron buzz, and a wing planform. Comparisons with inviscid calculations, other viscous calculation methods, and experimental data are presented. The results demonstrate that the present technique can economically and accurately calculate unsteady transonic flow fields that have viscous-inviscid interactions with mild flow separation.

  13. Evaluation of 3 numerical methods for propulsion integration studies on transonic transport configurations

    NASA Technical Reports Server (NTRS)

    Yaros, S. F.; Carlson, J. R.; Chandrasekaran, B.

    1986-01-01

    An effort has been undertaken at the NASA Langley Research Center to assess the capabilities of available computational methods for use in propulsion integration design studies of transonic transport aircraft, particularly of pylon/nacelle combinations which exhibit essentially no interference drag. The three computer codes selected represent state-of-the-art computational methods for analyzing complex configurations at subsonic and transonic flight conditions. These are: EULER, a finitie volume solution of the Euler equation; VSAERO, a panel solution of the Laplace equation; and PPW, a finite difference solution of the small disturbance transonic equations. In general, all three codes have certain capabilities that allow them to be of some value in predicting the flows about transport configurations, but all have limitations. Until more accurate methods are available, careful application and interpretation of the results of these codes are needed.

  14. 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.

  15. The design of supercritical wings by the use of three-dimensional transonic theory

    NASA Technical Reports Server (NTRS)

    Mann, M. J.

    1979-01-01

    A procedure was developed for the design of transonic wings by the iterative use of three dimensional, inviscid, transonic analysis methods. The procedure was based on simple principles of supersonic flow and provided the designer with a set of guidelines for the systematic alteration of wing profile shapes to achieve some desired pressure distribution. The method was generally applicable to wing design at conditions involving a large region of supercriterical flow. To illustrate the method, it was applied to the design of a wing for a supercritical maneuvering fighter that operates at high lift and transonic Mach number. The wing profiles were altered to produce a large region of supercritical flow which was terminated by a weak shock wave. The spanwise variation of drag of this wing and some principles for selecting the streamwise pressure distribution are also discussed.

  16. 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.

  17. Some recent progress in transonic flow computation. [flow distribution, numerical optimization, and airfoil design

    NASA Technical Reports Server (NTRS)

    Ballhaus, W. F.

    1976-01-01

    Although the development of a finite difference relaxation procedure to solve the steady form of equations of motion gave birth to the study of computational transonic aerodynamics and considerable progress has been made using the small disturbance theory, no general analytical solution method yet exists for transonic flows that include three dimensional unsteady, and viscous effects. Two techniques are described which are useful in computational transonic aerodynamics applications. The finite volume method simplifies the application of boundary conditions without introducing the constriction associated with small disturbance theory. Governing equations are solved in a Cartesian coordinate system using a body-oriented and shock-oriented mesh network. Only the volume and surface normal directions of the volume elements must be known. The other method, configuration design by numerical optimization, can be used by aircraft designers to develop configurations that satisfy specific geometric performance constraints. Two examples of airfoil design by numerical optimization are presented.

  18. Perturbation solutions for transonic flow on the blade-to-blade surface of compressor blade rows

    NASA Technical Reports Server (NTRS)

    Stahara, S. S.; Chaussee, D. S.; Spreiter, J. R.

    1978-01-01

    A preliminary investigation was conducted to establish the theoretical basis of perturbation techniques with the objective of minimizing computational requirements associated with parametric studies of transonic flows in turbomachines. The theoretical analysis involved the development of perturbation methods for determining first order changes in the flow solution due to variations of one or more geometrical or flow parameters. The formulation is primarily directed toward transonic flows on the blade to blade surface of a single blade row compressor. Two different perturbation approaches were identified and studied. Applications and results of these methods for various perturbations are presented for selected two dimensional transonic cascade flows to illustrate the advantages and disadvantages of each technique. Additionally, it was found that, for flows with shock waves, proper account of shock displacement was crucial.

  19. 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.

  20. 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.

  1. A computational transonic flutter boundary tracking procedure. M.S. Thesis

    NASA Technical Reports Server (NTRS)

    Gallman, J. W.; Batina, J. T.; Yang, T. Y.

    1986-01-01

    An automated flutter boundary tracking procedure for the efficient calculation of transonic flutter boundaries is presented. The procedure uses aeroelastic responses to march along the boundary by taking steps in speed and Mach number, thereby reducing the number of response calculations previously required to determine a transonic flutter boundary. Flutter boundary results are presented for a typical airfoil section oscillating with pitch and plunge degrees of freedom. These transonic flutter boundaries are in good agreement with exact boundaries calculated using the conventional time-marching method. The tracking procedure is extended to include static aeroelastic twist as a simulation of the static deformation of a wing and contains all of the essential features that are required to apply it to practical three-dimensional cases. The procedure is also applied to flutter boundaries as a function of structural parameters.

  2. 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.

  3. Transonic airfoil computation using the integral equation with and without embedded Euler domains

    NASA Technical Reports Server (NTRS)

    Kandil, Osama A.; Hu, Hong

    1987-01-01

    Two transonic computational schemes which are based on the Integral Equation Formulation of the full potential equation were presented. The first scheme is a Shock Capturing-Shock Fitting (SCSF) scheme which uses the full potential equation throughout with the exception of the shock wave where the Rankine-Hugoniot relations are used to cross and fit the shock. The second scheme is an Integral Equation with Embedded Euler (IEEE) scheme which uses the full potential equation with an embedded region where the Euler equations are used. The two schemes are applied to several transonic airfoil flows and the results were compared with numerous computational results and experimental domains with fine grids. The SCSF-scheme is restricted to flows with weak shock, while the IEEE-scheme can handle strong shocks. Currently, the IEEE scheme is applied to other transonic flows with strong shocks as well as to unsteady pitching oscillations.

  4. 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.

  5. Time-marching transonic flutter solutions including angle-of-attack effects

    NASA Technical Reports Server (NTRS)

    Edwards, J. W.; Bennett, R. M.; Whitlow, W., Jr.; Seidel, D. A.

    1982-01-01

    Transonic aeroelastic solutions based upon the transonic small perturbation potential equation were studied. Time-marching transient solutions of plunging and pitching airfoils were analyzed using a complex exponential modal identification technique, and seven alternative integration techniques for the structural equations were evaluated. The HYTRAN2 code was used to determine transonic flutter boundaries versus Mach number and angle-of-attack for NACA 64A010 and MBB A-3 airfoils. In the code, a monotone differencing method, which eliminates leading edge expansion shocks, is used to solve the potential equation. When the effect of static pitching moment upon the angle-of-attack is included, the MBB A-3 airfoil can have multiple flutter speeds at a given Mach number.

  6. Calculation of viscous effects on transonic flow for oscillating airfoils and comparisons with experiment

    NASA Technical Reports Server (NTRS)

    Howlett, James T.; Bland, Samuel R.

    1987-01-01

    A method is described for calculating unsteady transonic flow with viscous interaction by coupling a steady integral boundary-layer code with an unsteady, transonic, inviscid small-disturbance computer code in a quasi-steady fashion. Explicit coupling of the equations together with viscous -inviscid iterations at each time step yield converged solutions with computer times about double those required to obtain inviscid solutions. The accuracy and range of applicability of the method are investigated by applying it to four AGARD standard airfoils. The first-harmonic components of both the unsteady pressure distributions and the lift and moment coefficients have been calculated. Comparisons with inviscid calcualtions and experimental data are presented. The results demonstrate that accurate solutions for transonic flows with viscous effects can be obtained for flows involving moderate-strength shock waves.

  7. Design features and operational characteristics of the Langley 0.3-meter transonic cryogenic tunnel

    NASA Technical Reports Server (NTRS)

    Kilgore, R. A.

    1976-01-01

    Experience with the Langley 0.3 meter transonic cryogenic tunnel, which is fan driven, indicated that such a tunnel presents no unusual design difficulties and is simple to operate. Purging, cooldown, and warmup times were acceptable and were predicted with good accuracy. Cooling with liquid nitrogen was practical over a wide range of operating conditions at power levels required for transonic testing, and good temperature distributions were obtained by using a simple liquid nitrogen injection system. To take full advantage of the unique Reynolds number capabilities of the 0.3 meter transonic tunnel, it was designed to accommodate test sections other than the original, octagonal, three dimensional test section. A 20- by 60-cm two dimensional test section was recently installed and is being calibrated. A two dimensional test section with self-streamlining walls and a test section incorporating a magnetic suspension and balance system are being considered.

  8. 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.

