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Sample records for low-speed wind-tunnel tests

  1. 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.

  2. Tests of models equipped with TPS in low speed ONERA F1 pressurized wind tunnel

    NASA Astrophysics Data System (ADS)

    Leynaert, J.

    1992-09-01

    The particular conditions of tests of models equipped with a turbofan powered simulator (TPS) at high Reynolds numbers in a pressurized wind tunnel are presented. The high-pressure air supply system of the wind tunnel, the equipment of the balance with the high-pressure traversing flow and its calibration, and the thrust calibration method of the TPS and its verification in the wind tunnel are described.

  3. Low-speed wind tunnel test results of the Canard Rotor/Wing concept

    NASA Technical Reports Server (NTRS)

    Bass, Steven M.; Thompson, Thomas L.; Rutherford, John W.; Swanson, Stephen

    1993-01-01

    The Canard Rotor/Wing (CRW), a high-speed rotorcraft concept, was tested at the National Aeronautics and Space Administration (NASA) Ames Research Center's 40- by 80-Foot Wind Tunnel in Mountain View, California. The 1/5-scale model was tested to identify certain low-speed, fixed-wing, aerodynamic characteristics of the configuration and investigate the effectiveness of two empennages, an H-Tail and a T-Tail. The paper addresses the principal test objectives and the results achieved in the wind tunnel test. These are summarized as: i) drag build-up and differences between the H-Tail and T-Tail configuration, ii) longitudinal stability of the H-Tail and T-Tail configurations in the conversion and cruise modes, iii) control derivatives for the canard and elevator in the conversion and cruise modes, iv) aerodynamic characteristics of varying the rotor/wing azimuth position, and v) canard and tail lift/trim capability for conversion conditions.

  4. Infrared thermography for detection of laminar-turbulent transition in low-speed wind tunnel testing

    NASA Astrophysics Data System (ADS)

    Joseph, Liselle A.; Borgoltz, Aurelien; Devenport, William

    2016-05-01

    This work presents the details of a system for experimentally identifying laminar-to-turbulent transition using infrared thermography applied to large, metal models in low-speed wind tunnel tests. Key elements of the transition detection system include infrared cameras with sensitivity in the 7.5- to 14.0-µm spectral range and a thin, insulating coat for the model. The fidelity of the system was validated through experiments on two wind-turbine blade airfoil sections tested at Reynolds numbers between Re = 1.5 × 106 and 3 × 106. Results compare well with measurements from surface pressure distributions and stethoscope observations. However, the infrared-based system provides data over a much broader range of conditions and locations on the model. This paper chronicles the design, implementation and validation of the infrared transition detection system, a subject which has not been widely detailed in the literature to date.

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

    NASA Technical Reports Server (NTRS)

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

    1977-01-01

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

  6. Laser Velocimetry In Low-Speed Wind Tunnels

    NASA Technical Reports Server (NTRS)

    Orloff, Kenneth L.; Snyder, Philip K.; Reinath, Michael S.

    1990-01-01

    Design and performance of three-dimensional and two-dimensional backscatter laser velocimeter, both used in low-speed wind tunnels, described in report together with historical overview of development of laser velocimetry (LV). Provides measurements of airflow in wind-tunnel tests without perturbing effects of probes and probe-supporting structures. Applicable in such related fields as ventilation engineering and possibly in detection of wing vortexes from large aircraft at airports.

  7. Low-Speed Wind Tunnel Tests of Two Waverider Configuration Models

    NASA Technical Reports Server (NTRS)

    Pegg, Robert J.; Hahne, David E.; Cockrell,Charles E., Jr.

    1995-01-01

    A definitive measurement of the low-speed flight characteristics of waverider-based aircraft is required to augment the overall design database for this important class of vehicles which have great potential for efficient high-speed flight. Two separate waverider-derived vehicles were tested; one in the 14- by 22-Foot Tunnel and the other in the 12-Foot Low Speed Tunnel at Langley Research Center. These tests provided measurements of moments and forces about all three axes, control effectiveness, flow field characteristics and the effects of configuration changes. The results of these tunnel tests are summarized and the subsonic aerodynamic characteristics of the two configurations are shown.

  8. Pratt & Whitney Two Dimensional HSR Nozzle Test in the NASA Lewis 9- By 15- Foot Low Speed Wind Tunnel: Aerodynamic Results

    NASA Technical Reports Server (NTRS)

    Wolter, John D.; Jones, Christopher W.

    1999-01-01

    This paper discusses a test that was conducted jointly by Pratt & Whitney Aircraft Engines and NASA Lewis Research Center. The test was conducted in NASA's 9- by 15-Foot Low Speed Wind Tunnel (9x15 LSWT). The test setup, methods, and aerodynamic results of this test are discussed. Acoustical results are discussed in a separate paper by J. Bridges and J. Marino.

  9. Model-Scale Aerodynamic Performance Testing of Proposed Modifications to the NASA Langley Low Speed Aeroacoustic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Booth, Earl R., Jr.; Coston, Calvin W., Jr.

    2005-01-01

    Tests were performed on a 1/20th-scale model of the Low Speed Aeroacoustic Wind Tunnel to determine the performance effects of insertion of acoustic baffles in the tunnel inlet, replacement of the existing collector with a new collector design in the open jet test section, and addition of flow splitters to the acoustic baffle section downstream of the test section. As expected, the inlet baffles caused a reduction in facility performance. About half of the performance loss was recovered by addition the flow splitters to the downstream baffles. All collectors tested reduced facility performance. However, test chamber recirculation flow was reduced by the new collector designs and shielding of some of the microphones was reduced owing to the smaller size of the new collector. Overall performance loss in the facility is expected to be a 5 percent top flow speed reduction, but the facility will meet OSHA limits for external noise levels and recirculation in the test section will be reduced.

  10. Test data report, low speed wind tunnel tests of a full scale lift/cruise-fan inlet, with engine, at high angles of attack

    NASA Technical Reports Server (NTRS)

    Shain, W. M.

    1978-01-01

    A low speed wind tunnel test of a fixed lip inlet with engine, was performed. The inlet was close coupled to a Hamilton Standard 1.4 meter, variable pitch fan driven by a lycoming T55-L-11A engine. Tests were conducted with various combinations of inlet angle of attack freestream velocities, and fan airflows. Data were recorded to define the inlet airflow separation boundaries, performance characteristics, and fan blade stresses. The test model, installation, instrumentation, test, data reduction and final data are described.

  11. Low speed wind tunnel test of a propulsive wing/canard concept in the STOL configuration. Volume 2: Test data

    NASA Technical Reports Server (NTRS)

    Stewart, V. R.

    1987-01-01

    A propulsive wind/canard model was tested at STOL operating conditions in the NASA Langley Research Center 4 x 7 meter wind tunnel. Longitudinal and lateral/directional aerodynamic characteristics were measured for various flap deflections, angles of attack and sideslip, and blowing coefficients. Testing was conducted for several model heights to determine ground proximity effects on the aerodynamic characteristics. Flow field surveys of local flow angles and velocities were performed behind both the canard and the wing. This is volume 2 of a 2 volume report. All of the test data in three appendices are presented. Appendix A presented tabulated six component force and moment data, Appendix B presents tabulated wing pressure coefficients, and Appendix C presents the flow field data.

  12. Low-Speed Dynamic Wind Tunnel Test Analysis of a Generic 53 Degree Swept UCAV Configuration With Controls

    NASA Technical Reports Server (NTRS)

    Vicroy, Dan D.; Huber, Kerstin C.; Rohlf, Detlef; Loser, Thomas

    2014-01-01

    Several static and dynamic forced-motion wind tunnel tests have been conducted on a generic unmanned combat air vehicle (UCAV) configuration with a 53deg swept leading edge. These tests are part of an international research effort to assess and advance the state-of-art of computational fluid dynamics (CFD) methods to predict the static and dynamic stability and control characteristics for this type of configuration. This paper describes the dynamic forced motion data collected from two different models of this UCAV configuration as well as analysis of the control surface deflections on the dynamic forces and moments.

  13. Abe Silverstein 10- by 10-Foot Supersonic Wind Tunnel Validated for Low-Speed (Subsonic) Operation

    NASA Technical Reports Server (NTRS)

    Hoffman, Thomas R.

    2001-01-01

    The NASA Glenn Research Center and Lockheed Martin Corporation tested an aircraft model in two wind tunnels to compare low-speed (subsonic) flow characteristics. Objectives of the test were to determine and document the similarities and uniqueness of the tunnels and to validate that Glenn's 10- by 10-Foot Supersonic Wind Tunnel (10x10 SWT) is a viable low-speed test facility. Results from two of Glenn's wind tunnels compare very favorably and show that the 10x10 SWT is a viable low-speed wind tunnel. The Subsonic Comparison Test was a joint effort by NASA and Lockheed Martin using the Lockheed Martin's Joint Strike Fighter Concept Demonstration Aircraft model. Although Glenn's 10310 and 836 SWT's have many similarities, they also have unique characteristics. Therefore, test data were collected for multiple model configurations at various vertical locations in the test section, starting at the test section centerline and extending into the ceiling and floor boundary layers.

  14. Simulation of the flow past a model in the closed test section of a low-speed wind tunnel and in the free stream

    NASA Astrophysics Data System (ADS)

    Bui, V. T.; Lapygin, V. I.

    2015-05-01

    The flow around a model in the closed test section of a low-speed wind tunnel has been analyzed in 2D approximation. As the contour of the nozzle, test section, and diffuser, the contour of the T-324 wind tunnel, of the Khristianovich Institute of Theoretical and Applied Mechanics (ITAM SB RAS, Novosibirsk), in its symmetry plane was adopted. A comparison of experimental with calculated data on the distribution of velocities and dynamic pressures in the test section is given. The effect due to the sizes of a model installed in the test section on the values of the aerodynamic coefficients of the model is analyzed. As the aerodynamic model, the NASA0012 airfoil and the circular cylinder were considered. For the airfoil chord length b = 20 % of nozzle height, the values of the aerodynamic coefficients of the airfoil in the free stream and in the test section proved to be close to each other up to the angle of attack a = 7°, which configuration corresponds to blockage-factor value ξ ≈ 7 %. The obtained data are indicative of the expedience of taking into account, in choosing the model scale, not only the degree of flow passage area blockage by the model but, also, the length of the well-streamlined model. In the case of a strongly blunted body with a high drag-coefficient value, the admissible blockage factor ξ may reach a value of 10 %.

  15. Slotted-wall research with disk and parachute models in a low-speed wind tunnel

    SciTech Connect

    Macha, J.M.; Buffington, R.J.; Henfling, J.L. ); Every, D. Van; Harris, J.L. )

    1990-01-01

    An experimental investigation of slotted-wall blockage interference has been conducted using disk and parachute models in a low speed wind tunnel. Test section open area ratio, model geometric blockage ratio, and model location along the length of the test section were systematically varied. Resulting drag coefficients were compared to each other and to interference-free measurements obtained in a much larger wind tunnel where the geometric blockage ratio was less than 0.0025. 9 refs., 10 figs.

  16. An experimental study of several wind tunnel wall configurations using two V/STOL model configurations. [low speed wind tunnels

    NASA Technical Reports Server (NTRS)

    Binion, T. W., Jr.

    1975-01-01

    Experiments were conducted in the low speed wind tunnel using two V/STOL models, a jet-flap and a jet-in-fuselage configuration, to search for a wind tunnel wall configuration to minimize wall interference on V/STOL models. Data were also obtained on the jet-flap model with a uniform slotted wall configuration to provide comparisons between theoretical and experimental wall interference. A test section configuration was found which provided some data in reasonable agreement with interference-free results over a wide range of momentum coefficients.

  17. Pretest Report for the Full Span Propulsive Wing/Canard Model Test in the NASA Langley 4 x 7 Meter Low Speed Wind Tunnel Second Series Test

    NASA Technical Reports Server (NTRS)

    Stewart, V. R.

    1986-01-01

    A full span propulsive wing/canard model is to be tested in the NASA Langley Research Center (LaRC) 4 x 7 meter low speed wind tunnel. These tests are a continuation of the tests conducted in Feb. 1984, NASA test No.290, and are being conducted under NASA Contract NAS1-17171. The purpose of these tests is to obtain extensive lateral-directional data with a revised fuselage concept. The wings, canards, and vertical tail of this second test series model are the same as tested in the previous test period. The fuselage and internal flow path have been modified to better reflect an external configuration suitable for a fighter airplane. Internal ducting and structure were changed as required to provide test efficiency and blowing control. The model fuselage tested during the 1984 tests was fabricated with flat sides to provide multiple wing and canard placement variations. The locations of the wing and canard are important variables in configuration development. With the establishment of the desired relative placement of the lifting surfaces, a typically shaped fuselage has been fabricated for these tests. This report provides the information necessary for the second series tests of the propulsive wing/canard model. The discussion in this report is limited to that affected by the model changes and to the second series test program. The pretest report information for test 290 which is valid for the second series test was published in Rockwell report NR 83H-79. This report is presented as Appendix 1 and the modified fuselage stress report is presented as Appendix 2 to this pretest report.

  18. Low-Speed Wind-Tunnel Tests of a Pilotless Aircraft Having Horizontal and Vertical Wings and Cruciform Tail

    NASA Technical Reports Server (NTRS)

    Mastrocola, N; Assadourian, A

    1947-01-01

    Low-speed tests of a pilotless aircraft were conducted in the Langley propeller-research tunnel to provide information for the estimation of the longitudinal stability and. control, to measure the aileron effectiveness, and to calibrate the radome and the Machmeter pitot-static orifices. It was found that the model possessed a stEb.le variation of elevator angle required for trim throughout the speed range at the design angle of attack. A comparison of the airplane with and without JATO units and with an alternate rocket booster showed that a large loss in longitudinal stability and control resulting from the addition of the rocket booster to the aircraft was sufficient to make the rocket-booster assembly unsatisfactory as an alternate for the JATO units. Reversal of the aileron effectiveness was evident at positive deflections of the vertical wing flap indicating that the roll-stabilization system would produce roiling moments in a tight right turn contrary to its design purpose. Vertical-wing-flap deflections caused large errors in the static-pressure reading obtained by the original static-tube installation. A practical installation point on the fuselage was located which should yield reliable measurement of the free-stream static pressure.

  19. Evaluation of spray drift using low speed wind tunnel measurements and dispersion modeling

    Technology Transfer Automated Retrieval System (TEKTRAN)

    The objective of this work was to evaluate the EPA’s proposed Test Plan for the validation testing of pesticide spray drift reduction technologies (DRTs) for row and field crops, focusing on the evaluation of ground application systems using the low-speed wind tunnel protocols and processing the dat...

  20. Comparison of the 10x10 and the 8x6 Supersonic Wind Tunnels at the NASA Glenn Research Center for Low-Speed (Subsonic) Operation

    NASA Technical Reports Server (NTRS)

    Hoffman, Thomas R.; Johns, Albert L.; Bury, Mark E.

    2002-01-01

    NASA Glenn Research Center and Lockheed Martin tested an aircraft model in two wind tunnels to compare low-speed (subsonic) flow characteristics. Test objectives were to determine and document similarities and uniqueness of the tunnels and to verify that the 10- by 10-Foot Supersonic Wind Tunnel (10x10 SWT) is a viable low-speed test facility when compared to the 8- by 6-Foot Supersonic Wind Tunnel (8x6 SWT). Conclusions are that the data from the two facilities compares very favorably and that the 10-by 10-Foot Supersonic Wind Tunnel at NASA Glenn Research Center is a viable low-speed wind tunnel.

  1. Hot gas ingestion test results of a two-poster vectored thrust concept with flow visualization in the NASA Lewis 9- by 15-foot low speed wind tunnel

    NASA Technical Reports Server (NTRS)

    Johns, Albert L.; Neiner, George; Bencic, Timothy J.; Flood, Joseph D.; Amuedo, Kurt C.

    1990-01-01

    A 9.2 percent scale STOVL hot gas ingestion model was tested in the NASA Lewis 9 x 15-foot Low-Speed Wind Tunnel. Flow visualization from the Phase 1 test program, which evaluated the hot ingestion phenomena and control techniques, is covered. The Phase 2 test program evaluated the hot gas ingestion phenomena at higher temperatures and used a laser sheet to investigate the flow field. Hot gas ingestion levels were measured for the several forward nozzle splay configurations and with flow control/life improvement devices (LIDs) which reduced the hot gas ingestion. The test was conducted at full scale nozzle pressure ratios and inlet Mach numbers. Results are presented over a range of nozzle pressure ratios at a 10 kn headwind velocity. The Phase 2 program was conducted at exhaust nozzle temperatures up to 1460 R and utilized a sheet laser system for flow visualization of the model flow field in and out of ground effects. The results reported are for nozzle exhaust temperatures up to 1160 R and contain the compressor face pressure and temperature distortions, the total pressure recovery, the inlet temperature rise, and the environmental effects of the hot gas. The environmental effects include the ground plane contours, the model airframe heating, and the location of the ground flow separation.

  2. Hot gas ingestion testing of an advanced STOVL concept in the NASA Lewis 9- by 15-foot low speed wind tunnel with flow visualization

    NASA Technical Reports Server (NTRS)

    Johns, Albert L.; Flood, Joseph D.; Strock, Thomas W.; Amuedo, Kurt C.

    1988-01-01

    Advanced Short Takeoff/Vertical Landing (STOVL) aircraft capable of operating from remote sites, damaged runways, and small air capable ships are being pursued for deployment around the turn of the century. To achieve this goal, it is important that the technologies critical to this unique class of aircraft be developed. Recognizing this need, NASA Lewis Research Center, McDonnell Douglas Aircraft, and DARPA defined a cooperative program for testing in the NASA Lewis 9- by 15-Foot Low Speed Wind Tunnel (LSWT) to establish a database for hot gas ingestion, one of the technologies critical to STOVL. Results from a test program are presented along with a discussion of the facility modifications allowing this type of testing at model scale. These modifications to the tunnel include a novel ground plane, an elaborate model support which included 4 degrees of freedom, heated high pressure air for nozzle flow, a suction system exhaust for inlet flow, and tunnel sidewall modifications. Several flow visualization techniques were employed including water mist in the nozzle flows and tufts on the ground plane. Headwind (free-stream) velocity was varied from 8 to 23 knots.

  3. Hot gas ingestion testing of an advanced STOVL concept in the NASA Lewis 9- by 15-foot Low Speed Wind Tunnel with flow visualization

    NASA Technical Reports Server (NTRS)

    Johns, Albert L.; Flood, Joseph D.; Strock, Thomas W.; Amuedo, Kurt C.

    1988-01-01

    Advanced Short Takeoff/Vertical Landing (STOVL) aircraft capable of operating from remote sites, damaged runways, and small air capable ships are being pursued for deployment around the turn of the century. To achieve this goal, it is important that the technologies critical to this unique class of aircraft be developed. Recognizing this need, NASA Lewis Research Center, McDonnell Douglas Aircraft, and DARPA defined a cooperative program for testing in the NASA Lewis 9- by 15-foot Low Speed Wind Tunnel (LSWT) to establish a database for hot gas ingestion, one of the technologies critical to STOVL. Results from a test program are presented along with a discussion of the facility modifications allowing this type of testing at modal scale. These modifications to the tunnel include a novel ground plane, an elaborate model support which included 4 degrees of freedom, heated high pressure air for nozzle flow, a suction system exhaust for inlet flow, and tunnel sidewall modifications. Several flow visualization techniques were employed including water mist in the nozzle flows and tufts on the ground plane. Headwind (free-stream) velocity was varied from 8 to 23 knots.

  4. STOL and STOVL hot gas ingestion and airframe heating tests in the NASA Lewis 9- by 15-foot low-speed wind tunnel

    NASA Technical Reports Server (NTRS)

    Johns, Albert L.

    1989-01-01

    Short takeoff and landing (STOL) and advanced short takeoff and vertical landing (STOVL) aircraft are being pursued for deployment near the end of this century. These concepts offer unique capabilities not seen in conventional aircraft: for example, shorter takeoff distances and the ability to operate from damaged runways and remote sites. However, special technology is critical to the development of this unique class of aircraft. Some of the real issues that are associated with these concepts are hot gas ingestion and airframe heating while in ground effects. Over the past nine years, NASA Lewis Research Center has been involved in several cooperative programs in the 9- by 15 Foot Low-Speed Wind Tunnel (LSWT) to establish a database for hot gas ingestion and airframe heating. The modifications are presented that were made in the 9- by 15-Foot LSWT, including the evolution of the ground plane, model support system, and tunnel sidewalls; and flow visualization techniques, instrumentation, test procedures, and test results. The 9- by 15-Foot LSWT tests were conducted at full scale exhaust nozzle pressure ratios. The headwind velocities varied from 8 to 120 kn depending on the concept (STOL or STOVL). Typical compressor-face distortions (pressure and temperature), ground plane contours, and model surface temperature profiles are presented.

  5. Stability and control characteristics for the inner mold line configuration of the space shuttle orbiter (OA110). [tested in the low speed wind tunnel

    NASA Technical Reports Server (NTRS)

    Hughes, T.; Rogge, R.

    1974-01-01

    Experimental aerodynamic investigations were conducted on a sting mounted 0.0405-scale representation of the -140A/B inner mold line (IML) space shuttle orbiter in 7.75 x 11 foot low speed wind tunnel, during the time period from 18 March 1974 to 20 March 1974. The primary test objectives were to establish basic longitudinal and lateral-directional stability and control characteristics for the IML orbiter. Additional configurations investigated were sealed elevon hingeline gaps, sealed rudder split line and hingeline gaps, larger radius leading edge on the vertical tail, and sealed speedbrake base. Aerodynamic force and moment data for the orbiter were measured in the body-axis system by an internally mounted, six-component strain gage balance. The model was sting mounted with the center of rotation located at approximately the wing trailing edge. The nominal angle of attack range was from -4 to +30 degrees. Yaw polars were recorded over a nominal yaw angle range from -14 to +14 degrees at constant angles of attack of 0, + or - 5, 10, 15 and 20 degrees.

  6. Test data report: Low speed wind tunnel tests of a full scale, fixed geometry inlet, with engine, at high angles of attack

    NASA Technical Reports Server (NTRS)

    Shain, W. M.

    1976-01-01

    A full scale inlet test was to be done in the NASA-ARC 40' X 80' WT to demonstrate satisfactory inlet performance at high angles of attack. The inlet was designed to match a Hamilton-Standard 55 inch, variable pitch fan, driven by a Lycoming T55-L-11A gas generator. The test was installed in the wind tunnel on two separate occasions, but mechanical failures in the fan drive gear box early in each period terminated testing. A detailed description is included of the Model, installation, instrumentation and data reduction procedures.

  7. Low speed wind tunnel test of a propulsive wing/canard concept in the STOL configuration. Volume 1: Test description and discussion of results

    NASA Technical Reports Server (NTRS)

    Stewart, V. R.

    1987-01-01

    A propulsive wing/canard model was tested at STOL operating conditions in the NASA Langley Research Center 4 x 7 meter wind tunnel. Longitudinal and lateral/directional aerodynamic characteristics were measured for various flap deflections, angles of attack and sideslip, and blowing coefficients. Testing was conducted for several model heights to determine ground proximity effects on the aerodynamic characteristics. Flow field surveys of local flow angles and velocities were performed behind both the canard and the wing. This is volume 1 of a 2 volume report. The model, instrumentation, and test procedures are described. An analysis of the data is included.

  8. Hot gas ingestion test results of a two-poster vectored thrust concept with flow visualization in the NASA Lewis 9- x 15-foot low speed wind tunnel

    NASA Technical Reports Server (NTRS)

    Johns, Albert L.; Neiner, George; Bencic, Timothy J.; Flood, Joseph D.; Amuedo, Kurt C.; Strock, Thomas W.

    1990-01-01

    A 9.2 percent scale Short Takeoff and Vertical Landing (STOVL) hot gas ingestion model was designed and built by McDonnell Douglas Corporation (MCAIR) and tested in the Lewis Research Center 9 x 15 foot Low Speed Wind Tunnel (LSWT). Hot gas ingestion, the entrainment of heated engine exhaust into the inlet flow field, is a key development issure for advanced short takeoff and vertical landing aircraft. Flow visualization from the Phase 1 test program, which evaluated the hot ingestion phenomena and control techniques, is covered. The Phase 2 test program evaluated the hot gas ingestion phenomena at higher temperatures and used a laser sheet to investigate the flow field. Hot gas ingestion levels were measured for the several forward nozzle splay configurations and with flow control/life improvement devices (LIDs) which reduced the hot gas ingestion. The model support system had four degrees of freedom - pitch, roll, yaw, and vertical height variation. The model support system also provided heated high-pressure air for nozzle flow and a suction system exhaust for inlet flow. The test was conducted at full scale nozzle pressure ratios and inlet Mach numbers. Test and data analysis results from Phase 2 and flow visualization from both Phase 1 and 2 are documented. A description of the model and facility modifications is also provided. Headwind velocity was varied from 10 to 23 kn. Results are presented over a range of nozzle pressure ratios at a 10 kn headwind velocity. The Phase 2 program was conducted at exhaust nozzle temperatures up to 1460 R and utilized a sheet laser system for flow visualization of the model flow field in and out of ground effects. The results reported are for nozzle exhaust temperatures up to 1160 R. These results will contain the compressor face pressure and temperature distortions, the total pressure recovery, the inlet temperature rise, and the environmental effects of the hot gas. The environmental effects include the ground plane contours

  9. 9- by 15-Foot Low Speed Wind Tunnel Acoustic Improvements Expanded Overview

    NASA Technical Reports Server (NTRS)

    Stephens, David

    2016-01-01

    The 9- by 15-Foot Low Speed Wind Tunnel (9x15 LSWT) at NASA Glenn Research Center was built in 1969 in the return leg of the 8- by 6-Foot Supersonic Wind Tunnel (8x6 SWT). The 8x6 SWT was completed in 1949 and acoustically treated to mitigate community noise issues in 1950. This treatment included the addition of a large muffler downstream of the 8x6 SWT test section and diffuser. The 9x15 LSWT was designed for performance testing of V/STOL aircraft models, but with the addition of the current acoustic treatment in 1986 the tunnel been used principally for acoustic and performance testing of aircraft propulsion systems. The present document describes an anticipated acoustic upgrade to be completed in 2017.

  10. Additional Testing of the DHC-6 Twin Otter Tailplane Iced Airfoil Section in the Ohio State University 7x10 Low Speed Wind Tunnel. Volume 2

    NASA Technical Reports Server (NTRS)

    Gregorek, Gerald; Dresse, John J.; LaNoe, Karine; Ratvasky, Thomas (Technical Monitor)

    2000-01-01

    The need for fundamental research in Ice Contaminated Tailplane Stall (ICTS) was established through three international conferences sponsored by the FAA. A joint NASA/FAA Tailplane Icing Program was formed in 1994 with the Ohio State University playing a critical role for wind tunnel and analytical research. Two entries of a full-scale 2-dimensional tailplane airfoil model of a DHC-6 Twin Otter were made in The Ohio State University 7x10 ft wind tunnel. This report describes the second test entry that examined additional ice shapes and roughness, as well as airfoil section differences. The addition data obtained in this test fortified the original database of aerodynamic coefficients that permit a detailed analysis of flight test results with an OSU-developed analytical program. The testing encompassed a full range of angles of attack and elevator deflections at flight Reynolds number conditions. Aerodynamic coefficients, C(L), C(M), and C(He), were obtained by integrating static pressure coefficient, C(P), values obtained from surface taps. Comparisons of clean and iced airfoil results show a significant decrease in the tailplane aeroperformance (decreased C(Lmax), decreased stall angle, increased C(He)) for all ice shapes with the grit having the lease affect and the LEWICE shape having the greatest affect. All results were consistent with observed tailplane stall phenomena and constitute an effective set of data for comprehensive analysis of ICTS.

  11. Experimental study of the performance of clustered parachutes in a low speed wind tunnel

    SciTech Connect

    Baca, B.K.

    1984-04-01

    Increased interest in the use of clustered parachutes for high performance decelerator applications has led to a survey of cluster parachute data and theory. This examination indicted that very little well-documented data exists that is suitable for broad application to this type of problem. In order to remedy this situation, a low speed wind tunnel test program was devised to measure the effects of variations in four important parameters on the terminal drag of clustered parachutes. These parameters were canopy number, canopy type, riser length, and reefing. Wind tunnel blockage by the clustered parachutes was also studied and that effect was accounted for in the data reduction. Compensations for drag loss due to forebody wake were also made. In addition to drag data, information on stability and riser configuration was obtained. Possible explanations for various observed phenomena were proposed.

  12. The Acoustic Environment of the NASA Glenn 9- by 15-foot Low-Speed Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Stephens, David B.

    2015-01-01

    The 9- by 15-Foot Low Speed Wind Tunnel is an acoustic testing facility with a long history of aircraft propulsion noise research. Due to interest in renovating the facility to support future testing of advanced quiet engine designs, a study was conducted to document the background noise level in the facility and investigate the sources of contaminating noise. The anechoic quality of the facility was also investigated using an interrupted noise method. The present report discusses these aspects of the noise environment in this facility.

  13. NASA Lewis 9- by 15-foot low-speed wind tunnel user manual

    NASA Technical Reports Server (NTRS)

    Soeder, Ronald H.

    1993-01-01

    This manual describes the 9- by 15-Foot Low-Speed Wind Tunnel at the Lewis Research Center and provides information for users who wish to conduct experiments in this atmospheric facility. Tunnel variables such as pressures, temperatures, available tests section area, and Mach number ranges (0.05 to 0.20) are discussed. In addition, general support systems such as air systems, hydraulic system, hydrogen system, laser system, flow visualization system, and model support systems are described. Instrumentation and data processing and acquisition systems are also discussed.

  14. Large-Scale Wind-Tunnel Tests and Evaluation of the Low-Speed Performance of a 35 deg Sweptback Wing Jet Transport Model Equipped with a Blowing Boundary-Layer-Control Flap and Leading-Edge Slat

    NASA Technical Reports Server (NTRS)

    Hickey, David H.; Aoyagi, Kiyoshi

    1960-01-01

    A wind-tunnel investigation was conducted to determine the effect of trailing-edge flaps with blowing-type boundary-layer control and leading-edge slats on the low-speed performance of a large-scale jet transport model with four engines and a 35 deg. sweptback wing of aspect ratio 7. Two spanwise extents and several deflections of the trailing-edge flap were tested. Results were obtained with a normal leading-edge and with full-span leading-edge slats. Three-component longitudinal force and moment data and boundary-layer-control flow requirements are presented. The test results are analyzed in terms of possible improvements in low-speed performance. The effect on performance of the source of boundary-layer-control air flow is considered in the analysis.

  15. Binocular videogrammetric system for three-dimensional measurement in low-speed wind tunnel

    NASA Astrophysics Data System (ADS)

    Zhu, Ye; Gu, Yonggang; Zhai, Chao

    2014-11-01

    In order to avoid the defects of contact measurement, such as limited range, complex constructing and disability of 3-D parameter acquisition, we built a binocular videogrammetric system for measuring 3-D geometry parameters of wind tunnel test models, for instance, displacement, rotation angle and vibration, in low-speed wind tunnel. The system is based on the principles of close-range digital photogrammetry. As a non-contact system, it acquires parameters without interference in the experiments, and it has adjustable range and simple structure. It is worth mentioning that this is a Realtime measurement system, so that it can greatly compress the experiment period, furthermore, it is also able to provide some specific experiments with parameters for online adjustment. In this system, images are acquired through two industrial digital cameras and a PCI-E image acquisition card, and they are processed in a PC. The two cameras are triggered by signals come from a function signal generator, so that images of different cameras will have good temporal synchronization to ensure the accuracy of 3-D reconstruction. A two-step stereo calibration technique using planar pattern developed by Zhengyou Zhang is used to calibrate these cameras. Results of wind tunnel test indicate that the system can provide displacement accuracy better than 0.1% and rotation angle accuracy better than 0.1 degree, besides, the vibration frequency accuracy is superior to 0.1Hz in the low-frequency range.

  16. Landing pressure loads of the 140A/B space shuttle orbiter (model 43-0) determined in the Rockwell International low speed wind tunnel (OA69), volume 1. [wind tunnel tests

    NASA Technical Reports Server (NTRS)

    Soard, T. L.

    1975-01-01

    Wind tunnel tests of a 0.0405 scale model of the -1404A/B configuration of the Space Shuttle Vehicle Orbiter are presented. Pressure loads data were obtained from the orbiter in the landing configuration in the presence of the ground for structural strength analysis. This was accomplished by locating as many as 30 static pressure bugs at various locations on external model surfaces as each configuration was tested. A complete pressure loads survey was generated for each configuration by combining data from all bug locations, and these loads are described for the fuselage, wing, vertical tail, and landing gear doors. Aerodynamic force data was measured by a six component internal strain gage balance. This data was recorded to correct model angles of attack and sideslip for sting and balance deflections and to determine the aerodynamic effects of landing gear extension. All testing was conducted at a Mach number of 0.165 and a Reynolds number of 1.2 million per foot. Photographs of test configurations are shown.

  17. Low-speed aerodynamic characteristics from wind-tunnel tests of a large-scale advanced arrow-wing supersonic-cruise transport concept

    NASA Technical Reports Server (NTRS)

    Smith, P. M.

    1978-01-01

    Tests have been conducted to extend the existing low speed aerodynamic data base of advanced supersonic-cruise arrow wing configurations. Principle configuration variables included wing leading-edge flap deflection, wing trailing-edge flap deflection, horizontal tail effectiveness, and fuselage forebody strakes. A limited investigation was also conducted to determine the low speed aerodynamic effects due to slotted training-edge flaps. Results of this investigation demonstrate that deflecting the wing leading-edge flaps downward to suppress the wing apex vortices provides improved static longitudinal stability; however, it also results in significantly reduced static directional stability. The use of a selected fuselage forebody strakes is found to be effective in increasing the level of positive static directional stability. Drooping the fuselage nose, which is required for low-speed pilot vision, significantly improves the later-directional trim characteristics.

  18. Advancing Test Capabilities at NASA Wind Tunnels

    NASA Technical Reports Server (NTRS)

    Bell, James

    2015-01-01

    NASA maintains twelve major wind tunnels at three field centers capable of providing flows at 0.1 M 10 and unit Reynolds numbers up to 45106m. The maintenance and enhancement of these facilities is handled through a unified management structure under NASAs Aeronautics and Evaluation and Test Capability (AETC) project. The AETC facilities are; the 11x11 transonic and 9x7 supersonic wind tunnels at NASA Ames; the 10x10 and 8x6 supersonic wind tunnels, 9x15 low speed tunnel, Icing Research Tunnel, and Propulsion Simulator Laboratory, all at NASA Glenn; and the National Transonic Facility, Transonic Dynamics Tunnel, LAL aerothermodynamics laboratory, 8 High Temperature Tunnel, and 14x22 low speed tunnel, all at NASA Langley. This presentation describes the primary AETC facilities and their current capabilities, as well as improvements which are planned over the next five years. These improvements fall into three categories. The first are operations and maintenance improvements designed to increase the efficiency and reliability of the wind tunnels. These include new (possibly composite) fan blades at several facilities, new temperature control systems, and new and much more capable facility data systems. The second category of improvements are facility capability advancements. These include significant improvements to optical access in wind tunnel test sections at Ames, improvements to test section acoustics at Glenn and Langley, the development of a Supercooled Large Droplet capability for icing research, and the development of an icing capability for large engine testing. The final category of improvements consists of test technology enhancements which provide value across multiple facilities. These include projects to increase balance accuracy, provide NIST-traceable calibration characterization for wind tunnels, and to advance optical instruments for Computational Fluid Dynamics (CFD) validation. Taken as a whole, these individual projects provide significant

  19. A low speed wind tunnel test of a 0.050 scale model of shuttle orbiter (model 089B) to investigate the longitudinal and lateral directional effects of canard and tail configurational modifications in the LTV LSWT (MA14)

    NASA Technical Reports Server (NTRS)

    Chambliss, E. B.

    1976-01-01

    A low speed wind tunnel test was conducted to determine the effects of 6 canard configurations on the 0.050 scale model of shuttle orbiter 089B. In addition, two horizontal tail configurations were tested at two positions on the model as were two wing configurations. Since this test was restricted to 103 runs, only a limited number of permutations of the configurational changes could be tested. The testing was done in the 15 by 20 foot section of the LSWT and consisted of pitch polars, one yawed polar and several yaw runs. The pitch polars encompassed an alpha range from 0 to 28 deg; the yawed polar was run at beta = +2 degrees and the yaw runs covered a beta range from -6 to +6 deg at angles-of-attack of 0, 4, 10, 16, and 20 deg.

  20. Low-speed wind-tunnel tests of a 1/10-scale model of an advanced arrow-wing supersonic cruise configuration designed for cruise at Mach 2.2. [Langley Full Scale Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Yip, L. P.

    1979-01-01

    The low-speed longitudinal and lateral-directional characteristics of a scale model of an advanced arrow-wing supersonic cruise configuration were investigated in tests conducted at a Reynolds number of 4.19 x 10 to the 6th power based on the mean aerodynamic chord, with an angle of attack range from - 6 deg to 23 deg and sideslip angle range from -15 deg to 20 deg. The effects of segmented leading-edge flaps, slotted trailing-edge flaps, horizontal and vertical tails, and ailerons and spoilers were determined. Extensive pressure data and flow visualization pictures with non-intrusive fluorescent mini-tufts were obtained.

  1. NASA Now: Engineering Design: Wind Tunnel Testing

    NASA Video Gallery

    Dr. Norman W. Schaeffler, a NASA aerospace research engineer, describes how wind tunnels work and how aircraft designers use them to understand aerodynamic forces at low speeds. Learn the advantage...

  2. Low-speed wind tunnel tests of a 50.8-centimeter (20-in.) 1.15-pressure-ratio fan engine model

    NASA Technical Reports Server (NTRS)

    Wesoky, H. L.; Abbott, J. M.; Albers, J. A.; Dietrich, D. A.

    1974-01-01

    At a typical STOL aircraft takeoff and landing velocity, wind tunnel aerodynamic and acoustic measurements demonstrated that an inlet lip-area contraction ratio of 1.35 was superior to a ratio of 1.26 at high incidence angles. A 17 percent reduction in net thrust and an increase of 9 decibels in sound pressure level at the blade passing frequency resulted from inlet flow separation at an incidence angle of 50 deg with the 1.26-contraction-ratio inlet. Reverse-thrust forces obtained with blade rotation through the feathered angle were 1.8 times larger than with blade rotation through the flat angle. Reverse-thrust force was reduced from 30 to 50 percent and sound pressure level increased from 3 to 7 decibels at the blade passing frequency between the wind-tunnel-off condition and a typical STOL aircraft landing velocity.

  3. The results of a low-speed wind tunnel test to investigate the effects of the Refan JT8D engine target thrust reverser on the stability and control characteristics of the Boeing 727-200 airplane

    NASA Technical Reports Server (NTRS)

    Kupcis, E. A.

    1974-01-01

    The effects of the Refan JT8D side engine target thrust reverser on the stability and control characteristics of the Boeing 727-200 airplane were investigated using the Boeing-Vertol 20 x 20 ft Low-Speed Wind Tunnel. A powered model of the 727-200 was tested in groud effect in the landing configuration. The Refan target reverser configuration was evaluated relative to the basic production 727 airplane with its clamshell-deflector door thrust reverser design. The Refan configuration had slightly improved directional control characteristics relative to the basic airplane. Clocking the Refan thrust reversers 20 degrees outboard to direct the reverser flow away from the vertical tail, had little effect on directional control. However, clocking them 20 degrees inboard resulted in a complete loss of rudder effectiveness for speeds greater than 90 knots. Variations in Refan reverser lip/fence geometry had a minor effect on directional control.

  4. RITD – Wind tunnel testing

    NASA Astrophysics Data System (ADS)

    Haukka, Harri; Harri, Ari-Matti; Aleksashkin, Sergei; Koryanov, Valeri; Schmidt, Walter; Heilimo, Jyri; Finchenko, Valeri; Martynov, Maxim; Ponomarenko, Andrey; Kazakovtsev, Victor; Arruego, Ignazio

    2015-04-01

    An atmospheric re-entry and descent and landing system (EDLS) concept based on inflatable hypersonic decelerator techniques is highly promising for the Earth re-entry missions. We developed such EDLS for the Earth re-entry utilizing a concept that was originally developed for Mars. This EU-funded project is called RITD - Re-entry: Inflatable Technology Development - and it was to assess the bene¬fits of this technology when deploying small payloads from low Earth orbits to the surface of the Earth with modest costs. The principal goal was to assess and develope a preliminary EDLS design for the entire relevant range of aerodynamic regimes expected to be encountered in Earth's atmosphere during entry, descent and landing. The RITD entry and descent system utilizes an inflatable hypersonic decelerator. Development of such system requires a combination of wind tunnel tests and numerical simulations. This included wind tunnel tests both in transsonic and subsonic regimes. The principal aim of the wind tunnel tests was the determination of the RITD damping factors in the Earth atmosphere and recalculation of the results for the case of the vehicle descent in the Mars atmosphere. The RITD mock-up model used in the tests was in scale of 1:15 of the real-size vehicle as the dimensions were (midsection) diameter of 74.2 mm and length of 42 mm. For wind tunnel testing purposes the frontal part of the mock-up model body was manufactured by using a PolyJet 3D printing technology based on the light curing of liquid resin. The tail part of the mock-up model body was manufactured of M1 grade copper. The structure of the mock-up model placed th center of gravity in the same position as that of the real-size RITD. The wind tunnel test program included the defining of the damping factor at seven values of Mach numbers 0.85; 0.95; 1.10; 1.20; 1.25; 1.30 and 1.55 with the angle of attack ranging from 0 degree to 40 degrees with the step of 5 degrees. The damping characteristics of

  5. The cryogenic wind tunnel for high Reynolds number testing

    NASA Technical Reports Server (NTRS)

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

    1982-01-01

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

  6. Space Shuttle Orbiter Crew Hatch Jettison Test using a 0.0405-scale model (16-0) in the Texas A/M low speed wind tunnel (OA362). Space Shuttle aerothermodynamic data report

    NASA Technical Reports Server (NTRS)

    Mitchell, C. E.