  9. Reynolds number effects on the transonic aerodynamics of a slender wing-body configuration

    NASA Technical Reports Server (NTRS)

    Luckring, James M.; Fox, Charles H., Jr.; Cundiff, Jeffrey S.

    1989-01-01

    Aerodynamic forces and moments for a slender wing-body configuration are summarized from an investigation in the Langley National Transonic Facility (NTF). The results include both longitudinal and lateral-directional aerodynamic properties as well as slideslip derivatives. Results were selected to emphasize Reynolds number effects at a transonic speed although some lower speed results are also presented for context. The data indicate nominal Reynolds number effects on the longitudinal aerodynamic coefficients and more pronounced effects for the lateral-directional aerodynamic coefficients. The Reynolds number sensitivities for the lateral-directional coefficients were limited to high angles of attack.

  10. A Green's function formulation for a nonlinear potential flow solution applicable to transonic flow

    NASA Technical Reports Server (NTRS)

    Baker, A. J.; Fox, C. H., Jr.

    1977-01-01

    Routine determination of inviscid subsonic flow fields about wing-body-tail configurations employing a Green's function approach for numerical solution of the perturbation velocity potential equation is successfully extended into the high subsonic subcritical flow regime and into the shock-free supersonic flow regime. A modified Green's function formulation, valid throughout a range of Mach numbers including transonic, that takes an explicit accounting of the intrinsic nonlinearity in the parent governing partial differential equations is developed. Some considerations pertinent to flow field predictions in the transonic flow regime are discussed.

  11. 3D CFD modeling of subsonic and transonic flowing-gas DPALs with different pumping geometries

    NASA Astrophysics Data System (ADS)

    Yacoby, Eyal; Sadot, Oren; Barmashenko, Boris D.; Rosenwaks, Salman

    2015-10-01

    Three-dimensional computational fluid dynamics (3D CFD) modeling of subsonic (Mach number M ~ 0.2) and transonic (M ~ 0.9) diode pumped alkali lasers (DPALs), taking into account fluid dynamics and kinetic processes in the lasing medium is reported. The performance of these lasers is compared with that of supersonic (M ~ 2.7 for Cs and M ~ 2.4 for K) DPALs. The motivation for this study stems from the fact that subsonic and transonic DPALs require much simpler hardware than supersonic ones where supersonic nozzle, diffuser and high power mechanical pump (due to a drop in the gas total pressure in the nozzle) are required for continuous closed cycle operation. For Cs DPALs with 5 x 5 cm2 flow cross section pumped by large cross section (5 x 2 cm2) beam the maximum achievable power of supersonic devices is higher than that of the transonic and subsonic devices by only ~ 3% and ~ 10%, respectively. Thus in this case the supersonic operation mode has no substantial advantage over the transonic one. The main processes limiting the power of Cs supersonic DPALs are saturation of the D2 transition and large ~ 60% losses of alkali atoms due to ionization, whereas the influence of gas heating is negligible. For K transonic DPALs both the gas heating and ionization effects are shown to be unimportant. The maximum values of the power are higher than those in Cs transonic laser by ~ 11%. The power achieved in the supersonic and transonic K DPAL is higher than for the subsonic version, with the same resonator and K density at the inlet, by ~ 84% and ~ 27%, respectively, showing a considerable advantaged of the supersonic device over the transonic one. For pumping by rectangular beams of the same (5 x 2 cm2) cross section, comparison between end-pumping - where the laser beam and pump beam both propagate at along the same axis, and transverse-pumping - where they propagate perpendicularly to each other, shows that the output power and optical-to-optical efficiency are not

  12. 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.

  13. 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.

  14. Effects of Unsteady Flow Interactions on the Performance of a Highly-Loaded Transonic Compressor Stage

    NASA Technical Reports Server (NTRS)

    Hah, Chunill

    2015-01-01

    Effects of unsteady flow interactions on the aerodynamic performance of a highly-loaded transonic axial compressor are investigated in the present study. The primary focus of the study is to investigate how unsteady flow interactions between blade rows affect the aerodynamic performance of a highly-loaded transonic axial compressor. Recent experimental and numerical studies of current highly-loaded axial compressor performance indicated that predicting calculating the loss generation is very challenging with various analysis tools. In the present study, the effects of generation and transport of shock induced vortices on the compressor performance is investigated in detail.

  15. Finite-difference simulation of transonic separated flow using a full potential boundary layer interaction approach

    NASA Technical Reports Server (NTRS)

    Van Dalsem, W. R.; Steger, J. L.

    1983-01-01

    A new, fast, direct-inverse, finite-difference boundary-layer code has been developed and coupled with a full-potential transonic airfoil analysis code via new inviscid-viscous interaction algorithms. The resulting code has been used to calculate transonic separated flows. The results are in good agreement with Navier-Stokes calculations and experimental data. Solutions are obtained in considerably less computer time than Navier-Stokes solutions of equal resolution. Because efficient inviscid and viscous algorithms are used, it is expected this code will also compare favorably with other codes of its type as they become available.

  16. Transonic calculations for a flexible supercritical wing and comparison with experiment

    NASA Technical Reports Server (NTRS)

    Bennett, R. M.; Seidel, D. A.; Sandford, M. C.

    1985-01-01

    Pressure data measured on the flexible DAST ARW-2 wing are compared with results calculated using the transonic small perturbation code XTRAN3S. A brief description of the analysis is given and a recently-developed grid coordinate transformation is described. Calculations are presented for the rigid and flexible wing for Mach numbers from 0.60 to 0.90 and dynamic pressures from 0 to 1000 psf. Calculated and measured static pressures and wing deflections are compared, and calculated static aeroelastic trends are given. Attempts to calculate the transonic instability boundary of the wing are described.

  17. Compilation of Information on the Transonic Attachment of Flows at the Leading Edges of Airfoils

    NASA Technical Reports Server (NTRS)

    Lindsey, Walter F; Landrum, Emma Jean

    1958-01-01

    Schlieren photographs have been compiled of the two-dimensional flow at transonic speeds past 37 airfoils. These airfoils have variously shaped profiles, and some are related in thickness and camber. The data for these airfoils were analyzed to provide basic information on the flow changes involved and to determine factors affecting transonic-flow attachment, which is a transition from separated to unseparated flow at the leading edges of two-dimensional airfoils at fixed angles as the subsonic Mach number is increased.

  18. Measurements of flow quality in the Ames 2 x 2ft transonic wind tunnel

    NASA Technical Reports Server (NTRS)

    Owen, F. K.

    1981-01-01

    For decades, wind tunnel testing has been conducted in test section environments which have not been adequately documented. However, with the advent of the energy shortage, the need for improved fuel-efficient transports employing supercritical or LFC airfoils has increased the awareness of the possible influence of freestream turbulence on advanced experimental testing. This has already lead to detailed flow quality measurements in NASA transonic wind tunnels. The purpose of this paper is to present results of a study in the Ames 2 x 2 ft transonic wind tunnel.

  19. An efficient coordinate transformation technique for unsteady, transonic aerodynamic analysis of low aspect-ratio wings

    NASA Technical Reports Server (NTRS)

    Guruswamy, G. P.; Goorjian, P. M.

    1984-01-01

    An efficient coordinate transformation technique is presented for constructing grids for unsteady, transonic aerodynamic computations for delta-type wings. The original shearing transformation yielded computations that were numerically unstable and this paper discusses the sources of those instabilities. The new shearing transformation yields computations that are stable, fast, and accurate. Comparisons of those two methods are shown for the flow over the F5 wing that demonstrate the new stability. Also, comparisons are made with experimental data that demonstrate the accuracy of the new method. The computations were made by using a time-accurate, finite-difference, alternating-direction-implicit (ADI) algorithm for the transonic small-disturbance potential equation.

  20. 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

  1. Subsonic and Transonic Dynamic Stability Characteristics of the X-33

    NASA Technical Reports Server (NTRS)

    Tomek, D.; Boyden, R.