    1992-01-01

    This report contains post-test information for the Space Shuttle Orbiter Crew Hatch Jettison Test OA362 which was conducted in the Texas A&M Low Speed Wind Tunnel from 6/15/87 to 6/22/87. The test objective was to verify that the crew hatch, once jettisoned, would clear the orbiter under various simulated flight conditions. Several model hatches were used with the 0.0405-scale orbiter (Model 16-0). The model's angle of attack was set at 10, 15, and 20 degrees while the sideslip had values of minus 5, 0, and plus 5 degrees. The full scale Qbars that were simulated were 105, 128, 160, and 210 psf. In the hatch jettison mechanism itself, the plunger pressure was varied to achieve horizontal velocities of 3, 5, 7, and 20.1 feet per second model scale, and the plunger location was varied to achieve a variety of rotational velocities. The orbiter model was subjected to 122 runs with 13 different hatches. Of these, 60 were good runs.

  7. Wind tunnel force and pressure tests

    NASA Technical Reports Server (NTRS)

    Wentz, W. H., Jr.

    1981-01-01

    Force and surface pressure distributions were measured for a 13% medium speed (NASA MS(1)-0313) airfoil fitted with 20% aileron, 25% slotted flap and 10% slot lip spoiler. All tests were conducted in the Walter Beech Memorial Wind Tunnel at a Reynolds number of 2.2 million and a Mach number of 0.13. Results include lift, drag, pitching moments, control surface normal force and hinge moments, and surface pressure distributions. The basic airfoil exhibits low speed characteristics similar to the GA(W)-2 airfoil. Incremental aileron and spoiler performance are quite comparable to that obtained on the GA(W)-2 airfoil. Slotted flap performance on this section is reduced compared to the GA(W)-2, resulting in a highest c sub l max of 3.00 compared to 3.35 for the GA(W)-2.

  8. Low speed wind tunnel test of ground proximity and deck edge effects on a lift cruise fan V/STOL configuration, volume 1

    NASA Technical Reports Server (NTRS)

    Stewart, V. R.

    1979-01-01

    The characteristics were determined of a lift cruise fan V/STOL multi-mission configuration in the near proximity to the edge of a small flat surface representation of a ship deck. Tests were conducted at both static and forward speed test conditions. The model (0.12 scale) tested was a four fan configuration with modifications to represent a three fan configuration. Analysis of data showed that the deck edge effects were in general less critical in terms of differences from free air than a full deck (in ground effect) configuration. The one exception to this was when the aft edge of the deck was located under the center of gravity. This condition, representative of an approach from the rear, showed a significant lift loss. Induced moments were generally small compared to the single axis control power requirements, but will likely add to the pilot work load.

  9. Analysis of a Split-Plot Experimental Design Applied to a Low-Speed Wind Tunnel Investigation

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.

    2013-01-01

    A procedure to analyze a split-plot experimental design featuring two input factors, two levels of randomization, and two error structures in a low-speed wind tunnel investigation of a small-scale model of a fighter airplane configuration is described in this report. Standard commercially-available statistical software was used to analyze the test results obtained in a randomization-restricted environment often encountered in wind tunnel testing. The input factors were differential horizontal stabilizer incidence and the angle of attack. The response variables were the aerodynamic coefficients of lift, drag, and pitching moment. Using split-plot terminology, the whole plot, or difficult-to-change, factor was the differential horizontal stabilizer incidence, and the subplot, or easy-to-change, factor was the angle of attack. The whole plot and subplot factors were both tested at three levels. Degrees of freedom for the whole plot error were provided by replication in the form of three blocks, or replicates, which were intended to simulate three consecutive days of wind tunnel facility operation. The analysis was conducted in three stages, which yielded the estimated mean squares, multiple regression function coefficients, and corresponding tests of significance for all individual terms at the whole plot and subplot levels for the three aerodynamic response variables. The estimated regression functions included main effects and two-factor interaction for the lift coefficient, main effects, two-factor interaction, and quadratic effects for the drag coefficient, and only main effects for the pitching moment coefficient.

  10. Low-Speed Wind-Tunnel Test of an Unpowered High-Speed Stoppable Rotor Concept in Fixed-Wing Mode

    NASA Technical Reports Server (NTRS)

    Lance, Michael B.; Sung, Daniel Y.; Stroub, Robert H.

    1991-01-01

    An experimental investigation of the M85, a High Speed Rotor Concept, was conducted at the NASA Langley 14 x 22 foot Subsonic Tunnel, assisted by NASA-Ames. An unpowered 1/5 scale model of the XH-59A helicopter fuselage with a large circular hub fairing, two rotor blades, and a shaft fairing was used as a baseline configuration. The M85 is a rotor wing hybrid aircraft design, and the model was tested with the rotor blade in the fixed wing mode. Assessments were made of the aerodynamic characteristics of various model rotor configurations. Variation in configurations were produced by changing the rotor blade sweep angle and the blade chord length. The most favorable M85 configuration tested included wide chord blades at 0 deg sweep, and it attained a system lift to drag ratio of 8.4.

  11. The design of a low-speed wind tunnel for studying the flow field of insects' flight

    NASA Astrophysics Data System (ADS)

    Zhao, Hong-yan; Zhang, Peng-fei; Ma, Yun; Ning, Jian-guo

    2015-03-01

    In this paper, low-speed smoke wind tunnel has been designed and fabricated for the insects' flow field visualization. The test section and the contraction section of the tunnel are optimized and determined as to size by the method of computational fluid dynamics. And fairing devices are equipped in different sections to reduce the turbulence intensity and increase the flow uniformity in the experimental sections. For the smoke visualization of small insects, the smokeemitting equipment has been specially designed and carefully debugged. Composed of wind tunnel, light source and high-speed camera, experimental platform for visualization and filming of insect flight flow field has been established. Besides, the feasible and stable method for insect fixing has been designed. With the smoke wind tunnel, flow filed visualization experiment for the honeybee's flapping was conducted and smoke flow filed in the experiment was recorded and analyzed. Near-filed and far-filed vortex structure when the honeybee fly can be recorded clearly. The experimental results indicate that the experimental platform is appropriate for flow filed study on insects flapping.

  12. Slotted-wall research with disk and parachute models in the DSMA low-speed wind tunnel

    SciTech Connect

    Van Every, D.; Harris, J.L. )

    1990-06-01

    A test program investigated the effects of wall open area ratio (OAR) and model axial position on the measured drag of disk and parachute models in a low-speed wind tunnel. The data and discussion presented in this report provide new insight into the nature of slotted-wall interference for bluff bodies in steady flow and give the first quantitative information on nonsteady wall interference and airflow response during the inflation of a parachute. The report concludes that a fixed OAR of between 5% and 15% should eliminate wall interference during inflation and greatly reduce steady-flow interference for geometric blockages up to 15%. Preliminary arguments suggest that an optimum OAR may be found that alleviates wall interference for large models at low speeds while providing for acceptable testing of smaller models in the transonic speed range. 10 refs., 36 figs., 14 tabs.

  13. V/STOL wind-tunnel testing

    NASA Technical Reports Server (NTRS)

    Koenig, D. G.

    1984-01-01

    Factors influencing effective program planning for V/STOL wind-tunnel testing are discussed. The planning sequence itself, which includes a short checklist of considerations that could enhance the value of the tests, is also described. Each of the considerations, choice of wind tunnel, type of model installation, model development and test operations is discussed, and examples of appropriate past and current V/STOL test programs are provided. A short survey of the moderate to large subsonic wind tunnels is followed by a review of several model installations, from dimensional to large-scale models of complete aircraft configurations. Model sizing, power simulation, and planning are treated, including three areas in test operations: data acquisition systems, acoustic measurements in wind tunnels, and flow surveying.

  14. V/STOL wind-tunnel testing

    NASA Technical Reports Server (NTRS)

    Koenig, D. G.

    1984-01-01

    Factors influencing effective program planning for V/STOL wind-tunnel testing are discussed. The planning sequence itself, which includes a short checklist of considerations that could enhance the value of the tests, is also described. Each of the considerations, choice of wind tunnel, type of model installation, model development and test operations, is discussed, and examples of appropriate past and current V/STOL test programs are provided. A short survey of the moderate to large subsonic wind tunnels is followed by a review of several model installations, from two-dimensional to large-scale models of complete aircraft configurations. Model sizing, power simulation, and planning are treated, including three areas is test operations: data-acquisition systems, acoustic measurements in wind tunnels, and flow surveying.

  15. Gottingen Wind Tunnel for Testing Aircraft Models

    NASA Technical Reports Server (NTRS)

    Prandtl, L

    1920-01-01

    Given here is a brief description of the Gottingen Wind Tunnel for the testing of aircraft models, preceded by a history of its development. Included are a number of diagrams illustrating, among other things, a sectional elevation of the wind tunnel, the pressure regulator, the entrance cone and method of supporting a model for simple drag tests, a three-component balance, and a propeller testing device, all of which are discussed in the text.

  16. AMELIA Tests in NASA Wind Tunnel

    NASA Video Gallery

    This report from "This Week @ NASA" describes recent aerodynamic tests of a subscale model of the Advanced Model for Extreme Lift and Improved Aeroacoustics, or "AMELIA," in a NASA wind tunnel. The...

  17. Space Shuttle wind tunnel testing program

    NASA Technical Reports Server (NTRS)

    Whitnah, A. M.; Hillje, E. R.

    1984-01-01

    A major phase of the Space Shuttle Vehicle (SSV) Development Program was the acquisition of data through the space shuttle wind tunnel testing program. It became obvious that the large number of configuration/environment combinations would necessitate an extremely large wind tunnel testing program. To make the most efficient use of available test facilities and to assist the prime contractor for orbiter design and space shuttle vehicle integration, a unique management plan was devised for the design and development phase. The space shuttle program is reviewed together with the evolutional development of the shuttle configuration. The wind tunnel testing rationale and the associated test program management plan and its overall results is reviewed. Information is given for the various facilities and models used within this program. A unique posttest documentation procedure and a summary of the types of test per disciplines, per facility, and per model are presented with detailed listing of the posttest documentation.

  18. Investigation of Model Wake Blockage Effects at High Angles of Attack in Low-Speed Wind Tunnel

    NASA Astrophysics Data System (ADS)

    Shyu, Lih-Shyng; Chuang, Shu-Hao

    To improve the fidelity of measured aerodynamic characteristics at high angle of attack for modern jet fighters, this paper examines the model wake blockage effect. The wake blockage effect in a 2.2×3.1 m low-speed wind tunnel is investigated by analyzing drag and wall pressure measurements. Circular flat plates of different sizes are used to simulate a test model at high angles of attack. The present analysis results in simple formulas for corrections of model wake blockage effect. To verify the present correction formula, the NASA TP-1803 model is force-tested in the tunnel. The corrected test data agree well with the NASA TP-1803 data.

  19. Full-scale S-76 rotor performance and loads at low speeds in the NASA Ames 80- by 120-Foot Wind Tunnel. Vol. 1

    NASA Technical Reports Server (NTRS)

    Shinoda, Patrick M.

    1996-01-01

    A full-scale helicopter rotor test was conducted in the NASA Ames 80- by 120-Foot Wind Tunnel with a four-bladed S-76 rotor system. Rotor performance and loads data were obtained over a wide range of rotor shaft angles-of-attack and thrust conditions at tunnel speeds ranging from 0 to 100 kt. The primary objectives of this test were (1) to acquire forward flight rotor performance and loads data for comparison with analytical results; (2) to acquire S-76 forward flight rotor performance data in the 80- by 120-Foot Wind Tunnel to compare with existing full-scale 40- by 80-Foot Wind Tunnel test data that were acquired in 1977; (3) to evaluate the acoustic capability of the 80- by 120- Foot Wind Tunnel for acquiring blade vortex interaction (BVI) noise in the low speed range and compare BVI noise with in-flight test data; and (4) to evaluate the capability of the 80- by 120-Foot Wind Tunnel test section as a hover facility. The secondary objectives were (1) to evaluate rotor inflow and wake effects (variations in tunnel speed, shaft angle, and thrust condition) on wind tunnel test section wall and floor pressures; (2) to establish the criteria for the definition of flow breakdown (condition where wall corrections are no longer valid) for this size rotor and wind tunnel cross-sectional area; and (3) to evaluate the wide-field shadowgraph technique for visualizing full-scale rotor wakes. This data base of rotor performance and loads can be used for analytical and experimental comparison studies for full-scale, four-bladed, fully articulated rotor systems. Rotor performance and structural loads data are presented in this report.

  20. Enabling Advanced Wind-Tunnel Research Methods Using the NASA Langley 12-Foot Low Speed Tunnel

    NASA Technical Reports Server (NTRS)

    Busan, Ronald C.; Rothhaar, Paul M.; Croom, Mark A.; Murphy, Patrick C.; Grafton, Sue B.; O-Neal, Anthony W.

    2014-01-01

    Design of Experiment (DOE) testing methods were used to gather wind tunnel data characterizing the aerodynamic and propulsion forces and moments acting on a complex vehicle configuration with 10 motor-driven propellers, 9 control surfaces, a tilt wing, and a tilt tail. This paper describes the potential benefits and practical implications of using DOE methods for wind tunnel testing - with an emphasis on describing how it can affect model hardware, facility hardware, and software for control and data acquisition. With up to 23 independent variables (19 model and 2 tunnel) for some vehicle configurations, this recent test also provides an excellent example of using DOE methods to assess critical coupling effects in a reasonable timeframe for complex vehicle configurations. Results for an exploratory test using conventional angle of attack sweeps to assess aerodynamic hysteresis is summarized, and DOE results are presented for an exploratory test used to set the data sampling time for the overall test. DOE results are also shown for one production test characterizing normal force in the Cruise mode for the vehicle.

  1. Evaluation of the buoyancy drag on automobile models in low speed wind tunnels

    NASA Astrophysics Data System (ADS)

    Mokry, Miroslav

    Of the several sources of inaccuracy in interpreting wind tunnel data for automobile models, the most prominent is the blockage interference. Streamwise variation of the wall induced pressure gives, in addition, rise to buoyancy drag. Buoyancy drag is analyzed in closed, 3/4 open, and slotted wind tunnels. The disturbance velocity potential is represented by a simple layer distribution. A numerical solution is obtained by a first-order panel method, approximating the surface by an assembly of flat panels, with a piecewise constant source density. The increment of the pressure coefficient due to wall interference considers only the contributions of the wall panels. Examples of the calculated buoyancy drag are given for the generic car model of the Motor Industry Research Association. Judged by the magnitude of the buoyancy drag, experiments at high blockage ratios would be highly distorted if performed in a closed-wall test section. However, with 30 percent open area ratio slotted walls, the buoyancy drag is reduced to about the same magnitude as that for test sections with low blockage ratios.

  2. A low speed wind tunnel test of the 0.050 scale NASA-JSC shuttle orbiter 089B to determine the longitudinal and lateral directional effects of control surface modifications

    NASA Technical Reports Server (NTRS)

    Oldenbuttel, R. H.

    1973-01-01

    Wind tunnel tests to determine the longitudinal and lateral-directional effects of control surface modifications on the space shuttle orbiter aerodynamic characteristics are discussed. A total of 103 data runs were made which consisted of pitch runs through a range of zero to 28 degrees at a zero yaw angle and yaw runs from minus 6 to plus 6 degrees at various fixed pitch angles. At each data point, data from an internal strain gage balance was sampled with the digital data system. Also recorded were the model angles of pitch and yaw and the test section static pressure. Results are presented in the form of tabulated aerodynamic coefficient data about the model reference center.

  3. Wind tunnel results of the low-speed NLF(1)-0414F airfoil

    NASA Technical Reports Server (NTRS)

    Murri, Daniel G.; Mcghee, Robert J.; Jordan, Frank L., Jr.; Davis, Patrick J.; Viken, Jeffrey K.

    1987-01-01

    The large performance gains predicted for the Natural Laminar Flow (NLF)(1)-0414F airfoil were demonstrated in two-dimensional airfoil tests and in wind tunnel tests conducted with a full scale modified Cessna 210. The performance gains result from maintaining extensive areas of natural laminar flow, and were verified by flight tests conducted with the modified Cessna. The lift, stability, and control characteristics of the Cessna were found to be essentially unchanged when boundary layer transition was fixed near the wing leading edge. These characteristics are very desirable from a safety and certification view where premature boundary layer transition (due to insect contamination, etc.) must be considered. The leading edge modifications were found to enhance the roll damping of the Cessna at the stall, and were therefore considered effective in improving the stall/departure resistance. Also, the modifications were found to be responsible for only minor performance penalties.

  4. Low-speed wind tunnel performance of high-speed counterrotation propellers at angle-of-attack

    NASA Technical Reports Server (NTRS)

    Hughes, Christopher E.; Gazzaniga, John A.

    1989-01-01

    The low-speed aerodynamic performance characteristics of two advanced counterrotation pusher-propeller configurations with cruise design Mach numbers of 0.72 were investigated in the NASA Lewis 9- by 15-Foot Low-Speed Wind Tunnel. The tests were conducted at Mach number 0.20, which is representative of the aircraft take-off/landing flight regime. The investigation determined the effect of nonuniform inflow on the propeller performance characteristics for several blade angle settings and a range of rotational speeds. The inflow was varied by yawing the propeller model to angle-of-attack by as much as plus or minus 16 degrees and by installing on the counterrotation propeller test rig near the propeller rotors a model simulator of an aircraft engine support pylon and fuselage. The results of the investigation indicated that the low-speed performance of the counterrotation propeller configurations near the take-off target operating points were reasonable and were fairly insensitive to changes in model angle-of-attack without the aircraft pylon/fuselage simulators installed on the propeller test rig. When the aircraft pylon/fuselage simulators were installed, small changes in propeller performance were seen at zero angle-of-attack, but fairly large changes in total power coefficient and very large changes of aft-to-forward-rotor torque ratio were produced when the propeller model was taken to angle-of-attack. The propeller net efficiency, though, was fairly insensitive to any changes in the propeller flowfield conditions near the take-off target operating points.

  5. Study on Busemann Biplane Airfoil in Low-Speed Smoke Wind Tunnel

    NASA Astrophysics Data System (ADS)

    Kashitani, Masashi; Yamaguchi, Yutaka; Kai, Yoshiharu; Hirata, Kenichi; Kusunose, Kazuhiro

    The Busemann biplane airfoil is considered one of the candidates for reducing sonic boom. In aircraft designs utilizing the biplane concept, high-lift devices must be used for takeoff and landing in low-speed conditions. In this work, flow visualizations were performed around a Busemann biplane airfoil equipped with leading and trailing edge flaps in a smoke wind tunnel. The lift coefficient of the biplane airfoil was estimated by utilizing a method based on measurements of smoke line patterns. The aspect ratio of the baseline Busemann biplane model was 0.75, the thickness ratio of the single element was 5%, and the wave cancellation condition was designed for Mach number 1.7. The length of each of the flap chords was 30% of the baseline. The Reynolds number, which is based on the chord length of the airfoil, is about 2.8×105. The results of the study are summarized as follows. For the baseline Busemann airfoil without flaps, the lift coefficient increases linearly as the angle of attack increases. The slope of the lift coefficient cl is 0.062 (1/deg.), which is in good agreement with reference data. This indicates that measuring smoke line patterns is a valid method for estimating the lift coefficient of biplane airfoils. Based on the visualization of the flow around the biplane model equipped with deflected leading and trailing edge flaps, confirmed that the separation bubble is smaller than in the baseline model due to the effective increase in camber. When the deflection angle of the trailing edge flap is increased, the lift coefficient also increases. The trend of the increasing cl is similar to that of conventional monoplane airfoil models with trailing edge flaps. Therefore, such flaps can be considered effective high-lift devices for Busemann biplane airfoils.

  6. Atmospheric Probe Model: Construction and Wind Tunnel Tests

    NASA Technical Reports Server (NTRS)

    Vogel, Jerald M.

    1998-01-01

    The material contained in this document represents a summary of the results of a low speed wind tunnel test program to determine the performance of an atmospheric probe at low speed. The probe configuration tested consists of a 2/3 scale model constructed from a combination of hard maple wood and aluminum stock. The model design includes approximately 130 surface static pressure taps. Additional hardware incorporated in the baseline model provides a mechanism for simulating external and internal trailing edge split flaps for probe flow control. Test matrix parameters include probe side slip angle, external/internal split flap deflection angle, and trip strip applications. Test output database includes surface pressure distributions on both inner and outer annular wings and probe center line velocity distributions from forward probe to aft probe locations.

  7. Nano-ADEPT Aeroloads Wind Tunnel Test

    NASA Technical Reports Server (NTRS)

    Smith, Brandon; Yount, Bryan; Kruger, Carl; Brivkalns, Chad; Makino, Alberto; Cassell, Alan; Zarchi, Kerry; McDaniel, Ryan; Ross, James; Wercinski, Paul; Venkatapathy, Ethiraj; Swanson, Gregory; Gold, Nili

    2016-01-01

    A wind tunnel test of the Adaptable Deployable Entry and Placement Technology (ADEPT) was conducted in April 2015 at the US Army's 7 by10 Foot Wind Tunnel located at NASA Ames Research Center. Key geometric features of the fabric test article were a 0.7 meter deployed base diameter, a 70 degree half-angle forebody cone angle, eight ribs, and a nose-to-base radius ratio of 0.7. The primary objective of this wind tunnel test was to obtain static deflected shape and pressure distributions while varying pretension at dynamic pressures and angles of attack relevant to entry conditions at Earth, Mars, and Venus. Other objectives included obtaining aerodynamic force and moment data and determining the presence and magnitude of any dynamic aeroelastic behavior (buzz/flutter) in the fabric trailing edge. All instrumentation systems worked as planned and a rich data set was obtained. This paper describes the test articles, instrumentation systems, data products, and test results. Four notable conclusions are drawn. First, test data support adopting a pre-tension lower bound of 10 foot pounds per inch for Nano-ADEPT mission applications in order to minimize the impact of static deflection. Second, test results indicate that the fabric conditioning process needs to be reevaluated. Third, no flutter/buzz of the fabric was observed for any test condition and should also not occur at hypersonic speeds. Fourth, translating one of the gores caused ADEPT to generate lift without the need for a center of gravity offset. At hypersonic speeds, the lift generated by actuating ADEPT gores could be used for vehicle control.

  8. Results of the Low Speed Aeroelastic Buffet Test with a 0.046-scale Model (747-ax1322-d-3/orbiter 8-0) of the 747 Cam/orbiter in the University of Washington Wind Tunnel (CS 3)

    NASA Technical Reports Server (NTRS)

    Gillins, R. L.

    1976-01-01

    A series of wind tunnel studies designed to assess the potential buffet problems resulting from orbiter wake characteristics with its tailcone removed are presented to provide design loads and acceleration environments, and to develop data on buffet sensitivity to various aerodynamic configurations and flight parameters. Data are intended to support subsequent analyses of structural fatigue life, crew efficiency, and equipment vibrations.

  9. Wind Tunnel Tests of Wind Turbine Airfoils at High Reynolds Numbers

    NASA Astrophysics Data System (ADS)

    Llorente, E.; Gorostidi, A.; Jacobs, M.; Timmer, W. A.; Munduate, X.; Pires, O.

    2014-06-01

    Wind tunnel tests have been performed to measure the two-dimensional aerodynamic characteristics of two different airfoil families at high Reynolds numbers (from 3 to 12 millions) in the DNW High Pressure Wind Tunnel in Gottingen (HDG), Germany. Also, tests at a Reynolds number of 3 millions have been performed in the Low-Speed Low- Turbulence Wind Tunnel of Delft University, The Netherlands. The airfoils tested belong to two wind turbine dedicated families: the TU-Delft DU family and the ACCIONA Windpower AWA family that was designed in collaboration with CENER. Reynolds number effects on airfoil performance have been obtained in the range of 3 to 12 millions. The availability of data from two different wind tunnels has brought the opportunity to cross compare the results from the two facilities.

  10. Comparison between design and installed acoustic characteristics of NASA Lewis 9- by 15-foot low-speed wind tunnel acoustic treatment

    NASA Technical Reports Server (NTRS)

    Dahl, Milo D.; Woodward, Richard P.

    1990-01-01

    The test section of the NASA Lewis 9- by 15-Foot Low-Speed Wind Tunnel was acoustically treated to allow the measurement of sound under simulated free-field conditions. The treatment was designed for high sound absorption at frequencies above 250 Hz and for withstanding the environmental conditions in the test section. In order to achieve the design requirements, a fibrous, bulk-absorber material was packed into removable panel sections. Each section was divided into two equal-depth layers packed with material to different bulk densities. The lower density was next to the facing of the treatment. The facing consisted of a perforated plate and screening material layered together. Sample tests for normal-incidence acoustic absorption were also conducted in an impedance tube to provide data to aid in the treatment design. Tests with no airflow, involving the measurement of the absorptive properties of the treatment installed in the 9- by 15-foot wind tunnel test section, combined the use of time-delay spectrometry with a previously established free-field measurement method. This new application of time-delay spectrometry enabled these free-field measurements to be made in nonanechoic conditions. The results showed that the installed acoustic treatment had absorption coefficients greater than 0.95 over the frequency range 250 Hz to 4 kHz. The measurements in the wind tunnel were in good agreement with both the analytical prediction and the impedance tube test data.

  11. Investigation of space shuttle orbiter subsonic stability and control characteristics and determination of control surface hinge moments in the Rockwell International low speed wind tunnel (OA37)

    NASA Technical Reports Server (NTRS)

    Hughes, T.

    1974-01-01

    Experimental aerodynamic investigations were conducted on a string-mounted 0.030 scale representation of the 140A/B space shuttle orbiter in the 7.75- by 11-foot low speed wind tunnel. The primary test objectives were to establish basic longitudinal and lateral directional stability and control characteristics for the basic configuration plus control surface hinge moments. Aerodynamic force and moment data were measured in the body axis system by an internally mounted, six-component strain gage balance. Additional configurations investigated were sealed rudder hingeline gaps, sealed elevon gaps and compartmentized speedbrakes.

  12. Low-speed wind-tunnel tests of a one-tenth-scale model of a blended-arrow advanced supersonic transport. [conducted in Langley full-scale tunnel

    NASA Technical Reports Server (NTRS)

    Lemore, H. C.; Parett, L. P.

    1975-01-01

    Tests were conducted in the Langley full scale tunnel to determine the low-speed aerodynamic characteristics of a 1/10 scale model of a blended-arrow advanced supersonic transport. Tests were made for the clean configuration and a high-lift configuration with several combinations of leading- and trailing-edge flaps deflected for providing improved lift and longitudinal stability in the landing and takeoff modes. The tests were conducted for a range of angles of attack from about -6 deg to 30 deg, sideslip angles from -5 deg to 10 deg, and for Reynolds numbers from 6.78 x 1,000,000 to 13.85 x 1,000,000 corresponding to test velocities of 41 knots to 85 knots, respectively.

  13. An Auto-Tuning PI Control System for an Open-Circuit Low-Speed Wind Tunnel Designed for Greenhouse Technology

    PubMed Central

    Espinoza, Karlos; Valera, Diego L.; Torres, José A.; López, Alejandro; Molina-Aiz, Francisco D.

    2015-01-01

    Wind tunnels are a key experimental tool for the analysis of airflow parameters in many fields of application. Despite their great potential impact on agricultural research, few contributions have dealt with the development of automatic control systems for wind tunnels in the field of greenhouse technology. The objective of this paper is to present an automatic control system that provides precision and speed of measurement, as well as efficient data processing in low-speed wind tunnel experiments for greenhouse engineering applications. The system is based on an algorithm that identifies the system model and calculates the optimum PI controller. The validation of the system was performed on a cellulose evaporative cooling pad and on insect-proof screens to assess its response to perturbations. The control system provided an accuracy of <0.06 m·s−1 for airflow speed and <0.50 Pa for pressure drop, thus permitting the reproducibility and standardization of the tests. The proposed control system also incorporates a fully-integrated software unit that manages the tests in terms of airflow speed and pressure drop set points. PMID:26274962

  14. Advanced recovery systems wind tunnel test report

    NASA Technical Reports Server (NTRS)

    Geiger, R. H.; Wailes, W. K.

    1990-01-01

    Pioneer Aerospace Corporation (PAC) conducted parafoil wind tunnel testing in the NASA-Ames 80 by 120 test sections of the National Full-Scale Aerodynamic Complex, Moffett Field, CA. The investigation was conducted to determine the aerodynamic characteristics of two scale ram air wings in support of air drop testing and full scale development of Advanced Recovery Systems for the Next Generation Space Transportation System. Two models were tested during this investigation. Both the primary test article, a 1/9 geometric scale model with wing area of 1200 square feet and secondary test article, a 1/36 geometric scale model with wing area of 300 square feet, had an aspect ratio of 3. The test results show that both models were statically stable about a model reference point at angles of attack from 2 to 10 degrees. The maximum lift-drag ratio varied between 2.9 and 2.4 for increasing wing loading.

  15. Rocket Plume Scaling for Orion Wind Tunnel Testing

    NASA Technical Reports Server (NTRS)

    Brauckmann, Gregory J.; Greathouse, James S.; White, Molly E.

    2011-01-01

    A wind tunnel test program was undertaken to assess the jet interaction effects caused by the various solid rocket motors used on the Orion Launch Abort Vehicle (LAV). These interactions of the external flowfield and the various rocket plumes can cause localized aerodynamic disturbances yielding significant and highly non-linear control amplifications and attenuations. This paper discusses the scaling methodologies used to model the flight plumes in the wind tunnel using cold air as the simulant gas. Comparisons of predicted flight, predicted wind tunnel, and measured wind tunnel forces-and-moments and plume flowfields are made to assess the effectiveness of the selected scaling methodologies.

  16. Wind tunnel test of the S814 thick root airfoil

    SciTech Connect

    Somers, D.M.; Tangler, J.L.

    1996-11-01

    The objective of this wind-tunnel test was to verify the predictions of the Eppler Airfoil Design and Analysis Code for a very thick airfoil having a high maximum lift coefficient designed to be largely insensitive to leading-edge roughness effects. The 24 percent thick S814 airfoil was designed with these characteristics to accommodate aerodynamic and structural considerations for the root region of a wind-turbine blade. In addition, the airfoil`s maximum lift-to-drag ratio was designed to occur at a high lift coefficient. To accomplish the objective, a two-dimensional wind tunnel test of the S814 thick root airfoil was conducted in January 1994 in the low-turbulence wind tunnel of the Delft University of Technology Low Speed Laboratory, The Netherlands. Data were obtained with transition free and transition fixed for Reynolds numbers of 0.7, 1.0, 1.5, 2.0, and 3.0 {times} 10{sup 6}. For the design Reynolds number of 1.5 {times} 10{sup 6}, the maximum lift coefficient with transition free is 1.32, which satisfies the design specification. However, this value is significantly lower than the predicted maximum lift coefficient of almost 1.6. With transition fixed at the leading edge, the maximum lift coefficient is 1.22. The small difference in maximum lift coefficient between the transition-free and transition-fixed conditions demonstrates the airfoil`s minimal sensitivity to roughness effects. The S814 root airfoil was designed to complement existing NREL low maximum-lift-coefficient tip-region airfoils for rotor blades 10 to 15 meters in length.

  17. 40 CFR Table F-2 to Subpart F of... - Particle Sizes and Wind Speeds for Full Wind Tunnel Test, Wind Tunnel Inlet Aspiration Test, and...

    Code of Federal Regulations, 2012 CFR

    2012-07-01

    ... 40 Protection of Environment 6 2012-07-01 2012-07-01 false Particle Sizes and Wind Speeds for Full Wind Tunnel Test, Wind Tunnel Inlet Aspiration Test, and Static Chamber Test F Table F-2 to Subpart F... Part 53—Particle Sizes and Wind Speeds for Full Wind Tunnel Test, Wind Tunnel Inlet Aspiration...

  18. 40 CFR Table F-2 to Subpart F of... - Particle Sizes and Wind Speeds for Full Wind Tunnel Test, Wind Tunnel Inlet Aspiration Test, and...

    Code of Federal Regulations, 2014 CFR

    2014-07-01

    ... 40 Protection of Environment 6 2014-07-01 2014-07-01 false Particle Sizes and Wind Speeds for Full Wind Tunnel Test, Wind Tunnel Inlet Aspiration Test, and Static Chamber Test F Table F-2 to Subpart F... Part 53—Particle Sizes and Wind Speeds for Full Wind Tunnel Test, Wind Tunnel Inlet Aspiration...

  19. 40 CFR Table F-2 to Subpart F of... - Particle Sizes and Wind Speeds for Full Wind Tunnel Test, Wind Tunnel Inlet Aspiration Test, and...

    Code of Federal Regulations, 2010 CFR

    2010-07-01

    ... 40 Protection of Environment 5 2010-07-01 2010-07-01 false Particle Sizes and Wind Speeds for Full Wind Tunnel Test, Wind Tunnel Inlet Aspiration Test, and Static Chamber Test F Table F-2 to Subpart F... Part 53—Particle Sizes and Wind Speeds for Full Wind Tunnel Test, Wind Tunnel Inlet Aspiration...

  20. 40 CFR Table F-2 to Subpart F of... - Particle Sizes and Wind Speeds for Full Wind Tunnel Test, Wind Tunnel Inlet Aspiration Test, and...

    Code of Federal Regulations, 2011 CFR

    2011-07-01

    ... 40 Protection of Environment 5 2011-07-01 2011-07-01 false Particle Sizes and Wind Speeds for Full Wind Tunnel Test, Wind Tunnel Inlet Aspiration Test, and Static Chamber Test F Table F-2 to Subpart F... Part 53—Particle Sizes and Wind Speeds for Full Wind Tunnel Test, Wind Tunnel Inlet Aspiration...

  1. 40 CFR Table F-2 to Subpart F of... - Particle Sizes and Wind Speeds for Full Wind Tunnel Test, Wind Tunnel Inlet Aspiration Test, and...

    Code of Federal Regulations, 2013 CFR

    2013-07-01

    ... 40 Protection of Environment 6 2013-07-01 2013-07-01 false Particle Sizes and Wind Speeds for Full Wind Tunnel Test, Wind Tunnel Inlet Aspiration Test, and Static Chamber Test F Table F-2 to Subpart F... Part 53—Particle Sizes and Wind Speeds for Full Wind Tunnel Test, Wind Tunnel Inlet Aspiration...

  2. Build an Inexpensive Wind Tunnel to Test CO2 Cars

    ERIC Educational Resources Information Center

    McCormick, Kevin

    2012-01-01

    As part of the technology education curriculum, the author's eighth-grade students design, build, test, and race CO2 vehicles. To help them in refining their designs, they use a wind tunnel to test for aerodynamic drag. In this article, the author describes how to build a wind tunnel using inexpensive, readily available materials. (Contains 1…

  3. Low-speed wind tunnel investigation of the stability and control characteristics of a series of flying wings with sweep angles of 60 deg

    NASA Technical Reports Server (NTRS)

    Moul, Thomas M.; Fears, Scott P.; Ross, Holly M.; Foster, John V.

    1995-01-01

    A wind tunnel investigation was conducted in the Langley 12-Foot Low-Speed Wind Tunnel to study the low-speed stability and control characteristics of a series of four flying wings over an extended range of angle of attack (-8 deg to 48 deg). Because of the current emphasis on reducing the radar cross section of new military aircraft, the planform of each wing was composed of lines swept at a relatively high angle of 60 deg, and all the trailing-edge lines were aligned with one of the two leading edges. Three arrow planforms with different aspect ratios and one diamond planform were tested. The models incorporated leading-edge flaps for improved pitching-moment characteristics and lateral stability and had three sets of trailing-edge flaps that were deflected differentially for roll control, symmetrically for pitch control, and in a split fashion for yaw control. Top bodies of three widths and twin vertical tails of various sizes and locations were also tested on each model. A large aerodynamic database was compiled that could be used to evaluate some of the trade-offs involved in the design of a configuration with a reduced radar cross section and good flight dynamic characteristics.

  4. Shuttle derived launch vehicle wind tunnel tests

    NASA Technical Reports Server (NTRS)

    Tewell, J. R.; Buell, D. N.

    1985-01-01

    Studies are being conducted regarding new launch vehicle configurations which may effectively and economically share the delivery of payloads to orbit with the present Space Transportation System (STS). The role envisaged for these launch vehicles is related to the execution of missions whose requirements exceed the STS Shuttle capabilities, taking into account the delivery of much heavier or larger payloads. One class of advanced launch vehicles is configured to take advantage of the existing Shuttle hardware and facilities. Such vehicles are referred to as Shuttle Derived Vehicles (SDV). One version of an SDV consists of two STS elements, including the external tank (ET) and solid rocket boosters, and a cargo carrier. Attention is given to wind tunnel tests, which are being conducted with SDV sidemount configurations incorporating various size payload modules.

  5. Integral method of wall interference correction in low-speed wind tunnels

    NASA Technical Reports Server (NTRS)

    Zhou, Changhai

    1987-01-01

    The analytical solution of Poisson's equation, derived form the definition of vortex, was applied to the calculation of interference velocities due to the presence of wind tunnel walls. This approach, called the Integral Method, allows an accurate evaluation of wall interference for separated or more complicated flows without the need for considering any features of the model. All the information necessary for obtaining the wall correction is contained in wall pressure measurements. The correction is not sensitive to normal data-scatter, and the computations are fast enough for on-line data processing.

  6. 5. VIEW NORTH OF TEST SECTION IN FULLSCALE WIND TUNNEL ...

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

    5. VIEW NORTH OF TEST SECTION IN FULL-SCALE WIND TUNNEL WITH FREE-FLIGHT MODEL OF A BOEING 737 SUSPENDED FROM A SAFETY CABLE. - NASA Langley Research Center, Full-Scale Wind Tunnel, 224 Hunting Avenue, Hampton, Hampton, VA

  7. Modular asymmetric parachute for wind-tunnel testing

    SciTech Connect

    Klimas, P.C.; Widdows, H.E.; Croll, R.H.

    1981-01-01

    The construction of a series of asymmetrical wind tunnel model parachutes designed to a modular concept is described. The static force, inflation force, and dynamic force and motion time history wind tunnel testing of up to 123 different configurations is discussed.

  8. Short Takeoff and Vertical Landing Capability Upgraded in NASA Glenn's 9- by 15-Foot Low-Speed Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Stark, David E.

    2003-01-01

    The NASA Glenn Research Center supports short takeoff and vertical landing (STOVL) tests in its 9- by 15-Foot Low Speed Wind Tunnel (9 x 15 LSWT). As part of a facility capability upgrade, a dynamic actuation system (DAS) was fabricated to enhance the STOVL testing capabilities. The DAS serves as the mechanical interface between the 9 x 15 LSWT test section structure and the STOVL model to be tested. It provides vertical and horizontal translation of the model in the test section and maintains the model attitude (pitch, yaw, and roll) during translation. It also integrates a piping system to supply the model with exhaust and hot air to simulate the inlet suction and nozzle exhausts, respectively. Hot gas ingestion studies have been performed with the facility ground plane installed. The DAS provides vertical (ascent and descent) translation speeds of up to 48 in./s and horizontal translation speeds of up to 12 in./s. Model pitch variations of +/- 7, roll variations of +/- 5, and yaw variations of 0 to 180 deg can be accommodated and are maintained within 0.25 deg throughout the translation profile. The hot air supply, generated by the facility heaters and regulated by control valves, provides three separate temperature zones to the model for STOVL and hot gas ingestion testing. Channels along the supertube provide instrumentation paths from the model to the facility data system for data collection purposes. The DAS is supported by the 9 x 15 LSWT test section ceiling structure. A carriage that rides on two linear rails provides for horizontal translation of the system along the test section longitudinal axis. A vertical translation assembly, consisting of a cage and supertube, is secured to the carriage. The supertube traverses vertically through the cage on a set of linear rails. Both translation axes are hydraulically actuated and provide position and velocity profile control. The lower flange on the supertube serves as the model interface to the DAS. The

  9. Application of the adaptive-wall concept to three-dimensional low-speed wind tunnels

    NASA Technical Reports Server (NTRS)

    Erickson, J. C., Jr.

    1976-01-01

    Three methods for evaluating the functional relationships required to obtain interference-free flows about a model in a wind tunnel have been developed. The first, the original multipole expansion (MPE) procedure, is based on a series of point singularities which satisfy the governing Prandtl-Glauert equation. The second, the modified MPE, provides an improved representation of finite-span wings and thereby extends the range of validity of the original MPE to larger ratios of span-to-control-surface-width. The third method is more general and is based on source distributions over the control surface. Several numerical examples are presented to help establish the range of validity of these methods. An accuracy-assessment procedure, which combines the original MPE procedure with classical wall-correction theory, has been developed to estimate the degree of interference at the model if the functional relationships are not satisfied exactly. Several numerical examples are presented for representative wings and bodies.

  10. A low-speed wind tunnel study of vortex interaction control techniques on a chine-forebody/delta-wing configuration

    NASA Technical Reports Server (NTRS)

    Rao, Dhanvada M.; Bhat, M. K.

    1992-01-01

    A low speed wind tunnel evaluation was conducted of passive and active techniques proposed as a means to impede the interaction of forebody chine and delta wing vortices, when such interaction leads to undesirable aerodynamic characteristics particularly in the post stall regime. The passive method was based on physically disconnecting the chine/wing junction; the active technique employed deflection of inboard leading edge flaps. In either case, the intent was to forcibly shed the chine vortices before they encountered the downwash of wing vortices. Flow visualizations, wing pressures, and six component force/moment measurements confirmed the benefits of forced vortex de-coupling at post stall angles of attack and in sideslip, viz., alleviation of post stall zero beta asymmetry, lateral instability and twin tail buffet, with insignificant loss of maximum lift.

  11. Initial Investigation of the Acoustics of a Counter-Rotating Open Rotor Model with Historical Baseline Blades in a Low-Speed Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Elliott, David M.

    2012-01-01

    A counter-rotating open rotor scale model was tested in the NASA Glenn Research Center 9- by 15-Foot Low-Speed Wind Tunnel (LSWT). This model used a historical baseline blade set with which modern blade designs will be compared against on an acoustic and aerodynamic performance basis. Different blade pitch angles simulating approach and takeoff conditions were tested, along with angle-of-attack configurations. A configuration was also tested in order to determine the acoustic effects of a pylon. The shaft speed was varied for each configuration in order to get data over a range of operability. The freestream Mach number was also varied for some configurations. Sideline acoustic data were taken for each of these test configurations.

  12. Low-speed Wind-Tunnel Study of Reaction Control-jet Effectiveness for Hover and Transition of a STOVL Fighter Concept

    NASA Technical Reports Server (NTRS)

    Riley, Donald R.; Shah, Gautam H.; Kuhn, Richard E.

    1989-01-01

    A brief wind-tunnel study was conducted in the Langley 12-Foot Low-Speed Tunnel to determine reaction control-jet effectiveness and some associated aerodynamic characteristics of a 15 percent scale model of the General Dynamics E-7A STOVL fighter/attack aircraft concept applicable to hover and transition flight. Tests were made with the model at various attitude angles in the tunnel test section and at various tunnel airspeeds for a range of control-jet nozzle pressure ratios. Eight reaction control-jets were tested individually. Four jets were at the design baseline locations providing roll, pitch, and yaw control. Comparisons of measured data with values calculated using empirical methods were made where possible.

  13. NASA Langley Low Speed Aeroacoustic Wind Tunnel: Background Noise and Flow Survey Results Prior to FY05 Construction of Facilities Modifications

    NASA Technical Reports Server (NTRS)

    Booth, Earl R., Jr.; Henderson, Brenda S.

    2005-01-01

    The NASA Langley Research Center Low Speed Aeroacoustic Wind Tunnel is a premier facility for model-scale testing of jet noise reduction concepts at realistic flow conditions. However, flow inside the open jet test section is less than optimum. A Construction of Facilities project, scheduled for FY 05, will replace the flow collector with a new design intended to reduce recirculation in the open jet test section. The reduction of recirculation will reduce background noise levels measured by a microphone array impinged by the recirculation flow and will improve flow characteristics in the open jet tunnel flow. In order to assess the degree to which this modification is successful, background noise levels and tunnel flow are documented, in order to establish a baseline, in this report.

  14. Planar Doppler Velocimetry for Large-Scale Wind Tunnel Testing

    NASA Technical Reports Server (NTRS)

    McKenzie, Robert L.

    1997-01-01

    Recently, Planar Doppler Velocimetry (PDV) has been shown by several laboratories to offer an attractive means for measuring three-dimensional velocity vectors everywhere in a light sheet placed in a flow. Unlike other optical means of measuring flow velocities, PDV is particularly attractive for use in large wind tunnels where distances to the sample region may be several meters, because it does not require the spatial resolution and tracking of individual scattering particles or the alignment of crossed beams at large distances. To date, demonstrations of PDV have been made either in low speed flows without quantitative comparison to other measurements, or in supersonic flows where the Doppler shift is large and its measurement is relatively insensitive to instrumental errors. Moreover, most reported applications have relied on the use of continuous-wave lasers, which limit the measurement to time-averaged velocity fields. This work summarizes the results of two previous studies of PDV in which the use of pulsed lasers to obtain instantaneous velocity vector fields is evaluated. The objective has been to quantitatively define and demonstrate PDV capabilities for applications in large-scale wind tunnels that are intended primarily for the production testing of subsonic aircraft. For such applications, the adequate resolution of low-speed flow fields requires accurate measurements of small Doppler shifts that are obtained at distances of several meters from the sample region. The use of pulsed lasers provides the unique capability to obtain not only time-averaged fields, but also their statistical fluctuation amplitudes and the spatial excursions of unsteady flow regions such as wakes and separations. To accomplish the objectives indicated, the PDV measurement process is first modeled and its performance evaluated computationally. The noise sources considered include those related to the optical and electronic properties of Charge-Coupled Device (CCD) arrays and to

  15. Aeroservoelastic Wind-Tunnel Test of the SUGAR Truss Braced Wing Wind-Tunnel Model

    NASA Technical Reports Server (NTRS)

    Scott, Robert C.; Allen, Timothy J.; Funk, Christie J.; Castelluccio, Mark A.; Sexton, Bradley W.; Claggett, Scott; Dykman, John; Coulson, David A.; Bartels, Robert E.

    2015-01-01

    The Subsonic Ultra Green Aircraft Research (SUGAR) Truss-Braced Wing (TBW) aeroservoelastic (ASE) wind-tunnel test was conducted in the NASA Langley Transonic Dynamics Tunnel (TDT) and was completed in April, 2014. The primary goals of the test were to identify the open-loop flutter boundary and then demonstrate flutter suppression. A secondary goal was to demonstrate gust load alleviation (GLA). Open-loop flutter and limit cycle oscillation onset boundaries were identified for a range of Mach numbers and various angles of attack. Two sets of control laws were designed for the model and both sets of control laws were successful in suppressing flutter. Control laws optimized for GLA were not designed; however, the flutter suppression control laws were assessed using the TDT Airstream Oscillation System. This paper describes the experimental apparatus, procedures, and results of the TBW wind-tunnel test. Acquired system ID data used to generate ASE models is also discussed.2 study.

  16. Results of investigations on a 0.0405 scale model PRR version of the NR-SSV orbiter in the North American Aeronautical Laboratory low speed wind tunnel

    NASA Technical Reports Server (NTRS)

    Kingsland, R. B.; Vaughn, J. E.; Singellton, R.

    1973-01-01

    Experimental aerodynamic investigations were conducted in a low speed wind tunnel on a scale model space shuttle vehicle (SSV) orbiter. The purpose of the test was to investigate the longitudinal and lateral-directional aerodynamic characteristics of the space shuttle orbiter. Emphasis was placed on model component, wing-glove, and wing-body fairing effects, as well as elevon, aileron, and rudder control effectiveness. Angles of attack from - 5 deg to + 30 deg and angles of sideslip of - 5 deg, 0 deg, and + 5 deg were tested. Static pressures were recorded on base, fuselage, and wing surfaces. Tufts and talc-kerosene flow visualization techniques were also utilized. The aerodynamic force balance results are presented in plotted and tabular form.

  17. Space shuttle phase B wind tunnel test database

    NASA Technical Reports Server (NTRS)

    Glynn, J. L.; Poucher, D. E.

    1988-01-01

    Archived wind tunnel test data are available for flyback booster or other alternate recoverable configurations as well as reusable orbiters studied during initial development (Phase B) of the Space Shuttle. Considerable wind tunnel data were acquired by competing contractors and NASA centers for an extensive variety of configurations with an array of wing and body planforms. This wind tunnel test data has been compiled into a database and are available for application to current winged flyback or recoverable booster aerodynamic studies. The Space Shuttle Phase B Wind Tunnel Database is structured by vehicle component and configuration type. Basic components include the booster, the orbiter and the launch vehicle. Booster configuration types include straight and delta wings, canard, cylindrical, retro-glide and twin body. Orbiter configuration types include straight and delta wings, lifting body, drop tanks and double delta wings.

  18. SAMPSON smart inlet design overview and wind tunnel test: II. Wind tunnel test

    NASA Astrophysics Data System (ADS)

    Pitt, Dale M.; Dunne, James P.; White, Edward V.

    2002-07-01

    The Smart Aircraft and Marine System Projects Demonstration (SAMPSON) program was a DARPA funded effort conducted by the Boeing Company, General Dynamics - Electric Boat Division, and the Pennsylvania State University. NASA Langley Research Center (NASA LaRC) was technical monitor for the aircraft demonstration, while the Navy's Office of Naval Research (ONR) was technical monitor for the marine demonstration. Dr. Ephrahim Garcia, DARPA/DSO, acted as the DARPA program manager for SAMPSON. The SAMPSON program objectives were to demonstrate smart structures based systems on large/full scale structures in realistic environments. The SAMPSON aircraft demonstration was the wind tunnel testing of a full scale F-15 aircraft inlet that was capable of in-flight structural variations accomplished using smart materials, called the 'SAMPSON Smart Inlet'. The SAMPSON Smart Inlet was removed from an F-15E airframe and structurally modified to interface with the NASA LaRC 16-Foot Transonic Tunnel model support system. This is Part II of two works documenting the SAMPSON Smart Inlet design and testing. A discussion of the two wind tunnel tests will be presented here in Part II. The design of the shape changing components of the Smart Inlet is presented in a separate work, Part I.

  19. Low-speed wind tunnel investigation of a semispan STOL jet transport wing body with an upper surface blown jet flap

    NASA Technical Reports Server (NTRS)

    Phelps, A. E., III; Letko, W.; Henderson, R. L.

    1973-01-01

    An investigation of the static longitudinal aerodynamic characteristics of a semispan STOL jet transport wing-body with an upper-surface blown jet flap for lift augmentation was conducted in a low-speed wind tunnel having a 12-ft octagonal test section. The semispan swept wing had an aspect ratio of 3.92 (7.84 for the full span) and had two simulated turbofan engines mounted ahead of and above the wing in a siamese pod equipped with an exhaust deflector. The purpose of the deflector was to spread the engine exhaust into a jet sheet attached to the upper surface of the wing so that it would turn downward over the flap and provide lift augmentation. The wing also had optional boundary-layer control provided by air blowing through a thin slot over a full-span plain trailing-edge flap.

  20. Photogrammetry Applied to Wind Tunnel Testing

    NASA Technical Reports Server (NTRS)

    Liu, Tian-Shu; Cattafesta, L. N., III; Radeztsky, R. H.; Burner, A. W.

    2000-01-01

    In image-based measurements, quantitative image data must be mapped to three-dimensional object space. Analytical photogrammetric methods, which may be used to accomplish this task, are discussed from the viewpoint of experimental fluid dynamicists. The Direct Linear Transformation (DLT) for camera calibration, used in pressure sensitive paint, is summarized. An optimization method for camera calibration is developed that can be used to determine the camera calibration parameters, including those describing lens distortion, from a single image. Combined with the DLT method, this method allows a rapid and comprehensive in-situ camera calibration and therefore is particularly useful for quantitative flow visualization and other measurements such as model attitude and deformation in production wind tunnels. The paper also includes a brief description of typical photogrammetric applications to temperature- and pressure-sensitive paint measurements and model deformation measurements in wind tunnels.

  1. 40 CFR 53.62 - Test procedure: Full wind tunnel test.

    Code of Federal Regulations, 2013 CFR

    2013-07-01

    ... 40 Protection of Environment 6 2013-07-01 2013-07-01 false Test procedure: Full wind tunnel test... Performance Characteristics of Class II Equivalent Methods for PM 2.5 § 53.62 Test procedure: Full wind tunnel test. (a) Overview. The full wind tunnel test evaluates the effectiveness of the candidate sampler at...

  2. 40 CFR 53.62 - Test procedure: Full wind tunnel test.

    Code of Federal Regulations, 2010 CFR

    2010-07-01

    ... 40 Protection of Environment 5 2010-07-01 2010-07-01 false Test procedure: Full wind tunnel test... Performance Characteristics of Class II Equivalent Methods for PM2.5 § 53.62 Test procedure: Full wind tunnel test. (a) Overview. The full wind tunnel test evaluates the effectiveness of the candidate sampler at...

  3. Low-speed wind-tunnel investigation of the aerodynamic and acoustic performance of a translating grid choked flow inlet

    NASA Technical Reports Server (NTRS)

    Abbott, J. M.; Miller, B. A.; Golladay, R. L.

    1974-01-01

    The aerodynamic and acoustic performance of a translating grid choked-flow inlet was determined in a low-speed wind tunnel at free-stream velocities of 24, 32, and 45 m/sec and incidence angles of 0, 10, 20, 30, 35, 40, 45, and 50 deg. The inlet was sized to fit a 13.97- centimeter-diameter fan with a design weight flow of 2.49 kg/sec. Measurements were made to determine inlet total pressure recovery, flow distortion, and sound pressure level for both choked and unchoked geometries over a range of inlet weight flows. For the unchoked geometry, inlet total pressure recovery ranged from 0.983 to 0.989 at incidence angles less than 40 deg. At 40 deg incidence angle, inlet cowl separation was encountered which resulted in lower values of pressure recovery and higher levels of fan broadband noise. For the choked geometry, increasing total pressure losses occurred with increasing inlet weight flow that prevented the inlet from reaching full choked conditions with the particular fan used. These losses were attributed to the high Mach number drag rise characteristics of airfoil grid. At maximum attainable inlet weight flow, the total pressure recovery at static conditions was 0.935. The fan blade passing frequency and other fan generated pure tones were eliminated from the noise spectrum, but the broadband level was increased.

  4. Low-speed wind-tunnel investigation of the stability and control characteristics of a series of flying wings with sweep angles of 70 deg

    NASA Technical Reports Server (NTRS)

    Ross, Holly M.; Fears, Scott P.; Moul, Thomas M.

    1995-01-01

    A wind-tunnel investigation was conducted in the Langley 12-Foot Low-Speed Tunnel to study the low-speed stability and control characteristics of a series of four flying wings over an extended range of angle of attack (-8 deg to 48 deg). Because of the current emphasis on reducing the radar cross section (RCS) of new military aircraft, the planform of each wing was composed of lines swept at a relatively high angle of 70 deg, and all the trailing edges and control surface hinge lines were aligned with one of the two leading edges. Three arrow planforms with different aspect ratios and one diamond planform were tested. The models incorporated leading-edge flaps for improved longitudinal characteristics and lateral stability and had three sets of trailing-edge flaps that were deflected differentially for roll control, symmetrically for pitch control, and in a split fashion for yaw control. Three top body widths and two sizes of twin vertical tails were also tested on each model. A large aerodynamic database was compiled that could be used to evaluate some of the trade-offs involved in the design of a configuration with a reduced RCS and good flight dynamic characteristics.

  5. Wind tunnel tests of high-lift systems for advanced transports using high-aspect-ratio supercritical wings

    NASA Technical Reports Server (NTRS)

    Allen, J. B.; Oliver, W. R.; Spacht, L. A.

    1982-01-01

    The wind tunnel testing of an advanced technology high lift system for a wide body and a narrow body transport incorporating high aspect ratio supercritical wings is described. This testing has added to the very limited low speed high Reynolds number data base for this class or aircraft. The experimental results include the effects on low speed aerodynamic characteristics of various leading and trailing edge devices, nacelles and pylons, ailerons, and spoilers, and the effects of Mach and Reynolds numbers.

  6. Design techniques for developing a computerized instrumentation test plan. [for wind tunnel test data acquisition system

    NASA Technical Reports Server (NTRS)

    Burnett, S. Kay; Forsyth, Theodore J.; Maynard, Everett E.

    1987-01-01

    The development of a computerized instrumentation test plan (ITP) for the NASA/Ames Research Center National Full Scale Aerodynamics Complex (NFAC) is discussed. The objective of the ITP program was to aid the instrumentation engineer in documenting the configuration and calibration of data acquisition systems for a given test at any of four low speed wind tunnel facilities (Outdoor Aerodynamic Research Facility, 7 x 10, 40 x 80, and 80 x 120) at the NFAC. It is noted that automation of the ITP has decreased errors, engineering hours, and setup time while adding a higher level of consistency and traceability.

  7. SMART Rotor Development and Wind-Tunnel Test

    NASA Technical Reports Server (NTRS)

    Lau, Benton H.; Straub, Friedrich; Anand, V. R.; Birchette, Terry

    2009-01-01

    Boeing and a team from Air Force, NASA, Army, Massachusetts Institute of Technology, University of California at Los Angeles, and University of Maryland have successfully completed a wind-tunnel test of the smart material actuated rotor technology (SMART) rotor in the 40- by 80-foot wind-tunnel of the National Full-Scale Aerodynamic Complex at NASA Ames Research Center, figure 1. The SMART rotor is a full-scale, five-bladed bearingless MD 900 helicopter rotor modified with a piezoelectric-actuated trailing-edge flap on each blade. The development effort included design, fabrication, and component testing of the rotor blades, the trailing-edge flaps, the piezoelectric actuators, the switching power amplifiers, the actuator control system, and the data/power system. Development of the smart rotor culminated in a whirl-tower hover test which demonstrated the functionality, robustness, and required authority of the active flap system. The eleven-week wind tunnel test program evaluated the forward flight characteristics of the active-flap rotor, gathered data to validate state-of-the-art codes for rotor noise analysis, and quantified the effects of open- and closed-loop active-flap control on rotor loads, noise, and performance. The test demonstrated on-blade smart material control of flaps on a full-scale rotor for the first time in a wind tunnel. The effectiveness and the reliability of the flap actuation system were successfully demonstrated in more than 60 hours of wind-tunnel testing. The data acquired and lessons learned will be instrumental in maturing this technology and transitioning it into production. The development effort, test hardware, wind-tunnel test program, and test results will be presented in the full paper.

  8. Experience with scale effects in non-airplane wind tunnel testing

    NASA Technical Reports Server (NTRS)

    Ross, J. C.; Olson, M. E.

    1990-01-01

    The aerodynamics results of two tests performed in the 80- by 120-Foot Wind Tunnel at NASA Ames Research Center are discussed with particular emphasis on the effects of model scale. The tests are unusual for this facility in that they were performed on non-airplane configurations: a full-scale tractor/trailer and large ramair inflated wings. For the truck drag measurements, comparisons with 1/8th-scale drag data taken at the Low Speed Wind Tunnel at Texas A&M indicate that small scale measurements can provide adequate accuracy if care is taken to test at high enough Reynolds numbers and if large regions of separated flow and reattachment are avoided. Some of the important aerodynamic and structural aspects of parafoil testing are also discussed. These include the effects of Reynolds number and aeroelastic effects such as fabric and support line stretch.

  9. Ares I Aerodynamic Testing at the Boeing Polysonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Pinier, Jeremy T.; Niskey, Charles J.; Hanke, Jeremy L.; Tomek, William G.

    2011-01-01

    Throughout three full design analysis cycles, the Ares I project within the Constellation program has consistently relied on the Boeing Polysonic Wind Tunnel (PSWT) for aerodynamic testing of the subsonic, transonic and supersonic portions of the atmospheric flight envelope (Mach=0.5 to 4.5). Each design cycle required the development of aerodynamic databases for the 6 degree-of-freedom (DOF) forces and moments, as well as distributed line-loads databases covering the full range of Mach number, total angle-of-attack, and aerodynamic roll angle. The high fidelity data collected in this facility has been consistent with the data collected in NASA Langley s Unitary Plan Wind Tunnel (UPWT) at the overlapping condition ofMach=1.6. Much insight into the aerodynamic behavior of the launch vehicle during all phases of flight was gained through wind tunnel testing. Important knowledge pertaining to slender launch vehicle aerodynamics in particular was accumulated. In conducting these wind tunnel tests and developing experimental aerodynamic databases, some challenges were encountered and are reported as lessons learned in this paper for the benefit of future crew launch vehicle aerodynamic developments.

  10. Low speed wind tunnel flow field results for JT8D refan engines on the Boeing 727-200

    NASA Technical Reports Server (NTRS)

    Easterbrook, W. G.; Roberts, W. H.

    1974-01-01

    Low speed flow angularity results are presented showing flow direction at the nacelle locations on the Boeing 727-200. Flow angle probes (yawheads) were used for measurements at side and center inlet positions on the aft fuselage. A range of flap settings were tested with flap angles of 0 deg, 15 deg, and 40 deg selected for investigation.

  11. SOFIA 2 model telescope wind tunnel test report

    NASA Technical Reports Server (NTRS)

    Keas, Paul

    1995-01-01

    This document outlines the tests performed to make aerodynamic force and torque measurements on the SOFIA wind tunnel model telescope. These tests were performed during the SOFIA 2 wind tunnel test in the 14 ft wind tunnel during the months of June through August 1994. The test was designed to measure the dynamic cross elevation moment acting on the SOFIA model telescope due to aerodynamic loading. The measurements were taken with the telescope mounted in an open cavity in the tail section of the SOFIA model 747. The purpose of the test was to obtain an estimate of the full scale aerodynamic disturbance spectrum, by scaling up the wind tunnel results (taking into account differences in sail area, air density, cavity dimension, etc.). An estimate of the full scale cross elevation moment spectrum was needed to help determine the impact this disturbance would have on the telescope positioning system requirements. A model of the telescope structure, made of a light weight composite material, was mounted in the open cavity of the SOFIA wind tunnel model. This model was mounted via a force balance to the cavity bulkhead. Despite efforts to use a 'stiff' balance, and a lightweight model, the balance/telescope system had a very low resonant frequency (37 Hz) compared to the desired measurement bandwidth (1000 Hz). Due to this mechanical resonance of the balance/telescope system, the balance alone could not provide an accurate measure of applied aerodynamic force at the high frequencies desired. A method of measurement was developed that incorporated accelerometers in addition to the balance signal, to calculate the aerodynamic force.

  12. Noise measurements from an ejector suppressor nozzle in the NASA Lewis 9- by 15-foot low speed wind tunnel

    NASA Technical Reports Server (NTRS)

    Krejsa, Eugene A.; Cooper, Beth A.; Hall, David G.; Khavaran, Abbas

    1990-01-01

    Acoustic results are presented of a cooperative nozzle test program between NASA and Pratt and Whitney, conducted in the NASA-Lewis 9 x 15 ft Anechoic Wind Tunnel. The nozzle tested was the P and W Hypermix Nozzle concept, a 2-D lobed mixer nozzle followed by a short ejector section made to promote rapid mixing of the induced ejector nozzle flow. Acoustic and aerodynamic measurements were made to determine the amount of ejector pumping, degree of mixing, and noise reduction achieved. A series of tests were run to verify the acoustic quality of this tunnel. The results indicated that the tunnel test section is reasonably anechoic but that background noise can limit the amount of suppression observed from suppressor nozzles. Also, a possible internal noise was observed in the air supply system. The P and W ejector suppressor nozzle demonstrated the potential of this concept to significantly reduce jet noise. Significant reduction in low frequency noise was achieved by increasing the peak jet noise frequency. This was accomplished by breaking the jet into segments with smaller dimensions than those of the baseline nozzle. Variations in ejector parameters had little effect on the noise for the geometries and the range of temperatures and pressure ratios tested.

  13. Acoustic Performance of the GEAE UPS Research Fan in the NASA Glenn 9- by 15-Foot Low-Speed Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Woodward, Richard P.; Hughes, Christopher E.

    2012-01-01

    A model advanced turbofan was acoustically tested in the NASA Glenn 9- by 15-Foot Low-Speed Wind Tunnel in 1994. The Universal Propulsion Simulator fan was designed and manufactured by General Electric Aircraft Engines, and included an active core, as well as bypass, flow paths. The fan was tested with several rotors featuring unswept, forward-swept and aft-swept designs of both metal and composite construction. Sideline acoustic data were taken with both hard and acoustically treated walls in the flow passages. The fan was tested within an airflow at a Mach number of 0.20, which is representative of aircraft takeoff/approach conditions. All rotors showed similar aerodynamic performance. However, the composite rotors typically showed higher noise levels than did corresponding metal rotors. Aft and forward rotor sweep showed at most modest reductions of transonic multiple pure tone levels. However, rotor sweep often introduced increased rotor-stator interaction tone levels. Broadband noise was typically higher for the composite rotors and also for the aft-swept metal rotor. Transonic MPT generation was reduced with increasing fan axis angle of attack (AOA); however, higher downstream noise levels did increase with AOA resulting in higher overall Effective Perceived Noise Level.

  14. Continued investigations in the NAAL low speed wind tunnel into the effects of the air breathing propulsion system on orbiter subsonic stability and control characteristics (OA62A)

    NASA Technical Reports Server (NTRS)

    Mennell, R.

    1974-01-01

    Experimental aerodynamic investigations were conducted on a stingmounted 0.0405-scale representation (model 43-0) of the 140A/B Space Shuttle Orbiter in a Low Speed Wind Tunnel. The NASA designation for this test was 0A62A. The primary test objective was to continue studies, initiated on tests 0A16 and 0A71A and 0A71C, in optimizing the air breathing propulsion system (ABPS) and investigating the aerodynamic effects of various nacelle number/location configurations on the orbiter stability and control characteristics. Orbiter stability and control characteristics, both with and without ABPS, were investigated at elevon deflections of 0, + or -5, + or -19, + or -5, and -20 deg; aileron deflections of 0 and 10 deg (about 0 deg elevon); and rudder deflections of 0, -7.5, and -15 deg. Aerodynamic force and moment data was measured in the body axis system by a 2.5-inch task type internal balance. The model was sting supported through the base region with a nominal angle of attack range of -4 to 30 deg. Yaw polars were recorded over the beta range of -10 to 10 deg at fixed angles of attack of 0, 5, 10, and 15 deg.

  15. Low-speed wind-tunnel investigation of a porous forebody and nose strakes for yaw control of a multirole fighter aircraft

    NASA Technical Reports Server (NTRS)

    Fears, Scott P.

    1995-01-01

    Low-speed wind-tunnel tests were conducted in the Langley 12-Foot Low-Speed Tunnel on a model of the Boeing Multirole Fighter (BMRF) aircraft. This single-seat, single-engine configuration was intended to be an F-16 replacement that would incorporate many of the design goals and advanced technologies of the F-22. Its mission requirements included supersonic cruise without afterburner, reduced observability, and the ability to attack both air-to-air and air-to-ground targets. So that it would be effective in all phases of air combat, the ability to maneuver at angles of attack up to and beyond maximum lift was also desired. Traditional aerodynamic yaw controls, such as rudders, are typically ineffective at these higher angles of attack because they are usually located in the wake from the wings and fuselage. For this reason, this study focused on investigating forebody-mounted controls that produces yawing moments by modifying the strong vortex flowfield being shed from the forebody at high angles of attack. Two forebody strakes were tested that varied in planform and chordwise location. Various patterns of porosity in the forebody skin were also tested that differed in their radial coverage and chordwise location. The tests were performed at a dynamic pressure of 4 lb/ft(exp 2) over an angle-of-attack range of -4 deg to 72 deg and a sideslip range of -10 deg to 10 deg. Static force data, static pressures on the surface of the forebody, and videotapes of flow-visualization using laser-illuminated smoke were obtained.

  16. A New Method of Testing in Wind Tunnels

    NASA Technical Reports Server (NTRS)

    Margoulis, W

    1921-01-01

    Now, in existing wind tunnels, using a horsepower of 100 to 300, the models are generally made to a 1/10 scale and the speed is appreciably lower than the speeds currently attained by airplanes. The Reynolds number realized is thus 15 to 25 times smaller than that reached by airplanes in free flight, while the ratio of speed to the velocity of sound is between a third and three quarters of the true ratio. The necessary increases in either the diameter of the wind tunnel or the velocity of the airstream are too costly. However, the author shows that it is possible to have wind tunnels in which the Reynolds number will be greater than that now obtained by airplanes, and in which the ratio of the velocity to the velocity of sound will also be greater than that realized in practice, by employing a gas other than air, at a pressure and temperature different from those of the surrounding atmosphere. The gas is carbonic acid, a gas having a low coefficient of viscosity, high density, and a low ratio of specific heat. The positive results of using carbonic acid in wind tunnel tests are given.

  17. Investigation of space shuttle orbiter subsonic stability and control characteristics in the NAAL low speed wind tunnel (0A62b), volume 1

    NASA Technical Reports Server (NTRS)

    Mennell, R.; Hughes, T.

    1974-01-01

    Experimental aerodynamic investigations were conducted on a sting-mounted 0.0405 scale representation of the 140A/B space shuttle orbiter in a 7.75 ft by 11 ft low speed wind tunnel during the time period from November 14, 1973, to December 6, 1973, with the primary test objectives being to establish basic longitudinal stability characteristics in and out of ground effect, as well as lateral-directional stability characteristics in free air. Two dual podded nacelle configurations were also tested, one with three dual podded nacelles on the lower wing surface, and the other with a single dual nacelle on the lower centerline with dual nacelle pylons mounted above each wing. Stability and control characteristics were investigated at nominal elevon, rudder, aileron, and body flap deflections. Pressure bugs were used to determine pressures on the vertical tail at spanwise stations, and aerodynamic force and moment data were measured in the stability axis system by an internally mounted, six component strain gage balance.

  18. Blended-Wing-Body Low-Speed Flight Dynamics: Summary of Ground Tests and Sample Results

    NASA Technical Reports Server (NTRS)

    Vicroy, Dan D.

    2009-01-01

    A series of low-speed wind tunnel tests of a Blended-Wing-Body tri-jet configuration to evaluate the low-speed static and dynamic stability and control characteristics over the full envelope of angle of attack and sideslip are summarized. These data were collected for use in simulation studies of the edge-of-the-envelope and potential out-of-control flight characteristics. Some selected results with lessons learned are presented.

  19. Low-speed wind tunnel investigation of the aerodynamic and acoustic performance of a translating-centerbody choked-flow inlet

    NASA Technical Reports Server (NTRS)

    Miller, B. A.; Abbott, J. M.

    1973-01-01

    Low-speed wind-tunnel tests were conducted to determine the effects of free-stream velocity and incidence angle on the aerodynamic and acoustic performance of a translating centerbody choked-flow inlet. The inlet was sized to fit a 13.97 cm diameter fan with a design weight flow of 2.49 kg/sec. Performance was determined at free-stream velocities to 45 meters per second and incidence angles of 0 deg to 50 deg. The inlet was operated in both the choked and unchoked modes over a range of weight flows. Measurements were made of inlet total pressure recovery, flow distortion, surface static pressure distribution, and fan noise suppression. In the choked mode, increasing incidence angle tended to reduce the amount of inlet noise suppression for a given amount of inlet suction. This tendency was overcome by applying sufficient inlet suction to increase the flow Mach number. At 45 meters per second free-stream velocity, at least 22 decibels of suppression were measured at 35 deg incidence angle with a total pressure recovery of 0.985.

  20. Pre-Test Assessment of the Use Envelope of the Normal Force of a Wind Tunnel Strain-Gage Balance

    NASA Technical Reports Server (NTRS)

    Ulbrich, N.

    2016-01-01

    The relationship between the aerodynamic lift force generated by a wind tunnel model, the model weight, and the measured normal force of a strain-gage balance is investigated to better understand the expected use envelope of the normal force during a wind tunnel test. First, the fundamental relationship between normal force, model weight, lift curve slope, model reference area, dynamic pressure, and angle of attack is derived. Then, based on this fundamental relationship, the use envelope of a balance is examined for four typical wind tunnel test cases. The first case looks at the use envelope of the normal force during the test of a light wind tunnel model at high subsonic Mach numbers. The second case examines the use envelope of the normal force during the test of a heavy wind tunnel model in an atmospheric low-speed facility. The third case reviews the use envelope of the normal force during the test of a floor-mounted semi-span model. The fourth case discusses the normal force characteristics during the test of a rotated full-span model. The wind tunnel model's lift-to-weight ratio is introduced as a new parameter that may be used for a quick pre-test assessment of the use envelope of the normal force of a balance. The parameter is derived as a function of the lift coefficient, the dimensionless dynamic pressure, and the dimensionless model weight. Lower and upper bounds of the use envelope of a balance are defined using the model's lift-to-weight ratio. Finally, data from a pressurized wind tunnel is used to illustrate both application and interpretation of the model's lift-to-weight ratio.

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

    NASA Technical Reports Server (NTRS)

    Kuhlman, John M.; Brown, Christopher K.

    1989-01-01

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

  2. The role of wind tunnel testing in the development of advanced rotary-wing aircraft

    NASA Technical Reports Server (NTRS)

    Kelly, M. W.

    1973-01-01

    The relations of wind tunnel test objectives to wind tunnel test requirements are reviewed in an assessment of the current role of wind tunnel testing in the development of advanced rotary-wing aircraft. Elements of typical development programs are examined, and a comparison of fixed wing and rotary wing aircraft programs is presented. Proposed new test facilities for fixed wing aircraft and typical aircraft program costs are discussed, along with the use of wind tunnels for tilt rotor research aircraft and the role of 40 x 80 ft wind tunnels in tilt rotor aircraft development. Some changes in current programs and methods are outlined for bringing about desired improvements.

  3. Results of buffet tests in a cryogenic wind tunnel

    NASA Technical Reports Server (NTRS)

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

    1982-01-01

    Buffet tests on two semispan wing models with different leading edge sweep show that it is feasibile to use the standard dynamic wing root bending moment technique in a cryogenic wind tunnel. One model was a slender 65 deg swept delta wing with sharp leading edges. The other model was an unswept wing of aspect ratio 1.5 with a British NPL 9510 airfoil section. The results for the 65 deg swept delta wing indicate the importance of matching the reduced frequency parameter in model tests for planforms which are sensitive to reduced frequency parameter if quantitative buffet measurements are required. The unique ability of a pressurized cryogenic wind tunnel to separate the effects of Reynolds number and of static aeroelastic distortion by variations in the tunnel stagnation temperature and pressure were demonstrated.

  4. Incremental wind tunnel testing of high lift systems

    NASA Astrophysics Data System (ADS)

    Victor, Pricop Mihai; Mircea, Boscoianu; Daniel-Eugeniu, Crunteanu

    2016-06-01

    Efficiency of trailing edge high lift systems is essential for long range future transport aircrafts evolving in the direction of laminar wings, because they have to compensate for the low performance of the leading edge devices. Modern high lift systems are subject of high performance requirements and constrained to simple actuation, combined with a reduced number of aerodynamic elements. Passive or active flow control is thus required for the performance enhancement. An experimental investigation of reduced kinematics flap combined with passive flow control took place in a low speed wind tunnel. The most important features of the experimental setup are the relatively large size, corresponding to a Reynolds number of about 2 Million, the sweep angle of 30 degrees corresponding to long range airliners with high sweep angle wings and the large number of flap settings and mechanical vortex generators. The model description, flap settings, methodology and results are presented.

  5. Flow quality studies of the NASA Lewis Research Center 8- by 6-foot supersonic/9- by 15-foot low speed wind tunnel

    NASA Technical Reports Server (NTRS)

    Arrington, E. Allen; Pickett, Mark T.

    1992-01-01

    A series of studies were conducted to determine the existing flow quality in the NASA Lewis 8 by 6 Foot Supersonic/9 by 15 Foot Low speed Wind Tunnel. The information gathered from these studies was used to determine the types and designs of flow manipulators which can be installed to improve overall tunnel flow quality and efficiency. Such manipulators include honeycomb flow straighteners, turbulence reduction screens, corner turning vanes, and acoustic treatments. The types of measurements, instrumentation, and results obtained from experiments conducted at several locations throughout the tunnel loop are described.

  6. Nacelle Chine Installation Based on Wind-Tunnel Test Using Efficient Global Optimization

    NASA Astrophysics Data System (ADS)

    Kanazaki, Masahiro; Yokokawa, Yuzuru; Murayama, Mitsuhiro; Ito, Takeshi; Jeong, Shinkyu; Yamamoto, Kazuomi

    Design exploration of a nacelle chine installation was carried out. The nacelle chine improves stall performance when deploying multi-element high-lift devices. This study proposes an efficient design process using a Kriging surrogate model to determine the nacelle chine installation point in wind-tunnel tests. The design exploration was conducted in a wind-tunnel using the JAXA high-lift aircraft model at the JAXA Large-scale Low-speed Wind Tunnel. The objective was to maximize the maximum lift. The chine installation points were designed on the engine nacelle in the axial and chord-wise direction, while the geometry of the chine was fixed. In the design process, efficient global optimization (EGO) which includes Kriging model and genetic algorithm (GA) was employed. This method makes it possible both to improve the accuracy of the response surface and to explore the global optimum efficiently. Detailed observations of flowfields using the Particle Image Velocimetry method confirmed the chine effect and design results.

  7. 40 CFR 53.63 - Test procedure: Wind tunnel inlet aspiration test.

    Code of Federal Regulations, 2010 CFR

    2010-07-01

    ... 40 Protection of Environment 5 2010-07-01 2010-07-01 false Test procedure: Wind tunnel inlet... Testing Performance Characteristics of Class II Equivalent Methods for PM2.5 § 53.63 Test procedure: Wind... extracts an ambient aerosol at elevated wind speeds. This wind tunnel test uses a single-sized,...

  8. 40 CFR 53.63 - Test procedure: Wind tunnel inlet aspiration test.

    Code of Federal Regulations, 2014 CFR

    2014-07-01

    ... 40 Protection of Environment 6 2014-07-01 2014-07-01 false Test procedure: Wind tunnel inlet... Testing Performance Characteristics of Class II Equivalent Methods for PM 2.5 § 53.63 Test procedure: Wind... extracts an ambient aerosol at elevated wind speeds. This wind tunnel test uses a single-sized,...

  9. 40 CFR 53.63 - Test procedure: Wind tunnel inlet aspiration test.

    Code of Federal Regulations, 2013 CFR

    2013-07-01

    ... 40 Protection of Environment 6 2013-07-01 2013-07-01 false Test procedure: Wind tunnel inlet... Testing Performance Characteristics of Class II Equivalent Methods for PM 2.5 § 53.63 Test procedure: Wind... extracts an ambient aerosol at elevated wind speeds. This wind tunnel test uses a single-sized,...

  10. 40 CFR 53.63 - Test procedure: Wind tunnel inlet aspiration test.

    Code of Federal Regulations, 2012 CFR

    2012-07-01

    ... 40 Protection of Environment 6 2012-07-01 2012-07-01 false Test procedure: Wind tunnel inlet... Testing Performance Characteristics of Class II Equivalent Methods for PM2.5 § 53.63 Test procedure: Wind... extracts an ambient aerosol at elevated wind speeds. This wind tunnel test uses a single-sized,...

  11. 40 CFR 53.63 - Test procedure: Wind tunnel inlet aspiration test.

    Code of Federal Regulations, 2011 CFR

    2011-07-01

    ... 40 Protection of Environment 5 2011-07-01 2011-07-01 false Test procedure: Wind tunnel inlet... Testing Performance Characteristics of Class II Equivalent Methods for PM2.5 § 53.63 Test procedure: Wind... extracts an ambient aerosol at elevated wind speeds. This wind tunnel test uses a single-sized,...

  12. Tactical Defenses Against Systematic Variation in Wind Tunnel Testing

    NASA Technical Reports Server (NTRS)

    DeLoach, Richard

    2002-01-01

    This paper examines the role of unexplained systematic variation on the reproducibility of wind tunnel test results. Sample means and variances estimated in the presence of systematic variations are shown to be susceptible to bias errors that are generally non-reproducible functions of those variations. Unless certain precautions are taken to defend against the effects of systematic variation, it is shown that experimental results can be difficult to duplicate and of dubious value for predicting system response with the highest precision or accuracy that could otherwise be achieved. Results are reported from an experiment designed to estimate how frequently systematic variations are in play in a representative wind tunnel experiment. These results suggest that significant systematic variation occurs frequently enough to cast doubts on the common assumption that sample observations can be reliably assumed to be independent. The consequences of ignoring correlation among observations induced by systematic variation are considered in some detail. Experimental tactics are described that defend against systematic variation. The effectiveness of these tactics is illustrated through computational experiments and real wind tunnel experimental results. Some tutorial information describes how to analyze experimental results that have been obtained using such quality assurance tactics.

  13. Cryogenic wind tunnels for high Reynolds number testing

    NASA Technical Reports Server (NTRS)

    Lawing, P. L.; Kilgore, R. A.; Mcguire, P. D.