    2000-01-01

    Dynamic stability testing was conducted on a 2.5% scale model of the X-33 technology demonstrator sub-orbital flight-test vehicle. This testing was conducted at the NASA Langley Research Center (LaRC) l6-Foot Transonic Wind Tunnel with the LaRC High-speed Dynamic Stability system. Forced oscillation data were acquired for various configurations over a Mach number range of 0.3 to 1.15 measuring pitch, roll and yaw damping, as well as the normal force due to pitch rate and the cross derivatives. The test angle of attack range was from -2 to 24 degrees, except for those cases where load constraints limited the higher angles of attack at the higher Mach numbers. A variety of model configurations with and without control surfaces were employed, including a body alone configuration. Stable pitch damping is exhibited for the baseline configuration throughout the angle of attack range for Mach numbers 0.3, 0.8, and 1.15. Stable pitch damping is present for Mach numbers 0.9 and 0.6 with the exception of angles 2 and 16 degrees, respectively. Constant and stable roll damping were present for the baseline configuration over the range of Mach numbers up to an angle of attack of 16 degrees. The yaw damping for the baseline is somewhat stable and constant for the angle of attack range from -2 to 8 degrees, with the exception of Mach numbers 0.6 and 0.8. Yaw damping becomes highly unstable for all Mach numbers at angles of attack greater than 8 degrees.

  2. 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

  3. Aerodynamic development and investigation of turbine transonic rotor blade cascades

    NASA Astrophysics Data System (ADS)

    Mayorskiy, E. V.; Mamaev, B. I.

    2015-05-01

    An intricate nature of the pattern in which working fluid flows over transonic blade cascades generates the need for experimentally studying their characteristics in designing them. Three cascades having identical main geometrical parameters and differing from one another only in the suction side curvature in the outlet area between the throat and trailing edge were tested in optimizing the rotor blade cascade for the reduced flow outlet velocity λ2 ≈ 1. In initial cascade 1, its curvature decreased monotonically toward the trailing edge. In cascade 2, the suction side curvature near the trailing edge was decreased, but the section near the throat had a larger curvature. In cascade 3, a profile with inverse concavity near the trailing edge was used. The cascades were blown at λ2 = 0.7-1.2 and at different incidence angles. The distribution of pressure over the profiles, profile losses, and the outlet angle were measured. Cascade 1 showed efficient performance in the design mode and under the conditions of noticeable deviations from it with respect to the values of λ2 and incidence angle. In cascade 2, flow separation zones were observed at the trailing edge, as well as an increased level of losses. Cascade 3 was found to be the best one: it had reduced positive pressure gradients as compared with cascade 1, and the relative reduction of losses in the design mode was equal to 24%. The profiles with inverse concavity on the suction side near the trailing edge were recommended for being used in heavily loaded turbine stages.

  4. 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.

  5. 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.

  6. Unsteady hybrid vortex technique for transonic vortex flows and flutter application

    NASA Technical Reports Server (NTRS)

    Kandil, Osama A.

    1987-01-01

    Papers resulting from work performed from January 1, 1987 to July 31, 1987 are listed. Transonic computational schemes based on Integral Equation Formulation of the full potential equation were presented. Classical and zero-total pressure-loss sets of Euler equations applied to delta wings were examined.

  7. Three-Dimensional Transonic Flow Theory Applied to Slender Wings and Bodies

    NASA Technical Reports Server (NTRS)

    Heaslet, Max A; Spreiter, John R

    1957-01-01

    The present paper re-examines the derivation of the integral equations for transonic flow around slender wings and bodies of revolution, giving special attention to conditions resulting from the presence of shock waves and to the reduction of the relations to the special forms necessary for the discussion of sonic flow, that is, flow at free-stream Mach number 1.

  8. A finite element method for the computation of transonic flow past airfoils

    NASA Technical Reports Server (NTRS)

    Eberle, A.

    1980-01-01

    A finite element method for the computation of the transonic flow with shocks past airfoils is presented using the artificial viscosity concept for the local supersonic regime. Generally, the classic element types do not meet the accuracy requirements of advanced numerical aerodynamics requiring special attention to the choice of an appropriate element. A series of computed pressure distributions exhibits the usefulness of the method.

  9. TRANDESNF: A computer program for transonic airfoil design and analysis in nonuniform flow

    NASA Technical Reports Server (NTRS)

    Chang, J. F.; Lan, C. Edward

    1987-01-01

    The use of a transonic airfoil code for analysis, inverse design, and direct optimization of an airfoil immersed in propfan slipstream is described. A summary of the theoretical method, program capabilities, input format, output variables, and program execution are described. Input data of sample test cases and the corresponding output are given.

  10. Design study of test models of maneuvering aircraft configurations for the National Transonic Facility (NTF)

    NASA Technical Reports Server (NTRS)

    Griffin, S. A.; Madsen, A. P.; Mcclain, A. A.

    1984-01-01

    The feasibility of designing advanced technology, highly maneuverable, fighter aircraft models to achieve full scale Reynolds number in the National Transonic Facility (NTF) is examined. Each of the selected configurations are tested for aeroelastic effects through the use of force and pressure data. A review of materials and material processes is also included.

  11. Nonequilibrium turbulence modeling effects on transonic vortical flows about delta wings

    NASA Technical Reports Server (NTRS)

    Kaynak, Unver; Tu, Eugene; Dindar, Mustafa; Barlas, Remzi

    1991-01-01

    The Johnson-King turbulence model that is a viable method for calculating two dimensional transonic separated flows was extended into three dimensions. The implementation was done for Navier Stokes flow solvers written in general curvilinear coordinates. The present approach used in turbulence modeling is based on streamwise integration of an ordinary differential equation (o.d.e.) that governs the maximum Reynolds shear stress behavior. Streamwise integration of the o.d.e. approach was found to offer great mathematical simplicity and economy for three dimensional Navier Stokes methods. Thus, the new method is quick, simple, and very cheap. The new method was first checked against the data of a well known transonic axisymmetric bump experiment, and a good agreement was obtained. Later, the new method was used to compute the flow around a low aspect ratio wing in a transonic wind tunnel. Finally it was employed to study the nonequilibrium turbulence effects on the transonic vortical flows about a 65 deg sweep round leading edge delta wing.

  12. Uncertainty Quantification of Turbulence Model Closure Coefficients for Transonic Wall-Bounded Flows

    NASA Technical Reports Server (NTRS)

    Schaefer, John; West, Thomas; Hosder, Serhat; Rumsey, Christopher; Carlson, Jan-Renee; Kleb, William

    2015-01-01

    The goal of this work was to quantify the uncertainty and sensitivity of commonly used turbulence models in Reynolds-Averaged Navier-Stokes codes due to uncertainty in the values of closure coefficients for transonic, wall-bounded flows and to rank the contribution of each coefficient to uncertainty in various output flow quantities of interest. Specifically, uncertainty quantification of turbulence model closure coefficients was performed for transonic flow over an axisymmetric bump at zero degrees angle of attack and the RAE 2822 transonic airfoil at a lift coefficient of 0.744. Three turbulence models were considered: the Spalart-Allmaras Model, Wilcox (2006) k-w Model, and the Menter Shear-Stress Trans- port Model. The FUN3D code developed by NASA Langley Research Center was used as the flow solver. The uncertainty quantification analysis employed stochastic expansions based on non-intrusive polynomial chaos as an efficient means of uncertainty propagation. Several integrated and point-quantities are considered as uncertain outputs for both CFD problems. All closure coefficients were treated as epistemic uncertain variables represented with intervals. Sobol indices were used to rank the relative contributions of each closure coefficient to the total uncertainty in the output quantities of interest. This study identified a number of closure coefficients for each turbulence model for which more information will reduce the amount of uncertainty in the output significantly for transonic, wall-bounded flows.

  13. Investigation of Northrop F-5A wing buffet intensity in transonic flight

    NASA Technical Reports Server (NTRS)

    Chintsun, H.; Pi, W. S.

    1974-01-01

    A flight test and data processing program utilizing a Northrop F-5A aircraft instrumented to acquire buffet pressures and response data during transonic maneuvers is discussed. The data are presented in real-time format followed by spectral and statistical analyses. Also covered is a comparison of the aircraft response data with computed responses based on the measured buffet pressures.

  14. A Summary of Flight-Determined Transonic Lift and Drag Characteristics of Several Research Airplane Configurations

    NASA Technical Reports Server (NTRS)

    Bellman, Donald R.

    1959-01-01

    Flight-determined lift and drag data from transonic flights of seven research airplane configurations of widely varying characteristics are presented and compared with wind-tunnel and rocket-model data. The airplanes are the X-5 (590 wing sweep), XF-92A, YF-102 with cambered wing, YF-102 with symmetrical wing, D-558-ii, X-3, and X-LE. The effects of some of the basic configuration differences on the lift and drag characteristics are demonstrated. As indicated by transonic similarity laws, most of the configurations demonstrate a relationship between the transonic increase in zero-lift drag and the maximum cross-sectional area. No such relationship was found between the drag-rise Mach number and its normally related parameters. A comparison of flight and wind-tunnel data shows a generally reasonable agreement, but Reynolds number differences can cause considerable variations in the drag levels of the flight and wind-tunnel tests. Maximum lift-drag ratios vary widely in the subsonic region as would be expected from differences in aspect ratio and wing thickness ratio; however, the variations diminish as the Mach number is increased through the transonic region. The attainment of maximum lift-drag ratio in level flight by several of the airplanes was limited by engine performance, stability characteristics, and buffet boundaries.