    1986-01-01

    A compilation of lectures presented at various Universities over a span of several years is discussed. A central theme of these lectures has been to present the research facility in terms of the service it provides to, and its potential effect on, the entire community, rather than just the research community. This theme is preserved in this paper which deals with the cryogenic transonic wind tunnels at Langley Research Center. Transonic aerodynamics is a focus both because of its crucial role in determining the success of aeronautical systems and because cryogenic wind tunnels are especially applicable to the transonics problem. The paper also provides historical perspective and technical background for cryogenic tunnels, culminating in a brief review of cryogenic wind tunnel projects around the world. An appendix is included to provide up to date information on testing techniques that have been developed for the cryogenic tunnels at Langley Research Center. In order to be as inclusive and as current as possible, the appendix is less formal than the main body of the paper. It is anticipated that this paper will be of particular value to the technical layman who is inquisitive as to the value of, and need for, cryogneic tunnels.

  14. Variable Stiffness Spar Wind-Tunnel Model Development and Testing

    NASA Technical Reports Server (NTRS)

    Florance, James R.; Heeg, Jennifer; Spain, Charles V.; Ivanco, Thomas G.; Wieseman, Carol D.; Lively, Peter S.

    2004-01-01

    The concept of exploiting wing flexibility to improve aerodynamic performance was investigated in the wind tunnel by employing multiple control surfaces and by varying wing structural stiffness via a Variable Stiffness Spar (VSS) mechanism. High design loads compromised the VSS effectiveness because the aerodynamic wind-tunnel model was much stiffer than desired in order to meet the strength requirements. Results from tests of the model include stiffness and modal data, model deformation data, aerodynamic loads, static control surface derivatives, and fuselage standoff pressure data. Effects of the VSS on the stiffness and modal characteristics, lift curve slope, and control surface effectiveness are discussed. The VSS had the most effect on the rolling moment generated by the leading-edge outboard flap at subsonic speeds. The effects of the VSS for the other control surfaces and speed regimes were less. The difficulties encountered and the ability of the VSS to alter the aeroelastic characteristics of the wing emphasize the need for the development of improved design and construction methods for static aeroelastic models. The data collected and presented is valuable in terms of understanding static aeroelastic wind-tunnel model development.

  15. Results of design studies and wind tunnel tests of an advanced high lift system for an Energy Efficient Transport

    NASA Technical Reports Server (NTRS)

    Oliver, W. R.

    1980-01-01

    The development of an advanced technology high lift system for an energy efficient transport incorporating a high aspect ratio supercritical wing is described. This development is based on the results of trade studies to select the high lift system, analysis techniques utilized to design the high lift system, and results of a wind tunnel test program. The program included the first experimental low speed, high Reynolds number wind tunnel test for this class of aircraft. The experimental results include the effects on low speed aerodynamic characteristics of various leading and trailing edge devices, nacelles and pylons, aileron, spoilers, and Mach and Reynolds numbers. Results are discussed and compared with the experimental data and the various aerodynamic characteristics are estimated.

  16. Nano-ADEPT Aeroloads Wind Tunnel Test

    NASA Technical Reports Server (NTRS)

    Smith, Brandon; Cassell, A.; Yount, B.; Kruger, C.; Brivkalns, C.; Makino, A.; Zarchi, K.; McDaniel, R.; Venkatapathy, E.; Swanson, G.

    2015-01-01

    Analysis completed since the test suggests that all test objectives were met– This claim will be verified in the coming weeks as the data is examined further– Final disposition of test objective success will be documented in a final reportsubmitted to NASA stakeholders (early August 2015)– Expect conference paper in early 2016• Data products and observations made during testing will be used to refinecomputational models of Nano-ADEPT• Carbon fabric relaxed from its pre-test state during the test– System-level tolerance for relaxation will be driven by destination-specific andmission-specific aerothermal and aerodynamic requirements• Bonus experiment of asymmetric shape demonstrates that an asymmetricdeployable blunt body can be used to generate measureable lift– With a strut actuation system and a robust GN&C algorithm, this effect could beused to steer a blunt body at hypersonic speeds to aid precision landing

  17. Static Wind-Tunnel and Radio-Controlled Flight Test Investigation of a Remotely Piloted Vehicle Having a Delta Wing Planform

    NASA Technical Reports Server (NTRS)

    Yip, Long P.; Fratello, David J.; Robelen, David B.; Makowiec, George M.

    1990-01-01

    At the request of the United States Marine Corps, an exploratory wind-tunnel and flight test investigation was conducted by the Flight Dynamics Branch at the NASA Langley Research Center to improve the stability, controllability, and general flight characteristics of the Marine Corps Exdrone RPV (Remotely Piloted Vehicle) configuration. Static wind tunnel tests were conducted in the Langley 12 foot Low Speed Wind Tunnel to identify and improve the stability and control characteristics of the vehicle. The wind tunnel test resulted in several configuration modifications which included increased elevator size, increased vertical tail size and tail moment arm, increased rudder size and aileron size, the addition of vertical wing tip fins, and the addition of leading-edge droops on the outboard wing panel to improve stall departure resistance. Flight tests of the modified configuration were conducted at the NASA Plum Tree Test Site to provide a qualitative evaluation of the flight characteristics of the modified configuration.

  18. Combined Experiment Phase 1. [Horizontal axis wind turbines: wind tunnel testing versus field testing

    SciTech Connect

    Butterfield, C.P.; Musial, W.P.; Simms, D.A.

    1992-10-01

    How does wind tunnel airfoil data differ from the airfoil performance on an operating horizontal axis wind turbine (HAWT) The National Renewable Energy laboratory has been conducting a comprehensive test program focused on answering this question and understanding the basic fluid mechanics of rotating HAWT stall aerodynamics. The basic approach was to instrument a wind rotor, using an airfoil that was well documented by wind tunnel tests, and measure operating pressure distributions on the rotating blade. Based an the integrated values of the pressure data, airfoil performance coefficients were obtained, and comparisons were made between the rotating data and the wind tunnel data. Care was taken to the aerodynamic and geometric differences between the rotating and the wind tunnel models. This is the first of two reports describing the Combined Experiment Program and its results. This Phase I report covers background information such as test setup and instrumentation. It also includes wind tunnel test results and roughness testing.

  19. Wind tunnel tests on a 3 m diameter Musgrove windmill

    NASA Astrophysics Data System (ADS)

    Willmer, A. C.

    1980-01-01

    A 3 m diameter model of a two bladed Musgrove vertical axis windmill has been tested in the British Aerospace wind tunnel at Filton. Tunnel constraints were kept to a minimum by using a low flow blockage and appropriate corrections were applied to the measurements. The results of these tests demonstrate the good performance of this type of windmill. Comparison of the measured performance with predictions from a simple mathematical model show excellent agreement. Maximum loads measured on the windmill are not well predicted by the mathematical model. In order to reconcile measurement and prediction large induced crossflows must be postulated at some blade rotational positions.

  20. Wind-tunnel Tests of a Cyclogiro Rotor

    NASA Technical Reports Server (NTRS)

    Wheatley, John B; Windler, Ray

    1935-01-01

    During an extensive study of all types of rotating wings, the NACA examined the cyclogiro rotor and made an aerodynamic analysis of that system (reference 1). The examination disclosed that such a machine had sufficient promise to justify an experimental investigation; a model with a diameter and span of 8 feet was therefore constructed and tested in the 20-foot wind tunnel during 1934. The experimental work included tests of the effect of the motion upon the rotor forces during the static-lift and forward-flight conditions at several rotor speeds and the determination of the relations between the forces generated by the rotor and the power required by it.

  1. A wind tunnel test of newly developed personal bioaerosol samplers.

    PubMed

    Su, Wei-Chung; Tolchinsky, Alexander D; Sigaev, Vladimir I; Cheng, Yung Sung

    2012-07-01

    In this study the performance of two newly developed personal bioaerosol samplers was evaluated. The two test samplers are cyclone-based personal samplers that incorporate a recirculating liquid film. The performance evaluations focused on the physical efficiencies that a personal bioaerosol sampler could provide, including aspiration, collection, and capture efficiencies. The evaluation tests were carried out in a wind tunnel, and the test personal samplers were mounted on the chest of a full-size manikin placed in the test chamber of the wind tunnel. Monodisperse fluorescent aerosols ranging from 0.5 to 20 microm were used to challenge the samplers. Two wind speeds of 0.5 and 2.0 m/sec were employed as the test wind speeds in this study. The test results indicated that the aspiration efficiency of the two test samplers closely agreed with the ACGIH inhalable convention within the size range of the test aerosols. The aspiration efficiency was found to be independent of the sampling orientation. The collection efficiency acquired from these two samplers showed that the 50% cutoff diameters were both around 0.6 microm. However the wall loss of these two test samplers increased as the aerosol size increased, and the wall loss of PAS-4 was considerably higher than that of PAS-5, especially in the aerosol size larger than 5 microm, which resulted in PAS-4 having a relatively lower capture efficiency than PAS-5. Overall, the PAS-5 is considered a better personal bioaerosol sampler than the PAS-4.

  2. Evaluation of hydrogen as a cryogenic wind tunnel test gas

    NASA Technical Reports Server (NTRS)

    Haut, R. C.

    1977-01-01

    The nondimensional ratios used to describe various flow situations in hydrogen were determined and compared with the corresponding ideal diatomic gas ratios. The results were used to examine different inviscid flow configurations. The relatively high value of the characteristic rotational temperature causes the behavior of hydrogen, under cryogenic conditions, to deviate substantially from the behavior of an ideal diatomic gas in the compressible flow regime. Therefore, if an idea diatomic gas is to be modeled, cryogenic hydrogen is unacceptable as a wind tunnel test gas in a compressible flow situation.

  3. Wind tunnel testing of low-drag airfoils

    NASA Technical Reports Server (NTRS)

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

    1986-01-01

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

  4. An introduction to testing parachutes in wind tunnels

    SciTech Connect

    Macha, J.

    1991-01-01

    This paper reviews some of the technical considerations and current practices for testing parachutes in conventional wind tunnels. Special challenges to the experimentalist caused by the fabric construction, flexible geometry, and buff shape of parachutes are discussed. In particular, the topics of measurement technique, similarity considerations, and wall interference are addressed in a summary manner. Many references are cited which provide detailed coverage of the state of the art in testing methods. From the discussions presented, it is obvious that there are some serious problems with state of the art methods, especially in the area of canopy instrumentation and when working with reduced-scale models. But if the experimentalist is informed about the relative importance of the various factors for a specific test objective, it is usually possible to design a test that will yield meaningful results. The lower cost and the more favorable measurement environment of wind tunnels make their use an attractive alternative to flight testing whenever possible. 26 refs., 5 figs., 1 tab.

  5. Wind tunnel test IA300 analysis and results, volume 1

    NASA Technical Reports Server (NTRS)

    Kelley, P. B.; Beaufait, W. B.; Kitchens, L. L.; Pace, J. P.

    1987-01-01

    The analysis and interpretation of wind tunnel pressure data from the Space Shuttle wind tunnel test IA300 are presented. The primary objective of the test was to determine the effects of the Space Shuttle Main Engine (SSME) and the Solid Rocket Booster (SRB) plumes on the integrated vehicle forebody pressure distributions, the elevon hinge moments, and wing loads. The results of this test will be combined with flight test results to form a new data base to be employed in the IVBC-3 airloads analysis. A secondary objective was to obtain solid plume data for correlation with the results of gaseous plume tests. Data from the power level portion was used in conjunction with flight base pressures to evaluate nominal power levels to be used during the investigation of changes in model attitude, eleveon deflection, and nozzle gimbal angle. The plume induced aerodynamic loads were developed for the Space Shuttle bases and forebody areas. A computer code was developed to integrate the pressure data. Using simplified geometrical models of the Space Shuttle elements and components, the pressure data were integrated to develop plume induced force and moments coefficients that can be combined with a power-off data base to develop a power-on data base.

  6. Wind Tunnel Aerodynamic Tests of Six Airfoils for Use on Small Wind Turbines; Period of Performance: October 31, 2002--January 31, 2003

    SciTech Connect

    Selig, M. S.; McGranahan, B. D.

    2004-10-01

    Wind Tunnel Aerodynamic Tests of Six Airfoils for Use on Small Wind Turbinesrepresents the fourth installment in a series of volumes documenting the ongoing work of th University of Illinois at Urbana-Champaign Low-Speed Airfoil Tests Program. This particular volume deals with airfoils that are candidates for use on small wind turbines, which operate at low Reynolds numbers.

  7. Practical application of RINO, a smartphone-based dynamic displacement sensing application for wind tunnel tests

    NASA Astrophysics Data System (ADS)

    Lee, Seung-Woo; Jeong, Jong-Hyun; Knez, Kyle P.; Min, Jae-Hong; Jo, Hongki

    2016-04-01

    Dynamic displacement is one of the most important measurands in wind tunnel tests of structures. Laser sensors or optical sensors are usually used in wind tunnel tests to measure displacements. However, these commercial sensors have limitations in its use, cost and installation despite of their good performance in accuracy. RINO (Real-time Image- processing for Non-contact monitoring), an iOS software application for dynamic displacement monitoring, has been developed in the previous study. In this study, feasibility of RINO in practical use for wind tunnel tests is explored. Series of wind tunnel tests show that performances of RINO are comparable with those of conventional displacement sensors.

  8. Tests of a protective shell passive release mechanism for hypersonic wind-tunnel models

    NASA Technical Reports Server (NTRS)

    Puster, R. L.; Dunn, J. E.

    1979-01-01

    A protective shell mechanism for wind tunnel models was developed and tested. The mechanism is passive in operation, reliable, and imposes no new structural design changes for wind tunnel models. Methods of predicting the release time and the measured loads associated with the release of the shell are given. The mechanism was tested in a series of wind tunnel tests to validate the removal process and measure the pressure loads on the model. The protective shell can be used for wind tunnel models that require a step input of heating and loading such as a thin skin heat transfer model. The mechanism may have other potential applications.

  9. Testing a Parachute for Mars in World's Largest Wind Tunnel

    NASA Technical Reports Server (NTRS)

    2007-01-01

    The team developing the landing system for NASA's Mars Science Laboratory tested the deployment of an early parachute design in mid-October 2007 inside the world's largest wind tunnel, at NASA Ames Research Center, Moffett Field, California.

    In this image, two engineers are dwarfed by the parachute, which holds more air than a 280-square-meter (3,000-square-foot) house and is designed to survive loads in excess of 36,000 kilograms (80,000 pounds).

    The parachute, built by Pioneer Aerospace, South Windsor, Connecticut, has 80 suspension lines, measures more than 50 meters (165 feet) in length, and opens to a diameter of nearly 17 meters (55 feet). It is the largest disk-gap-band parachute ever built and is shown here inflated in the test section with only about 3.8 meters (12.5 feet) of clearance to both the floor and ceiling.

    The wind tunnel, which is 24 meters (80 feet) tall and 37 meters (120 feet) wide and big enough to house a Boeing 737, is part of the National Full-Scale Aerodynamics Complex, operated by the U.S. Air Force, Arnold Engineering Development Center.

    NASA's Jet Propulsion Laboratory, Pasadena, California, is building and testing the Mars Science Laboratory spacecraft for launch in 2009. The mission will land a roving analytical laboratory on the surface of Mars in 2010. JPL is a division of the California Institute of Technology.

  10. Further buffeting tests in a cryogenic wind tunnel

    NASA Technical Reports Server (NTRS)

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

    1992-01-01

    Further measurements of buffeting, using wing-root strain gauges, were made in the NASA Langley 0.3 m Cryogenic Wind Tunnel to refine techniques which will be used in larger cryogenic facilities such as the United States National Transonic Facility (NTF) and European Transonic Wind Tunnel (ETW). The questions addressed included the relative importance of variations in frequency parameter and Reynolds number, the choice of model material (considering both stiffness and damping) and the effects of static aeroelastic distortion. The main series of tests was made on half models of slender 65 deg delta wings with a sharp leading edge. The three delta wings had the same planform but widely different bending stiffness and frequencies (obtained by varying both the material and the thickness of the wings). It was known that the flow on this configuration would be insensitive to variations in Reynold number. Additional tests were made on one unswept half-wing of aspect ratio 1.5 with an NPL 9510 aerofoil section, known to be sensitive to variations in Reynolds number at transonic speeds. For brevity the test Mach numbers were restricted to M = 0.21 and 0.35 for the delta wings and to M = 0.30 for the unswept wing.

  11. Wind Tunnel Testing for the Stratospheric Observatory for Infrared Astronomy

    NASA Technical Reports Server (NTRS)

    Schenberger, Deborah; Alvarez, Teresa (Technical Monitor)

    1994-01-01

    NASA Ames Research Center is pursuing the development of SOFIA, the Stratospheric Observatory For Infrared Astronomy. SOFIA will consist of a 2.5 meter telescope mounted aft of the wing of a Boeing 747 aircraft. Since a large portion of the infrared spectrum is not visible at ground level due to absorption by water vapor in the atmosphere below 40,000 feet, it is highly desirable to make observations above this altitude. SOFIA will provide the opportunity for astronomers to conduct high-altitude research for extended periods of time. Current study is focused on wind tunnel testing for the open cavity. If not controlled, air would create resonance and damage the telescope. For this reason, SOFIA will design a boundary layer control device to achieve laminar flow over the cavity. This also provides a clearer flow for seeing, thus improving resolution on infrared sources. Other effects being tested in the wind tunnel are aerodynamic torque loads on the telescope, and flutter loads on the tail.

  12. A wind-tunnel investigation of parameters affecting helicopter directional control at low speeds in ground effect

    NASA Technical Reports Server (NTRS)

    Yeager, W. T., Jr.; Young, W. H., Jr.; Mantay, W. R.

    1974-01-01

    An investigation was conducted in the Langley full-scale tunnel to measure the performance of several helicopter tail-rotor/fin configurations with regard to directional control problems encountered at low speeds in ground effect. Tests were conducted at wind azimuths of 0 deg to 360 deg in increments of 30 deg and 60 deg and at wind speeds from 0 to 35 knots. The results indicate that at certain combinations of wind speed and wind azimuth, large increases in adverse fin force require correspondingly large increases in the tail-rotor thrust, collective pitch, and power required to maintain yaw trim. Changing the tail-rotor direction of rotation to top blade aft for either a pusher tail rotor (tail-rotor wake blowing away from fin) or a tractor tail rotor (tail-rotor wake blowing against fin) will alleviate this problem. For a pusher tail rotor at 180 deg wind azimuth, increases in the fin/tail-rotor gap were not found to have any significant influence on the overall vehicle directional control capability. Changing the tail rotor to a higher position was found to improve tail-rotor performance for a fin-off configuration at a wind azimuth of 180 deg. A V-tail configuration with a pusher tail rotor with top blade aft direction of rotation was found to be the best configuration with regard to overall directional control capability.

  13. 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.

  14. Wind Tunnel Test of an RPV with Shape-Change Control Effector and Sensor Arrays

    NASA Technical Reports Server (NTRS)

    Raney, David L.; Cabell, Randolph H.; Sloan, Adam R.; Barnwell, William G.; Lion, S. Todd; Hautamaki, Bret A.

    2004-01-01

    A variety of novel control effector concepts have recently emerged that may enable new approaches to flight control. In particular, the potential exists to shift the composition of the typical aircraft control effector suite from a small number of high authority, specialized devices (rudder, aileron, elevator, flaps), toward larger numbers of smaller, less specialized, distributed device arrays. The concept envisions effector and sensor networks composed of relatively small high-bandwidth devices able to simultaneously perform a variety of control functions using feedback from disparate data sources. To investigate this concept, a remotely piloted flight vehicle has been equipped with an array of 24 trailing edge shape-change effectors and associated pressure measurements. The vehicle, called the Multifunctional Effector and Sensor Array (MESA) testbed, was recently tested in NASA Langley's 12-ft Low Speed wind tunnel to characterize its stability properties, control authorities, and distributed pressure sensitivities for use in a dynamic simulation prior to flight testing. Another objective was to implement and evaluate a scheme for actively controlling the spanwise pressure distribution using the shape-change array. This report describes the MESA testbed, design of the pressure distribution controller, and results of the wind tunnel test.

  15. Lessons learned from wind tunnel testing of a droop-nose morphing wingtip

    NASA Astrophysics Data System (ADS)

    Vasista, Srinivas; Riemenschneider, Johannes; van de Kamp, Bram; Monner, Hans Peter; Cheung, Ronald C. M.; Wales, Christopher; Cooper, Jonathan

    2016-04-01

    This work presents the lessons learned from wind tunnel tests of a droop-nose morphing wingtip as part of the EU project NOVEMOR. The design followed a sequential chain and was largely driven through optimization tools, including a glass-fiber composite skin optimization tool and a topology optimization tool for the design of internal super-elastic and aluminium compliant mechanisms. The device was tested in the low speed tunnel at the University of Bristol to determine the structural response under aerodynamic loading. Measurements of strain from strain gauges show that the structure is capable of handing the aerodynamic loads though also show an imbalance of strain between the components. Measurements of surface pressures show a small variation of cp with the 2° droop morphing variation as per the target. The wind tunnel testing showed that further developments to the design chain are necessary, in particular the need for a concurrent as opposed to sequential chain for the design of the various components. Considerations of other problem formulations, the inclusion of nonlinear finite element analysis, and ways to interpret the structural boundary of the topology optimization results with more confidence are required. The utilization of super-elastic materials in morphing structures may also prove to be highly beneficial for their performance.

  16. Wind-tunnel test of the S814 thick root airfoil

    SciTech Connect

    Somers, D.M.; Tangler, J.L.

    1995-01-01

    The objective of this wind-tunnel test was to verify the predictions of the Eppler Airfoil Design and Analysis Code for a very thick airfoil having a high maximum lift coefficient (c{sub 1,max} designed to be largely insensitive to leading edge roughness effects. The 24-percent-thick S814 airfoil was designed with these characteristics to accommodate aerodynamic and structural considerations for the root region of a wind-turbine blade. In addition, the airfoil`s maximum lift-to-drag ratio was designed to occur it a high lift coefficient. To accomplish the objective, a two-dimensional wind-tunnel test of the S814 thick root airfog was conducted in January 1994 in the low-turbulence wind tunnel of the Delft University of Technology Low Speed Laboratory. Data were obtained for transition-free and transition-fixed conditions at Reynolds numbers of 0.7, 1.0, 1.5, 2.0, and 3.0 {times} 10{sup 6}. For the design Reynolds numbers of 1.5 {times} l0{sup 6}, the transition-free c{sub 1,max} is 1.3 which satisfies the design specification. However, this value is significantly lower than the predicted c{sub 1,max} of almost l.6. With transition-fixed at the is 1.2. The difference in c{sub 1,max} between the transition-free and transition-fixed conditions demonstrates the airfoil`s minimal sensitivity to roughness effects. The S814 root airfoil was designed to complement existing NREL low c{sub 1,max} tip-region airfoils for rotor blades 10 to 15 meters in length.

  17. Overview of Low-Speed Aerodynamic Tests on a 5.75% Scale Blended-Wing-Body Twin Jet Configuration

    NASA Technical Reports Server (NTRS)

    Vicroy, Dan D.; Dickey, Eric; Princen, Norman; Beyar, Michael D.

    2016-01-01

    The NASA Environmentally Responsible Aviation (ERA) Project sponsored a series of computational and experimental investigations of the propulsion and airframe integration issues associated with Hybrid-Wing-Body (HWB) or Blended-Wing-Body (BWB) configurations. NASA collaborated with Boeing Research and Technology (BR&T) to conduct this research on a new twin-engine Boeing BWB transport configuration. The experimental investigations involved a series of wind tunnel tests with a 5.75-percent scale model conducted in two low-speed wind tunnels. This testing focused on the basic aerodynamics of the configuration and selection of the leading edge Krueger slat position for takeoff and landing. This paper reviews the results and analysis of these low-speed wind tunnel tests.

  18. Standardization Tests of NACA No. 1 Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Reid, Elliott G

    1925-01-01

    The tests described in this report were made in the 5-foot atmospheric wind tunnel of the National Advisory Committee for Aeronautics, at Langley Field. The primary objective of collecting data on the characteristics of this tunnel for comparison with those of others throughout the world, in order that, in the future, the results of tests made in all the principle laboratories may be interpreted, compared, and coordinated on a basis of scientifically established relationships, a process hitherto impossible due to the lack of comparable data. The work includes tests of a disk, spheres, cylinders, and airfoils, explorations of the test section for static pressure and velocity distribution, and determination of the variations of air flow direction throughout the operating range of the tunnel. (author)

  19. Key Topics for High-Lift Research: A Joint Wind Tunnel/Flight Test Approach

    NASA Technical Reports Server (NTRS)

    Fisher, David; Thomas, Flint O.; Nelson, Robert C.

    1996-01-01

    Future high-lift systems must achieve improved aerodynamic performance with simpler designs that involve fewer elements and reduced maintenance costs. To expeditiously achieve this, reliable CFD design tools are required. The development of useful CFD-based design tools for high lift systems requires increased attention to unresolved flow physics issues. The complex flow field over any multi-element airfoil may be broken down into certain generic component flows which are termed high-lift building block flows. In this report a broad spectrum of key flow field physics issues relevant to the design of improved high lift systems are considered. It is demonstrated that in-flight experiments utilizing the NASA Dryden Flight Test Fixture (which is essentially an instrumented ventral fin) carried on an F-15B support aircraft can provide a novel and cost effective method by which both Reynolds and Mach number effects associated with specific high lift building block flows can be investigated. These in-flight high lift building block flow experiments are most effective when performed in conjunction with coordinated ground based wind tunnel experiments in low speed facilities. For illustrative purposes three specific examples of in-flight high lift building block flow experiments capable of yielding a high payoff are described. The report concludes with a description of a joint wind tunnel/flight test approach to high lift aerodynamics research.

  20. The cryogenic wind tunnel for high Reynolds number testing. Ph.D. Thesis

    NASA Technical Reports Server (NTRS)

    Kilgore, R. A.

    1974-01-01

    Experiments performed at the NASA Langley Research Center in a cryogenic low-speed continuous-flow tunnel and in a cryogenic transonic continuous-flow pressure tunnel have demonstrated the predicted changes in Reynolds number, drive power, and fan speed with temperature, while operating with nitrogen as the test gas. The experiments have also demonstrated that cooling to cryogenic temperatures by spraying liquid nitrogen directly into the tunnel circuit is practical and that tunnel temperature can be controlled within very close limits. Whereas most types of wind tunnel could operate with advantage at cryogenic temperatures, the continuous-flow fan-driven tunnel is particularly well suited to take full advantage of operating at these temperatures. A continuous-flow fan-driven cryogenic tunnel to satisfy current requirements for test Reynolds number can be constructed and operated using existing techniques. Both capital and operating costs appear acceptable.

  1. Wind Tunnel Test of Mach 5 Class Hypersonic Airplane

    NASA Astrophysics Data System (ADS)

    Nakatani, Hiroki; Taguchi, Hideyuki; Fujita, Kazuhisa; Shindo, Shigemi; Honami, Shinji

    JAXA is currently performing studies on a Hypersonic Turbojet Experimental Vehicle, which involve a hypersonic flight test of a Small Pre-cooled Turbojet Engine. The aerodynamic performance of this airplane was examined at the JAXA hypersonic, supersonic, and transonic wind tunnel facilities. The 6-degrees-of-freedom forces and pressure distribution around the model were measured and evaluated. This airplane satisfies the lift-to-drag ratio requirement for a flight test at Mach 5. In addition, the results indicate that this airplane has longitudinal and directional static stability if the moment reference point is x/l smaller than 0.35. A separation occurs at the external expanding nozzle. Therefore, a redesign is necessary to solve these problems.

  2. Wind tunnel tests of a free yawing downwind wind turbine

    NASA Astrophysics Data System (ADS)

    Verelst, D. R. S.; Larsen, T. J.; van Wingerden, J. W.

    2014-12-01

    This research paper presents preliminary results on a behavioural study of a free yawing downwind wind turbine. A series of wind tunnel tests was performed at the TU Delft Open Jet Facility with a three bladed downwind wind turbine and a rotor radius of 0.8 meters. The setup includes an off the shelf three bladed hub, nacelle and generator on which relatively flexible blades are mounted. The tower support structure has free yawing capabilities provided at the base. A short overview on the technical details of the experiment is given as well as a brief summary of the design process. The discussed test cases show that the turbine is stable while operating in free yawing conditions. Further, the effect of the tower shadow passage on the blade flapwise strain measurement is evaluated. Finally, data from the experiment is compared with preliminary simulations using DTU Wind Energy's aeroelastic simulation program HAWC2.

  3. Results of an investigation elevon hinge moments and dual panel elevon effectiveness using an .0405-scale model (16-0) of the configuration 140C space shuttle orbiter in the Rockwell International NAAL low speed wind tunnel (OA119B)

    NASA Technical Reports Server (NTRS)

    Mennell, R. C.

    1975-01-01

    Experimental aerodynamic investigations were conducted on a sting mounted .0405-scale representation of the 140C outer mold line space shuttle orbiter in a 7.75 x 11.00 foot low speed wind tunnel during the time period from August 22, 1974 to September 6, 1974. The primary test objectives were to define dual panel elevon/aileron effectiveness and to investigate elevon hinge-moments for the 140C orbiter configuration with wing/elevon upper hingeline sealing flapper doors. The elevon parametric variations, consisting of the basic elevons with 6 inch gaps and flapper doors, elevons with no flapper doors and completely open upper hingeline gap, and an entirely sealed solid elevon, were tested with elevon deflections from +20 to -35 deg at various aileron deflections. Aerodynamic force and moment data were measured in the body axis system by a 2.5 inch task type internal strain gage balance.

  4. X-34 Ceramic Model Tested In NASA Langley Wind Tunnel

    NASA Technical Reports Server (NTRS)

    1997-01-01

    A ceramic aerothermodynamic heat transfer model of the X-34 is being prepared for testing in the 31-Inch Mach 10 Wind Tunnel located at NASA Langley Research Center in Hampton, Va. The model is a 0.018-scale version of the X-34 advanced technology flight demonstration vehicle. The X-34 is an autonomous rocket-powered, air-launched winged concept that is smaller, lighter and much less expensive than other lifting body vehicles, for example, the X-33. The X-34 is intended for hypersonic flight up to speeds of Mach 8, and would be used to demonstrate technologies for future reusable space transportation vehicles. The first flight of the X-34 vehicle is currently scheduled for early 1999.

  5. Flow-Visualization Techniques Used at High Speed by Configuration Aerodynamics Wind-Tunnel-Test Team

    NASA Technical Reports Server (NTRS)

    Lamar, John E. (Editor)

    2001-01-01

    This paper summarizes a variety of optically based flow-visualization techniques used for high-speed research by the Configuration Aerodynamics Wind-Tunnel Test Team of the High-Speed Research Program during its tenure. The work of other national experts is included for completeness. Details of each technique with applications and status in various national wind tunnels are given.

  6. Correlation of full-scale drag predictions with flight measurements on the C-141A aircraft : Phase 2: Wind tunnel tests, analysis, and prediction techniques. Volume 2: Wind tunnel test and basic data

    NASA Technical Reports Server (NTRS)

    Macwilkinson, D. G.; Blackerby, W. T.; Paterson, J. H.

    1974-01-01

    A research program has been conducted to determine the degree of cruise drag correlation on the C-141A aircraft between predictions based on wind tunnel test data, and flight test results. Information is presented on the wind tunnel test program and basic aerodynamic data on the C-141A wind-tunnel model used in the correlation studies.

  7. Development of Doppler Global Velocimetry for Wind Tunnel Testing

    NASA Technical Reports Server (NTRS)

    Meyers, James F.

    1994-01-01

    The development of Doppler global velocimetry is described. Emphasis is placed on the modifications necessary to advance this nonintrusive laser based measurement technique from a laboratory prototype to a viable wind tunnel flow diagnostics tool. Several example wind tunnel flow field investigations are described to illustrate the versatility of the technique. Flow conditions ranged from incompressible to Mach 2.8 with measurement distances extending from 1 to 15 m.

  8. Investigation of rotor blade element airloads for a teetering rotor in the blade stall regime (second wind tunnel test)

    NASA Technical Reports Server (NTRS)

    Dadone, L. U.; Fukushima, T.

    1975-01-01

    A test was conducted in the NASA-Ames 7 x 10 ft low speed wind tunnel on a seven-foot diameter model of a teetering rotor. The objectives of the test were: (1) acquire pressure data for correlation with laser and flow visualization measurements; (2) explore rotor propulsive force limits by varying the advance ratio at constant lift and propulsive force coefficients; (3) obtain additional data to define the differences between teetering and articulated rotors; and (4) verify the acceleration sensitivity of experimental transducers. Results are presented.

  9. Ares I Upper Stage Pressure Tests in Wind Tunnel

    NASA Technical Reports Server (NTRS)

    2007-01-01

    Under the goals of the Vision for Space Exploration, Ares I is a chief component of the cost-effective space transportation infrastructure being developed by NASA's Constellation Program. This transportation system will safely and reliably carry human explorers back to the moon, and then onward to Mars and other destinations in the solar system. The Ares I effort includes multiple project element teams at NASA centers and contract organizations around the nation, and is managed by the Exploration Launch Projects Office at NASA's Marshall Space Flight Center (MFSC). ATK Launch Systems near Brigham City, Utah, is the prime contractor for the first stage booster. ATK's subcontractor, United Space Alliance of Houston, is designing, developing and testing the parachutes at its facilities at NASA's Kennedy Space Center in Florida. NASA's Johnson Space Center in Houston hosts the Constellation Program and Orion Crew Capsule Project Office and provides test instrumentation and support personnel. Together, these teams are developing vehicle hardware, evolving proven technologies, and testing components and systems. Their work builds on powerful, reliable space shuttle propulsion elements and nearly a half-century of NASA space flight experience and technological advances. Ares I is an inline, two-stage rocket configuration topped by the Crew Exploration Vehicle, its service module, and a launch abort system. In this HD video image, the first stage reentry 1/2% model is undergoing pressure measurements inside the wind tunnel testing facility at MSFC. (Highest resolution available)

  10. Numerical Simulation of Selecting Model Scale of Cable in Wind Tunnel Test

    NASA Astrophysics Data System (ADS)

    Huang, Yifeng; Yang, Jixin

    The numerical simulation method based on computational Fluid Dynamics (CFD) provides a possible alternative means of physical wind tunnel test. Firstly, the correctness of the numerical simulation method is validated by one certain example. In order to select the minimum length of the cable as to a certain diameter in the numerical wind tunnel tests, the numerical wind tunnel tests based on CFD are carried out on the cables with several different length-diameter ratios (L/D). The results show that, when the L/D reaches to 18, the drag coefficient is stable essentially.

  11. Preliminary Low-Speed Wind-Tunnel Investigation of Some Aspects of the Aerodynamic Problems Associated with Missiles Carried Externally in Positions Near Airplane Wings

    NASA Technical Reports Server (NTRS)

    Alford, William J., Jr.; Silvers, H. Norman; King, Thomas J., Jr.

    1954-01-01

    A low-speed wind-tunnel investigation has been made of some aspects of the aerodynamic problems associated with the use of air-to-air missiles when carried externally on aircraft. Measurements of the forces and moments on a missile model for a range of positions under the mid-semispan location of a 45deg sweptback wing indicated longitudinal and lateral forces with regard to both carriage and release of the missiles. Surveys of the characteristics of the flow field in the region likely to be traversed by the missiles showed abrupt gradients in both flow angularity and in local dynamic pressure. Through the use of aerodynamic data on the isolated missile and the measured flow-field characteristics, the longitudinal forces and moments acting on the missile while in the presence of the wing-fuselage combination could be estimated with fair accuracy. Although the lateral forces and moments predicted were qualitatively correct, there existed some large discrepancies in absolute magnitude.

  12. Exploratory flutter test in a cryogenic wind tunnel

    NASA Technical Reports Server (NTRS)

    Cole, S. R.

    1985-01-01

    A model consisting of a rigid wing with an integral, flexible beam support that was cantilever mounted from the wall in the NASA LaRC 0.3-m transonic cryogenic tunnel was used in a flutter analysis study. The wing had a rectangular planform of aspect ratio 1.5 and a 64A010 airfoil. Various considerations and procedures for conducting flutter tests in a cryogenic wind tunnel were evaluated. Flutter onset conditions were established from extrapolated subcritical response measurements. A flutter boundary was determined at cryogenic temperatures over a Mach number M range from 0.5 to 0.9. Flutter was obtained at two different Reynolds numbers R at M = 0.5 (R = 4.4 and 18.4 x 10 to the 6th power) and at M = 0.8 (R = 5.0 and 10.4 x 10 to the 6th power). Flutter analyses using subsonic lifting surface (kernel function) aerodynamics were made over the range of test conditions. To evaluate the Reynolds number effects at M = 0.5 and 0.8, the experimental results were adjusted using analytical trends to account for differences in the model test temperatures and mass ratios. The adjusted experimental results indicate that increasing Reynolds number from 5.0 to 20.0 x 10 to the 6th power decreased the dynamic pressure by 4.0 to 6.5 percent at M = 0.5 and 0.8.

  13. Aerodynamic and Aeroacoustic Wind Tunnel Testing of the Orion Spacecraft

    NASA Technical Reports Server (NTRS)

    Ross, James C.

    2011-01-01

    The Orion aerodynamic testing team has completed more than 40 tests as part of developing the aerodynamic and loads databases for the vehicle. These databases are key to achieving good mechanical design for the vehicle and to ensure controllable flight during all potential atmospheric phases of a mission, including launch aborts. A wide variety of wind tunnels have been used by the team to document not only the aerodynamics but the aeroacoustic environment that the Orion might experience both during nominal ascents and launch aborts. During potential abort scenarios the effects of the various rocket motor plumes on the vehicle must be accurately understood. The Abort Motor (AM) is a high-thrust, short duration motor that rapidly separates Orion from its launch vehicle. The Attitude Control Motor (ACM), located in the nose of the Orion Launch Abort Vehicle, is used for control during a potential abort. The 8 plumes from the ACM interact in a nonlinear manner with the four AM plumes which required a carefully controlled test to define the interactions and their effect on the control authority provided by the ACM. Techniques for measuring dynamic stability and for simulating rocket plume aerodynamics and acoustics were improved or developed in the course of building the aerodynamic and loads databases for Orion.

  14. Comparison of Angle of Attack Measurements for Wind Tunnel Testing

    NASA Technical Reports Server (NTRS)

    Jones, Thomas, W.; Hoppe, John C.

    2001-01-01

    Two optical systems capable of measuring model attitude and deformation were compared to inertial devices employed to acquire wind tunnel model angle of attack measurements during the sting mounted full span 30% geometric scale flexible configuration of the Northrop Grumman Unmanned Combat Air Vehicle (UCAV) installed in the NASA Langley Transonic Dynamics Tunnel (TDT). The overall purpose of the test at TDT was to evaluate smart materials and structures adaptive wing technology. The optical techniques that were compared to inertial devices employed to measure angle of attack for this test were: (1) an Optotrak (registered) system, an optical system consisting of two sensors, each containing a pair of orthogonally oriented linear arrays to compute spatial positions of a set of active markers; and (2) Video Model Deformation (VMD) system, providing a single view of passive targets using a constrained photogrammetric solution whose primary function was to measure wing and control surface deformations. The Optotrak system was installed for this test for the first time at TDT in order to assess the usefulness of the system for future static and dynamic deformation measurements.

  15. The estimation of airplane performance from wind tunnel tests on conventional airplane models

    NASA Technical Reports Server (NTRS)

    Warner, Edward P; Ober, Shatswell

    1925-01-01

    Calculations of the magnitude of the correction factors and the range of their variations for wind tunnel models used in making aircraft performance predictions were made for 23 wind tunnel models. Calculated performances were compared with those actually determined for such airplanes as have been built and put through flight test. Except as otherwise noted, all the models have interplane struts and diagonal struts formed to streamwise shape. Wires were omitted in all cases. All the models were about 18 inches in span and were tested in a 4-foot wind tunnel. Results are given in tabular form.

  16. Wind Tunnel Tests on Aerodynamic Characteristics of Advanced Solid Rocket

    NASA Astrophysics Data System (ADS)

    Kitamura, Keiichi; Fujimoto, Keiichiro; Nonaka, Satoshi; Irikado, Tomoko; Fukuzoe, Moriyasu; Shima, Eiji

    The Advanced Solid Rocket is being developed by JAXA (Japan Aerospace Exploration Agency). Since its configuration has been changed very recently, its aerodynamic characteristics are of great interest of the JAXA Advanced Solid Rocket Team. In this study, we carried out wind tunnel tests on the aerodynamic characteristics of the present configuration for Mach 1.5. Six test cases were conducted with different body configurations, attack angles, and roll angles. A six component balance, oilflow visualization, Schlieren images were used throughout the experiments. It was found that, at zero angle-of-attack, the flow around the body were perturbed and its drag (axial force) characteristics were significantly influenced by protruding body components such as flanges, cable ducts, and attitude control units of SMSJ (Solid Motor Side Jet), while the nozzle had a minor role. With angle-of-attack of five degree, normal force of CNα = 3.50±0.03 was measured along with complex flow features observed in the full-component model; whereas no crossflow separations were induced around the no-protuberance model with CNα = 2.58±0.10. These values were almost constant with respect to the angle-of-attack in both of the cases. Furthermore, presence of roll angle made the flow more complicated, involving interactions of separation vortices. These data provide us with fundamental and important aerodynamic insights of the Advanced Solid Rocket, and they will be utilized as reference data for the corresponding numerical analysis.