  15. Numerical calculation of the transonic potential flow past a cranked wing

    NASA Technical Reports Server (NTRS)

    Chang, I. C.; Tauber, M.

    1983-01-01

    The widely transonic swept-wing code, FL022, was found to have an error in the transformed flow equation in the computational domain. The revised version of the code correctly accounted for the non-straight leading edge geometry and its effect on the pressure distribution.

  16. Breakdown of the conservative potential equation. [for study of transonic flow

    NASA Technical Reports Server (NTRS)

    Salas, M. D.; Gumbert, C. R.

    1985-01-01

    The conservative full-potential equation is used to study transonic flow over five airfoil sections. The results of the study indicate that once shock waves are present in the flow, the qualitative approximation is different from that observed with the Euler equations. The difference in behavior of the potential eventually leads to multiple solutions.

  17. Experimental investigation of a transonic potential flow around a symmetric airfoil

    NASA Technical Reports Server (NTRS)

    Hiller, W. J.; Meier, G. E. A.

    1981-01-01

    Experimental flow investigations on smooth airfoils were done using numerical solutions for transonic airfoil streaming with shockless supersonic range. The experimental flow reproduced essential sections of the theoretically computed frictionless solution. Agreement is better in the expansion part of the of the flow than in the compression part. The flow was nearly stationary in the entire velocity range investigated.

  18. Aftbody Closure Effects on the Reference H Configuration at Subsonic and Transonic Speeds

    NASA Technical Reports Server (NTRS)

    Wahls, Richard A.; Owens, Lewis R., Jr.; Londenberg, W. Kelly

    1999-01-01

    Experience with afterbody closure effects and accompanying test techniques issues on a High Speed Civil Transport (HSCT)-class configuration is described. An experimental data base has been developed which includes force, moment, and surface pressure data for the High Speed Research (HSR) Reference H configuration with a closed afterbody at subsonic and transonic speeds, and with a cylindrical afterbody at transonic and supersonic speeds. A supporting computational study has been performed using the USM3D unstructured Euler solver for the purposes of computational fluid dynamics (CFD) method assessment and model support system interference assessment with a focus on lower blade mount effects on longitudinal data at transonic speeds. Test technique issues related to a lower blade sting mount strategy are described based on experience in the National Transonic Facility (NTF). The assessment and application of the USM3D code to the afterbody/sting interference problem is discussed. Finally, status and plans to address critical test technique issues and for continuation of the computational study are presented.

  19. Operation of the ISL transonic shock tube in a high subsonic flow regime

    NASA Astrophysics Data System (ADS)

    Seiler, F.; Havermann, M.; Boller, F.; Mangold, P.; Takayama, K.

    The transonic flow regime plays an important role in experimental aerodynamic research. Modern civil aircraft fly up to a Mach number of M ≈ 0.9 in the high subsonic speed regime, as, for example, the Boeing or Airbus passenger aircraft. Nearly sonic Mach numbers are foreseen for innovative airplane concepts like the sonic cruiser by Boeing. In the military domain, guided missiles like the cruise missile also fly in the high subsonic flow regime. For testing purposes, transonic wind tunnels are mainly used for sub- as well as supersonic design applications. These wind tunnels have normally very large dimensions, which makes their operation quite expensive. If only small scale tests are required, a cheap working facility turns out to be more beneficial. For this purpose, a conventional shock tube operated at transonic flow conditions has been put into operation at the ISL. In the transonic flow regime, however, the reduction of the tube cross section by the model can produce severe distortions followed by a choking of the shock tube flow in the test section. Extensive experimental investigations were performed to determine the subsonic choking Mach number as a function of the model size. These results are compared with theoretical estimations and, more in detail, with CFD calculations.

  20. An Overview of National Transonic Facility Investigations for High Performance Military Aerodynamics (Invited)

    NASA Technical Reports Server (NTRS)

    Luckring, J. M.

    2001-01-01

    A review of National Transonic Facility (NTF) investigations for high-performance military aerodynamics has been completed. The review spans the entire operational period of the tunnel, and includes configurations ranging from full aircraft to basic research geometries. The intent for this document is to establish a comprehensive summary of these experiments with selected technical results

  1. A new consistent spatial differencing scheme for the transonic full-potential equation

    NASA Technical Reports Server (NTRS)

    Flores, J.; Holst, T. L.; Kwak, D.; Batiste, D. M.

    1983-01-01

    A new spatial differencing scheme for the transonic full-potential equation in conservative form has been developed. This scheme guarantees zero truncation error on any curvilinear mesh for freestream flows in either two- or three-space dimensions. Solutions obtained with this new differencing scheme, away from freestream regions, exhibit greatly improved accuracy, especially for nonsmooth or singular meshes.

  2. Steady and unsteady transonic small disturbance analysis of realistic aircraft configurations

    NASA Technical Reports Server (NTRS)

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

    1989-01-01

    A transonic unsteady aerodynamic and aeroelastic code called CAP-TSD (Computational Aeroelasticity Program - Transonic Small Disturbance) was developed for application to realistic aircraft configurations. It permits the calculation of steady and unsteady flows about complete aircraft configurations for aeroelastic analysis of the flutter critical transonic speed range. The CAP-TSD code uses a time accurate approximate factorization algorithm for solution of the unsteady transonic small disturbance potential equation. An overview is given of the CAP-TSD code development effort along with recent algorithm modifications which are listed and discussed. Calculations are presented for several configurations including the General Dynamics 1/9th scale F-16C aircraft model to evaluate the algorithm and hence the reliability of the CAP-TSD code in general. Calculations are also presented for a flutter analysis of a 45 deg sweptback wing which agree well with the experimental data. Descriptions are presented of the CAP-TSD code and algorithm details along with results and comparisons which demonstrate the stability, accuracy, efficiency, and utility of CAP-TSD.

  3. Numerical computation of transonic flows by finite-element and finite-difference methods

    NASA Technical Reports Server (NTRS)

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

    1978-01-01

    Studies on applications of the finite element approach to transonic flow calculations are reported. Different discretization techniques of the differential equations and boundary conditions are compared. Finite element analogs of Murman's mixed type finite difference operators for small disturbance formulations were constructed and the time dependent approach (using finite differences in time and finite elements in space) was examined.

  4. Hightip-speed, low-loading transonic fan stage. Part 2: Data compilation

    NASA Technical Reports Server (NTRS)

    Ware, T. C.; Kobayashi, R. J.; Jackson, R. J.

    1974-01-01

    Tests were conducted on a high-tip-speed low-loading transonic fan stage to determine the performance and inlet flow distortion tolerance of the design. Test data were recorded for overall and blade element performance with both uniform and distorted inlet flows. A tabular summary of the data and a representative selection of the computer data reduction sheets are presented.

  5. GRUMFOIL: A computer code for the viscous transonic flow over airfoils

    NASA Technical Reports Server (NTRS)

    Mead, H. R.; Melnik, R. E.

    1985-01-01

    A user's manual which describes the operation of the computer program, GRUMFOIL is presented. The program computes the viscous transonic flow over two dimensional airfoils using a boundary layer type viscid-inviscid interaction approach. The inviscid solution is obtained by a multigrid method for the full potential equation. The boundary layer solution is based on integral entrainment methods.

  6. Euler solutions for transonic oscillating cascade flows using dynamic triangular meshes

    SciTech Connect

    Hwang, C.J.; Yang, S.Y.