  17. Documentation and archiving of the Space Shuttle wind tunnel test data base. Volume 1: Background and description

    NASA Technical Reports Server (NTRS)

    Romere, Paul O.; Brown, Steve Wesley

    1995-01-01

    Development of the space shuttle necessitated an extensive wind tunnel test program, with the cooperation of all the major wind tunnels in the United States. The result was approximately 100,000 hours of space shuttle wind tunnel testing conducted for aerodynamics, heat transfer, and structural dynamics. The test results were converted into Chrysler DATAMAN computer program format to facilitate use by analysts, a very cost effective method of collecting the wind tunnel test results from many test facilities into one centralized location. This report provides final documentation of the space shuttle wind tunnel program. The two-volume set covers evolution of space shuttle aerodynamic configurations and gives wind tunnel test data, titles of wind tunnel data reports, sample data sets, and instructions for accessing the digital data base.

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

    NASA Technical Reports Server (NTRS)

    Deloach, Richard

    2012-01-01

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

  19. Integration of computational methods into automotive wind tunnel testing

    SciTech Connect

    Katz, J.

    1989-01-01

    This paper discusses the aerodynamics of a generic, enclosed-wheel racing-car shape without wheels investigated numerically and compared with one-quarter scale wind-tunnel data. Because both methods lack perfection in simulating actual road conditions, a complementary application of these methods was studied. The computations served for correcting the high-blockage wind-tunnel results and provided detailed pressure data which improved the physical understanding of the flow field. The experimental data was used here mainly to provide information on the location of flow-separation lines and on the aerodynamic loads; these in turn were used to validate and to calibrate the computations.

  20. Wind Tunnel and Propulsion Test Facilities: An Assessment of NASA's Capabilities to Serve National Needs

    NASA Technical Reports Server (NTRS)

    Anton, Philip S.; Gritton, Eugene C.; Mesic, Richard; Steinberg, Paul; Johnson, Dana J.

    2004-01-01

    This monograph reveals and discusses the National Aeronautics and Space Administration's (NASA's) wind tunnel and propulsion test facility management issues that are creating real risks to the United States' competitive aeronautics advantage.

  1. Wind Tunnel Tests of Parabolic Trough Solar Collectors: March 2001--August 2003

    SciTech Connect

    Hosoya, N.; Peterka, J. A.; Gee, R. C.; Kearney, D.

    2008-05-01

    Conducted extensive wind-tunnel tests on parabolic trough solar collectors to determine practical wind loads applicable to structural design for stress and deformation, and local component design for concentrator reflectors.

  2. Calibration of the NASA Glenn 8- by 6-Foot Supersonic Wind Tunnel (1996 and 1997 Tests)

    NASA Technical Reports Server (NTRS)

    Arrington, E. Allen

    2012-01-01

    There were several physical and operational changes made to the NASA Glenn Research Center 8- by 6-Foot Supersonic Wind Tunnel during the period of 1992 through 1996. Following each of these changes, a facility calibration was conducted to provide the required information to support the research test programs. Due to several factors (facility research test schedule, facility downtime and continued facility upgrades), a full test section calibration was not conducted until 1996. This calibration test incorporated all test section configurations and covered the existing operating range of the facility. However, near the end of that test entry, two of the vortex generators mounted on the compressor exit tailcone failed causing minor damage to the honeycomb flow straightener. The vortex generators were removed from the facility and calibration testing was terminated. A follow-up test entry was conducted in 1997 in order to fully calibrate the facility without the effects of the vortex generators and to provide a complete calibration of the newly expanded low speed operating range. During the 1997 tunnel entry, all planned test points required for a complete test section calibration were obtained. This data set included detailed in-plane and axial flow field distributions for use in quantifying the test section flow quality.

  3. 40 CFR 53.42 - Generation of test atmospheres for wind tunnel tests.

    Code of Federal Regulations, 2013 CFR

    2013-07-01

    ... 40 Protection of Environment 6 2013-07-01 2013-07-01 false Generation of test atmospheres for wind... Testing Performance Characteristics of Methods for PM 10 § 53.42 Generation of test atmospheres for wind... particle delivery system shall consist of a blower system and a wind tunnel having a test section...

  4. 40 CFR 53.42 - Generation of test atmospheres for wind tunnel tests.

    Code of Federal Regulations, 2014 CFR

    2014-07-01

    ... 40 Protection of Environment 6 2014-07-01 2014-07-01 false Generation of test atmospheres for wind... Testing Performance Characteristics of Methods for PM 10 § 53.42 Generation of test atmospheres for wind... particle delivery system shall consist of a blower system and a wind tunnel having a test section...

  5. 40 CFR 53.42 - Generation of test atmospheres for wind tunnel tests.

    Code of Federal Regulations, 2010 CFR

    2010-07-01

    ... 40 Protection of Environment 5 2010-07-01 2010-07-01 false Generation of test atmospheres for wind... Testing Performance Characteristics of Methods for PM10 § 53.42 Generation of test atmospheres for wind... particle delivery system shall consist of a blower system and a wind tunnel having a test section...

  6. 40 CFR 53.42 - Generation of test atmospheres for wind tunnel tests.

    Code of Federal Regulations, 2011 CFR

    2011-07-01

    ... 40 Protection of Environment 5 2011-07-01 2011-07-01 false Generation of test atmospheres for wind... Testing Performance Characteristics of Methods for PM10 § 53.42 Generation of test atmospheres for wind... particle delivery system shall consist of a blower system and a wind tunnel having a test section...

  7. Data Fusion in Wind Tunnel Testing; Combined Pressure Paint and Model Deformation Measurements (Invited)

    NASA Technical Reports Server (NTRS)

    Bell, James H.; Burner, Alpheus W.

    2004-01-01

    As the benefit-to-cost ratio of advanced optical techniques for wind tunnel measurements such as Video Model Deformation (VMD), Pressure-Sensitive Paint (PSP), and others increases, these techniques are being used more and more often in large-scale production type facilities. Further benefits might be achieved if multiple optical techniques could be deployed in a wind tunnel test simultaneously. The present study discusses the problems and benefits of combining VMD and PSP systems. The desirable attributes of useful optical techniques for wind tunnels, including the ability to accommodate the myriad optical techniques available today, are discussed. The VMD and PSP techniques are briefly reviewed. Commonalties and differences between the two techniques are discussed. Recent wind tunnel experiences and problems when combining PSP and VMD are presented, as are suggestions for future developments in combined PSP and deformation measurements.

  8. Wind Tunnel Testing of Various Disk-Gap-Band Parachutes

    NASA Technical Reports Server (NTRS)

    Cruz, Juan R.; Mineck, Raymond E.; Keller, Donald F.; Bobskill, Maria V.

    2003-01-01

    Two Disk-Gap-Band model parachute designs were tested in the NASA Langley Transonic Dynamics Tunnel. The purposes of these tests were to determine the drag and static stability coefficients of these two model parachutes at various subsonic Mach numbers in support of the Mars Exploration Rover mission. The two model parachute designs were designated 1.6 Viking and MPF. These model parachute designs were chosen to investigate the tradeoff between drag and static stability. Each of the parachute designs was tested with models fabricated from MIL-C-7020 Type III or F-111 fabric. The reason for testing model parachutes fabricated with different fabrics was to evaluate the effect of fabric permeability on the drag and static stability coefficients. Several improvements over the Viking-era wind tunnel tests were implemented in the testing procedures and data analyses. Among these improvements were corrections for test fixture drag interference and blockage effects, and use of an improved test fixture for measuring static stability coefficients. The 1.6 Viking model parachutes had drag coefficients from 0.440 to 0.539, while the MPF model parachutes had drag coefficients from 0.363 to 0.428. The 1.6 Viking model parachutes had drag coefficients 18 to 22 percent higher than the MPF model parachute for equivalent fabric materials and test conditions. Model parachutes of the same design tested at the same conditions had drag coefficients approximately 11 to 15 percent higher when manufactured from F-111 fabric as compared to those fabricated from MIL-C-7020 Type III fabric. The lower fabric permeability of the F-111 fabric was the source of this difference. The MPF model parachutes had smaller absolute statically stable trim angles of attack as compared to the 1.6 Viking model parachutes for equivalent fabric materials and test conditions. This was attributed to the MPF model parachutes larger band height to nominal diameter ratio. For both designs, model parachutes

  9. NTS-spill test facility wind tunnel exhaust plume characterization

    SciTech Connect

    Kerr, R.; Goldwire, H.; Smith, D.; Rawlings, J.; Schaffer, T.; Robson, J.

    1994-07-01

    The exhaust plume of the NTS-STF wind tunnel has been characterized to demonstrate its suitability as a target for CALIOPE experiments. Smoke from grenades has been released in multiple quantities and at different positions inside the tunnel. The smoke plumes have been recorded on video tape. The wind velocity profile has also been determined with a moveable array of miniature vane anemometers. These measurements will be used to determine the vapor concentration pathlength as part of the ground truth.

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

    NASA Technical Reports Server (NTRS)

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

    1990-01-01

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

  11. Low-speed wind-tunnel investigation of the flight dynamic characteristics of an advanced turboprop business/commuter aircraft configuration

    NASA Technical Reports Server (NTRS)

    Coe, Paul L., Jr.; Turner, Steven G.; Owens, D. Bruce

    1990-01-01

    An investigation was conducted to determine the low-speed flight dynamic behavior of a representative advanced turboprop business/commuter aircraft concept. Free-flight tests were conducted in the NASA Langley Research Center's 30- by 60-Foot Tunnel. In support of the free-flight tests, conventional static, dynamic, and free-to-roll oscillation tests were performed. Tests were intended to explore normal operating and post stall flight conditions, and conditions simulating the loss of power in one engine.

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

    NASA Astrophysics Data System (ADS)

    Petronevich, V. V.

    2016-10-01

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

  13. Aeroelastic Deformation: Adaptation of Wind Tunnel Measurement Concepts to Full-Scale Vehicle Flight Testing

    NASA Technical Reports Server (NTRS)

    Burner, Alpheus W.; Lokos, William A.; Barrows, Danny A.

    2005-01-01

    The adaptation of a proven wind tunnel test technique, known as Videogrammetry, to flight testing of full-scale vehicles is presented. A description is presented of the technique used at NASA's Dryden Flight Research Center for the measurement of the change in wing twist and deflection of an F/A-18 research aircraft as a function of both time and aerodynamic load. Requirements for in-flight measurements are compared and contrasted with those for wind tunnel testing. The methodology for the flight-testing technique and differences compared to wind tunnel testing are given. Measurement and operational comparisons to an older in-flight system known as the Flight Deflection Measurement System (FDMS) are presented.

  14. Plans and Status of Wind-Tunnel Testing Employing an Aeroservoelastic Semispan Model

    NASA Technical Reports Server (NTRS)

    Perry, Boyd, III; Silva, Walter A.; Florance, James R.; Wieseman, Carol D.; Pototzky, Anthony S.; Sanetrik, Mark D.; Scott, Robert C.; Keller, Donald F.; Cole, Stanley R.; Coulson, David A.

    2007-01-01

    This paper presents the research objectives, summarizes the pre-wind-tunnel-test experimental results to date, summarizes the analytical predictions to date, and outlines the wind-tunnel-test plans for an aeroservoelastic semispan wind-tunnel model. The model is referred to as the Supersonic Semispan Transport (S4T) Active Controls Testbed (ACT) and is based on a supersonic cruise configuration. The model has three hydraulically-actuated surfaces (all-movable horizontal tail, all-movable ride control vane, and aileron) for active controls. The model is instrumented with accelerometers, unsteady pressure transducers, and strain gages and will be mounted on a 5-component sidewall balance. The model will be tested twice in the Langley Transonic Dynamics Tunnel (TDT). The first entry will be an "open-loop" model-characterization test; the second entry will be a "closed-loop" test during which active flutter suppression, gust load alleviation and ride quality control experiments will be conducted.

  15. Analysis of wind-tunnel stability and control tests in terms of flying qualities of full-scale airplanes

    NASA Technical Reports Server (NTRS)

    Kayten, Gerald G

    1945-01-01

    The analysis of results of wind-tunnel stability and control tests of powered airplane models in terms of the flying qualities of full-scale airplanes is advocated. In order to indicated the topics upon which comments are considered desirable in the report of a wind-tunnel stability and control investigation and to demonstrate the nature of the suggested analysis, the present NACA flying-qualities requirements are discussed in relation to wind-tunnel tests. General procedures for the estimation of flying qualities from wind-tunnel tests are outlined.

  16. Rotary Balance Wind Tunnel Testing for the FASER Flight Research Aircraft

    NASA Technical Reports Server (NTRS)

    Denham, Casey; Owens, D. Bruce

    2016-01-01

    Flight dynamics research was conducted to collect and analyze rotary balance wind tunnel test data in order to improve the aerodynamic simulation and modeling of a low-cost small unmanned aircraft called FASER (Free-flying Aircraft for Sub-scale Experimental Research). The impetus for using FASER was to provide risk and cost reduction for flight testing of more expensive aircraft and assist in the improvement of wind tunnel and flight test techniques, and control laws. The FASER research aircraft has the benefit of allowing wind tunnel and flight tests to be conducted on the same model, improving correlation between wind tunnel, flight, and simulation data. Prior wind tunnel tests include a static force and moment test, including power effects, and a roll and yaw damping forced oscillation test. Rotary balance testing allows for the calculation of aircraft rotary derivatives and the prediction of steady-state spins. The rotary balance wind tunnel test was conducted in the NASA Langley Research Center (LaRC) 20-Foot Vertical Spin Tunnel (VST). Rotary balance testing includes runs for a set of given angular rotation rates at a range of angles of attack and sideslip angles in order to fully characterize the aircraft rotary dynamics. Tests were performed at angles of attack from 0 to 50 degrees, sideslip angles of -5 to 10 degrees, and non-dimensional spin rates from -0.5 to 0.5. The effects of pro-spin elevator and rudder deflection and pro- and anti-spin elevator, rudder, and aileron deflection were examined. The data are presented to illustrate the functional dependence of the forces and moments on angle of attack, sideslip angle, and angular rate for the rotary contributions to the forces and moments. Further investigation is necessary to fully characterize the control effectors. The data were also used with a steady state spin prediction tool that did not predict an equilibrium spin mode.

  17. Wind Tunnel and Hover Performance Test Results for Multicopter UAS Vehicles

    NASA Technical Reports Server (NTRS)

    Russell, Carl R.; Jung, Jaewoo; Willink, Gina; Glasner, Brett

    2016-01-01

    There is currently a lack of published data for the performance of multicopter unmanned aircraft system (UAS) vehicles, such as quadcopters and octocopters, often referred to collectively as drones. With the rapidly increasing popularity of multicopter UAS, there is interest in better characterizing the performance of this type of aircraft. By studying the performance of currently available vehicles, it will be possible to develop models for vehicles at this scale that can accurately predict performance and model trajectories. This paper describes a wind tunnel test that was recently performed in the U.S. Army's 7- by 10-ft Wind Tunnel at NASA Ames Research Center. During this wind tunnel entry, five multicopter UAS vehicles were tested to determine forces and moments as well as electrical power as a function of wind speed, rotor speed, and vehicle attitude. The test is described here in detail, and a selection of the key results from the test is presented.

  18. Validation of US3D for Capsule Aerodynamics using 05-CA Wind Tunnel Test Data

    NASA Technical Reports Server (NTRS)

    Schwing, Alan

    2012-01-01

    Several comparisons of computational fluid dynamics to wind tunnel test data are shown for the purpose of code validation. The wind tunnel test, 05-CA, uses a 7.66% model of NASA's Multi-Purpose Crew Vehicle in the 11-foot test section of the Ames Unitary Plan Wind tunnel. A variety of freestream conditions over four Mach numbers and three angles of attack are considered. Test data comparisons include time-averaged integrated forces and moments, time-averaged static pressure ports on the surface, and Strouhal Number. The applicability of the US3D code to subsonic and transonic flow over a bluff body is assessed on a comprehensive data set. With close comparison, this work validates US3D for highly separated flows similar to those examined here.

  19. Solid rocket booster sting interference wind tunnel test analysis, appendix D

    NASA Technical Reports Server (NTRS)

    Conine, B.; Boyle, W.

    1982-01-01

    Additional analyses of wind tunnel test results from SRB sting interference test TWT 660 and HRWT 042 were conducted to evaluate the sting interference that may be present in the Space Shuttle SRB reentry aerodynamic math model. Additional wind tunnel data was obtained at higher angles of attack from test program TWT 660 and test program HRWT 042. The additional data were analyzed to evaluate the procedures used to fair the data in the development of the SRB reentry aerodynamic data Tape no. 5.

  20. CFD wind tunnel test: Field velocity patterns of wind on a building with a refuge floor

    NASA Astrophysics Data System (ADS)

    Cheng, C. K.; Yuen, K. K.; Lam, K. M.; Lo, S. M.

    2005-10-01

    This paper reports a CFD wind tunnel study of wind patterns on a square-plan building with a refuge floor at its mid-height level. In this study, a technique of using calibrated power law equations of velocity and turbulent intensity applied as the boundary conditions in CFD wind tunnel test is being evaluated by the physical wind tunnel data obtained by the Principal Author with wind blowing perpendicularly on the building without a refuge floor. From the evaluated results, an optimised domain of flow required to produce qualitative agreement between the wind tunnel data and simulated results is proposed in this paper. Simulated results with the evaluated technique are validated by the wind tunnel data obtained by the Principal Author. The results contribute to an understanding of the fundamental behaviour of wind flow in a refuge floor when wind is blowing perpendicularly on the building. Moreover, the results reveal that the designed natural ventilation of a refuge floor may not perform desirably when the wind speed on the level is low. Under this situation, the refuge floor may become unsafe if smoke was dispersed in the leeward side of the building at a level immediately below the refuge floor.

  1. Orbiter/shuttle carrier aircraft separation: Wind tunnel, simulation, and flight test overview and results

    NASA Technical Reports Server (NTRS)

    Homan, D. J.; Denison, D. E.; Elchert, K. C.

    1980-01-01

    A summary of the approach and landing test phase of the space shuttle program is given from the orbiter/shuttle carrier aircraft separation point of view. The data and analyses used during the wind tunnel testing, simulation, and flight test phases in preparation for the orbiter approach and landing tests are reported.

  2. Results of design studies and wind tunnel tests of high-aspect-ratio supercritical wings for an energy efficient transport

    NASA Technical Reports Server (NTRS)

    Steckel, D. K.; Dahlin, J. A.; Henne, P. A.

    1980-01-01

    These basic characteristics of critical wings included wing area, aspect ratio, average thickness, and sweep as well as practical constraints on the planform and thickness near the wing root to allow for the landing gear. Within these constraints, a large matrix of wing designs was studied with spanwise variations in the types of airfoils and distribution of lift as well as some small planform changes. The criteria by which the five candidate wings were chosen for testing were the cruise and buffet characteristics in the transonic regime and the compatibility of the design with low speed (high-lift) requirements. Five wing-wide-body configurations were tested in the NASA Ames 11-foot transonic wind tunnel. Nacelles and pylons, flap support fairings, tail surfaces, and an outboard aileron were also tested on selected configurations.

  3. The cryogenic wind tunnel

    NASA Technical Reports Server (NTRS)

    Kilgore, R. A.

    1976-01-01

    Based on theoretical studies and experience with a low speed cryogenic tunnel and with a 1/3-meter transonic cryogenic tunnel, the cryogenic wind tunnel concept was shown to offer many advantages with respect to the attainment of full scale Reynolds number at reasonable levels of dynamic pressure in a ground based facility. The unique modes of operation available in a pressurized cryogenic tunnel make possible for the first time the separation of Mach number, Reynolds number, and aeroelastic effects. By reducing the drive-power requirements to a level where a conventional fan drive system may be used, the cryogenic concept makes possible a tunnel with high productivity and run times sufficiently long to allow for all types of tests at reduced capital costs and, for equal amounts of testing, reduced total energy consumption in comparison with other tunnel concepts.

  4. Introduction to cryogenic wind tunnels

    NASA Technical Reports Server (NTRS)

    Goodyer, M. J.

    1985-01-01

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

  5. Flutter suppression digital control law design and testing for the AFW wind-tunnel model

    NASA Technical Reports Server (NTRS)

    Mukhopadhyay, Vivek

    1992-01-01

    Design of a control law for simultaneously suppressing the symmetric and antisymmetric flutter modes of a string mounted fixed-in-roll aeroelastic wind tunnel model is described. The flutter suppression control law was designed using linear quadratic Gaussian theory and involved control law order reduction, a gain root-locus study, and the use of previous experimental results. A 23 percent increase in open-loop flutter dynamic pressure was demonstrated during the wind tunnel test. Rapid roll maneuvers at 11 percent above the symmetric flutter boundary were also performed when the model was in a free-to-roll configuration.

  6. Flutter suppression digital control law design and testing for the AFW wind tunnel model

    NASA Technical Reports Server (NTRS)

    Mukhopadhyay, Vivek

    1994-01-01

    The design of a control law for simultaneously suppressing the symmetric and antisymmetric flutter modes of a sting mounted fixed-in-roll aeroelastic wind-tunnel model is described. The flutter suppression control law was designed using linear quadratic Gaussian theory, and it also involved control law order reduction, a gain root-locus study, and use of previous experimental results. A 23 percent increase in the open-loop flutter dynamic pressure was demonstrated during the wind-tunnel test. Rapid roll maneuvers at 11 percent above the symmetric flutter boundary were also performed when the model was in a free-to-roll configuration.

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

    NASA Technical Reports Server (NTRS)

    Lewis, M. C.

    1984-01-01

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

  8. Analysis of the high Reynolds number 2D tests on a wind turbine airfoil performed at two different wind tunnels

    NASA Astrophysics Data System (ADS)

    Pires, O.; Munduate, X.; Ceyhan, O.; Jacobs, M.; Madsen, J.; Schepers, J. G.

    2016-09-01

    2D wind tunnel tests at high Reynolds numbers have been done within the EU FP7 AVATAR project (Advanced Aerodynamic Tools of lArge Rotors) on the DU00-W-212 airfoil and at two different test facilities: the DNW High Pressure Wind Tunnel in Gottingen (HDG) and the LM Wind Power in-house wind tunnel. Two conditions of Reynolds numbers have been performed in both tests: 3 and 6 million. The Mach number and turbulence intensity values are similar in both wind tunnels at the 3 million Reynolds number test, while they are significantly different at 6 million Reynolds number. The paper presents a comparison of the data obtained from the two wind tunnels, showing good repeatability at 3 million Reynolds number and differences at 6 million Reynolds number that are consistent with the different Mach number and turbulence intensity values.

  9. Analysis of flight and wind-tunnel tests on Udet airplanes with reference to spinning characteristics

    NASA Technical Reports Server (NTRS)

    Herrmann, H

    1929-01-01

    This report presents an analysis of results of wind-tunnel tests conducted at the D.V.L. Values were determined for the effectiveness of all the controls at various angles of attack. The autorotation was studied by subjecting the rotating model to an air blast.

  10. Turbulence Factors of NACA Wind Tunnels as Determined by Sphere Tests

    NASA Technical Reports Server (NTRS)

    Platt, Robert C

    1937-01-01

    Report presents the results of drag and pressure tests of spheres having diameters of 2, 4, 6, 8, 10, and 12 inches in eight NACA wind tunnels, in the air ahead of the carriage in the NACA tank, and beneath an autogiro in flight .

  11. Unsteady Aerodynamics Experiment Phase VI: Wind Tunnel Test Configurations and Available Data Campaigns

    SciTech Connect

    Hand, M. M.; Simms, D. A.; Fingersh, L. J.; Jager, D. W.; Cotrell, J. R.; Schreck, S.; Larwood, S. M.

    2001-12-01

    The primary objective of the insteady aerodynamics experiment was to provide information needed to quantify the full-scale, three-dimensional aerodynamic behavior of horizontal-axis wind turbines. This report is intended to familiarize the user with the entire scope of the wind tunnel test and to support the use of the resulting data.

  12. Advanced Background Subtraction Applied to Aeroacoustic Wind Tunnel Testing

    NASA Technical Reports Server (NTRS)

    Bahr, Christopher J.; Horne, William C.

    2015-01-01

    An advanced form of background subtraction is presented and applied to aeroacoustic wind tunnel data. A variant of this method has seen use in other fields such as climatology and medical imaging. The technique, based on an eigenvalue decomposition of the background noise cross-spectral matrix, is robust against situations where isolated background auto-spectral levels are measured to be higher than levels of combined source and background signals. It also provides an alternate estimate of the cross-spectrum, which previously might have poor definition for low signal-to-noise ratio measurements. Simulated results indicate similar performance to conventional background subtraction when the subtracted spectra are weaker than the true contaminating background levels. Superior performance is observed when the subtracted spectra are stronger than the true contaminating background levels. Experimental results show limited success in recovering signal behavior for data where conventional background subtraction fails. They also demonstrate the new subtraction technique's ability to maintain a proper coherence relationship in the modified cross-spectral matrix. Beam-forming and de-convolution results indicate the method can successfully separate sources. Results also show a reduced need for the use of diagonal removal in phased array processing, at least for the limited data sets considered.

  13. Background Pressure Profiles for Sonic Boom Vehicle Testing in the NASA Glenn 8- by 6-Foot Supersonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Castner, Raymond; Shaw, Stephen; Adamson, Eric; Simerly, Stephanie

    2013-01-01

    In an effort to identify test facilities that offer sonic boom measurement capabilities, an exploratory test program was initiated using wind tunnels at NASA research centers. The subject of this report is the sonic boom pressure rail data collected in the Glenn Research Center 8- by 6-Foot Supersonic Wind Tunnel. The purpose is to summarize the lessons learned based on the test activity, specifically relating to collecting sonic boom data which has a large amount of spatial pressure variation. The wind tunnel background pressure profiles are presented as well as data which demonstrated how both wind tunnel Mach number and model support-strut position affected the wind tunnel background pressure profile. Techniques were developed to mitigate these effects and are presented.

  14. Validation of the Lockheed Martin Morphing Concept with Wind Tunnel Testing

    NASA Technical Reports Server (NTRS)

    Ivanco, Thomas G.; Scott, Robert C.; Love, Michael H.; Zink Scott; Weisshaar, Terrence A.

    2007-01-01

    The Morphing Aircraft Structures (MAS) program is a Defense Advanced Research Projects Agency (DARPA) led effort to develop morphing flight vehicles capable of radical shape change in flight. Two performance parameters of interest are loiter time and dash speed as these define the persistence and responsiveness of an aircraft. The geometrical characteristics that optimize loiter time and dash speed require different geometrical planforms. Therefore, radical shape change, usually involving wing area and sweep, allows vehicle optimization across many flight regimes. The second phase of the MAS program consisted of wind tunnel tests conducted at the NASA Langley Transonic Dynamics Tunnel to demonstrate two morphing concepts and their enabling technologies with large-scale semi-span models. This paper will focus upon one of those wind tunnel tests that utilized a model developed by Lockheed Martin Aeronautics Company (LM). Wind tunnel success criteria were developed by NASA to support the DARPA program objectives. The primary focus of this paper will be the demonstration of the DARPA objectives by systematic evaluation of the wind tunnel model performance relative to the defined success criteria. This paper will also provide a description of the LM model and instrumentation, and document pertinent lessons learned. Finally, as part of the success criteria, aeroelastic characteristics of the LM derived MAS vehicle are also addressed. Evaluation of aeroelastic characteristics is the most detailed criterion investigated in this paper. While no aeroelastic instabilities were encountered as a direct result of the morphing design or components, several interesting and unexpected aeroelastic phenomenon arose during testing.

  15. Experimental Results from the Active Aeroelastic Wing Wind Tunnel Test Program

    NASA Technical Reports Server (NTRS)

    Heeg, Jennifer; Spain, Charles V.; Florance, James R.; Wieseman, Carol D.; Ivanco, Thomas G.; DeMoss, Joshua; Silva, Walter A.; Panetta, Andrew; Lively, Peter; Tumwa, Vic

    2005-01-01

    The Active Aeroelastic Wing (AAW) program is a cooperative effort among NASA, the Air Force Research Laboratory and the Boeing Company, encompassing flight testing, wind tunnel testing and analyses. The objective of the AAW program is to investigate the improvements that can be realized by exploiting aeroelastic characteristics, rather than viewing them as a detriment to vehicle performance and stability. To meet this objective, a wind tunnel model was crafted to duplicate the static aeroelastic behavior of the AAW flight vehicle. The model was tested in the NASA Langley Transonic Dynamics Tunnel in July and August 2004. The wind tunnel investigation served the program goal in three ways. First, the wind tunnel provided a benchmark for comparison with the flight vehicle and various levels of theoretical analyses. Second, it provided detailed insight highlighting the effects of individual parameters upon the aeroelastic response of the AAW vehicle. This parameter identification can then be used for future aeroelastic vehicle design guidance. Third, it provided data to validate scaling laws and their applicability with respect to statically scaled aeroelastic models.

  16. Space shuttle phase B wind tunnel model and test information. Volume 2: Orbiter configuration

    NASA Technical Reports Server (NTRS)

    Glynn, J. L.; Poucher, D. E.

    1988-01-01

    Archived wind tunnel test data are available for flyback booster or other alternative recoverable configurations as well as reusable orbiters studied during initial development (Phase B) of the Space Shuttle. Considerable wind tunnel data was acquired by the competing contractors and the NASA centers for an extensive variety of configurations with an array of wing and body planforms. All contractor and NASA wind tunnel test data acquired in the Phase B development have been compiled into a data base and are available for applying to current winged flyback or recoverable booster aerodynamic studies. The Space Shuttle Phase B Wind Tunnel Data Base is structured by vehicle component and configuration type. Basic components include the booster, the orbiter, and the launch vehicle. Booster configuration types include straight and delta wings, canard, cylindrical, retro-glide and twin body. Orbiter configuration types include straight and delta wings, lifting body, drop tanks, and double delta wings. Launch configuration types include booster and orbiter components in various stacked and tandem combinations.

  17. Space shuttle phase B wind tunnel model and test information. Volume 1: Booster configuration

    NASA Technical Reports Server (NTRS)

    Glynn, J. L.; Poucher, D. E.

    1988-01-01

    Archived wind tunnel test data are available for flyback booster or other alternative recoverable configurations as well as reusable orbiters studied during initial development (Phase B) of the Space Shuttle. Considerable wind tunnel data was acquired by the competing contractors and the NASA Centers for an extensive variety of configurations with an array of wing and body planforms. All contractor and NASA wind tunnel test data acquired in the Phase B development have been compiled into a database and are available for application to current winged flyback or recoverable booster aerodynamic studies. The Space Shuttle Phase B Wind Tunnel Database is structured by vehicle component and configuration type. Basic components include the booster, the orbiter and the launch vehicle. Booster configuration types include straight and delta wings, canard, cylindrical, retroglide and twin body. Orbiter configuration types include straight and delta wings, lifting body, drop tanks, and double delta wings. Launch configurations include booster and orbiter components in various stacked and tandem combinations. This is Volume 1 (Part 1) of the report -- Booster Configuration.

  18. Space shuttle phase B wind tunnel model and test information. Volume 2: Orbiter configuration

    NASA Technical Reports Server (NTRS)

    Glynn, J. L.; Poucher, D. E.

    1988-01-01

    Archived wind tunnel test data are available for flyback booster or other alternate recoverable configurations as well as reusable orbiters studied during initial development (Phase B) of the Space Shuttle. Considerable wind tunnel data was acquired by the competing contractors and the NASA centers for an extensive variety of configurations with an array of wing and body planforms. All contractor and NASA wind tunnel test data acquiredin the Phase B development have been compiled into a database and are available for applying to current winged flyback or recoverable booster aerodynamic studies. The Space Shuttle Phase B Wind Tunnel Database is structured by vehicle component and configuration type. Basic components include the booster, the orbiter, and the launch vehicle. Booster configuration types include straight and delta wings, canard, cylindrical, retroglide, and twin body. Orbiter configuration types include straight and delta wings, lifting body, drop tanks, and double delta wings. Launch configration types include booster and orbiter components in various stacked and tandom combinations. The digital database consists of 220 files of data containing basic tunnel recorded data.

  19. Space shuttle phase B wind tunnel model and test information. Volume 1: Booster configuration

    NASA Technical Reports Server (NTRS)

    Glynn, J. L.; Poucher, D. E.

    1988-01-01

    Archived wind tunnel test data are available for flyback booster or other alternative recoverable configurations as well as reusable orbiters studied during initial development (Phase B) of the Space Shuttle. Considerable wind tunnel data was acquired by the competing contractors and the NASA Centers for an extensive variety of configurations with an array of wing and body planforms. All contractor and NASA wind tunnel test data acquired in the Phase B development have been compiled into a database and are available for application to current winged flyback or recoverable booster aerodynamic studies. The Space Shuttle Phase B Wind Tunnel Database is structured by vehicle component and configuration type. Basic components include the booster, the orbiter, and the launch vehicle. Booster configuration types include straight and delta wings, canard, cylindrical, retroglide and twin body. Orbiter configuration types include straight and delta wings, lifting body, drop tanks and double delta wings. Launch configurations include booster and orbiter components in various stacked and tandem combinations. This is Volume 1 (Part 2) of the report -- Booster Configuration.

  20. Comparison of options for reduction of noise in the test section of the NASA Langley 4x7m wind tunnel, including reduction of nozzle area

    NASA Technical Reports Server (NTRS)

    Hayden, R. E.

    1984-01-01

    The acoustically significant features of the NASA 4X7m wind tunnel and the Dutch-German DNW low speed tunnel are compared to illustrate the reasons for large differences in background noise in the open jet test sections of the two tunnels. Also introduced is the concept of reducing test section noise levels through fan and turning vane source reductions which can be brought about by reducing the nozzle cross sectional area, and thus the circuit mass flow for a particular exit velocity. The costs and benefits of treating sources, paths, and changing nozzle geometry are reviewed.

  1. Low-speed wind-tunnel study of the high-angle-of-attack stability and control characteristics of a cranked-arrow-wing fighter configuration

    NASA Technical Reports Server (NTRS)

    Grafton, S. B.

    1984-01-01

    The low-speed, high-angle-of-attack stability and control characteristics of a fighter configuration incorporating a cranked arrow wing were investigated in the Langley 30- by 60-foot tunnel as part of a NASA/General Dynamics cooperative research program to investigate the application of advanced wing designs to combat aircraft. Tests were conducted on a baseline configuration and on several modified configurations. The results show that the baseline configuration exhibited a high level of maximum lift but displayed undesirable longitudinal and lateral-directional stability characteristics at high angles of attack. Various wing modifications were made which improved the longitudinal and lateral-directional stability characteristics of the configuration at high angles of attack. However, most of the modifications were detrimental to maximum lift.

  2. Wind Tunnel Interference Effects on Tilt Rotor Testing Using Computational Fluid Dynamics

    NASA Technical Reports Server (NTRS)

    Koning, Witold J. F.

    2015-01-01

    Experimental techniques to measure rotorcraft aerodynamic performance are widely used. However, most of them are either unable to capture interference effects from bodies, or require an extremely large computational budget. The objective of the present research is to develop an XV-15 Tilt Rotor Research Aircraft rotor model for investigation of wind tunnel wall interference using a novel Computational Fluid Dynamics (CFD) solver for rotorcraft, RotCFD. In RotCFD, a mid-fidelity URANS solver is used with an incompressible flow model and a realizable k-e turbulence model. The rotor is, however, not modeled using a computationally expensive, unsteady viscous body-fitted grid, but is instead modeled using a blade element model with a momentum source approach. Various flight modes of the XV-15 isolated rotor, including hover, tilt and airplane mode, have been simulated and correlated to existing experimental and theoretical data. The rotor model is subsequently used for wind tunnel wall interference simulations in the National Full-Scale Aerodynamics Complex (NFAC) at NASA Ames Research Center in California. The results from the validation of the isolated rotor performance showed good correlation with experimental and theoretical data. The results were on par with known theoretical analyses. In RotCFD the setup, grid generation and running of cases is faster than many CFD codes, which makes it a useful engineering tool. Performance predictions need not be as accurate as high-fidelity CFD codes, as long as wall effects can be properly simulated. For both test sections of the NFAC wall interference was examined by simulating the XV-15 rotor in the test section of the wind tunnel and with an identical grid but extended boundaries in free field. Both cases were also examined with an isolated rotor or with the rotor mounted on the modeled geometry of the Tiltrotor Test Rig (TTR). A 'quasi linear trim' was used to trim the thrust for the rotor to compare the power as a unique

  3. Wind-tunnel development of an SR-71 aerospike rocket flight test configuration

    NASA Technical Reports Server (NTRS)

    Smith, Stephen C.; Shirakata, Norm; Moes, Timothy R.; Cobleigh, Brent R.; Conners, Timothy H.

    1996-01-01

    A flight experiment has been proposed to investigate the performance of an aerospike rocket motor installed in a lifting body configuration. An SR-71 airplane would be used to carry the aerospike configuration to the desired flight test conditions. Wind-tunnel tests were completed on a 4-percent scale SR-71 airplane with the aerospike pod mounted in various locations on the upper fuselage. Testing was accomplished using sting and blade mounts from Mach 0.6 to Mach 3.2. Initial test objectives included assessing transonic drag and supersonic lateral-directional stability and control. During these tests, flight simulations were run with wind-tunnel data to assess the acceptability of the configurations. Early testing demonstrated that the initial configuration with the aerospike pod near the SR-71 center of gravity was unsuitable because of large nosedown pitching moments at transonic speeds. The excessive trim drag resulting from accommodating this pitching moment far exceeded the excess thrust capability of the airplane. Wind-tunnel testing continued in an attempt to find a configuration suitable for flight test. Multiple configurations were tested. Results indicate that an aft-mounted model configuration possessed acceptable performance, stability, and control characteristics.

  4. Methodology for the Assessment of 3D Conduction Effects in an Aerothermal Wind Tunnel Test

    NASA Technical Reports Server (NTRS)

    Oliver, Anthony Brandon

    2010-01-01

    This slide presentation reviews a method for the assessment of three-dimensional conduction effects during test in a Aerothermal Wind Tunnel. The test objectives were to duplicate and extend tests that were performed during the 1960's on thermal conduction on proturberance on a flat plate. Slides review the 1D versus 3D conduction data reduction error, the analysis process, CFD-based analysis, loose coupling method that simulates a wind tunnel test run, verification of the CFD solution, Grid convergence, Mach number trend, size trends, and a Sumary of the CFD conduction analysis. Other slides show comparisons to pretest CFD at Mach 1.5 and 2.16 and the geometries of the models and grids.

  5. Shake test results of the MDHC test stand in the 40- by 80-foot wind tunnel

    NASA Technical Reports Server (NTRS)

    Lau, Benton H.; Peterson, Randall

    1994-01-01

    A shake test was conducted to determine the modal properties of the MDHC (McDonnell Douglas Helicopter Company) test stand installed in the 40- by 80- Foot Wind Tunnel at Ames Research Center. The shake test was conducted for three wind-tunnel balance configurations with and without balance dampers, and with the snubber engagement to lock the balance frame. A hydraulic shaker was used to apply random excitation at the rotor hub in the longitudinal and lateral directions. A GenRad 2515 computer-aided test system computed the frequency response functions at the rotor hub and support struts. From these response functions, the modal properties, including the natural frequency, damping ratio, and mode shape were calculated. The critical modes with low damping ratios are identified as the test-stand second longitudinal mode for the dampers-off configuration, the test-stand yaw mode for the dampers-on configuration, and the test stand first longitudinal mode for the balance-frame locked configuration.

  6. Space shuttle solid rocket booster sting interference wind tunnel test analysis

    NASA Technical Reports Server (NTRS)

    Conine, B.; Boyle, W.

    1981-01-01

    Wind tunnel test results from shuttle solid rocket booster (SRB) sting interference tests were evaluated, yielding the general influence of the sting on the normal force and pitching moment coefficients and the side force and yawing moment coefficients. The procedures developed to determine the sting interference, the development of the corrected aerodynamic data, and the development of a new SRB aerodynamic mathematical model are documented.

  7. Evaluation of a wind-tunnel gust response technique including correlations with analytical and flight test results

    NASA Technical Reports Server (NTRS)

    Redd, L. T.; Hanson, P. W.; Wynne, E. C.

    1979-01-01

    A wind tunnel technique for obtaining gust frequency response functions for use in predicting the response of flexible aircraft to atmospheric turbulence is evaluated. The tunnel test results for a dynamically scaled cable supported aeroelastic model are compared with analytical and flight data. The wind tunnel technique, which employs oscillating vanes in the tunnel throat section to generate a sinusoidally varying flow field around the model, was evaluated by use of a 1/30 scale model of the B-52E airplane. Correlation between the wind tunnel results, flight test results, and analytical predictions for response in the short period and wing first elastic modes of motion are presented.

  8. A Free-flight Wind Tunnel for Aerodynamic Testing at Hypersonic Speeds

    NASA Technical Reports Server (NTRS)

    Seiff, Alvin

    1954-01-01

    The supersonic free-flight wind tunnel is a facility at the Ames Laboratory of the NACA in which aerodynamic test models are gun-launched at high speed and directed upstream through the test section of a supersonic wind tunnel. In this way, test Mach numbers up to 10 have been attained and indications are that still higher speeds will be realized. An advantage of this technique is that the air and model temperatures simulate those of flight through the atmosphere. Also the Reynolds numbers are high. Aerodynamic measurements are made from photographic observation of the model flight. Instruments and techniques have been developed for measuring the following aerodynamic properties: drag, initial lift-curve slope, initial pitching-moment-curve slope, center of pressure, skin friction, boundary-layer transition, damping in roll, and aileron effectiveness. (author)

  9. Comparative wind tunnel tests of NACA 23024 airfoils with several aileron and spoiler configurations

    NASA Technical Reports Server (NTRS)

    Wentz, W. H., Jr.; Snyder, M. H.

    1995-01-01

    This paper reviews research efforts at Wichita State University sponsored by NASA Lewis Research Center to design and evaluate aerodynamic braking devices which will be smaller and lighter than full-chord blade pitch control. Devices evaluated include a variety of aileron configurations, and spoilers located at both trailing edge and near the leading edge. The paper discusses analytical modeling, wind tunnel tests, and for some configurations, full-scale rotor tests. Current designs have not provided adequate control power at high angles of attack (low tip-speed-ratios). The reasons for these limitations are discussed. Analysis and wind tunnel test data indicate that several options are available to the designer to provide aerodynamic slowdown without full-chord pitch control. Three options are suggested; adding venting in front of the control surface hingeline, using spoilers located near the leading edge, and using a two-piece control combining downward deflection inboard with upward deflection outboard.

  10. Wind tunnel test of a variable-diameter tiltrotor (VDTR) model

    NASA Technical Reports Server (NTRS)

    Matuska, David; Dale, Allen; Lorber, Peter

    1994-01-01

    This report documents the results from a wind tunnel test of a 1/6th scale Variable Diameter Tiltrotor (VDTR). This test was a joint effort of NASA Ames and Sikorsky Aircraft. The objective was to evaluate the aeroelastic and performance characteristics of the VDTR in conversion, hover, and cruise. The rotor diameter and nacelle angle of the model were remotely changed to represent tiltrotor operating conditions. Data is presented showing the propulsive force required in conversion, blade loads, angle of attack stability and simulated gust response, and hover and cruise performance. This test represents the first wind tunnel test of a variable diameter rotor applied to a tiltrotor concept. The results confirm some of the potential advantages of the VDTR and establish the variable diameter rotor a viable candidate for an advanced tiltrotor. This wind tunnel test successfully demonstrated the feasibility of the Variable Diameter rotor for tilt rotor aircraft. A wide range of test points were taken in hover, conversion, and cruise modes. The concept was shown to have a number of advantages over conventional tiltrotors such as reduced hover downwash with lower disk loading and significantly reduced longitudinal gust response in cruise. In the conversion regime, a high propulsive force was demonstrated for sustained flight with acceptable blade loads. The VDTR demonstrated excellent gust response capabilities. The horizontal gust response correlated well with predictions revealing only half the response to turbulence of the conventional civil tiltrotor.

  11. Wind Tunnel Test Technique and Instrumentation Development at LaRC

    NASA Technical Reports Server (NTRS)

    Putnam, Lawrence E.

    1999-01-01

    LaRC has an aggressive test technique development program underway. This program has been developed using 3rd Generation R&D management techniques and is a closely coordinated program between suppliers and wind tunnel operators- wind tunnel customers' informal input relative to their needs has been an essential ingredient in developing the research portfolio. An attempt has been made to balance this portfolio to meet near term and long term test technique needs. Major efforts are underway to develop techniques for determining model wing twist and location of boundary layer transition in the NTF (National Transonic Facility). The foundation of all new instrumentation developments, procurements, and upgrades will be based on uncertainty analysis.

  12. A New Position Measurement System Using a Motion-Capture Camera for Wind Tunnel Tests

    PubMed Central

    Park, Hyo Seon; Kim, Ji Young; Kim, Jin Gi; Choi, Se Woon; Kim, Yousok

    2013-01-01

    Considering the characteristics of wind tunnel tests, a position measurement system that can minimize the effects on the flow of simulated wind must be established. In this study, a motion-capture camera was used to measure the displacement responses of structures in a wind tunnel test, and the applicability of the system was tested. A motion-capture system (MCS) could output 3D coordinates using two-dimensional image coordinates obtained from the camera. Furthermore, this remote sensing system had some flexibility regarding lab installation because of its ability to measure at relatively long distances from the target structures. In this study, we performed wind tunnel tests on a pylon specimen and compared the measured responses of the MCS with the displacements measured with a laser displacement sensor (LDS). The results of the comparison revealed that the time-history displacement measurements from the MCS slightly exceeded those of the LDS. In addition, we confirmed the measuring reliability of the MCS by identifying the dynamic properties (natural frequency, damping ratio, and mode shape) of the test specimen using system identification methods (frequency domain decomposition, FDD). By comparing the mode shape obtained using the aforementioned methods with that obtained using the LDS, we also confirmed that the MCS could construct a more accurate mode shape (bending-deflection mode shape) with the 3D measurements. PMID:24064600

  13. Aeroservoelastic Testing of a Sidewall Mounted Free Flying Wind-Tunnel Model

    NASA Technical Reports Server (NTRS)

    Scott, Robert C.; Vetter, Travis K.; Penning, Kevin B.; Coulson, David A.; Heeg, Jennifer

    2008-01-01

    A team comprised of the Air Force Research Laboratory (AFRL), Northrop Grumman, Lockheed Martin, and the NASA Langley Research Center conducted three j wind-tunnel tests in the Transonic Dynamics Tunnel to demonstrate active control technologies relevant to large, exible vehicles. In the rst of these three tests, a semispan, aeroelastically scaled, wind-tunnel model of a ying wing SensorCraft vehi- cle was mounted to a force balance to demonstrate gust load alleviation. In the second and third tests, the same wing was mated to a new, multi-degree-of-freedom, sidewall mount. This mount allowed the half-span model to translate vertically and pitch at the wing root, allowing better simulation of the full span vehicle's rigid-body modes. Gust Load Alleviation (GLA) and Body Freedom Flutter (BFF) suppression were successfully demonstrated. The rigid body degrees-of-freedom required that the model be own in the wind tunnel using an active control system. This risky mode of testing necessitated that a model arrestment system be integrated into the new mount. The safe and successful completion of these free-flying tests required the development and integration of custom hardware and software. This paper describes the many systems, software, and procedures that were developed as part of this effort.

  14. Fractional Factorial Experiment Designs to Minimize Configuration Changes in Wind Tunnel Testing

    NASA Technical Reports Server (NTRS)

    DeLoach, Richard; Cler, Daniel L.; Graham, Albert B.

    2002-01-01

    This paper serves as a tutorial to introduce the wind tunnel research community to configuration experiment designs that can satisfy resource constraints in a configuration study involving several variables, without arbitrarily eliminating any of them from the experiment initially. The special case of a configuration study featuring variables at two levels is examined in detail. This is the type of study in which each configuration variable has two natural states - 'on or off', 'deployed or not deployed', 'low or high', and so forth. The basic principles are illustrated by results obtained in configuration studies conducted in the Langley National Transonic Facility and in the ViGYAN Low Speed Tunnel in Hampton, Virginia. The crucial role of interactions among configuration variables is highlighted with an illustration of difficulties that can be encountered when they are not properly taken into account.

  15. Comparison of Tests on Air Propellers in Flight with Wind Tunnel Model Tests on Similar Forms

    NASA Technical Reports Server (NTRS)

    Durand, W F; Lesley, E P

    1926-01-01

    The purpose of this investigation was to determine the performance, characteristics, and coefficients of full-sized air propellers in flight and to compare these results with those derived from wind-tunnel tests on reduced scale models of similar geometrical form. The full-scale equipment comprised five propellers in combination with a VE-7 airplane and Wright E-4 engine. This part of the work was carried out at the Langley Memorial Aeronautical Laboratory, between May 1 and August 24, 1924, and was under the immediate charge of Mr. Lesley. The model or wind-tunnel part of the investigation was carried out at the Aerodynamic Laboratory of Stanford University and was under the immediate charge of Doctor Durand. A comparison of the curves for full-scale results with those derived from the model tests shows that while the efficiencies realized in flight are close to those derived from model tests, both thrust developed and power absorbed in flight are from 6 to 10 per cent greater than would be expected from the results of model tests.

  16. Low-speed wind tunnel investigation of the static stability and control characteristics of an advanced turboprop configuration with the propellers placed over the tail. M.S. Thesis

    NASA Technical Reports Server (NTRS)

    Rhodes, Graham Scott

    1990-01-01

    An exploratory wind tunnel investigation was performed in the 30 x 60 foot wind tunnel to determine the low speed static stability and control characteristics into the deep stall regime of an advanced turboprop aircraft with the propellers located over the horizontal tail. By this arrangement, the horizontal tail could potentially provide acoustic shielding to reduce the high community noise caused by the propeller blades. The current configuration was a generic turboprop model equipped with 1 foot diameter single rotating eight bladed propellers that were designed for efficient cruise operation at a Mach number of 0.8. The data presented is static force data. The effects of power on the configuration characteristics were generally favorable. An arrangement with the propellers rotating with the outboard blades moving down was found to have significantly higher installed thrust than an arrangement with the propellers rotating with the inboard blades moving down. The primary unfavorable effect was a large pitch trim change which occurred with power, but the trim change could be minimized with a proper configuration design.

  17. The characteristics of 78 related airfoil sections from tests in the variable-density wind tunnel

    NASA Technical Reports Server (NTRS)

    Jacobs, Eastman N; Ward, Kenneth E; Pinkerton, Robert M

    1933-01-01

    An investigation of a large group of related airfoils was made in the NACA variable-density wind tunnel at a large value of the Reynolds number. The tests were made to provide data that may be directly employed for a rational choice of the most suitable airfoil section for a given application. The variation of the aerodynamic characteristics with variations in thickness and mean-line form were systematically studied. (author)

  18. Evaluation tests of platinum resistance thermometers for a cryogenic wind tunnel application

    NASA Technical Reports Server (NTRS)

    Germain, E. F.; Compton, E. C.

    1984-01-01

    Thirty-one commercially designed platinum resistance thermometers were evaluated for applicability to stagnation temperature measurements between -190 C and +65 C in the Langley Research Center's National Transonic Facility. Evaluation tests included X-ray shadowgraphs, calibrations before and after aging, and time constant measurements. Two wire-wound low thermal mass probes of a conventional design were chosen as most suitable for this cryogenic wind tunnel application.

  19. Aerodynamic characteristics of forebody and nose strakes based on F-16 wind tunnel test experience. Volume 1: Summary and analysis

    NASA Technical Reports Server (NTRS)

    Smith, C. W.; Ralston, J. N.; Mann, H. W.

    1979-01-01

    The YF-16 and F-16 developmental wind tunnel test program was reviewed. Geometrical descriptions, general comments, representative data, and the initial efforts toward the development of design guides for the application of strakes to future aircraft are presented.

  20. Aerodynamic characteristics of forebody and nose strakes based on F-16 wind tunnel test experience. Volume 2: Data base

    NASA Technical Reports Server (NTRS)

    Smith, C. W.; Bhateley, I. C.

    1978-01-01

    The YF-16 and F-16 developmental wind tunnel test program was reviewed and all force data pertinent to the design of forebody and nose strakes extracted. A complete set of these data is presented without analysis.

  1. Comparison of Resource Requirements for a Wind Tunnel Test Designed with Conventional vs. Modern Design of Experiments Methods

    NASA Technical Reports Server (NTRS)

    DeLoach, Richard; Micol, John R.

    2011-01-01

    The factors that determine data volume requirements in a typical wind tunnel test are identified. It is suggested that productivity in wind tunnel testing can be enhanced by managing the inference error risk associated with evaluating residuals in a response surface modeling experiment. The relationship between minimum data volume requirements and the factors upon which they depend is described and certain simplifications to this relationship are realized when specific model adequacy criteria are adopted. The question of response model residual evaluation is treated and certain practical aspects of response surface modeling are considered, including inference subspace truncation. A wind tunnel test plan developed by using the Modern Design of Experiments illustrates the advantages of an early estimate of data volume requirements. Comparisons are made with a representative One Factor At a Time (OFAT) wind tunnel test matrix developed to evaluate a surface to air missile.

  2. Smart-actuated continuous moldline technology (CMT) mini wind tunnel test

    NASA Astrophysics Data System (ADS)

    Pitt, Dale M.; Dunne, James P.; Kilian, Kevin J.

    1999-07-01

    The Smart Aircraft and Marine Propulsion System Demonstration (SAMPSON) Program will culminate in two separate demonstrations of the application of Smart Materials and Structures technology. One demonstration will be for an aircraft application and the other for marine vehicles. The aircraft portion of the program will examine the application of smart materials to aircraft engine inlets which will deform the inlet in-flight in order to regulate the airflow rate into the engine. Continuous Moldline Technology (CMT), a load-bearing reinforced elastomer, will enable the use of smart materials in this application. The capabilities of CMT to withstand high-pressure subsonic and supersonic flows were tested in a sub-scale mini wind- tunnel. The fixture, used as the wind-tunnel test section, was designed to withstand pressure up to 100 psi. The top and bottom walls were 1-inch thick aluminum and the side walls were 1-inch thick LEXAN. High-pressure flow was introduced from the Boeing St. Louis poly-sonic wind tunnel supply line. CMT walls, mounted conformal to the upper and lower surfaces, were deflected inward to obtain a converging-diverging nozzle. The CMT walls were instrumented for vibration and deflection response. Schlieren photography was used to establish shock wave motion. Static pressure taps, embedded within one of the LEXAN walls, monitored pressure variation in the mini-wind tunnel. High mass flow in the exit region. This test documented the response of CMT technology in the presence of high subsonic flow and provided data to be used in the design of the SAMPSON Smart Inlet.

  3. Static and wind tunnel model tests for the development of externally blown flap noise reduction techniques

    NASA Technical Reports Server (NTRS)

    Pennock, A. P.; Swift, G.; Marbert, J. A.

    1975-01-01

    Externally blown flap models were tested for noise and performance at one-fifth scale in a static facility and at one-tenth scale in a large acoustically-treated wind tunnel. The static tests covered two flap designs, conical and ejector nozzles, third-flap noise-reduction treatments, internal blowing, and flap/nozzle geometry variations. The wind tunnel variables were triple-slotted or single-slotted flaps, sweep angle, and solid or perforated third flap. The static test program showed the following noise reductions at takeoff: 1.5 PNdB due to treating the third flap; 0.5 PNdB due to blowing from the third flap; 6 PNdB at flyover and 4.5 PNdB in the critical sideline plane (30 deg elevation) due to installation of the ejector nozzle. The wind tunnel program showed a reduction of 2 PNdB in the sideline plane due to a forward speed of 43.8 m/s (85 kn). The best combination of noise reduction concepts reduced the sideline noise of the reference aircraft at constant field length by 4 PNdB.

  4. Application of Rapid Prototyping Methods to High-Speed Wind Tunnel Testing

    NASA Technical Reports Server (NTRS)

    Springer, A. M.

    1998-01-01

    This study was undertaken in MSFC's 14-Inch Trisonic Wind Tunnel to determine if rapid prototyping methods could be used in the design and manufacturing of high speed wind tunnel models in direct testing applications, and if these methods would reduce model design/fabrication time and cost while providing models of high enough fidelity to provide adequate aerodynamic data, and of sufficient strength to survive the test environment. Rapid prototyping methods utilized to construct wind tunnel models in a wing-body-tail configuration were: fused deposition method using both ABS plastic and PEEK as building materials, stereolithography using the photopolymer SL-5170, selective laser sintering using glass reinforced nylon, and laminated object manufacturing using plastic reinforced with glass and 'paper'. This study revealed good agreement between the SLA model, the metal model with an FDM-ABS nose, an SLA nose, and the metal model for most operating conditions, while the FDM-ABS data diverged at higher loading conditions. Data from the initial SLS model showed poor agreement due to problems in post-processing, resulting in a different configuration. A second SLS model was tested and showed relatively good agreement. It can be concluded that rapid prototyping models show promise in preliminary aerodynamic development studies at subsonic, transonic, and supersonic speeds.

  5. Supersonic Retropropulsion CFD Validation with Ames Unitary Plan Wind Tunnel Test Data

    NASA Technical Reports Server (NTRS)

    Schauerhamer, Daniel G.; Zarchi, Kerry A.; Kleb, William L.; Edquist, Karl T.

    2013-01-01

    A validation study of Computational Fluid Dynamics (CFD) for Supersonic Retropropulsion (SRP) was conducted using three Navier-Stokes flow solvers (DPLR, FUN3D, and OVERFLOW). The study compared results from the CFD codes to each other and also to wind tunnel test data obtained in the NASA Ames Research Center 90 70 Unitary PlanWind Tunnel. Comparisons include surface pressure coefficient as well as unsteady plume effects, and cover a range of Mach numbers, levels of thrust, and angles of orientation. The comparisons show promising capability of CFD to simulate SRP, and best agreement with the tunnel data exists for the steadier cases of the 1-nozzle and high thrust 3-nozzle configurations.

  6. Dynamic response tests of inertial and optical wind-tunnel model attitude measurement devices

    NASA Technical Reports Server (NTRS)

    Buehrle, R. D.; Young, C. P., Jr.; Burner, A. W.; Tripp, J. S.; Tcheng, P.; Finley, T. D.; Popernack, T. G., Jr.

    1995-01-01

    Results are presented for an experimental study of the response of inertial and optical wind-tunnel model attitude measurement systems in a wind-off simulated dynamic environment. This study is part of an ongoing activity at the NASA Langley Research Center to develop high accuracy, advanced model attitude measurement systems that can be used in a dynamic wind-tunnel environment. This activity was prompted by the inertial model attitude sensor response observed during high levels of model vibration which results in a model attitude measurement bias error. Significant bias errors in model attitude measurement were found for the measurement using the inertial device during wind-off dynamic testing of a model system. The amount of bias present during wind-tunnel tests will depend on the amplitudes of the model dynamic response and the modal characteristics of the model system. Correction models are presented that predict the vibration-induced bias errors to a high degree of accuracy for the vibration modes characterized in the simulated dynamic environment. The optical system results were uncorrupted by model vibration in the laboratory setup.

  7. Data correlation and analysis of arc tunnel and wind tunnel tests of RSI joints and gaps. Volume 2: Data base

    NASA Technical Reports Server (NTRS)

    Christensen, H. E.; Kipp, H. W.

    1974-01-01

    Wind tunnel tests were conducted to determine the aerodynamic heating created by gaps in the reusable surface insulation (RSI) thermal protection system (TPS) for the space shuttle. The effects of various parameters of the RSI on convective heating characteristics are described. The wind tunnel tests provided a data base for accurate assessment of gap heating. Analysis and correlation of the data provide methods for predicting heating in the RSI gaps on the space shuttle.

  8. Assessing Videogrammetry for Static Aeroelastic Testing of a Wind-Tunnel Model

    NASA Technical Reports Server (NTRS)

    Spain, Charles V.; Heeg, Jennifer; Ivanco, Thomas G.; Barrows, Danny A.; Florance, James R.; Burner, Alpheus W.; DeMoss, Joshua; Lively, Peter S.

    2004-01-01

    The Videogrammetric Model Deformation (VMD) technique, developed at NASA Langley Research Center, was recently used to measure displacements and local surface angle changes on a static aeroelastic wind-tunnel model. The results were assessed for consistency, accuracy and usefulness. Vertical displacement measurements and surface angular deflections (derived from vertical displacements) taken at no-wind/no-load conditions were analyzed. For accuracy assessment, angular measurements were compared to those from a highly accurate accelerometer. Shewhart's Variables Control Charts were used in the assessment of consistency and uncertainty. Some bad data points were discovered, and it is shown that the measurement results at certain targets were more consistent than at other targets. Physical explanations for this lack of consistency have not been determined. However, overall the measurements were sufficiently accurate to be very useful in monitoring wind-tunnel model aeroelastic deformation and determining flexible stability and control derivatives. After a structural model component failed during a highly loaded condition, analysis of VMD data clearly indicated progressive structural deterioration as the wind-tunnel condition where failure occurred was approached. As a result, subsequent testing successfully incorporated near- real-time monitoring of VMD data in order to ensure structural integrity. The potential for higher levels of consistency and accuracy through the use of statistical quality control practices are discussed and recommended for future applications.

  9. Aeroservoelastic Testing of Free Flying Wind Tunnel Models Part 2: A Centerline Supported Fullspan Model Tested for Gust Load Alleviation

    NASA Technical Reports Server (NTRS)

    Scott, Robert C.; Vetter, Travis K.; Penning, Kevin B.; Coulson, David A.; Heeg, Jennifer

    2014-01-01

    This is part 2 of a two part document. Part 1 is titled: "Aeroservoelastic Testing of Free Flying Wind Tunnel Models Part 1: A Sidewall Supported Semispan Model Tested for Gust Load Alleviation and Flutter Suppression." A team comprised of the Air Force Research Laboratory (AFRL), Boeing, and the NASA Langley Research Center conducted three aeroservoelastic wind tunnel tests in the Transonic Dynamics Tunnel to demonstrate active control technologies relevant to large, flexible vehicles. In the first of these three tests, a full-span, aeroelastically scaled, wind tunnel model of a joined wing SensorCraft vehicle was mounted to a force balance to acquire a basic aerodynamic data set. In the second and third tests, the same wind tunnel model was mated to a new, two degree of freedom, beam mount. This mount allowed the full-span model to translate vertically and pitch. Trimmed flight at10 percent static margin and gust load alleviation were successfully demonstrated. The rigid body degrees of freedom required that the model be flown in the wind tunnel using an active control system. This risky mode of testing necessitated that a model arrestment system be integrated into the new mount. The safe and successful completion of these free-flying tests required the development and integration of custom hardware and software. This paper describes the many systems, software, and procedures that were developed as part of this effort. The balance and free flying wind tunnel tests will be summarized. The design of the trim and gust load alleviation control laws along with the associated results will also be discussed.

  10. Comparison of aircraft noise measured in flight test and in the NASA Ames 40- by 80-foot wind tunnel.

    NASA Technical Reports Server (NTRS)

    Atencio, A., Jr.; Soderman, P. T.

    1973-01-01

    A method to determine free-field aircraft noise spectra from wind-tunnel measurements has been developed. The crux of the method is the correction for reverberations. Calibrated loud speakers are used to simulate model sound sources in the wind tunnel. Corrections based on the difference between the direct and reverberant field levels are applied to wind-tunnel data for a wide range of aircraft noise sources. To establish the validity of the correction method, two research aircraft - one propeller-driven (YOV-10A) and one turbojet-powered (XV-5B) - were flown in free field and then tested in the wind tunnel. Corrected noise spectra from the two environments agree closely.

  11. Wind tunnel tests of a zero length, slotted-lip engine air inlet for a fixed nacelle V/STOL aircraft

    NASA Technical Reports Server (NTRS)

    Woollett, R. R.; Beck, W. E., Jr.; Glasgow, E. R.

    1982-01-01

    Zero length, slotted lip inlet performance and associated fan blade stresses were determined during model tests using a 20 inch diameter fan simulator in the NASA-LeRC 9 by 15 foot low speed wind tunnel. The model configuration variables consisted of inlet contraction ratio, slot width, circumferential extent of slot fillers, and length of a constant area section between the inlet throat and fan face. The inlet performance was dependent on slot gap width and relatively independent of inlet throat/fan face spacer length and slot flow blockage created by 90 degree slot fillers. Optimum performance was obtained at a slot gap width of 0.36 inch. The zero length, slotted lip inlet satisfied all critical low speed inlet operating requirements for fixed horizontal nacelles subsonic V/STOL aircraft.

  12. The application of cryogenics to high Reynolds number testing in wind tunnels. I - Evolution, theory, and advantages

    NASA Technical Reports Server (NTRS)

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

    1984-01-01

    During the time which has passed since the construction of the first wind tunnel in 1870, wind tunnels have been developed to a high degree of sophistication. However, their development has consistently failed to keep pace with the demands placed on them. One of the more serious problems to be found with existing transonic wind tunnels is their inability to test subscale aircraft models at Reynolds numbers sufficiently near full-scale values to ensure the validity of using the wind tunnel data to predict flight characteristics. The Reynolds number capability of a wind tunnel may be increased by a number of different approaches. However, the best solution in terms of model, balance, and model support loads, as well as in terms of capital and operating cost appears to be related to the reduction of the temperature of the test gas to cryogenic temperatures. The present paper has the objective to review the evolution of the cryogenic wind tunnel concept and to describe its more important advantages.

  13. The Beginner's Guide to Wind Tunnels with TunnelSim and TunnelSys

    NASA Technical Reports Server (NTRS)

    Benson, Thomas J.; Galica, Carol A.; Vila, Anthony J.

    2010-01-01

    The Beginner's Guide to Wind Tunnels is a Web-based, on-line textbook that explains and demonstrates the history, physics, and mathematics involved with wind tunnels and wind tunnel testing. The Web site contains several interactive computer programs to demonstrate scientific principles. TunnelSim is an interactive, educational computer program that demonstrates basic wind tunnel design and operation. TunnelSim is a Java (Sun Microsystems Inc.) applet that solves the continuity and Bernoulli equations to determine the velocity and pressure throughout a tunnel design. TunnelSys is a group of Java applications that mimic wind tunnel testing techniques. Using TunnelSys, a team of students designs, tests, and post-processes the data for a virtual, low speed, and aircraft wing.

  14. V/STOL Tandem Fan transition section model test. [in the Lewis Research Center 10-by-10 foot wind tunnel

    NASA Technical Reports Server (NTRS)

    Simpkin, W. E.

    1982-01-01

    An approximately 0.25 scale model of the transition section of a tandem fan variable cycle engine nacelle was tested in the NASA Lewis Research Center 10-by-10 foot wind tunnel. Two 12-inch, tip-turbine driven fans were used to simulate a tandem fan engine. Three testing modes simulated a V/STOL tandem fan airplane. Parallel mode has two separate propulsion streams for maximum low speed performance. A front inlet, fan, and downward vectorable nozzle forms one stream. An auxilliary top inlet provides air to the aft fan - supplying the core engine and aft vectorable nozzle. Front nozzle and top inlet closure, and removal of a blocker door separating the two streams configures the tandem fan for series mode operations as a typical aircraft propulsion system. Transition mode operation is formed by intermediate settings of the front nozzle, blocker door, and top inlet. Emphasis was on the total pressure recovery and flow distortion at the aft fan face. A range of fan flow rates were tested at tunnel airspeeds from 0 to 240 knots, and angles-of-attack from -10 to 40 deg for all three modes. In addition to the model variables for the three modes, model variants of the top inlet were tested in the parallel mode only. These lip variables were: aft lip boundary layer bleed holes, and Three position turning vane. Also a bellmouth extension of the top inlet side lips was tested in parallel mode.

  15. Experimental Study of Slat Noise from 30P30N Three-Element High-Lift Airfoil in JAXA Hard-Wall Low-Speed Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Murayama, Mitsuhiro; Nakakita, Kazuyuki; Yamamoto, Kazuomi; Ura, Hiroki; Ito, Yasushi; Choudhari, Meelan M.

    2014-01-01

    Aeroacoustic measurements associated with noise radiation from the leading edge slat of the canonical, unswept 30P30N three-element high-lift airfoil configuration have been obtained in a 2 m x 2 m hard-wall wind tunnel at the Japan Aerospace Exploration Agency (JAXA). Performed as part of a collaborative effort on airframe noise between JAXA and the National Aeronautics and Space Administration (NASA), the model geometry and majority of instrumentation details are identical to a NASA model with the exception of a larger span. For an angle of attack up to 10 degrees, the mean surface Cp distributions agree well with free-air computational fluid dynamics predictions corresponding to a corrected angle of attack. After employing suitable acoustic treatment for the brackets and end-wall effects, an approximately 2D noise source map is obtained from microphone array measurements, thus supporting the feasibility of generating a measurement database that can be used for comparison with free-air numerical simulations. Both surface pressure spectra obtained via KuliteTM transducers and the acoustic spectra derived from microphone array measurements display a mixture of a broad band component and narrow-band peaks (NBPs), both of which are most intense at the lower angles of attack and become progressively weaker as the angle of attack is increased. The NBPs exhibit a substantially higher spanwise coherence in comparison to the broadband portion of the spectrum and, hence, confirm the trends observed in previous numerical simulations. Somewhat surprisingly, measurements show that the presence of trip dots between the stagnation point and slat cusp enhances the NBP levels rather than mitigating them as found in a previous experiment.

  16. Langley Research Center's Unitary Plan Wind Tunnel: Testing Capabilities and Recent Modernization Activities

    NASA Technical Reports Server (NTRS)

    Micol, John R.

    2001-01-01

    Description, capabilities, initiatives, and utilization of the NASA Langley Research Center's Unitary Plan Wind Tunnel are presented. A brief overview of the facility's operational capabilities and testing techniques is provided. A recent Construction of Facilities (CoF) project to improve facility productivity and efficiency through facility automation has been completed and is discussed. Several new and maturing thrusts are underway that include systematic efforts to provide credible assessment for data quality, modifications to the new automation control system for increased compatibility with the Modern Design Of Experiments (MDOE) testing methodology, and process improvements for better test coordination, planning, and execution.

  17. Langley Research Center's Unitary Plan Wind Tunnel: Testing Capabilities and Recent Modernization Activities

    NASA Technical Reports Server (NTRS)

    Micol, John R.

    2001-01-01

    Description, capabilities, initiatives, and utilization of the NASA Langley Research Center's Unitary Plan Wind Tunnel are presented. A brief overview of the facility's operational capabilities and testing techniques is provided. A recent Construction of Facilities (Car) project to improve facility productivity and efficiency through facility automation has been completed and is discussed. Several new and maturing thrusts are underway that include systematic efforts to provide credible assessment for data quality, modifications to the new automation control system for increased compatibility with the Modern Design of Experiments (MDOE) testing methodology, and process improvements for better test coordination, planning, and execution.

  18. Orbiter Model in Wind Tunnel

    NASA Technical Reports Server (NTRS)

    1980-01-01

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

  19. Mixed-Phase Icing Simulation and Testing at the Cox Icing Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Al-Khalil, Kamel; Irani, Eddie; Miller, Dean

    2003-01-01

    A new capability was developed for indoor simulation of snow and mixed-phase icing conditions. This capability is useful for year-round testing in the Cox closed-loop Icing Wind Tunnel. Certification of aircraft for flight into these types of icing conditions is only required by the JAA in Europe. In an effort to harmonize certification requirements, the FAA in the US sponsored a preliminary program to study the effects of mixed-phase and fully glaciated icing conditions on the performance requirements of thermal ice protection systems. This paper describes the test program and the associated results.

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

    NASA Technical Reports Server (NTRS)

    Lewis, M. C.

    1988-01-01

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

  1. Wind-Tunnel Tests of 10-foot-diameter Autogiro Rotors

    NASA Technical Reports Server (NTRS)

    Wheatley, John B; Bioletti, Carlton

    1937-01-01

    Report presents the results of a series of 10-foot-diameter autogiro rotor models tested in the NACA 20-foot wind tunnel. Four of the models differed only in the airfoil sections of the blades, the sections used being the NACA 0012, 0018, 4412, and 4418. Three additional models employing the NACA 0012 section were tested, in which a varying portion of the blade near the hub was replaced by a streamline tube with a chord of about one-fourth the blade chord.

  2. Summary report of the second wind tunnel test of the Boeing LFC model

    NASA Technical Reports Server (NTRS)

    George-Falvy, D.

    1978-01-01

    An 8-ft span, 20-ft chord, 30 deg swept wing section having provisions for laminar boundary control over the first 30% of the upper surface and the first 15% of the lower surface was tested in a 5-ft by 8-ft wind tunnel to explore the sensitivity of laminar flow to various forms of disturbances such as surface imperfections, contamination, off-design pressure distribution (increased crossflow), and imposed noise. The test equipment used and instrumentation of the model are described. Typical results obtained from configurations with spanwise ridges and spanwise rows of disks are discussed as well as suction flow characteristics at reduced incidence.

  3. Error propagation equations for estimating the uncertainty in high-speed wind tunnel test results

    SciTech Connect

    Clark, E.L.

    1994-07-01

    Error propagation equations, based on the Taylor series model, are derived for the nondimensional ratios and coefficients most often encountered in high-speed wind tunnel testing. These include pressure ratio and coefficient, static force and moment coefficients, dynamic stability coefficients, and calibration Mach number. The error equations contain partial derivatives, denoted as sensitivity coefficients, which define the influence of free-steam Mach number, M{infinity}, on various aerodynamic ratios. To facilitate use of the error equations, sensitivity coefficients are derived and evaluated for five fundamental aerodynamic ratios which relate free-steam test conditions to a reference condition.

  4. Comparison of field and wind-tunnel Darrieus wind-turbine data

    NASA Astrophysics Data System (ADS)

    Sheldahl, R. E.

    1981-08-01

    A 2-m-diam Darrieus vertical axis wind turbine with NACA-0012 airfoil blades was tested in the field and in a 4.6 x 6.1-m low speed wind tunnel for a direct comparison. Comparisons were made with field data of equivalent chord Reynolds number, and at equivalent rotational field. Maximum values of the power coefficients compared favorably, and an examination of performance coefficients showed complete agreement between wind tunnel and field data. Due to excellent agreement in the first two comparisons, no further field testing was done, and the accuracy of the wind-tunnel test data was believed verified.

  5. Case Studies for the Statistical Design of Experiments Applied to Powered Rotor Wind Tunnel Tests

    NASA Technical Reports Server (NTRS)

    Overmeyer, Austin D.; Tanner, Philip E.; Martin, Preston B.; Commo, Sean A.

    2015-01-01

    The application of statistical Design of Experiments (DOE) to helicopter wind tunnel testing was explored during two powered rotor wind tunnel entries during the summers of 2012 and 2013. These tests were performed jointly by the U.S. Army Aviation Development Directorate Joint Research Program Office and NASA Rotary Wing Project Office, currently the Revolutionary Vertical Lift Project, at NASA Langley Research Center located in Hampton, Virginia. Both entries were conducted in the 14- by 22-Foot Subsonic Tunnel with a small portion of the overall tests devoted to developing case studies of the DOE approach as it applies to powered rotor testing. A 16-47 times reduction in the number of data points required was estimated by comparing the DOE approach to conventional testing methods. The average error for the DOE surface response model for the OH-58F test was 0.95 percent and 4.06 percent for drag and download, respectively. The DOE surface response model of the Active Flow Control test captured the drag within 4.1 percent of measured data. The operational differences between the two testing approaches are identified, but did not prevent the safe operation of the powered rotor model throughout the DOE test matrices.

  6. V/STOL tilt rotor aircraft study: Wind tunnel tests of a full scale hingeless prop/rotor designed for the Boeing Model 222 tilt rotor aircraft

    NASA Technical Reports Server (NTRS)

    Magee, J. P.; Alexander, H. R.

    1973-01-01

    The rotor system designed for the Boeing Model 222 tilt rotor aircraft is a soft-in-plane hingeless rotor design, 26 feet in diameter. This rotor has completed two test programs in the NASA Ames 40' X 80' wind tunnel. The first test was a windmilling rotor test on two dynamic wing test stands. The rotor was tested up to an advance ratio equivalence of 400 knots. The second test used the NASA powered propeller test rig and data were obtained in hover, transition and low speed cruise flight. Test data were obtained in the areas of wing-rotor dynamics, rotor loads, stability and control, feedback controls, and performance to meet the test objectives. These data are presented.

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

    NASA Technical Reports Server (NTRS)

    Murman, E. M.

    1982-01-01

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

  8. Development, simulation validation, and wind tunnel testing of a digital controller system for flutter suppression

    NASA Technical Reports Server (NTRS)

    Hoadley, Sherwood Tiffany; Buttrill, Carey S.; Mcgraw, Sandra M.; Houck, Jacob A.

    1991-01-01

    Flutter suppression (FS) is one of the active control concepts being investigated by the AFW program. The design goal for FS control laws was to increase the passive flutter dynamic pressure by 30 percent. In order to meet this goal, the FS control laws had to be capable of suppressing both symmetric and antisymmetric flutter instabilities simultaneously. In addition, the FS control laws had to be practical and low-order, robust and capable of real time execution within the 200 hz. sampling time. The purpose here is to present an overview of the development, simulation validation, and wind tunnel testing of a digital controller system for flutter suppression.

  9. Large-scale Advanced Prop-fan (LAP) high speed wind tunnel test report

    NASA Technical Reports Server (NTRS)

    Campbell, William A.; Wainauski, Harold S.; Arseneaux, Peter J.

    1988-01-01

    High Speed Wind Tunnel testing of the SR-7L Large Scale Advanced Prop-Fan (LAP) is reported. The LAP is a 2.74 meter (9.0 ft) diameter, 8-bladed tractor type rated for 4475 KW (6000 SHP) at 1698 rpm. It was designated and built by Hamilton Standard under contract to the NASA Lewis Research Center. The LAP employs thin swept blades to provide efficient propulsion at flight speeds up to Mach .85. Testing was conducted in the ONERA S1-MA Atmospheric Wind Tunnel in Modane, France. The test objectives were to confirm that the LAP is free from high speed classical flutter, determine the structural and aerodynamic response to angular inflow, measure blade surface pressures (static and dynamic) and evaluate the aerodynamic performance at various blade angles, rotational speeds and Mach numbers. The measured structural and aerodynamic performance of the LAP correlated well with analytical predictions thereby providing confidence in the computer prediction codes used for the design. There were no signs of classical flutter throughout all phases of the test up to and including the 0.84 maximum Mach number achieved. Steady and unsteady blade surface pressures were successfully measured for a wide range of Mach numbers, inflow angles, rotational speeds and blade angles. No barriers were discovered that would prevent proceeding with the PTA (Prop-Fan Test Assessment) Flight Test Program scheduled for early 1987.

  10. DHC-6 Twin Otter Tailplane Airfoil Section Testing in the Ohio State University 7x10 Wind Tunnel. Volume 1

    NASA Technical Reports Server (NTRS)

    Hiltner, Dale; McKee, Michael; LaNoe, Karine; Gregorek, Gerald; Ratvasky, Thomas (Technical Monitor)

    2000-01-01

    Ice contaminated tailplane stall (ICTS) has been found to be responsible for 16 accidents with 139 fatalities over the last three decades, and is suspected to have played a role in other accidents and incidents. The need for fundamental research in this area has been recognized at three international conferences sponsored by the FAA since 1991. In order to conduct such research, a joint NASA/FAA Tailplane Icing Program was formed in 1994: the Ohio State University has played an important role in this effort. The program employs icing tunnel testing, dry wind tunnel testing, flight testing, and analysis using a six-degrees-of-freedom computer code tailored to this problem. A central goal is to quantify the effect of tailplane icing on aircraft stability and control to aid in the analysis of flight test procedures to identify aircraft susceptibility to ICTS. This report contains the results ot testing of a full scale 2D model of a tailplane section of NASA's Icing Research Aircraft, with and without ice shapes, in an Ohio State University 7 x 10 Low Speed wind tunnel in 1994. The results have been integrated into a comprehensive database of aerodynamic coefficients and stability and control derivatives that will permit detailed analysis of flight test results with the analytical computer program. The testing encompassed a full range of angles of attack and elevator deflections, as well as two velocities to evaluate Reynolds number effects. Lift, drag, pitching moment, and hinge moment coefficients were obtained. In addition. instrumentation for use during flight testing was verified to be effective, all components showing acceptable fidelity. Comparison of clean and iced airfoil results show the ice shapes causing a significant decrease in the magnitude of CLmax (from -1.3 to -0.64) and associated stall angle (from -18.6 deg to -8.2 deg). Furthermore, the ice shapes caused an increase in hinge moment coefficient of approximately 0.02, the change being markedly abrupt

  11. Phase 2 and 3 wind tunnel tests of the J-97 powered, external augmentor V/STOL model. [at Ames 40 by 80 wind tunnel

    NASA Technical Reports Server (NTRS)

    Garland, D. B.; Harris, J. L.

    1980-01-01

    Static and forward speed tests were made in a 40 multiplied by 80 foot wind tunnel of a large-scale, ejector-powered V/STOL aircraft model. Modifications were made to the model following earlier tests primarily to improve longitudinal acceleration capability during transition from hovering to wingborne flight. A rearward deflection of the fuselage augmentor thrust vector was shown to be beneficial in this regard. Other augmentor modifications were tested, notably the removal of both endplates, which improved acceleration performance at the higher transition speeds. The model tests again demonstrated minimal interference of the fuselage augmentor on aerodynamic lift. A flapped canard surface also showed negligible influence on the performance of the wing and of the fuselage augmentor.

  12. RSRA sixth scale wind tunnel test. [of scale model of Sikorsky Whirlwind Helicopter

    NASA Technical Reports Server (NTRS)

    Flemming, R.; Ruddell, A.