    1995-07-01

    The modified total-variation-diminishing scheme and an improved dynamic triangular mesh algorithm are presented to investigate the transonic oscillating cascade flows. In a Cartesian coordinate system, the unsteady Euler equations are solved. To validate the accuracy of the present approach, transonic flow around a single NACA 0012 airfoil pitching harmonically about the quarter chord is computed first. The calculated instantaneous pressure coefficient distribution during a cycle of motion compare well with the related numerical and experimental data. To evaluate further the present approach involving nonzero interblade phase angle, the calculations of transonic flow around an oscillating cascade of two unstaggered NACA 0006 blades with interblade phase angle equal to 180 deg are performed. From the instantaneous pressure coefficient distributions and time history of lift coefficient, the present approach, where a simple spatial treatment is utilized on the periodic boundaries, gives satisfactory results. By using this solution procedure, transonic flows around an oscillating cascade of four biconvex blades with different oscillation amplitudes, reduced frequencies, and interblade phase angles are investigated. From the distributions of magnitude and phase angle of the dynamic pressure difference coefficient, the present numerical results show better agreement with the experimental data than those from the linearized theory in most of the cases. For every quarter of one cycle, the pressure contours repeat and proceed one pitch distance in the upward or downward direction for interblade phase angle equal to {minus}90 deg or 90 deg, respectively. The unsteady pressure wave and shock behaviors are observed. From the lift coefficient distributions, it is further confirmed that the oscillation amplitude, interblade phase angle, and reduced frequency all have significant effects on the transonic oscillating cascade flows.

  7. Application of Reduced Order Transonic Aerodynamic Influence Coefficient Matrix for Design Optimization

    NASA Technical Reports Server (NTRS)

    Pak, Chan-gi; Li, Wesley W.

    2009-01-01

    Supporting the Aeronautics Research Mission Directorate guidelines, the National Aeronautics and Space Administration [NASA] Dryden Flight Research Center is developing a multidisciplinary design, analysis, and optimization [MDAO] tool. This tool will leverage existing tools and practices, and allow the easy integration and adoption of new state-of-the-art software. Today s modern aircraft designs in transonic speed are a challenging task due to the computation time required for the unsteady aeroelastic analysis using a Computational Fluid Dynamics [CFD] code. Design approaches in this speed regime are mainly based on the manual trial and error. Because of the time required for unsteady CFD computations in time-domain, this will considerably slow down the whole design process. These analyses are usually performed repeatedly to optimize the final design. As a result, there is considerable motivation to be able to perform aeroelastic calculations more quickly and inexpensively. This paper will describe the development of unsteady transonic aeroelastic design methodology for design optimization using reduced modeling method and unsteady aerodynamic approximation. The method requires the unsteady transonic aerodynamics be represented in the frequency or Laplace domain. Dynamically linear assumption is used for creating Aerodynamic Influence Coefficient [AIC] matrices in transonic speed regime. Unsteady CFD computations are needed for the important columns of an AIC matrix which corresponded to the primary modes for the flutter. Order reduction techniques, such as Guyan reduction and improved reduction system, are used to reduce the size of problem transonic flutter can be found by the classic methods, such as Rational function approximation, p-k, p, root-locus etc. Such a methodology could be incorporated into MDAO tool for design optimization at a reasonable computational cost. The proposed technique is verified using the Aerostructures Test Wing 2 actually designed

  8. 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

  9. 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.

  10. 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.

  11. On the structure, interaction, and breakdown characteristics of slender wing vortices at subsonic, transonic, and supersonic speeds

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.; Schreiner, John A.; Rogers, Lawrence W.

    1989-01-01

    Slender wing vortex flows at subsonic, transonic, and supersonic speeds were investigated in a 6 x 6 ft wind tunnel. Test data obtained include off-body and surface flow visualizations, wing upper surface static pressure distributions, and six-component forces and moments. The results reveal the transition from the low-speed classical vortex regime to the transonic regime, beginning at a freestream Mach number of 0.60, where vortices coexist with shock waves. It is shown that the onset of core breakdown and the progression of core breakdown with the angle of attack were sensitive to the Mach number, and that the shock effects at transonic speeds were reduced by the interaction of the wing and the lead-edge extension (LEX) vortices. The vortex strengths and direct interaction of the wing and LEX cores (cores wrapping around each other) were found to diminish at transonic and supersonic speeds.

  12. Calculation of steady and unsteady pressures on wings at supersonic speeds with a transonic small disturbance code

    NASA Technical Reports Server (NTRS)

    Bennett, Robert M.; Bland, Samuel R.; Batina, John T.; Gibbons, Michael D.; Mabey, Dennis G.

    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 algorithm for solution of the unsteady transonic small-disturbance equation that is efficient for solution of steady and unsteady transonic flow problems including supersonic freestream flows. The new code can treat complete aircraft geometries with multiple lifting surfaces and bodies. Applications to wings in supersonic freestream flow are presented. Comparisons with selected exact solutions from linear theory are presented showing generally favorable results. Calculations for both steady and oscillatory cases for the F-5 and RAE tailplane models are compared with experimental data and also show good overall agreement. Selected steady calculations are further compared with a steady flow Euler code.

  13. Use of the PARC code to estimate the off-design transonic performance of an over/under turboramjet nozzle

    NASA Technical Reports Server (NTRS)

    Lam, David W.

    1995-01-01

    The transonic performance of a dual-throat, single-expansion-ramp nozzle (SERN) was investigated with a PARC computational fluid dynamics (CFD) code, an external flow Navier-Stokes solver. The nozzle configuration was from a conceptual Mach 5 cruise aircraft powered by four air-breathing turboramjets. Initial test cases used the two-dimensional version of PARC in Euler mode to investigate the effect of geometric variation on transonic performance. Additional cases used the two-dimensional version in viscous mode and the three-dimensional version in both Euler and viscous modes. Results of the analysis indicate low nozzle performance and a highly three-dimensional nozzle flow at transonic conditions. In another comparative study using the PARC code, a single-throat SERN configuration for which experimental data were available at transonic conditions was used to validate the results of the over/under turboramjet nozzle.

  14. 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.

  15. Calculation of transonic steady and oscillatory pressures on a low aspect ratio model and comparison with experiment

    NASA Technical Reports Server (NTRS)

    Bennett, R. M.; Wynne, E. C.; Mabey, D. G.

    1985-01-01

    Pressure data measured by the British Royal Aircraft Establishment for the AGARD SMP tailplane are compared with results calculated using the transonic small perturbation code XTRAN3S. A brief description of the analysis is given and a recently developed finite difference grid is described. Results are presented for five steady and nine harmonically oscillating cases near zero angle of attack and for a range of subsonic and transonic Mach numbers.

  16. Three-dimensional, transonic rotor flow field reconstructed from holographic interferogram data

    NASA Technical Reports Server (NTRS)

    Kittleson, J. K.; Yu, Y. H.; Becker, F.

    1985-01-01

    Holographic interferometry and computer-assisted tomography (CAT) are used to determine the transonic flow field of a model rotor blade in hover. A pulsed ruby laser records 40 interferograms with a 61 cm-diam view field near the model rotor-blade tip operating at a tip Mach number of 0.90. After digitizing the interferograms and extracting fringe-order functions, the data are transferred to a CAT code. The CAT code then calculates pressure coefficients in several planes above the blade surface. The values from the holography-CAT method compare favorably with previously obtained numerical computations and laser velocimeter measurements at most locations near the blade tip. The results demonstrate the technique's potential for three-dimensional transonic rotor flow studies.

  17. Reconstruction of a three-dimensional, transonic rotor flow field from holographic interferogram data

    NASA Technical Reports Server (NTRS)

    Kittleson, John K.; Yu, Yung H.

    1987-01-01

    Holographic interferometry and computerized aided tomography (CAT) are used to determine the transonic velocity field of a model rotor blade in hover. A pulsed ruby laser recorded 40 interferograms with a 2 ft dia view field near the model rotor blade tip operating at a tip Mach number of 0.90. After digitizing the interferograms and extracting the fringe order functions, the data are transferred to a CAT code. The CAT code then calculates the perturbation velocity in several planes above the blade surface. The values from the holography-CAT method compare favorably with previously obtained numerical computations in most locations near the blade tip. The results demonstrate the technique's potential for three dimensional transonic rotor flow studies.

  18. Reconstruction of a 3-dimensional transonic rotor flow field from holographic interferogram data

    NASA Technical Reports Server (NTRS)

    Yu, Y. H.; Kittleson, J. K.; Becker, F.

    1985-01-01

    Holographic interferometry and computer-assisted tomography (CAT) are used to determine the transonic velocity field of a model rotor blade in hover. A pulsed ruby laser recorded 40 interferograms with a 2-ft-diam view field near the model rotor-blade tip operating at a tip Mach number of 0.90. After digitizing the interferograms and extracting fringe-order functions, the data are transferred to a CAT code. The CAT code then calculates the perturbation velocity in seeral planes above the blade surface. The values from the holography-CAT method compare favorably with previously obtained numerical computations in most locations near the blade tip. The results demonstrate the technique's potential for three-dimensional transonic rotor flow studies.

  19. 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.