    1974-01-01

    The sixth scale model of the Sikorsky/NASA/Army rotor systems research aircraft was tested in an 18-foot section of a large subsonic wind tunnel for the purpose of obtaining basic data in the areas of performance, stability, and body surface loads. The model was mounted in the tunnel on the struts arranged in tandem. Basic testing was limited to forward flight with angles of yaw from -20 to +20 degrees and angles of attack from -20 to +25 degrees. Tunnel test speeds were varied up to 172 knots (q = 96 psf). Test data were monitored through a high speed static data acquisition system, linked to a PDP-6 computer. This system provided immediate records of angle of attack, angle of yaw, six component force and moment data, and static and total pressure information. The wind tunnel model was constructed of aluminum structural members with aluminum, fiberglass, and wood skins. Tabulated force and moment data, flow visualization photographs, tabulated surface pressure data are presented for the basic helicopter and compound configurations. Limited discussions of the results of the test are included.

  13. Static and Wind Tunnel Aero-Performance Tests of NASA AST Separate Flow Nozzle Noise Reduction Configurations

    NASA Technical Reports Server (NTRS)

    Mikkelsen, Kevin L.; McDonald, Timothy J.; Saiyed, Naseem (Technical Monitor)

    2001-01-01

    This report presents the results of cold flow model tests to determine the static and wind tunnel performance of several NASA AST separate flow nozzle noise reduction configurations. The tests were conducted by Aero Systems Engineering, Inc., for NASA Glenn Research Center. The tests were performed in the Channels 14 and 6 static thrust stands and the Channel 10 transonic wind tunnel at the FluiDyne Aerodynamics Laboratory in Plymouth, Minnesota. Facility checkout tests were made using standard ASME long-radius metering nozzles. These tests demonstrated facility data accuracy at flow conditions similar to the model tests. Channel 14 static tests reported here consisted of 21 ASME nozzle facility checkout tests and 57 static model performance tests (including 22 at no charge). Fan nozzle pressure ratio varied from 1.4 to 2.0, and fan to core total pressure ratio varied from 1.0 to 1.19. Core to fan total temperature ratio was 1.0. Channel 10 wind tunnel tests consisted of 15 tests at Mach number 0.28 and 31 tests at Mach 0.8. The sting was checked out statically in Channel 6 before the wind tunnel tests. In the Channel 6 facility, 12 ASME nozzle data points were taken and 7 model data points were taken. In the wind tunnel, fan nozzle pressure ratio varied from 1.73 to 2.8, and fan to core total pressure ratio varied from 1.0 to 1.19. Core to fan total temperature ratio was 1.0. Test results include thrust coefficients, thrust vector angle, core and fan nozzle discharge coefficients, total pressure and temperature charging station profiles, and boat-tail static pressure distributions in the wind tunnel.

  14. Wind Tunnel Testing of Powered Lift, All-Wing STOL Model

    NASA Technical Reports Server (NTRS)

    Collins, Scott W.; Westra, Bryan W.; Lin, John C.; Jones, Gregory S.; Zeune, Cal H.

    2008-01-01

    Short take-off and landing (STOL) systems can offer significant capabilities to warfighters and, for civil operators thriving on maximizing efficiencies they can improve airspace use while containing noise within airport environments. In order to provide data for next generation systems, a wind tunnel test of an all-wing cruise efficient, short take-off and landing (CE STOL) configuration was conducted in the National Aeronautics and Space Administration (NASA) Langley Research Center (LaRC) 14- by 22-foot Subsonic Wind Tunnel. The test s purpose was to mature the aerodynamic aspects of an integrated powered lift system within an advanced mobility configuration capable of CE STOL. The full-span model made use of steady flap blowing and a lifting centerbody to achieve high lift coefficients. The test occurred during April through June of 2007 and included objectives for advancing the state-of-the-art of powered lift testing through gathering force and moment data, on-body pressure data, and off-body flow field measurements during automatically controlled blowing conditions. Data were obtained for variations in model configuration, angles of attack and sideslip, blowing coefficient, and height above ground. The database produced by this effort is being used to advance design techniques and computational tools for developing systems with integrated powered lift technologies.

  15. Testing of the Crew Exploration Vehicle in NASA Langley's Unitary Plan Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Murphy, Kelly J.; Borg, Stephen E.; Watkins, Anthony N.; Cole, Daniel R.; Schwartz, Richard J.

    2007-01-01

    As part of a strategic, multi-facility test program, subscale testing of NASA s Crew Exploration Vehicle was conducted in both legs of NASA Langley s Unitary Plan Wind Tunnel. The objectives of these tests were to generate aerodynamic and surface pressure data over a range of supersonic Mach numbers and reentry angles of attack for experimental and computational validation and aerodynamic database development. To provide initial information on boundary layer transition at supersonic test conditions, transition studies were conducted using temperature sensitive paint and infrared thermography optical techniques. To support implementation of these optical diagnostics in the Unitary Wind Tunnel, the experiment was first modeled using the Virtual Diagnostics Interface software. For reentry orientations of 140 to 170 degrees (heat shield forward), windward surface flow was entirely laminar for freestream unit Reynolds numbers equal to or less than 3 million per foot. Optical techniques showed qualitative evidence of forced transition on the windward heat shield with application of both distributed grit and discreet trip dots. Longitudinal static force and moment data showed the largest differences with Mach number and angle of attack variations. Differences associated with Reynolds number variation and/or laminar versus turbulent flow on the heat shield were very small. Static surface pressure data supported the aforementioned trends with Mach number, Reynolds number, and angle of attack.

  16. Jet noise results from static, wind tunnel, and flight tests of conical and mechanical suppressor nozzles

    NASA Astrophysics Data System (ADS)

    McKinnon, R. A.; Johnson, E. S.; Atencio, A., Jr.

    1981-10-01

    Results of jet noise suppression tests conducted on a Rolls-Royce Viper 601 turbojet engine are reported. Seven exhaust nozzle configurations are tested, including two conical nozzles, two suppressor mixers, and three treated ejector configurations with different ejector inlets. Tests are conducted at the NASA Ames outdoor static test facility and the 40- by 80-ft wind tunnel facility at minimum tunnel flow velocity and normal flow velocities of 230 and 290 ft/sec. Near-field multiple sideline noise levels are projected to the far fields to compare far-field fixed microphone outdoor static noise levels, and wind tunnel near-field noise data are projected to the far field and flight distances to compare with noise levels recorded from an Hs-125 aircraft. Near-field outdoor noise data duplicate the far-field data recorded from fixed microphones within 2 PNdB, and the Douglas mechanical jet noise suppressor/treated ejector exhaust system achieves a noise reduction of 12 EPNdB relative to a conic reference nozzle at equal thrust in flight.

  17. Jet noise results from static, wind tunnel, and flight tests of conical and mechanical suppressor nozzles

    NASA Technical Reports Server (NTRS)

    Mckinnon, R. A.; Johnson, E. S.; Atencio, A., Jr.

    1981-01-01

    Results of jet noise suppression tests conducted on a Rolls-Royce Viper 601 turbojet engine are reported. Seven exhaust nozzle configurations are tested, including two conical nozzles, two suppressor mixers, and three treated ejector configurations with different ejector inlets. Tests are conducted at the NASA Ames outdoor static test facility and the 40- by 80-ft wind tunnel facility at minimum tunnel flow velocity and normal flow velocities of 230 and 290 ft/sec. Near-field multiple sideline noise levels are projected to the far fields to compare far-field fixed microphone outdoor static noise levels, and wind tunnel near-field noise data are projected to the far field and flight distances to compare with noise levels recorded from an Hs-125 aircraft. Near-field outdoor noise data duplicate the far-field data recorded from fixed microphones within 2 PNdB, and the Douglas mechanical jet noise suppressor/treated ejector exhaust system achieves a noise reduction of 12 EPNdB relative to a conic reference nozzle at equal thrust in flight.

  18. Wind tunnel validation of AeroDyn within LIFES50+ project: imposed Surge and Pitch tests

    NASA Astrophysics Data System (ADS)

    Bayati, I.; Belloli, M.; Bernini, L.; Zasso, A.

    2016-09-01

    This paper presents the first set of results of the steady and unsteady wind tunnel tests, performed at Politecnico di Milano wind tunnel, on a 1/75 rigid scale model of the DTU 10 MW wind turbine, within the LIFES50+ project. The aim of these tests is the validation of the open source code AeroDyn developed at NREL. Numerical and experimental steady results are compared in terms of thrust and torque coefficients, showing good agreement, as well as for unsteady measurements gathered with a 2 degree-of-freedom test rig, capable of imposing the displacements at the base of the model, and providing the surge and pitch motion of the floating offshore wind turbine (FOWT) scale model. The measurements of the unsteady test configuration are compared with AeroDyn/Dynin module results, implementing the generalized dynamic wake (GDW) model. Numerical and experimental comparison showed similar behaviours in terms of non linear hysteresis, however some discrepancies are herein reported and need further data analysis and interpretations about the aerodynamic integral quantities, with a special attention to the physics of the unsteady phenomenon.

  19. Wind tunnel wall interference in V/STOL and high lift testing: A selected, annotated bibliography

    NASA Technical Reports Server (NTRS)

    Tuttle, M. H.; Mineck, R. E.; Cole, K. L.

    1986-01-01

    This bibliography, with abstracts, consists of 260 citations of interest to persons involved in correcting aerodynamic data, from high lift or V/STOL type configurations, for the interference arising from the wind tunnel test section walls. It provides references which may be useful in correcting high lift data from wind tunnel to free air conditions. References are included which deal with the simulation of ground effect, since it could be viewed as having interference from three tunnel walls. The references could be used to design tests from the standpoint of model size and ground effect simulation, or to determine the available testing envelope with consideration of the problem of flow breakdown. The arrangement of the citations is chronological by date of publication in the case of reports or books, and by date of presentation in the case of papers. Included are some documents of historical interest in the development of high lift testing techniques and wall interference correction methods. Subject, corporate source, and author indices, by citation numbers, have been provided to assist the users. The appendix includes citations of some books and documents which may not deal directly with high lift or V/STOL wall interference, but include additional information which may be helpful.

  20. SAMPSON smart inlet design overview and wind tunnel test: I. Design overview

    NASA Astrophysics Data System (ADS)

    Pitt, Dale M.; Dunne, James P.; White, Edward V.

    2002-07-01

    The Smart Aircraft and Marine System Projects Demonstration (SAMPSON) program was a DARPA funded effort conducted by the Boeing Company, General Dynamics - Electric Boat Division, and the Pennsylvania State University. NASA Langley Research Center (NASA LaRC) was technical monitor for the aircraft demonstration, while the Navy's Office of Naval Research (ONR) was technical monitor for the marine demonstration. Dr. Ephrahim Garcia, DARPA/DSO, acted as the DARPA program manager for SAMPSON. The SAMPSON program objectives were to demonstrate smart structures based systems on large/full scale structures in realistic environments. The SAMPSON aircraft demonstration was the wind tunnel testing of a full scale F-15 aircraft inlet that was capable of in-flight structural variations accomplished using smart materials, called the 'SAMPSON Smart Inlet'. The SAMPSON Smart Inlet was removed from an F-15E airframe and structurally modified to interface with the NASA LaRC 16-Foot Transonic Tunnel model support system. This is Part I of two works documenting the SAMPSON Smart Inlet design and testing. A discussion of the design aspects and constraints will be presented here in Part I. The ground and wind tunnel testing of the Smart Inlet is presented in a separate work, Part II.

  1. March 1971 wind tunnel tests of the Dorand DH 2011 jet flap rotor, volume 1

    NASA Technical Reports Server (NTRS)

    Kretz, M.; Aubrun, J.; Larche, M.

    1973-01-01

    The results of wind tunnel tests, second series of tests performed in the NASA Ames 40 x 80 foot wind tunnel, of the DH 2011 jet-flap rotor are presented and analyzed. The tests have been focused on multicyclic effects and the capability of this rotor to reduce the vibratory loads and stresses in the blades. The reductions of the vibrations and stresses at tip speed ratio of 0.4 have attained 50%. The theory shows further reductions possible, reaching 80%. The results show that the performance characteristics after the modifications introduced since 1965 remained unchanged. The domain of investigation has been enlarged to include the tip speed ratios of 0.6 and 0.7. To analyze the complex aeroelastic phenomena a new analytical technique has been utilized to represent the mathematical model of the rotor. This technique, based on transfer matrices and transfer functions, appears very simple and it is believed that this analysis is applicable to many kinds of investigations involving large numbers of variables.

  2. Flight and wind tunnel test results of the mechanical jet noise suppressor nozzle

    NASA Astrophysics Data System (ADS)

    Fitzsimmons, R. D.; McKinnon, R. A.; Johnson, E. S.; Brooks, J. R.

    1980-01-01

    Comprehensive acoustic and propulsion data are presented, based on flight and wind tunnel tests, of a mechanical jet noise suppressor designed to satisfy the requirements of an advanced supersonic transport (AST) under study by the McDonnell Douglas Corporation. The flight program was conducted jointly by MDC, Rolls-Royce Ltd., and the British Aerospace Corporation, using an HS-125 aircraft modified to accept an upgraded RR Viper 601 engine with conical reference and mechanical suppressor nozzles and an acoustically treated ejector. The nacelle, engine and nozzle configurations from the HS-125 were also tested in one of NASA's wind tunnels to obtain thrust performance at forward velocity and acoustic data. The acoustic flight test data, when scaled to an AST engine nozzle size and projected to a typical sideline distance, indicate reduction in effective perceived noise level of 16 EPNdB at the takeoff power setting. It is estimated that the in-flight thrust loss for a typical AST suppressor/ejector nozzle configuration (37.5 inch equivalent diameter) would be 5.4 percent at takeoff power settings and 6.6 percent at cutback power settings.

  3. Space shuttle phase B wind tunnel model and test information. Volume 3: Launch configuration

    NASA Technical Reports Server (NTRS)

    Glynn, J. L.; Poucher, D. E.

    1988-01-01

    Archived wind tunnel test data are available for flyback booster or other alternate recoverable configuration as well as reusable orbiters studied during initial development (Phase B) of the Space Shuttle, including contractor data for an extensive variety of configurations with an array of wing and body planforms. The test data have been compiled into a database and are available for application to current winged flyback or recoverable booster aerodynamic studies. The Space Shuttle Phase B Wind Tunnel Database is structured by vehicle component and configuration. Basic components include booster, orbiter, and launch vehicle. Booster configuration types include straight and delta wings, canard, cylindrical, retroglide and twin body. Orbiter configurations include straight and delta wings, lifting body, drop tanks and double delta wings. Launch configurations include booster and orbiter components in various stacked and tandem combinations. The digital database consists of 220 files containing basic tunnel data. Database structure is documented in a series of reports which include configuration sketches for the various planforms tested. This is Volume 3 -- launch configurations.

  4. Aerodynamic control of NASP-type vehicles through Vortex manipulation. Volume 2: Static wind tunnel tests

    NASA Technical Reports Server (NTRS)

    Suarez, Carlos J.; Kramer, Brian R.; Smith, Brooke C.; Malcolm, Gerald N.

    1993-01-01

    Forebody Vortex Control (FVC) was explored in this research program for potential application to a NASP-type configuration. Wind tunnel tests were conducted to evaluate a number of jet blowing schemes. The configuration tested has a slender forebody and a 78 deg swept delta wing. Blowing jets were implemented on the leeward side of the forebody with small circular tubes tangential to the surface that could be directed aft, forward, or at angles in between. The effects of blowing are observed primarily in the yawing and rolling moments and are highly dependent on the jet configuration and the angle of attack. Results show that the baseline flow field, without blowing activated, is quite sensitive to the geometry differences of the various protruding jets, as well as being sensitive to the blowing, particularly in the angle of attack range where the forebody vortices are naturally asymmetric. The time lag of the flow field response to the initiation of blowing was also measured. The time response was very short, on the order of the time required for the flow disturbance to travel the distance from the nozzle to the specific airframe location of interest at the free stream velocity. Overall, results indicate that sizable yawing and rolling moments can be induced with modest blowing levels. However, direct application of this technique on a very slender forebody would require thorough wind tunnel testing to optimize the jet location and configuration.

  5. Advanced Capabilities for Wind Tunnel Testing in the 21st Century

    NASA Technical Reports Server (NTRS)

    Kegelman, Jerome T.; Danehy, Paul M.; Schwartz, Richard J.

    2010-01-01

    Wind tunnel testing methods and test technologies for the 21st century using advanced capabilities are presented. These capabilities are necessary to capture more accurate and high quality test results by eliminating the uncertainties in testing and to facilitate verification of computational tools for design. This paper discusses near term developments underway in ground testing capabilities, which will enhance the quality of information of both the test article and airstream flow details. Also discussed is a selection of new capability investments that have been made to accommodate such developments. Examples include advanced experimental methods for measuring the test gas itself; using efficient experiment methodologies, including quality assurance strategies within the test; and increasing test result information density by using extensive optical visualization together with computed flow field results. These points could be made for both major investments in existing tunnel capabilities or for entirely new capabilities.

  6. Design and Development of a Deep Acoustic Lining for the 40-by 80-Foot Wind Tunnel Test Section

    NASA Technical Reports Server (NTRS)

    Soderman, Paul T.; Schmitz, Fredric H.; Allen, Christopher S.; Jaeger, Stephen M.; Sacco, Joe N.; Mosher, Marianne; Hayes, Julie A.

    2002-01-01

    The work described in this report has made effective use of design teams to build a state-of-the-art anechoic wind-tunnel facility. Many potential design solutions were evaluated using engineering analysis, and computational tools. Design alternatives were then evaluated using specially developed testing techniques, Large-scale coupon testing was then performed to develop confidence that the preferred design would meet the acoustic, aerodynamic, and structural objectives of the project. Finally, designs were frozen and the final product was installed in the wind tunnel. The result of this technically ambitious project has been the creation of a unique acoustic wind tunnel. Its large test section (39 ft x 79 ft x SO ft), potentially near-anechoic environment, and medium subsonic speed capability (M = 0.45) will support a full range of aeroacoustic testing-from rotorcraft and other vertical takeoff and landing aircraft to the take-off/landing configurations of both subsonic and supersonic transports.

  7. Ground simulation with moving belt and tangential blowing for full-scale automotive testing in a wind tunnel

    SciTech Connect

    Mercker, E.; Knape, H.W.

    1989-01-01

    This paper describes full-scale vehicle tests made on a standard-type passenger car in a wind tunnel and on the road in order to evaluate different moving-ground simulation techniques for wind tunnels. The test was first executed over a moving belt, supporting the car with a rear sting and measuring the aerodynamic forces with an internal balance. The test was then repeated with the same support arrangement over a fixed test-section floor, and moving-ground simulation was attained with boundary layer control by tangential blowing. Besides force measurements, the surface pressure distribution underneath the vehicle and at the base were also measured.

  8. Design and fabrication of forward-swept counterrotation blade configuration for wind tunnel testing

    NASA Technical Reports Server (NTRS)

    Nichols, G. H.

    1994-01-01

    Work performed by GE Aircraft on advanced counterrotation blade configuration concepts for high speed turboprop system is described. Primary emphasis was placed on theoretically and experimentally evaluating the aerodynamic, aeromechanical, and acoustic performance of GE-defined counterrotating blade concepts. Several blade design concepts were considered. Feasibility studies were conducted to evaluate a forward-swept versus an aft-swept blade application and how the given blade design would affect interaction between rotors. Two blade designs were initially selected. Both designs involved in-depth aerodynamic, aeromechanical, mechanical, and acoustic analyses followed by the fabrication of forward-swept, forward rotor blade sets to be wind tunnel tested with an aft-swept, aft rotor blade set. A third blade set was later produced from a NASA design that was based on wind tunnel test results from the first two blade sets. This blade set had a stiffer outer ply material added to the original blade design, in order to reach the design point operating line. Detailed analyses, feasibility studies, and fabrication procedures for all blade sets are presented.

  9. DARPA/AFRL/NASA Smart Wing Second Wind Tunnel Test Results

    NASA Technical Reports Server (NTRS)

    Scherer, L. B.; Martin, C. A.; West, M.; Florance, J. P.; Wieseman, C. D.; Burner, A. W.; Fleming, G. A.

    2001-01-01

    To quantify the benefits of smart materials and structures adaptive wing technology, Northrop Grumman Corp. (NGC) built and tested two 16% scale wind tunnel models (a conventional and a "smart" model) of a fighter/attack aircraft under the DARPA/AFRL/NASA Smart Materials and Structures Development - Smart Wing Phase 1. Performance gains quantified included increased pitching moment (C(sub M)), increased rolling moment (C(subl)) and improved pressure distribution. The benefits were obtained for hingeless, contoured trailing edge control surfaces with embedded shape memory alloy (SMA) wires and spanwise wing twist effected by SMA torque tube mechanisms, compared to conventional hinged control surfaces. This paper presents an overview of the results from the second wind tunnel test performed at the NASA Langley Research Center s (LaRC) 16ft Transonic Dynamic Tunnel (TDT) in June 1998. Successful results obtained were: 1) 5 degrees of spanwise twist and 8-12% increase in rolling moment utilizing a single SMA torque tube, 2) 12 degrees of deflection, and 10% increase in rolling moment due to hingeless, contoured aileron, and 3) demonstration of optical techniques for measuring spanwise twist and deflected shape.

  10. The Impact of Truth Surrogate Variance on Quality Assessment/Assurance in Wind Tunnel Testing

    NASA Technical Reports Server (NTRS)

    DeLoach, Richard

    2016-01-01

    Minimum data volume requirements for wind tunnel testing are reviewed and shown to depend on error tolerance, response model complexity, random error variance in the measurement environment, and maximum acceptable levels of inference error risk. Distinctions are made between such related concepts as quality assurance and quality assessment in response surface modeling, as well as between precision and accuracy. Earlier research on the scaling of wind tunnel tests is extended to account for variance in the truth surrogates used at confirmation sites in the design space to validate proposed response models. A model adequacy metric is presented that represents the fraction of the design space within which model predictions can be expected to satisfy prescribed quality specifications. The impact of inference error on the assessment of response model residuals is reviewed. The number of sites where reasonably well-fitted response models actually predict inadequately is shown to be considerably less than the number of sites where residuals are out of tolerance. The significance of such inference error effects on common response model assessment strategies is examined.

  11. Space Launch System Booster Separation Aerodynamic Testing in the NASA Langley Unitary Plan Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Wilcox, Floyd J., Jr.; Pinier, Jeremy T.; Chan, David T.; Crosby, William A.

    2016-01-01

    A wind-tunnel investigation of a 0.009 scale model of the Space Launch System (SLS) was conducted in the NASA Langley Unitary Plan Wind Tunnel to characterize the aerodynamics of the core and solid rocket boosters (SRBs) during booster separation. High-pressure air was used to simulate plumes from the booster separation motors (BSMs) located on the nose and aft skirt of the SRBs. Force and moment data were acquired on the core and SRBs. These data were used to corroborate computational fluid dynamics (CFD) calculations that were used in developing a booster separation database. The SRBs could be remotely positioned in the x-, y-, and z-direction relative to the core. Data were acquired continuously while the SRBs were moved in the axial direction. The primary parameters varied during the test were: core pitch angle; SRB pitch and yaw angles; SRB nose x-, y-, and z-position relative to the core; and BSM plenum pressure. The test was conducted at a free-stream Mach number of 4.25 and a unit Reynolds number of 1.5 million per foot.

  12. Evaluation of electrolytic tilt sensors for measuring model angle of attack in wind tunnel tests

    NASA Technical Reports Server (NTRS)

    Wong, Douglas T.

    1992-01-01

    The results of a laboratory evaluation of electrolytic tilt sensors as potential candidates for measuring model attitude or angle of attack in wind tunnel tests are presented. The performance of eight electrolytic tilt sensors was compared with that of typical servo accelerometers used for angle-of-attack measurements. The areas evaluated included linearity, hysteresis, repeatability, temperature characteristics, roll-on-pitch interaction, sensitivity to lead-wire resistance, step response time, and rectification. Among the sensors being evaluated, the Spectron model RG-37 electrolytic tilt sensors have the highest overall accuracy in terms of linearity, hysteresis, repeatability, temperature sensitivity, and roll sensitivity. A comparison of the sensors with the servo accelerometers revealed that the accuracy of the RG-37 sensors was on the average about one order of magnitude worse. Even though a comparison indicates that the cost of each tilt sensor is about one-third the cost of each servo accelerometer, the sensors are considered unsuitable for angle-of-attack measurements. However, the potential exists for other applications such as wind tunnel wall-attitude measurements where the errors resulting from roll interaction, vibration, and response time are less and sensor temperature can be controlled.

  13. Reynolds number effects on boattail drag of exhaust nozzles from wind tunnel and flight tests

    NASA Technical Reports Server (NTRS)

    Wilcox, F. A.; Chamberlin, R.

    1974-01-01

    A family of nacelle mounted high angle boattail nozzles was tested to investigate Reynolds number effects on drag. The nozzles were flown on a modified F-106B and mounted on scale models of a F-106 in a wind tunnel. A 19- to 1-range of Reynolds number was covered as a result of the large size differences between models and by flying over a range of altitude. In flight, the nozzles were mounted behind J-85 turbojet engines. Jet boundary simulators and a powered turbojet engine simulator were used on the wind tunnel models. Data were taken at Mach numbers of 0.6 and 0.9. Boattail drag was found to be affected by boattail number. The effect is a complex relationship dependent upon boundary layer thickness and nozzle boattail shape. As Reynolds number was increased from the lowest values obtained with scale models, boattail drag first increased to a maximum at the lowest flight Reynolds number and then decreased.

  14. Supersonic Retropropulsion Test 1853 in NASA LaRC Unitary Plan Wind Tunnel Test Section 2

    NASA Technical Reports Server (NTRS)

    Berry, Scott A.; Rhode, Matthew N.

    2014-01-01

    A supersonic retropropulsion experiment was conducted in the Langley Research Center Unitary Plan Wind Tunnel Test Section 2 at Mach numbers of 2.4, 3.5, and 4.6. Intended as a code validation effort, this study used pretest computations to size and refine the model such that tunnel blockage and internal flow separations were minimized. A 5-in diameter 70 degree sphere-cone forebody, which can accommodate up to four 4:1 area ratio nozzles, followed by a 9.55 inches long cylindrical aft body was selected for this test after computational maturation. The primary measurements for this experiment were high spatial-density surface pressures. In addition, high speed schlieren video and internal pressures and temperatures were acquired. The test included parametric variations in the number of nozzles utilized, thrust coefficients (roughly 0 to 4), and angles of attack (-8 to 20 degrees). The run matrix was developed to also allow quantification of various sources of experimental uncertainty, such as random errors due to run-to-run variations and systematic errors due to flowfield or model misalignments. To accommodate the uncertainty assessment, many runs and replicates were conducted with the model at various locations within the tunnel and with model roll angles of 0, 60, 120, and 180 degrees. This test report provides operational details of the experiment, contains a review of trends, and provides all schlieren and pressure results within appendices.

  15. Fastener load tests and retention systems tests for cryogenic wind-tunnel models

    NASA Technical Reports Server (NTRS)

    Wallace, J. W.

    1984-01-01

    A-286 stainless steel screws were tested to determine the tensile load capability and failure mode of various screw sizes and types at both cryogenic and room temperature. Additionally, five fastener retention systems were tested by using A-286 screws with specimens made from the primary metallic alloys that are currently used for cryogenic models. The locking system effectiveness was examined by simple no-load cycling to cryogenic temperatures (-275 F) as well as by dynamic and static loading at cryogenic temperatures. In general, most systems were found to be effective retention devices. There are some differences between the various devices with respect to ease of application, cleanup, and reuse. Results of tests at -275 F imply that the cold temperatures act to improve screw retention. The improved retention is probably the result of differential thermal contraction and/or increased friction (thread-binding effects). The data provided are useful in selecting screw sizes, types, and locking devices for model systems to be tested in cryogenic wind tunnels.

  16. Low-Speed Wind-Tunnel Investigation of Blowing Boundary-Layer Control on Leading- and Trailing-Edge Flaps of a Large-Scale, Low-Aspect-Ratio, 45 Swept-wing Airplane Configuration

    NASA Technical Reports Server (NTRS)

    Maki, Ralph L.

    1959-01-01

    Blowing boundary-layer control was applied to the leading- and trailing-edge flaps of a 45 deg sweptback-wing complete model in a full-scale low-speed wind-tunnel study. The principal purpose of the study was to determine the effects of leading-edge flap deflection and boundary-layer control on maximum lift and longitudinal stability. Leading-edge flap deflection alone was sufficient to maintain static longitudinal stability without trailing-edge flaps. However, leading-edge flap blowing was required to maintain longitudinal stability by delaying leading-edge flow separation when trailing-edge flaps were deflected either with or without blowing. Partial-span leading-edge flaps deflected 60 deg with moderate blowing gave the major increase in maximum lift, although higher deflection and additional blowing gave some further increase. Inboard of 0.4 semispan leading-edge flap deflection could be reduced to 40 deg and/or blowing could be omitted with only small loss in maximum lift. Trailing-edge flap lift increments were increased by boundary-layer control for deflections greater than 45 deg. Maximum lift was not increased with deflected trailing-edge flaps with blowing.

  17. New Model Exhaust System Supports Testing in NASA Lewis' 10- by 10-Foot Supersonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Roeder, James W., Jr.

    1998-01-01

    In early 1996, the ability to run NASA Lewis Research Center's Abe Silverstein 10- by 10- Foot Supersonic Wind Tunnel (10x10) at subsonic test section speeds was reestablished. Taking advantage of this new speed range, a subsonic research test program was scheduled for the 10x10 in the fall of 1996. However, many subsonic aircraft test models require an exhaust source to simulate main engine flow, engine bleed flows, and other phenomena. This was also true of the proposed test model, but at the time the 10x10 did not have a model exhaust capability. So, through an in-house effort over a period of only 5 months, a new model exhaust system was designed, installed, checked out, and made ready in time to support the scheduled test program.

  18. Analysis of high Reynolds numbers effects on a wind turbine airfoil using 2D wind tunnel test data

    NASA Astrophysics Data System (ADS)

    Pires, O.; Munduate, X.; Ceyhan, O.; Jacobs, M.; Snel, H.

    2016-09-01

    The aerodynamic behaviour of a wind turbine airfoil has been measured in a dedicated 2D wind tunnel test at the DNW High Pressure Wind Tunnel in Gottingen (HDG), Germany. The tests have been performed on the DU00W212 airfoil at different Reynolds numbers: 3, 6, 9, 12 and 15 million, and at low Mach numbers (below 0.1). Both clean and tripped conditions of the airfoil have been measured. An analysis of the impact of a wide Reynolds number variation over the aerodynamic characteristics of this airfoil has been performed.

  19. Wind tunnel and ground static tests of a .094 scale powered model of a modified T-39 lift/cruise fan V/STOL research airplane

    NASA Technical Reports Server (NTRS)

    Hunt, D.; Clinglan, J.; Salemann, V.; Omar, E.

    1977-01-01

    Ground static and wind tunnel test of a scale model modified T-39 airplane are reported. The configuration in the nose and replacement of the existing nacelles with tilting lift/cruise fans. The model was powered with three 14 cm diameter tip driven turbopowered simulators. Forces and moments were measured by an internal strain guage balance. Engine simulator thrust and mass flow were measured by calibrated pressure and temperature instrumentation mounted downstream of the fans. The low speed handling qualities and general aerodynamic characteristics of the modified T-39 were defined. Test variables include thrust level and thrust balance, forward speed, model pitch and sideslip angle at forward speeds, model pitch, roll, and ground height during static tests, lift/cruise fan tilt angle, flap and aileron deflection angle, and horizonal stabilizer angle. The effects of removing the landing gear, the lift/cruise fans, and the tail surfaces were also investigated.

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

    NASA Technical Reports Server (NTRS)

    Struzynski, N. A.

    1975-01-01

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

  1. Analysis of the wind tunnel test of a tilt rotor power force model

    NASA Technical Reports Server (NTRS)

    Marr, R. L.; Ford, D. G.; Ferguson, S. W.

    1974-01-01

    Two series of wind tunnel tests were made to determine performance, stability and control, and rotor wake interaction on the airframe, using a one-tenth scale powered force model of a tilt rotor aircraft. Testing covered hover (IGE/OCE), helicopter, conversion, and airplane flight configurations. Forces and moments were recorded for the model from predetermined trim attitudes. Control positions were adjusted to trim flight (one-g lift, pitching moment and drag zero) within the uncorrected test data balance accuracy. Pitch and yaw sweeps were made about the trim attitudes with the control held at the trimmed settings to determine the static stability characteristics. Tail on, tail off, rotors on, and rotors off configurations were testes to determine the rotor wake effects on the empennage. Results are presented and discussed.

  2. Forebody/Inlet of the Joint Strike Fighter Tested at Low Speeds

    NASA Technical Reports Server (NTRS)

    Johns, Albert L.

    1998-01-01

    As part of a national cooperative effort to develop a multinational fighter aircraft, a model of a Joint Strike Fighter concept was tested in several NASA Lewis Research Center wind tunnels at low speeds over a range of headwind velocities and model attitudes. This Joint Strike Fighter concept, which is scheduled to go into production in 2005, will greatly improve the range, capability, maneuverability, and survivability of fighter aircraft, and the production program could ultimately be worth $100 billion. The test program was a team effort between Lewis and Lockheed Martin Tactical Aircraft Systems. Testing was completed in September 1997, several weeks ahead of schedule, allowing Lockheed additional time to review the results and analysis data before the next test and resulting in significant cost savings for Lockheed. Several major milestones related to dynamic and steady-state data acquisition and overall model performance were reached during this model test. Results from this program will contribute to both the concept demonstration phase and the production aircraft.

  3. A numerical study of the effects of wind tunnel wall proximity on an airfoil model

    NASA Technical Reports Server (NTRS)

    Potsdam, Mark; Roberts, Leonard

    1990-01-01

    A procedure was developed for modeling wind tunnel flows using computational fluid dynamics. Using this method, a numerical study was undertaken to explore the effects of solid wind tunnel wall proximity and Reynolds number on a two-dimensional airfoil model at low speed. Wind tunnel walls are located at varying wind tunnel height to airfoil chord ratios and the results are compared with freestream flow in the absence of wind tunnel walls. Discrepancies between the constrained and unconstrained flows can be attributed to the presence of the walls. Results are for a Mach Number of 0.25 at angles of attack through stall. A typical wind tunnel Reynolds number of 1,200,000 and full-scale flight Reynolds number of 6,000,000 were investigated. At this low Mach number, wind tunnel wall corrections to Mach number and angle of attack are supported. Reynolds number effects are seen to be a consideration in wind tunnel testing and wall interference correction methods. An unstructured grid Navier-Stokes code is used with a Baldwin-Lomax turbulence model. The numerical method is described since unstructured flow solvers present several difficulties and fundamental differences from structured grid codes, especially in the area of turbulence modeling and grid generation.

  4. Free-to-Roll Testing of Airplane Models in Wind Tunnels

    NASA Technical Reports Server (NTRS)

    Capone, Francis J.; Owens, D. Bruce; Hall, Robert M.

    2007-01-01

    A free-to-roll (FTR) test technique and test rig make it possible to evaluate both the transonic performance and the wingdrop/ rock behavior of a high-strength airplane model in a single wind-tunnel entry. The free-to-roll test technique is a single degree-of-motion method in which the model is free to roll about the longitudinal axis. The rolling motion is observed, recorded, and analyzed to gain insight into wing-drop/rock behavior. Wing-drop/rock is one of several phenomena symptomatic of abrupt wing stall. FTR testing was developed as part of the NASA/Navy Abrupt Wing Stall Program, which was established for the purposes of understanding and preventing significant unexpected and uncommanded (thus, highly undesirable) lateral-directional motions associated with wing-drop/rock, which have been observed mostly in fighter airplanes under high-subsonic and transonic maneuvering conditions. Before FTR testing became available, wingrock/ drop behavior of high-performance airplanes undergoing development was not recognized until flight testing. FTR testing is a reliable means of detecting, and evaluating design modifications for reducing or preventing, very complex abrupt wing stall phenomena in a ground facility prior to flight testing. The FTR test rig was designed to replace an older sting attachment butt, such that a model with its force balance and support sting could freely rotate about the longitudinal axis. The rig (see figure) includes a rotary head supported in a stationary head with a forward spherical roller bearing and an aft needle bearing. Rotation is amplified by a set of gears and measured by a shaft-angle resolver; the roll angle can be resolved to within 0.067 degrees at a rotational speed up to 1,000 degrees/s. An assembly of electrically actuated brakes between the rotary and stationary heads can be used to hold the model against a rolling torque at a commanded roll angle. When static testing is required, a locking bar is used to fix the rotating

  5. Wind Tunnel Testing of Microtabs and Microjets for Active Load Control of Wind Turbine Blades

    NASA Astrophysics Data System (ADS)

    Cooperman, Aubryn Murray

    Increases in wind turbine size have made controlling loads on the blades an important consideration for future turbine designs. One approach that could reduce extreme loads and minimize load variation is to incorporate active control devices into the blades that are able to change the aerodynamic forces acting on the turbine. A wind tunnel model has been constructed to allow testing of different active aerodynamic load control devices. Two such devices have been tested in the UC Davis Aeronautical Wind Tunnel: microtabs and microjets. Microtabs are small surfaces oriented perpendicular to an airfoil surface that can be deployed and retracted to alter the lift coefficient of the airfoil. Microjets produce similar effects using air blown perpendicular to the airfoil surface. Results are presented here for both static and dynamic performance of the two devices. Microtabs, located at 95% chord on the lower surface and 90% chord on the upper surface, with a height of 1% chord, produce a change in the lift coefficient of 0.18, increasing lift when deployed on the lower surface and decreasing lift when deployed on the upper surface. Microjets with a momentum coefficient of 0.006 at the same locations produce a change in the lift coefficient of 0.19. The activation time for both devices is less than 0.3 s, which is rapid compared to typical gust rise times. The potential of active device to mitigate changes in loads was tested using simulated gusts. The gusts were produced in the wind tunnel by accelerating the test section air speed at rates of up to 7 ft/s 2. Open-loop control of microtabs was tested in two modes: simultaneous and sequential tab deployment. Activating all tabs along the model span simultaneously was found to produce a change in the loads that occurred more rapidly than a gust. Sequential tab deployment more closely matched the rates of change due to gusts and tab deployment. A closed-loop control system was developed for the microtabs using a simple

  6. Correaltion of full-scale drag predictions with flight measurements on the C-141A aircraft. Phase 2: Wind tunnel test, analysis, and prediction techniques. Volume 1: Drag predictions, wind tunnel data analysis and correlation

    NASA Technical Reports Server (NTRS)

    Macwilkinson, D. G.; Blackerby, W. T.; Paterson, J. H.

    1974-01-01

    The degree of cruise drag correlation on the C-141A aircraft is determined between predictions based on wind tunnel test data, and flight test results. An analysis of wind tunnel tests on a 0.0275 scale model at Reynolds number up to 3.05 x 1 million/MAC is reported. Model support interference corrections are evaluated through a series of tests, and fully corrected model data are analyzed to provide details on model component interference factors. It is shown that predicted minimum profile drag for the complete configuration agrees within 0.75% of flight test data, using a wind tunnel extrapolation method based on flat plate skin friction and component shape factors. An alternative method of extrapolation, based on computed profile drag from a subsonic viscous theory, results in a prediction four percent lower than flight test data.

  7. ViDI: Virtual Diagnostics Interface. Volume 1; The Future of Wind Tunnel Testing

    NASA Technical Reports Server (NTRS)

    Fleming, Gary A. (Technical Monitor); Schwartz, Richard J.

    2004-01-01

    The quality of data acquired in a given test facility ultimately resides within the fidelity and implementation of the instrumentation systems. Over the last decade, the emergence of robust optical techniques has vastly expanded the envelope of measurement possibilities. At the same time the capabilities for data processing, data archiving and data visualization required to extract the highest level of knowledge from these global, on and off body measurement techniques have equally expanded. Yet today, while the instrumentation has matured to the production stage, an optimized solution for gaining knowledge from the gigabytes of data acquired per test (or even per test point) is lacking. A technological void has to be filled in order to possess a mechanism for near-real time knowledge extraction during wind tunnel experiments. Under these auspices, the Virtual Diagnostics Interface, or ViDI, was developed.

  8. Nozzle diffuser for use with an open test section of a wind tunnel

    NASA Technical Reports Server (NTRS)

    Barna, P. Stephen (Inventor)

    1993-01-01

    The nozzle diffuser has an inlet in fluid communication with the narrowed inlet of an open test chamber in a conventional wind tunnel. The nozzle diffuser has a passageway extending from its inlet to an outlet in communication with the open test section. The passageway has an internal cross sectional area which increases from its inlet to its outlet and which may be defined by top and bottom isosceles trapezoid walls of a particular flare angle and by isosceles trapezoid side walls of a different flare angle. In addition, a collector having a decreasing internal cross sectional area from inlet to outlet may be provided at the opposite end of the test chamber such that its outlet is in communication with a diffuser located at this outlet.