  20. A Computer Program for the Calculation of Three-Dimensional Transonic Nacelle/Inlet Flowfields

    NASA Technical Reports Server (NTRS)

    Vadyak, J.; Atta, E. H.

    1983-01-01

    A highly efficient computer analysis was developed for predicting transonic nacelle/inlet flowfields. This algorithm can compute the three dimensional transonic flowfield about axisymmetric (or asymmetric) nacelle/inlet configurations at zero or nonzero incidence. The flowfield is determined by solving the full-potential equation in conservative form on a body-fitted curvilinear computational mesh. The difference equations are solved using the AF2 approximate factorization scheme. This report presents a discussion of the computational methods used to both generate the body-fitted curvilinear mesh and to obtain the inviscid flow solution. Computed results and correlations with existing methods and experiment are presented. Also presented are discussions on the organization of the grid generation (NGRIDA) computer program and the flow solution (NACELLE) computer program, descriptions of the respective subroutines, definitions of the required input parameters for both algorithms, a brief discussion on interpretation of the output, and sample cases to illustrate application of the analysis.

  1. 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.

  2. Theoretical study of the transonic lift of a double-wedge profile with detached bow wave

    NASA Technical Reports Server (NTRS)

    Vincenti, Walter G; Wagoner, Cleo B

    1954-01-01

    A theoretical study is described of the aerodynamic characteristics at small angle of attack of a thin, double-wedge profile in the range of supersonic flight speed in which the bow wave is detached. The analysis is carried out within the framework of the transonic (nonlinear) small-disturbance theory, and the effects of angle of attack are regarded as a small perturbation on the flow previously calculated at zero angle. The mixed flow about the front half of the profile is calculated by relaxation solution of a suitably defined boundary-value problem for transonic small-disturbance equation in the hodograph plane (i.e., the Tricomi equation). The purely supersonic flow about the rear half is found by an extension of the usual numerical method of characteristics. Analytical results are also obtained, within the framework of the same theory, for the range of speed in which the bow wave is attached and the flow is completely supersonic.

  3. Numerical computation of transonic flow about wing-fuselage configurations on a vector computer

    NASA Technical Reports Server (NTRS)

    Thomas, S. D.; Holst, T. L.

    1983-01-01

    The transonic wing analysis code TWING, which uses the AF2 relaxation algorithm, has been vectorized to run on the Cray-1S computer. Vectorization of this code improved computational efficiency over that of the CDC 7600 computer by factors of 11 to 13. The improvement compares favorably with the prediction of a theoretical performance model. A convenient generalization now permits the treatment of rudimentary wing-fuselage combinations. Flow predictions for a transport configuration in both isolated-wing and wing-fuselage modes show the expected trends in shock strength and position when compared with wind-tunnel results. An isolated fighter wing is examined in terms of execution time on three different computers and in comparison with experimental data. The computational fluid dynamics code produced during this study is a careful union of an efficient three-dimensional, transonic, numerical algorithm and the vector features presently available on modern computers.

  4. 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.

  5. Finite volume computation of unsteady inviscid rotational transonic flows past airfoils in rigid body motion

    NASA Technical Reports Server (NTRS)

    Damodaran, Murali

    1988-01-01

    Unsteady inviscid transonic flow over airfoils in arbitrary rigid body motion is analyzed numerically by solving the two-dimensional unsteady Euler equations in integral form using a finite volume scheme. The solution procedure is based on an explicit Runge-Kutta time-stepping scheme wherein the spatial terms are central-differenced and a combination of second- and fourth-differences in the flow variables are used to form the numerical dissipation terms to stabilize the scheme. Unsteady calculations are started from converged steady-state solutions as initial conditions. Nonreflective boundary conditions are imposed on the far-field boundaries. Results are presented and, where possible, validated against available numerical and experimental data for airfoils subjected to a step change in angle of attack, airfoils oscillating and plunging in transonic flow, and airfoils immersed in a time-varying free stream.

  6. 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.

  7. 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.

  8. Unsteady transonic flow simulation on a full-span-wing-body configuration

    NASA Technical Reports Server (NTRS)

    Guruswamy, Guru P.; Goorjian, Peter M.

    1987-01-01

    The presence of a body influences both the aerodynamic and aeroelastic performance of wings. Such effects are more pronounced in the transonic regime. To accurately account for the effect of the body, particularly when the wings are experiencing asymmetric modal motions, it is necessary to model the full configuration in the nonlinear transonic regime. In this study, full-span-wing-body configurations are simulated for the first time by a theoretical method that uses the unsteady potential equations based on the small-disturbance theory. The body geometry is modeled exactly as the physical shape, instead of as a rectangular box, which has been done in the past. Steady pressure computations for wing-body configurations compare well with the available experimental data. Unsteady pressure computations when the wings are oscillating in asymmetric modes show significant influence of the body.

  9. Improved method for calculating transonic velocities on blade-to-blade stream surfaces of a turbomachine

    NASA Technical Reports Server (NTRS)

    Wood, J. R.

    1981-01-01

    A method was developed to improve the accuracy of an existing computer program used to calculate transonic velocities on a blade-to-blade surface of a turbomachine. The method eliminates problems encountered in obtaining solutions with the velocity gradient equation when large gradients in velocity occur through the blade row. With the improved method, results indicate that the transonic solution can be obtained by scaling the velocities obtained at the reduced mass flow rate where all velocities are subsonic thereby eliminating the need for a solution of the velocity gradient equation. Solutions obtained with the scaling method on a two dimensional compressor cascade and an axial turbine stator show good agreement with experimental data. The results obtained for the stationary blade rows and comparison of analytical results obtained with and without the present method suggest that the method will yield an improved solution for centrifugal compressor impellers.

  10. Numerical simulation on the dynamical stall process of airfoils in transonic flow

    SciTech Connect

    Guo, G.L.; Yang, Y.N.; Ye, Z.Y.

    1994-12-31

    In this paper, a method is presented to simulate the dynamic stall process of airfoils in transonic flow. The flowfield around the oscillating airfoil is analyzed by solving the two-dimensional time averaged compressible Navier-Stokes equations with the Baldwin-Lomax turbulence model. An implicit Lower-Upper-factorized algorithm is constructed in a body-fitted coordinate system. In the algorithm, an improved NND (Non-oscillatory, Non-free--parameter, Dissipative) scheme which is a kind of the TVD (Total Variation Diminishing) schemes is adopted. The computation grid is generated by an algebraic method. To save computation time, the grid is rigidly attached to the airfoil. The dynamic stall process is well simulated on NACA 0012 airfoil oscillating in pitch at a high incidence angle in a transonic flow. The pressure distributions and pressure contours at different moment are given to show the movement of shock wave and the change of pressure distributions.

  11. 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.

    1977-01-01

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

  12. 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.

  13. Complex Flow Separation Pattern on Transonic Fan Airfoils Revealed by Flow Visualization

    NASA Technical Reports Server (NTRS)

    Lepicovsky, Jan

    2001-01-01

    Modern turbofan engines employ a highly loaded fan stage with transonic or low-supersonic velocities in the blade-tip region. The fan blades are often prone to flutter at off-design conditions. Flutter is a highly undesirable and dangerous self-excited mode of blade oscillations that can result in high-cycle fatigue blade failure. The origins of blade flutter are not fully understood yet. The latest view is that the blade oscillations are triggered by high-frequency changes in the extent of the partially separated area on the airfoil suction side. There is a lack of experimental data describing the separated flow characteristics of modern airfoils for transonic fans.

  14. 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.

  15. 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.

  16. Computational fluid dynamics drag prediction: Results from the Viscous Transonic Airfoil Workshop

    NASA Technical Reports Server (NTRS)

    Holst, Terry L.

    1988-01-01

    Results from the Viscous Transonic Airfoil Workshop are compared with each other and with experimental data. Test cases used include attached and separated transonic flows for the NACA 0012 airfoil. A total of 23 sets of numerical results from 15 different author groups are included. The numerical method used vary widely and include: 16 Navier-Stokes methods, 2 Euler boundary layer methods, and 5 potential boundary layer methods. The results indicate a high degree of sophistication among the numerical methods with generally good agreement between the various computed and experimental results for attached or moderately separated cases. The agreement for cases with larger separation is only fair and suggests additional work is required in this area.

  17. Results from tests on a high work transonic turbine for an energy efficient engine

    NASA Technical Reports Server (NTRS)

    Crow, D. E.; Welna, H.; Singer, I. D.; Vanco, M. R.

    1980-01-01

    The experimental results of the evaluation of two high work, transonic, single-stage turbines investigated under the Energy Efficient Engine (E3) Program are presented. The objective of the E3 program is to provide an advanced technology base for a new generation of fuel-conservative turbofan engines. A single-stage turbine required fewer cooled airfoils, a reduced number of leakage paths and no interstage seals. These advanced energy efficient engines require high engine pressure ratios resulting in high expansion ratio, transonic, turbine designs which must have high aerodynamic efficiency. The goal of the turbine program is to develop a high pressure turbine that is compatible with the overall engine design and has an uncooled efficiency of 90.8 percent.

  18. Effect of Large Bulk Viscosity on Two-Dimensional Transonic Flow

    NASA Astrophysics Data System (ADS)

    Cramer, Mark

    2012-11-01

    We examine steady two-dimensional transonic flows over a thin airfoil or turbine blade. The wing Reynolds number is taken to be large and the fluid is described by the classical Navier-Stokes equations. The bulk viscosity is taken to be large compared to the shear viscosity. We use the Method of Matched Asymptotic Expansions to give the conditions under which the effects of large bulk viscosity are no longer negligible. We show that longitudinal viscous effects must be considered at lowest order when the ratio of bulk to shear viscosity is on the order of the product of the conventional Reynolds number times the two-thirds power of the non-dimensional airfoil thickness. Under these conditions the flow is shown to be frictional, irrotational, and governed by the viscous form of the transonic small disturbance equation. This work was supported by NSF Grant CBET-0625015.

  19. Reduction of Unsteady Forcing in a Vaned, Contra-Rotating Transonic Turbine Configuration

    NASA Technical Reports Server (NTRS)

    Clark, John

    2010-01-01

    HPT blade unsteadiness in the presence of a downstream vane consistent with contra-rotation is characterized by strong interaction at the first harmonic of downstream vane passing. E An existing stage-and-one-half transonic turbine rig design was used as a baseline to investigate means of reducing such a blade-vane interaction. E Methods assessed included: Aerodynamic shaping of HPT blades 3D stacking of the downstream vane Steady pressure-side blowing E Of the methods assessed, a combination of vane bowing and steady pressure-side blowing produced the most favorable result. E Transonic turbine experiments are planned to assess predictive accuracy for the baseline turbine and any design improvements.

  20. Effect of Stator and Rotor Aspect Ratio on Transonic-Turbine Performance

    NASA Technical Reports Server (NTRS)

    Wong, Robert Y.; Monroe, Daniel E.

    1959-01-01

    The effect of stator and rotor aspect ratio on transonic-turbine performance was experimentally investigated. The stator aspect ratios covered were 1.6. 0.8, and 0.4, while the rotor aspect ratios investigated were 1.46 and 0.73. It was found that the observed variation in turbine design-point efficiency was negligible. Thus, within the range of aspect ratio investigated, these results verify for turbines operating in the transonic flow range the finding of a reference report, which showed analytically that, if blade shape and solidity are held constant, the aspect ratio may be varied over a wide range without appreciable change in turbine efficiency.

  1. The application of CFD for military aircraft design at transonic speeds

    NASA Technical Reports Server (NTRS)

    Smith, C. W.; Braymen, W. W.; Bhateley, I. C.; Londenberg, W. K.

    1989-01-01

    Numerous computational fluid dynamics (CFD) codes are available that solve any of several variations of the transonic flow equations from small disturbance to full Navier-Stokes. The design philosophy at General Dynamics Fort Worth Division involves use of all these levels of codes, depending on the stage of configuration development. Throughout this process, drag calculation is a central issue. An overview is provided for several transonic codes and representative test-to-theory comparisons for fighter-type configurations are presented. Correlations are shown for lift, drag, pitching moment, and pressure distributions. The future of applied CFD is also discussed, including the important task of code validation. With the progress being made in code development and the continued evolution in computer hardware, the routine application of these codes for increasingly more complex geometries and flow conditions seems apparent.

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

    NASA Technical Reports Server (NTRS)

    Ray, E. J.; Ladson, C. L.; Adcock, J. B.; Lawing, P. L.; Hall, R. M.

    1979-01-01

    The past 6 years of operation with the NASA Langley 0.3 m transonic cryogenic tunnel (TCT) show that there are no insurmountable problems associated with cryogenic testing with gaseous nitrogen at transonic Mach numbers. The fundamentals of the concept were validated both analytically and experimentally and the 0.3 m TCT, with its unique Reynolds number capability, was used for a wide variety of aerodynamic tests. Techniques regarding real-gas effects were developed and cryogenic tunnel conditions can be set and maintained accurately. Cryogenic cooling by injecting liquid nitrogen directly into the tunnel circuit imposes no problems with temperature distribution or dynamic response characteristics. Experience with the 0.3 m TCT, indicates that there is a significant learning process associated with cryogenic, high Reynolds number testing. Many of the questions have already been answered; however, factors such as tunnel control, run logic, economics, instrumentation, and model technology present many new and challenging problems.

  3. Comparison of the full potential and Euler formulations for computing transonic airfoil flows

    NASA Technical Reports Server (NTRS)

    Flores, J.; Barton, J.; Holst, T. L.; Pulliam, T.

    1984-01-01

    A quantitative comparison between the Euler and full potential formulations with respect to speed and accuracy is presented. The robustness of the codes used is tested by a number of transonic airfoil cases. The computed results are from four transonic airfoil computer codes. The full potential codes use fully implicit iteration algorithms. The first Euler code uses a fully implicit ADI iteration scheme. The second Euler code uses an explicit Runge Kutta time stepping algorithm which is enhanced by a multigrid convergence acceleration scheme. Quantitative comparisons are made using various plots of lift coefficient versus the average mesh spacing along the airfoil. Besides yielding an asymptotic limit to the lift coefficient, these results also demonstrate the truncation error behavior of the various codes. Quantitative conclusions regarding the full potential and Euler formulations with respect to accuracy, speed, and robustness can be presented.

  4. Transonic solutions for a multielement airfoil using the full-potential equation

    NASA Technical Reports Server (NTRS)

    Flores, J.; Holst, T. L.; Sorenson, R. L.

    1984-01-01

    Transonic flow solutions are obtained over a multielement airfoil (augmentor-wing) using the full-potential equation. Solutions obtained for a subcritical case and a strong shock case show good quantitative agreement with experiment in regions not dominated by viscous effects. In those regions where viscous effects are dominant, the results are still in good qualitative agreement. For the strong shock case, Mach number and angle-of-attack corrections were necessary to match experimental coefficient of lift. Typical results from the transonic augmentor-wing Potential Code on the Cray-1S computer require about 10 sec of CPU time for a three-order-of-magnitude drop in the maximum residual. The speed with which solutions can be generated, and the associated low cost, will make this code a practical tool for the design aerodynamicist.

  5. 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.

  6. Numerical investigation of tip clearance effects in an axial transonic compressor

    NASA Astrophysics Data System (ADS)

    Ciorciari, R.; Lesser, A.; Blaim, F.; Niehuis, R.

    2012-04-01

    Numerical investigations of the Darmstadt transonic single stage compressor (DTC), in the Rotor1-Stator1 configuration, aimed at advancing the understanding of the effect of different rotor tip gaps and transition modelling on the blade surfaces are presented. Steady three dimensional Reynolds Averaged Navier Stokes (RANS) simulations were performed to obtain the flow fields for the different configurations at different operating conditions using the RANS-Solver TRACE. The stage geometry and the multi-block structured grid were generated by G3DMESH and a grid sensitivity analysis was conducted. For the clearance gap region, a fully gridded special H-grid was chosen. Comparisons were made between the flow characteristic at design speed, representative for a transonic flow regime, and at 65% speed, representative for a subsonic flow regime. The computations were used to analyse the flow phenomena through the tip clearance region for the different configurations and their impact on the performance of the compressor stage.

  7. Wing flutter calculations with the CAP-TSD unsteady transonic small disturbance program

    NASA Technical Reports Server (NTRS)

    Bennett, Robert M.; Batina, John T.; Cunningham, Herbert J.

    1988-01-01

    The application and assessment is described of CAP-TSD (Computational Aeroelasticity Program - Transonic Small Disturbance) code for flutter prediction. The CAP-TSD program was developed for aeroelastic analysis of complete aircraft configurations and was previously applied to the calculation of steady and unsteady pressures. Flutter calculations are presented for two thin, swept-and-tapered wing planforms with well defined modal properties. The calculations are for Mach numbers from low subsonic to low supersonic values, including the transonic range, and are compared with subsonic linear theory and experimental flutter data. The CAP-TSD flutter results are generally in good agreement with the experimental values and are in good agreement with subsonic linear theory when wing thickness is neglected.

  8. Predicting the aeroelastic behavior of a wind-tunnel model using transonic small disturbance theory

    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 AFW wind-tunnel experiments. Dynamic results are presented in the form of root loci at different Mach numbers for a heavy gas and for air test mediums. The resultant flutter boundaries for both gases, and the effects of viscous damping and angle of attack on the flutter boundary in air, are also presented.

  9. High-Tip-Speed, Low-Loading Transonic Fan Stage. Part 1: Aerodynamic and Mechanical Design

    NASA Technical Reports Server (NTRS)

    Wright, L. C.; Vitale, N. G.; Ware, T. C.; Erwin, J. R.

    1973-01-01

    A high-tip-speed, low-loading transonic fan stage was designed to deliver an overall pressure ratio of 1.5 with an adiabatic efficiency of 86 percent. The design flow per unit annulus area is 42.0 pounds per square foot. The fan features a hub/tip ratio of 0.46, a tip diameter of 28.74 in. and operates at a design tip speed of 1600 fps. For these design conditions, the rotor blade tip region operates with supersonic inlet and supersonic discharge relative velocities. A sophisticated quasi-three-dimensional characteristic section design procedure was used for the all-supersonic sections and the inlet of the midspan transonic sections. For regions where the relative outlet velocities are supersonic, the blade operates with weak oblique shocks only.

  10. Effect of Moist Air on Transonic Internal Flow around a Plate

    NASA Astrophysics Data System (ADS)

    Hasan, A. B. M. Toufique; Matsuo, Shigeru; Setoguchi, Toshiaki; Kim, Heuy Dong

    The unsteady phenomena in the transonic flow around airfoils are observed in the flow field of fan, compressor blades and butterfly valves, and this causes often serious problems such as the aeroacoustic noise and the vibration. In the transonic or supersonic flow where vapor is contained in the main flow, the rapid expansion of the flow may give rise to a non-equilibrium condensation. In the present study, the effect of non-equilibrium condensation of moist air on the shock induced flow field oscillation around a plate was investigated numerically. The results showed that in the case with non-equilibrium condensation, the flow field aerodynamic unsteadiness is reduced significantly compared with those without the non-equilibrium condensation.

  11. Study of boundary-layer transition using transonic-cone preston tube data

    NASA Technical Reports Server (NTRS)

    Reed, T. D.; Moretti, P. M.

    1980-01-01

    The laminar boundary layer on a 10 degree cone in a transonic wind tunnel was studied. The inviscid flow and boundary layer development were simulated by computer programs. The effects of pitch and yaw angles on the boundary layer were examined. Preston-tube data, taken on the boundary-layer-transition cone in the NASA Ames 11 ft transonic wind tunnel, were used to develope a correlation which relates the measurements to theoretical values of laminar skin friction. The recommended correlation is based on a compressible form of the classical law-of-the-wall. The computer codes successfully simulates the laminar boundary layer for near-zero pitch and yaw angles. However, in cases of significant pitch and/or yaw angles, the flow is three dimensional and the boundary layer computer code used here cannot provide a satisfactory model. The skin-friction correlation is thought to be valid for body geometries other than cones.

  12. 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.

  13. Nine percent nickel steel heavy forging weld repair study. [National Transonic Wind Tunnel fan components

    NASA Technical Reports Server (NTRS)

    Young, C. P., Jr.; Gerringer, A. H.; Brooks, T. G.; Berry, R. F., Jr.

    1978-01-01

    The feasibility of making weld repairs on heavy section 9% nickel steel forgings such as those being manufactured for the National Transonic Facility fan disk and fan drive shaft components was evaluated. Results indicate that 9% nickel steel in heavy forgings has very good weldability characteristics for the particular weld rod and weld procedures used. A comparison of data for known similar work is included.

  14. Design of transonic airfoils and wings using a hybrid design algorithm

    NASA Technical Reports Server (NTRS)

    Campbell, Richard L.; Smith, Leigh A.

    1987-01-01

    A method has been developed for designing airfoils and wings at transonic speeds. It utilizes a hybrid design algorithm in an iterative predictor/corrector approach, alternating between analysis code and a design module. This method has been successfully applied to a variety of airfoil and wing design problems, including both transport and highly-swept fighter wing configurations. An efficient approach to viscous airfoild design and the effect of including static aeroelastic deflections in the wing design process are also illustrated.

  15. Additional flow quality measurements in the Langley Research Center 8-Foot Transonic Pressure Tunnel

    NASA Technical Reports Server (NTRS)

    Brooks, J. D.; Stainback, P. C.; Brooks, C. W., Jr.

    1980-01-01

    Additional tests were conducted to further define the disturbance characteristics of the Langley 8-Foot Transonic Pressure Tunnel. Measurements were made in the settling chamber with hot wire probes and in the test section with pressure transducers when various methods were used to choke the flow. In addition to presenting rms values measured at various locations and tunnel condition, autocorrelations and cross correlation data are also presented.

  16. Modeling transonic aerodynamic response using nonlinear systems theory for use with modern control theory

    NASA Technical Reports Server (NTRS)

    Silva, Walter A.

    1993-01-01

    The presentation begins with a brief description of the motivation and approach that has been taken for this research. This will be followed by a description of the Volterra Theory of Nonlinear Systems and the CAP-TSD code which is an aeroelastic, transonic CFD (Computational Fluid Dynamics) code. The application of the Volterra theory to a CFD model and, more specifically, to a CAP-TSD model of a rectangular wing with a NACA 0012 airfoil section will be presented.

  17. A Study of Grid Resolution, Transition and Turbulence Model Using the Transonic Simple Straked Delta Wing

    NASA Technical Reports Server (NTRS)

    Bartels, Robert E.

    2001-01-01

    Three-dimensional transonic flow over a delta wing is investigated using several turbulence models. The performance of linear eddy viscosity models and an explicit algebraic stress model is assessed at the start of vortex flow, and the results compared with experimental data. To assess the effect of transition location, computations that either fix transition aft of the leading edge or are fully turbulent are performed. These computations show that grid resolution, transition location and turbulence model significantly affect the 3D flowfield.

  18. Performance of a High-solidity High-pressure-ratio Transonic Rotor

    NASA Technical Reports Server (NTRS)

    Neumann, Harvey E

    1955-01-01

    A high-solidity low-aspect-ratio transonic rotor with an inlet hub-tip ratio of 0.52 was investigated experimentally. The rotor developed a total-pressure tatio of 1.93 and an efficiency of 92 percent at an equivalent wheel tip speed of 1036 feet per second. The feasibility of using this rotor as a component of an inlet stage of an axial-flow compressor was investigated.

  19. 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.

  20. Investigation of Unsteady Flow Behavior in Transonic Compressor Rotors with LES and PIV Measurements

    NASA Technical Reports Server (NTRS)

    Hah, Chunill; Voges, Melanie; Mueller, Martin; Schiffer, Heinz-Peter

    2009-01-01

    In the present study, unsteady flow behavior in a modern transonic axial compressor rotor is studied in detail with large eddy simulation (LES) and particle image velocimetry (PIV). The main purpose of the study is to advance the current understanding of the flow field near the blade tip in an axial transonic compressor rotor near the stall and peak-efficiency conditions. Flow interaction between the tip leakage vortex and the passage shock is inherently unsteady in a transonic compressor. Casing-mounted unsteady pressure transducers have been widely applied to investigate steady and unsteady flow behavior near the casing. Although many aspects of flow have been revealed, flow structures below the casing cannot be studied with casing-mounted pressure transducers. In the present study, unsteady velocity fields are measured with a PIV system and the measured unsteady flow fields are compared with LES simulations. The currently applied PIV measurements indicate that the flow near the tip region is not steady even at the design condition. This self-induced unsteadiness increases significantly as the compressor rotor operates near the stall condition. Measured data from PIV show that the tip clearance vortex oscillates substantially near stall. The calculated unsteady characteristics of the flow from LES agree well with the PIV measurements. Calculated unsteady flow fields show that the formation of the tip clearance vortex is intermittent and the concept of vortex breakdown from steady flow analysis does not seem to apply in the current flow field. Fluid with low momentum near the pressure side of the blade close to the leading edge periodically spills over into the adjacent blade passage. The present study indicates that stall inception is heavily dependent on unsteady behavior of the flow field near the leading edge of the blade tip section for the present transonic compressor rotor.