  9. Improving Large-Scale Testing Capability by Modifying the 40- by 80-ft Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Mort, Kenneth W.; Soderman, Paul T.; Eckert, William T.

    1979-01-01

    Interagency studies conducted during the last several years have indicated the need to Improve full-scale testing capabilities. The studies showed that the most effective trade between test capability and facility cost was provided by re-powering the existing Ames Research Center 40- by 80-ft Wind Tunnel to Increase the maximum speed from about 100 m/s (200 knots) lo about 150 m/s (300 knots) and by adding a new 24- by 37-m (80- by 120-ft) test section powered for about a 50-m/s (100-knot) maximum speed. This paper reviews the design of the facility, a few or its capabilities, and some of its unique features.

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

    NASA Technical Reports Server (NTRS)

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

    1975-01-01

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

  11. Tests of the NACA 0025 and 0035 Airfoils in the Full-Scale Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Bullivant, W Kenneth

    1941-01-01

    This report presents the results of an investigation conducted in the NACA full-scale wind tunnel to determine the aerodynamic characteristics of the 6 by 36-foot rectangular NACA 0025 and 0035 airfoils. The aerodynamic characteristics of the plain airfoils with rounded and square tips were determined by force tests through a complete angle-of-attack range, in addition, the profile drag was determined by the momentum method. The transition points on the airfoils were located by boundary-layer determinations with small total-head and static tubes. Each airfoil was also tested with a 0.20c full-span split flap. Tuft surveys were included to show the progressive breakdown of flow with increasing angles of attack. Previously published data from tests of the NACA 0009, 0012, and 0018 airfoils in the full-scale tunnel have been included in the summary curves.

  12. Two-dimensional wind tunnel test of an oscillating rotor airfoil, volume 1

    NASA Technical Reports Server (NTRS)

    Dadone, L. U.

    1977-01-01

    A two dimensional wind tunnel test was conducted to obtain the quasisteady and unsteady characteristics of an advanced airfoil designed for helicopter rotor applications. Differential pressures were measured at 17 locations along the chord of the airfoil model. The airfoil motions were sinusoidal forced-pitch oscillations about the quarter chord at amplitudes varying from 2.5 to 10.0 degrees and at frequencies from 23 Hz to 90 Hz. The quasisteady tests were conducted at Mach numbers from 0.2 to 0.9, and the oscillatory tests between M = 0.2 and M = 0.7. At quasisteady conditions a limited number of drag measurements were made with a wake-traversing probe.

  13. On the use of freon-12 for increasing Reynolds number in wind-tunnel testing of three dimensional aircraft models at subcritical and supercritical Mach numbers

    NASA Technical Reports Server (NTRS)

    Treon, S. L.; Hofstetter, W. R.; Abbott, F. T.

    1971-01-01

    The aerodynamic suitability of Freon-12 for general wind-tunnel testing was investigated at low and high subsonic speeds. Static aerodynamic characteristics of two transport airplane models were determined from strain gage balance measurements in both air and Freon-12 at several Reynolds numbers. A low-speed high-lift configuration was evaluated at Mach number 0.25, and a high-speed cruise wing-fuselage combination was tested at Mach numbers up to 0.825. The data obtained in air and in Freon-12 agree well, even in stalled flow, until compressibility effects evidently become significant in air and in Freon-12 agree well, even in stalled flow, until compressibility effects evidently become significant in air.

  14. DDS-Suite - A Dynamic Data Acquisition, Processing, and Analysis System for Wind Tunnel Testing

    NASA Technical Reports Server (NTRS)

    Burnside, Jathan J.

    2012-01-01

    Wind Tunnels have optimized their steady-state data systems for acquisition and analysis and even implemented large dynamic-data acquisition systems, however development of near real-time processing and analysis tools for dynamic-data have lagged. DDS-Suite is a set of tools used to acquire, process, and analyze large amounts of dynamic data. Each phase of the testing process: acquisition, processing, and analysis are handled by separate components so that bottlenecks in one phase of the process do not affect the other, leading to a robust system. DDS-Suite is capable of acquiring 672 channels of dynamic data at rate of 275 MB / s. More than 300 channels of the system use 24-bit analog-to-digital cards and are capable of producing data with less than 0.01 of phase difference at 1 kHz. System architecture, design philosophy, and examples of use during NASA Constellation and Fundamental Aerodynamic tests are discussed.

  15. Evaluation and Analysis of F-16XL Wind Tunnel Data from Dynamic Tests

    NASA Technical Reports Server (NTRS)

    Kim, Sungwan; Murphy, Patrick C.; Klein, Vladislav

    2003-01-01

    A series of wind tunnel tests were conducted in the NASA Langley Research Center as part of an ongoing effort to develop and test mathematical models for aircraft rigid-body aerodynamics in nonlinear unsteady flight regimes. Analysis of measurement accuracy, especially for nonlinear dynamic systems that may exhibit complicated behaviors, is an essential component of this ongoing effort. In this paper, tools for harmonic analysis of dynamic data and assessing measurement accuracy are presented. A linear aerodynamic model is assumed that is appropriate for conventional forced-oscillation experiments, although more general models can be used with these tools. Application of the tools to experimental data is demonstrated and results indicate the levels of uncertainty in output measurements that can arise from experimental setup, calibration procedures, mechanical limitations, and input errors.

  16. Evaluation and Analysis of F-16XL Wind Tunnel Data From Static and Dynamic Tests

    NASA Technical Reports Server (NTRS)

    Kim, Sungwan; Murphy, Patrick C.; Klein, Vladislav

    2004-01-01

    A series of wind tunnel tests were conducted in the NASA Langley Research Center as part of an ongoing effort to develop and test mathematical models for aircraft rigid-body aerodynamics in nonlinear unsteady flight regimes. Analysis of measurement accuracy, especially for nonlinear dynamic systems that may exhibit complicated behaviors, is an essential component of this ongoing effort. In this report, tools for harmonic analysis of dynamic data and assessing measurement accuracy are presented. A linear aerodynamic model is assumed that is appropriate for conventional forced-oscillation experiments, although more general models can be used with these tools. Application of the tools to experimental data is demonstrated and results indicate the levels of uncertainty in output measurements that can arise from experimental setup, calibration procedures, mechanical limitations, and input errors.

  17. RSRA sixth scale wind tunnel test. Tabulated balance data, volume 2

    NASA Technical Reports Server (NTRS)

    Ruddell, A.; Flemming, R.

    1974-01-01

    Summaries are presented of all the force and moment data acquired during the RSRA Sixth Scale Wind Tunnel Test. These data include and supplement the data presented in curve form in previous reports. Each summary includes the model configuration, wing and empennage incidences and deflections, and recorded balance data. The first group of data in each summary presents the force and moment data in full scale parametric form, the dynamic pressure and velocity in the test section, and the powered nacelle fan speed. The second and third groups of data are the balance data in nondimensional coefficient form. The wind axis coefficient data corresponds to the parametric data divided by the wing area for forces and divided by the product of the wing area and wing span or mean aerodynamic chord for moments. The stability axis data resolves the wind axis data with respect to the angle of yaw.

  18. Hypersonic Wind Tunnel Test of a Flare-type Membrane Aeroshell for Atmospheric Entry Capsules

    NASA Astrophysics Data System (ADS)

    Yamada, Kazuhiko; Koyama, Masashi; Kimura, Yusuke; Suzuki, Kojiro; Abe, Takashi; Koichi Hayashi, A.

    A flexible aeroshell for atmospheric entry vehicles has attracted attention as an innovative space transportation system. In this study, hypersonic wind tunnel tests were carried out to investigate the behavior, aerodynamic characteristics and aerodynamic heating environment in hypersonic flow for a previously developed capsule-type vehicle with a flare-type membrane aeroshell made of ZYLON textile sustained by a rigid torus frame. Two different models with different flare angles (45º and 60º) were tested to experimentally clarify the effect of flare angle. Results indicate that flare angle of aeroshell has significant and complicate effect on flow field and aerodynamic heating in hypersonic flow at Mach 9.45 and the flare angle is very important parameter for vehicle design with the flare-type membrane aeroshell.

  19. Wind Tunnel Test Results for Gas Flows Inside Axisymmetric Cavities on Cylindric Bodies with Nose Cones

    NASA Technical Reports Server (NTRS)

    Shvets, A. L.; Gilinsky, M.; Blankson, I. M.

    2004-01-01

    Experimental test results of air flow inside and at the cylindrical cavity located on axisymmetric body are presented. These tests were conducted in the wind tunnel A-7 of Institute of Mechanics at Moscow State University. Pressure distribution along the cavities and optical measurements were obtained. Dependence of these characteristics of length of a cavity in the range: L/D = 0.5 - 14 and free stream Mach in the range: M(sub infinity) = 0.6 - 3.0 was determined. Flow structure inside the cavity, cause of flow regime change, separation zones geometry and others were studied. In particular, the flow modes of with open and closed separation zones are determined.

  20. Development of an Active Twist Rotor for Wind: Tunnel Testing (NLPN97-310

    NASA Technical Reports Server (NTRS)

    Cesnik, Carlos E. S.; Shin, SangJoon; Hagood, Nesbitt W., IV

    1998-01-01

    The development of the Active Twist Rotor prototype blade for hub vibration and noise reduction studies is presented in this report. Details of the modeling, design, and manufacturing are explored. The rotor blade is integrally twisted by direct strain actuation. This is accomplished by distributing embedded piezoelectric fiber composites along the span of the blade. The development of the analysis framework for this type of active blade is presented. The requirements for the prototype blade, along with the final design results are also presented. A detail discussion on the manufacturing aspects of the prototype blade is described. Experimental structural characteristics of the prototype blade compare well with design goals, and preliminary bench actuation tests show lower performance than originally predicted. Electrical difficulties with the actuators are also discussed. The presented prototype blade is leading to a complete fully articulated four-blade active twist rotor system for future wind tunnel tests.

  1. Experimental Data from the Benchmark SuperCritical Wing Wind Tunnel Test on an Oscillating Turntable

    NASA Technical Reports Server (NTRS)

    Heeg, Jennifer; Piatak, David J.

    2013-01-01

    The Benchmark SuperCritical Wing (BSCW) wind tunnel model served as a semi-blind testcase for the 2012 AIAA Aeroelastic Prediction Workshop (AePW). The BSCW was chosen as a testcase due to its geometric simplicity and flow physics complexity. The data sets examined include unforced system information and forced pitching oscillations. The aerodynamic challenges presented by this AePW testcase include a strong shock that was observed to be unsteady for even the unforced system cases, shock-induced separation and trailing edge separation. The current paper quantifies these characteristics at the AePW test condition and at a suggested benchmarking test condition. General characteristics of the model's behavior are examined for the entire available data set.

  2. Wind-tunnel Tests of a Hall High-life Wing

    NASA Technical Reports Server (NTRS)

    Weick, Fred E; Sanders, Robert

    1932-01-01

    Wind-tunnel tests have been made to find the lift, drag, and center-of-pressure characteristics of a Hall high-lift wing model. The Hall wing is essentially a split-flap airfoil with an internal air passage. Air enters the passage through an opening in the lower surface somewhat back of and parallel to the leading edge, and flows out through an opening made by deflecting the rear portion of the under surface downward as a flap. For ordinary flight conditions the front opening and the rear flap can be closed, providing in effect a conventional airfoil (the Clark Y in this case). The tests were made with various flap settings and with the entrance to the passage both open and closed. The highest lift coefficient found, C(sub L) = 2.08, was obtained with the passage closed.

  3. Test technique development in interference free testing, flow visualization, and remote control model technology at Langley's Unitary Plan wind tunnel

    NASA Technical Reports Server (NTRS)

    Corlett, W. A.

    1979-01-01

    A metric half-span model is considered as a means of mechanical support for a wind-tunnel model which allows measurement of aerodynamic forces and moments without support interference or model distortion. This technique can be applied to interference-free propulsion models. The vapor screen method of flow visualization at supersonic Mach numbers is discussed. The use of smoke instead of water vapor as a medium to produce the screen is outlined. Vapor screen data are being used in the development of analytical vortex tracking programs. Test results for a remote control model system are evaluated. Detailed control effectiveness and cross-coupling data were obtained with a single run. For the afterbody tail configuration, tested control boundaries at several roll orientations were established utilizing the facility's on-line capability to 'fly' the model in the wind tunnel.

  4. Dry wind tunnel system

    NASA Technical Reports Server (NTRS)

    Chen, Ping-Chih (Inventor)

    2013-01-01

    This invention is a ground flutter testing system without a wind tunnel, called Dry Wind Tunnel (DWT) System. The DWT system consists of a Ground Vibration Test (GVT) hardware system, a multiple input multiple output (MIMO) force controller software, and a real-time unsteady aerodynamic force generation software, that is developed from an aerodynamic reduced order model (ROM). The ground flutter test using the DWT System operates on a real structural model, therefore no scaled-down structural model, which is required by the conventional wind tunnel flutter test, is involved. Furthermore, the impact of the structural nonlinearities on the aeroelastic stability can be included automatically. Moreover, the aeroservoelastic characteristics of the aircraft can be easily measured by simply including the flight control system in-the-loop. In addition, the unsteady aerodynamics generated computationally is interference-free from the wind tunnel walls. Finally, the DWT System can be conveniently and inexpensively carried out as a post GVT test with the same hardware, only with some possible rearrangement of the shakers and the inclusion of additional sensors.

  5. Results of a landing gear loads test using a 0.0405-scale model (16-0) of the space shuttle orbiter in the Rockwell International NAAL wind tunnel (OA163), volume 1

    NASA Technical Reports Server (NTRS)

    Mennell, R. C.

    1976-01-01

    Experimental aerodynamic investigations were conducted on a sting mounted scale representation of the 140C outer mold line space shuttle orbiter configuration in the low speed wind tunnel. The primary test objectives were to define the orbiter landing gear system pressure loading and to record landing gear door and strut hingemoment levels. Secondary objectives included recording the aerodynamic influence of various landing gear configurations on orbiter force data as well as investigating 40 x 80 ft. Ames Wind Tunnel strut simulation effects on both orbiter landing gear loads and aerodynamic characteristics. Testing was conducted at a Mach number of 0.17, free stream dynamic pressure of 42.5 PSF, and Reynolds number per unit length of 1.2 million per foot. Angle of attack variation was 0 to 20 while yaw angles ranged from -10 to 10 deg.

  6. Experimental study on the flight dynamics of a bioinspired ornithopter: free flight testing and wind tunnel testing

    NASA Astrophysics Data System (ADS)

    Lee, Jun-Seong; Han, Jae-Hung

    2012-09-01

    This study experimentally shows the flight dynamics of a bioinspired ornithopter using two different types of approach: (1) free flight testing, and (2) wind tunnel testing. An ornithopter is flown in straight and level flight with a fixed wingbeat frequency and tail elevation angle. A three-dimensional visual tracking system is applied to follow the retro-reflective markers on the ornithopter and record the flight trajectories. The unique oscillatory behavior of the body in the longitudinal plane is observed in the free flight testing and the detailed wing and tail deformations are also obtained. Based on the trim flight data, a specially devised tether device is designed and employed to emulate the free flight conditions in the wind tunnel. The tether device provides minimal mechanical interference and longitudinal flight dynamic characteristics similar to those of free flight. On introducing a pitching moment disturbance to the body, the oscillation recovered to the original trajectory turns out to be a stable limit-cycle oscillation (LCO). During the wind tunnel testing, the magnitude of LCO is effectively suppressed by active tail motion.

  7. An engineering study of hybrid adaptation of wind tunnel walls for three-dimensional testing

    NASA Technical Reports Server (NTRS)

    Brown, Clinton; Kalumuck, Kenneth; Waxman, David

    1987-01-01

    Solid wall tunnels having only upper and lower walls flexing are described. An algorithm for selecting the wall contours for both 2 and 3 dimensional wall flexure is presented and numerical experiments are used to validate its applicability to the general test case of 3 dimensional lifting aircraft models in rectangular cross section wind tunnels. The method requires an initial approximate representation of the model flow field at a given lift with wallls absent. The numerical methods utilized are derived by use of Green's source solutions obtained using the method of images; first order linearized flow theory is employed with Prandtl-Glauert compressibility transformations. Equations are derived for the flexed shape of a simple constant thickness plate wall under the influence of a finite number of jacks in an axial row along the plate centerline. The Green's source methods are developed to provide estimations of residual flow distortion (interferences) with measured wall pressures and wall flow inclinations as inputs.

  8. Dynamic Wind-Tunnel Testing of a Sub-Scale Iced S-3B Viking

    NASA Technical Reports Server (NTRS)

    Lee, Sam; Barnhart, Billy; Ratvasky, Thomas P.

    2012-01-01

    The effect of ice accretion on a 1/12-scale complete aircraft model of S-3B Viking was studied in a rotary-balance wind tunnel. Two types of ice accretions were considered: ice protection system failure shape and runback shapes that form downstream of the thermal ice protection system. The results showed that the ice shapes altered the stall characteristics of the aircraft. The ice shapes also reduced the control surface effectiveness, but mostly near the stall angle of attack. There were some discrepancies with the data with the flaps deflected that were attributed to the low Reynolds number of the test. Rotational and forced-oscillation studies showed that the effects of ice were mostly in the longitudinal forces, and the effects on the lateral forces were relatively minor.

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

    NASA Technical Reports Server (NTRS)

    Kemp, W. B., Jr.

    1986-01-01

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

  10. Comparison of Force and Moment Coefficients for the Same Test Article in Multiple Wind Tunnels

    NASA Technical Reports Server (NTRS)

    Deloach, Richard

    2013-01-01

    This paper compares the results of force and moment measurements made on the same test article and with the same balance in three transonic wind tunnels. Comparisons are made for the same combination of Reynolds number, Mach number, sideslip angle, control surface configuration, and angle of attack range. Between-tunnel force and moment differences are quantified. An analysis of variance was performed at four unique sites in the design space to assess the statistical significance of between-tunnel variation and any interaction with angle of attack. Tunnel to tunnel differences too large to attribute to random error were detected were observed for all forces and moments. In some cases these differences were independent of angle of attack and in other cases they changed with angle of attack.

  11. Wind Tunnel Testing on Crosswind Aerodynamic Forces Acting on Railway Vehicles

    NASA Astrophysics Data System (ADS)

    Kwon, Hyeok-Bin; Nam, Seong-Won; You, Won-Hee

    This study is devoted to measure the aerodynamic forces acting on two railway trains, one of which is a high-speed train at 300km/h maximum operation speed, and the other is a conventional train at the operating speed 100km/h. The three-dimensional train shapes have been modeled as detailed as possible including the inter-car, the upper cavity for pantograph, and the bogie systems. The aerodynamic forces on each vehicle of the trains have been measured in the subsonic wind tunnel with 4m×3m test section of Korea Aerospace Research Institute at Daejeon, Korea. The aerodynamic forces and moments of the train models have been plotted for various yaw angles and the characteristics of the aerodynamic coefficients has been discussed relating to the experimental conditions.

  12. Wind tunnel tests of a free-wing/free-trimmer model

    NASA Technical Reports Server (NTRS)

    Sandlin, D. R.

    1982-01-01

    The riding qualities of an aircraft with low wing loading can be improved by freeing the wing to rotate about its spanwise axis. A trimming surface also free to rotate about its spanwise axis can be added at the wing tips to permit the use of high lift devices. Wind tunnel tests of the free wing/free trimmer model with the trimmer attached to the wing tips aft of the wing chord were conducted to validate a mathematical model developed to predict the dynamic characteristics of a free wing/free trimmer aircraft. A model consisting of a semispan wing with the trimmer mounted on with the wing on an air bearing and the trimmer on a ball bearing was displaced to various angles of attack and released. The damped oscillations of the wing and trimmer were recorded. Real and imaginary parts of the characteristic equations of motion were determined and compared to values predicted using the mathematical model.

  13. Wind-tunnel Tests of the Fowler Variable-area Wing

    NASA Technical Reports Server (NTRS)

    Weick, Fred E; Platt, Robert C

    1932-01-01

    The lift, drag, and center of pressure characteristics of a model of the Fowler variable-area wing were measured in the NACA 7 by 10 foot wind tunnel. The Fowler wing consists of a combination of a main wing and an extension surface, also of airfoil section. The extension surface can be entirely retracted within the lower rear portion of the main wing or it can be moved to the rear and downward. The tests were made with the nose of the extension airfoil in various positions near the trailing edge of the main wing and with the surface at various angular deflections. The highest lift coefficient obtained was C(sub L) = 3.17 as compared with 1.27 for the main wing alone.

  14. Wind tunnel tests for a flapping wing model with a changeable camber using macro-fiber composite actuators

    NASA Astrophysics Data System (ADS)

    Kim, Dae-Kwan; Han, Jae-Hung; Kwon, Ki-Jung

    2009-02-01

    In the present study, a biomimetic flexible flapping wing was developed on a real ornithopter scale by using macro-fiber composite (MFC) actuators. With the actuators, the maximum camber of the wing can be linearly changed from -2.6% to +4.4% of the maximum chord length. Aerodynamic tests were carried out in a low-speed wind tunnel to investigate the aerodynamic characteristics, particularly the camber effect, the chordwise flexibility effect and the unsteady effect. Although the chordwise wing flexibility reduces the effective angle of attack, the maximum lift coefficient can be increased by the MFC actuators up to 24.4% in a static condition. Note also that the mean values of the perpendicular force coefficient rise to a value of considerably more than 3 in an unsteady aerodynamic flow region. Additionally, particle image velocimetry (PIV) tests were performed in static and dynamic test conditions to validate the flexibility and unsteady effects. The static PIV results confirm that the effective angle of attack is reduced by the coupling of the chordwise flexibility and the aerodynamic force, resulting in a delay in the stall phenomena. In contrast to the quasi-steady flow condition of a relatively high advance ratio, the unsteady aerodynamic effect due to a leading edge vortex can be found along the wing span in a low advance ratio region. The overall results show that the chordwise wing flexibility can produce a positive effect on flapping aerodynamic characteristics in quasi-steady and unsteady flow regions; thus, wing flexibility should be considered in the design of efficient flapping wings.

  15. Results of investigations on a 0.0405 scale model ATP version of the NR-SSV orbiter in the North American Aeronautical Laboratory low speed wind tunnel

    NASA Technical Reports Server (NTRS)

    Mennell, R.; Vaughn, J. E.; Singellton, R.

    1973-01-01

    Experimental aerodynamic investigations were conducted on a scale model space shuttle vehicle (SSV) orbiter. The purpose of the test was to investigate the longitudinal and lateral-directional aerodynamic characteristics. Emphasis was placed on model component, wing-glove, and wing-body fairing effects, as well as elevon, aileron, and rudder control effectiveness. Angles of attack from - 5 deg to + 30 deg and angles of sideslip from - 5 deg to + 10 deg were tested. Static pressures were recorded on base, fuselage, and wing surfaces. Tufts and talc-kerosene flow visualization techniques were also utilized. The aerodynamic force balance results are presented in plotted and tabular form.

  16. Results of two tests in the MSFC 14 by 14-inch trisonic wind tunnel, FA 27 (TWT-655) and FA 28 (TWT-656)

    NASA Technical Reports Server (NTRS)

    Braddock, W. F.

    1979-01-01

    Wind tunnel tests were conducted in a 14- inch wind tunnel with a 0.004 scale model of the space shuttle launch vehicle in order to (1) determine the cause and possible aerodynamic alterations required to eliminate the Orbiter rolling moment couple; (2) determine configuration alterations to alleviate the forward Orbiter external tank loads; and (3) provide data to verify previous data.

  17. Supersonic Aftbody Closure Wind-Tunnel Testing, Data Analysis, and Computational Results

    NASA Technical Reports Server (NTRS)

    Allen, Jerry; Martin, Grant; Kubiatko, Paul

    1999-01-01

    This paper reports on the model, test, and results from the Langley Supersonic Aftbody Closure wind tunnel test. This project is an experimental evaluation of the 1.5% Technology Concept Aircraft (TCA) aftbody closure model (Model 23) in the Langley Unitary Plan Wind Tunnel. The baseline TCA design is the result of a multidisciplinary, multipoint optimization process and was developed using linear design and analysis methods, supplemented with Euler and Navier-Stokes numerical methods. After a thorough design review, it was decided to use an upswept blade attached to the forebody as the mounting system. Structural concerns dictated that a wingtip support system would not be feasible. Only the aftbody part of the model is metric. The metric break was chosen to be at the fuselage station where prior aft-sting supported models had been truncated. Model 23 is thus a modified version of Model 20. The wing strongback, flap parts, and nacelles from Model 20 were used, whereas new aftbodies, a common forebody, and some new tails were fabricated. In summary, significant differences in longitudinal and direction stability and control characteristics between the ABF and ABB aftbody geometries were measured. Correcting the experimental data obtained for the TCA configuration with the flared aftbody to the representative of the baseline TCA closed aftbody will result in a significant reduction in longitudinal stability, a moderate reduction in stabilizer effectiveness and directional stability, and a moderate to significant reduction in rudder effectiveness. These reductions in the stability and control effectiveness levels of the baseline TCA closed aftbody are attributed to the reduction in carry-over area.

  18. Tests of Wing Machine-Gun and Cannon Installations in the NACA Full-Scale Wind Tunnel, Special Report

    NASA Technical Reports Server (NTRS)

    Czarnecki, K. R.; Guryansky, Eugene R.

    1941-01-01

    At the request of the Bureau of Aeronautics, an investigation was conducted in the full-scale wind tunnel of wing installations of .50-caliber machine guns and 20-millimeter cannons. The tests were made to determine the effect of various gun installations on the maximum lift and the high-speed drag of the airplane.

  19. Modification of the Ames 40- by 80-foot wind tunnel for component acoustic testing for the second generation supersonic transport

    NASA Technical Reports Server (NTRS)

    Schmitz, F. H.; Allmen, J. R.; Soderman, P. T.

    1994-01-01

    The development of a large-scale anechoic test facility where large models of engine/airframe/high-lift systems can be tested for both improved noise reduction and minimum performance degradation is described. The facility development is part of the effort to investigate economically viable methods of reducing second generation high speed civil transport noise during takeoff and climb-out that is now under way in the United States. This new capability will be achieved through acoustic modifications of NASA's second largest subsonic wind tunnel: the 40-by 80-Foot Wind Tunnel at the NASA Ames Research Center. Three major items are addressed in the design of this large anechoic and quiet wind tunnel: a new deep (42 inch (107 cm)) test section liner, expansion of the wind tunnel drive operating envelope at low rpm to reduce background noise, and other promising methods of improving signal-to-noise levels of inflow microphones. Current testing plans supporting the U.S. high speed civil transport program are also outlined.

  20. Design and preliminary test results at Mach 5 of an axisymmetric slotted sound shield. [for supersonic wind tunnels (noise reduction in wind tunnel nozzles)

    NASA Technical Reports Server (NTRS)

    Beckwith, I. E.; Spokowski, A. J.; Harvey, W. D.; Stainback, P. C.

    1975-01-01

    The basic theory and sound attenuation mechanisms, the design procedures, and preliminary experimental results are presented for a small axisymmetric sound shield for supersonic wind tunnels. The shield consists of an array of small diameter rods aligned nearly parallel to the entrance flow with small gaps between the rods for boundary layer suction. Results show that at the lowest test Reynolds number (based on rod diameter) of 52,000 the noise shield reduced the test section noise by about 60 percent ( or 8 db attenuation) but no attenuation was measured for the higher range of test reynolds numbers from 73,000 to 190,000. These results are below expectations based on data reported elsewhere on a flat sound shield model. The smaller attenuation from the present tests is attributed to insufficient suction at the gaps to prevent feedback of vacuum manifold noise into the shielded test flow and to insufficient suction to prevent transition of the rod boundary layers to turbulent flow at the higher Reynolds numbers. Schlieren photographs of the flow are shown.

  1. Development and Operation of an Automatic Rotor Trim Control System for the UH-60 Individual Blade Control Wind Tunnel Test

    NASA Technical Reports Server (NTRS)

    Theodore, Colin R.; Tischler, Mark B.

    2010-01-01

    An automatic rotor trim control system was developed and successfully used during a wind tunnel test of a full-scale UH-60 rotor system with Individual Blade Control (IBC) actuators. The trim control system allowed rotor trim to be set more quickly, precisely and repeatably than in previous wind tunnel tests. This control system also allowed the rotor trim state to be maintained during transients and drift in wind tunnel flow, and through changes in IBC actuation. The ability to maintain a consistent rotor trim state was key to quickly and accurately evaluating the effect of IBC on rotor performance, vibration, noise and loads. This paper presents details of the design and implementation of the trim control system including the rotor system hardware, trim control requirements, and trim control hardware and software implementation. Results are presented showing the effect of IBC on rotor trim and dynamic response, a validation of the rotor dynamic simulation used to calculate the initial control gains and tuning of the control system, and the overall performance of the trim control system during the wind tunnel test.

  2. Wind-tunnel investigation of the OMAC canard configuration

    NASA Technical Reports Server (NTRS)

    Ingram, W. C.; Yip, L. P.; Cook, E. L.

    1986-01-01

    Wind-tunnel tests were conducted on a 0.175-scale model of the OMAC Laser 300 canard configuration in the NASA Langley 12-Foot Low-Speed Wind Tunnel to determine its low-speed high angel-of-attack aerodynamic characteristics. The Laser 300 is a general aviation turboprop pusher aircraft utilizing a canard configuration. The design incorporates a low forward wing and a high main wing with a leading-edge droop installed on the outboard panel and tip fins mounted on the wing tips. The model was tested over a range of -6 to 50-deg angle-of-attack and 20 to -20 deg sideslip. Static force and moment data were measured, and the longitudinal and lateral-directional characteristics were determined.

  3. Low-speed stability and control wind-tunnel investigations of effects of spanwise blowing on fighter flight characteristics at high angles of attack. [Langely 12-ft low-speed tunnel and 30- by 60-ft tunnel

    NASA Technical Reports Server (NTRS)

    Satran, D. R.; Gilbert, W. P.; Anglin, E. L.

    1985-01-01

    The effects of spanwise blowing on two configurations representative of current fighter airplanes were investigated. The two configurations differed only in wing planform, with one incorporating a trapezoidal wing and the other a 60 delta wing. Emphasis was on determining the lateral-directional characteristics, particularly in the stall/departure angle-of-attack range; however, the effects of spanwise blowing on the longitudinal aerodynamics were also determined. The-tunnel tests included measurement of static force and forced-oscillation aerodynamic data, visualization of the airflow changes created by the spanwise blowing, and free-flight model tests. The effects of blowing rate, chordwise location of the blowing ports, asymmetric blowing, and blowing on the conventional aerodynamic control characteristics were investigated. In the angle-of-attack regions in which the spanwise blowing substantially improved the wing upper-surface flow field (i.e., provided reattachment of the flow aft of the leading-edge vortex), improvements in both static and dynamic lateral-directional stability were observed. Blowing effects on stability could be proverse or adverse depending on blowing rate, blowing port loaction, and wing planform. Free-flight model tests of the trapezoidal wing confirmed the beneficial effects of spanwise blowing measured in the static and dynamic force tests.

  4. Wind-tunnel investigation of the validity of a sonic-boom-minimization concept. [Langley Unitary Plan Wind Tunnel tests for supersonic transport design

    NASA Technical Reports Server (NTRS)

    Mack, R. J.; Darden, C. M.

    1979-01-01

    The Langley unitary plan unitary plan wind tunnel was used to determine the validity of a sonic-boom-minimization theory. Five models - two reference and three low-boom constrained - were tested at design Mach numbers of 1.5 and 2.7. Results show that the pressure signatures generated by the low-boom models had significantly lower overpressure levels than those produced by the reference models and that small changes in the Mach number and/or the lift caused relatively small changes in the signature shape and overpressure level. Boundary-layer effects were found in the signature shape and overpressure level. Boundary-layer effects were found to be sizable on the low-boom models, and when viscous corrections were included in the analysis, improved agreement between the predicted and the measured signatures was noted. Since this agreement was better at Mach 1.5 than at Mach 2.7, it was concluded that the minimization method was definitely valid at Mach 1.5 and was probably valid at Mach 2.7, with further work needed to resolve the uncertainty.

  5. Fast response vanes for sensing flow patterns in helicopter rotor environment. [wind tunnel tests of modified helicopter rotary wing

    NASA Technical Reports Server (NTRS)

    Barna, P. S.; Crossman, G. R.

    1974-01-01

    Wind tunnel experiments were conducted on four small-scale flow-direction vanes for the determination of aerodynamic response. The tests were further extended to include a standard sized low-inertia vane currently employed in aircraft flight testing. The four test vanes had different aspect ratios and were about 35 percent of the surface area of the standard vane. The test results indicate satisfactory damping and frequency response for all vanes tested and compare favorably with the standard design.

  6. Large-scale V/STOL testing. [conducted in the Ames 40- by 80-foot wind tunnel

    NASA Technical Reports Server (NTRS)

    Koenig, D. G.; Aiken, T. N.; Aoyagi, K.; Falarshi, M. D.

    1977-01-01

    Several facets of large-scale testing of V/STOL aircraft configurations are discussed with particular emphasis on test experience in the Ames 40- by 80-Foot Wind Tunnel. Examples of powered-lift test programs are presented in order to illustrate tradeoffs confronting the planner of V/STOL test programs. Large-scale V/STOL wind-tunnel testing can sometimes compete with small-scale testing in the effort required (overall test time) and program costs because of the possibility of conducting a number of different tests with a single large-scale model where several small-scale models would be required. The benefits of both high- or full-scale Reynolds numbers, more detailed configuration simulation, and number and type of onboard measurements are studied.

  7. Wind Tunnel Test of Subscale Ringsail and Disk-Gap-Band Parachutes

    NASA Technical Reports Server (NTRS)

    Zumwalt, Carlie H.; Cruz, Juan R.; Keller, Donald F.; O'Farrell, Clara

    2016-01-01

    A subsonic wind tunnel test was conducted to determine the drag and static aerodynamic coefficients, as well as to capture the dynamic motions of a new Supersonic Ringsail parachute developed by the Low Density Supersonic Decelerator Project. To provide a comparison against current Mars parachute technology, the Mars Science Laboratory's Disk-Gap-Band parachute was also included in the test. To account for the effect of fabric permeability, two fabrics ("low" and "standard" permeability) were used to fabricate each parachute canopy type, creating four combinations of canopy type and fabric material. A wide range of test conditions were covered during the test, spanning Mach numbers from 0.09 to 0.5, and static pressures from 103 to 2116 pounds per square inch (psf) (nominal values). The fabric permeability is shown to have a first-order effect on the aerodynamic coefficients and dynamic motions of the parachutes. For example, for a given parachute type and test condition, models fabricated from "low" permeability fabric always have a larger drag coefficient than models fabricated from "standard" permeability material. This paper describes the test setup and conditions, how the results were analyzed, and presents and discusses a sample of the results. The data collected during this test is being used to create and improve parachute aerodynamic databases for use in flight dynamics simulations for missions to Mars.

  8. Propfan test assessment testbed aircraft stability and control/performance 1/9-scale wind tunnel tests

    NASA Technical Reports Server (NTRS)

    Little, B. H., Jr.; Tomlin, K. H.; Aljabri, A. S.; Mason, C. A.

    1988-01-01

    One-ninth scale wind tunnel model tests of the Propfan Test Assessment (PTA) aircraft were performed in three different NASA facilities. Wing and propfan nacelle static pressures, model forces and moments, and flow field at the propfan plane were measured in these tests. Tests started in June 1985 and were completed in January 1987. These data were needed to assure PTA safety of flight, predict PTA performance, and validate analytical codes that will be used to predict flow fields in which the propfan will operate.

  9. An Application of CFD to Guide Forced Boundary-Layer Transition for Low-Speed Tests of a Hybrid Wing-Body Configuration

    NASA Technical Reports Server (NTRS)

    Luckring, James M.; Deere, Karen A.; Childs, Robert E.; Stremel, Paul M.; Long, Kurtis R.

    2016-01-01

    A hybrid transition trip-dot sizing and placement test technique was developed in support of recent experimental research on a hybrid wing-body configuration under study for the NASA Environmentally Responsible Aviation project. The approach combines traditional methods with Computational Fluid Dynamics. The application had three-dimensional boundary layers that were simulated with either fully turbulent or transitional flow models using established Reynolds-Averaged Navier-Stokes methods. Trip strip effectiveness was verified experimentally using infrared thermography during a low-speed wind tunnel test. Although the work was performed on one specific configuration, the process was based on fundamental flow physics and could be applicable to other configurations.

  10. Check Calibration of the NASA Glenn 10- by 10-Foot Supersonic Wind Tunnel (2014 Test Entry)

    NASA Technical Reports Server (NTRS)

    Johnson, Aaron; Pastor-Barsi, Christine; Arrington, E. Allen

    2016-01-01

    A check calibration of the 10- by 10-Foot Supersonic Wind Tunnel (SWT) was conducted in May/June 2014 using an array of five supersonic wedge probes to verify the 1999 Calibration. This check calibration was necessary following a control systems upgrade and an integrated systems test (IST). This check calibration was required to verify the tunnel flow quality was unchanged by the control systems upgrade prior to the next test customer beginning their test entry. The previous check calibration of the tunnel occurred in 2007, prior to the Mars Science Laboratory test program. Secondary objectives of this test entry included the validation of the new Cobra data acquisition system (DAS) against the current Escort DAS and the creation of statistical process control (SPC) charts through the collection of series of repeated test points at certain predetermined tunnel parameters. The SPC charts secondary objective was not completed due to schedule constraints. It is hoped that this effort will be readdressed and completed in the near future.

  11. New Test Section Installed in NASA Lewis' 1- by 1-Foot Supersonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Bauman, Steven W.

    1998-01-01

    NASA Lewis Research Center's 1- by 1-Foot Supersonic Wind Tunnel (1x1) is a critical facility that fulfills the needs of important national programs. This tunnel supports supersonic and hypersonic research test projects for NASA, for other Government agencies, and for industry, such as the High Speed Research (HSR) and Space Transportation Technologies (STT) programs. The 1x1, which is located in Lewis' Building 37, Cell 1NW, was built in 1954 and was upgraded to provide Mach 6.0 capability in 1989. Since 1954, only minor improvements had been made to the test section. To improve the 1x1's capabilities and meet the needs of these programs, Lewis recently redesigned and replaced the test section. The new test section has interchangeable window and wall inserts that allow easier and faster test configuration changes, thereby improving the adaptability and productivity of this highly utilized facility. In addition, both the wall and window areas are much larger. The larger walls provide more flexibility in how models are mounted and instrumented. The new window design vastly increases optical access to the research test hardware, which makes the use of advanced flow-visualization systems more effective.

  12. Propulsion and airframe aerodynamic interactions of supersonic V/STOL configurations. Volume 1: Wind tunnel test pressure data report

    NASA Technical Reports Server (NTRS)

    Zilz, D. E.; Devereaux, P. A.

    1985-01-01

    A wind tunnel model of a supersonic V/STOL fighter configuration has been tested to measure the aerodynamic interaction effects which can result from geometrically close-coupled propulsion system/airframe components. The approach was to configure the model to represent two different test techniques. One was a conventional test technique composed of two test modes. In the Flow-Through mode, absolute configuration aerodynamics are measured, including inlet/airframe interactions. In the Jet-Effects mode, incremental nozzle/airframe interactions are measured. The other test technique is a propulsion simulator approach, where a sub-scale, externally powered engine is mounted in the model. This allows proper measurement of inlet/airframe and nozzle/airframe interactions simultaneously. This is Volume 1 of 2: Wind Tunnel Test Pressure Data Report.

  13. Wind tunnel tests of a 5.3 m diameter yaw controlled turbine

    NASA Astrophysics Data System (ADS)

    Ulen, Eskil

    1993-09-01

    In order to meet the need for accurate wind turbine operating data from a variety of situations, a wind turbine was designed, built and tested. A description of the equipment and test results are given from a control point of view. The intention of the tests was to get a better understanding of the performance of a wind power station which from a system point of view is characterized by: two blades with fixed pitch angle, variable speed, yaw control of max rotor speed, hub with teeter, and induction generator. Some tests in stall region were also included. The intention was to get practical experiences by performing tests in a wind tunnel as realistically as possible. Of special interest are the following points: the load and aerodynamic torque of the blades with realistic values of rotor speed and yaw rate and angles; magnitude of teeter angle under yaw control; the required yaw motor torque; the coupling between yaw movements and tower lateral resonance ('fish tail' movement); disturbances, especially around yaw axis caused by blade rotation; and the performance of a hub with teeter in stall region. A small complete wind power station with a realistic yaw servo was used.

  14. Inlet Unstart Propulsion Integration Wind Tunnel Test Program Completed for High-Speed Civil Transport

    NASA Technical Reports Server (NTRS)

    Porro, A. Robert

    2000-01-01

    One of the propulsion system concepts to be considered for the High-Speed Civil Transport (HSCT) is an underwing, dual-propulsion, pod-per-wing installation. Adverse transient phenomena such as engine compressor stall and inlet unstart could severely degrade the performance of one of these propulsion pods. The subsequent loss of thrust and increased drag could cause aircraft stability and control problems that could lead to a catastrophic accident if countermeasures are not in place to anticipate and control these detrimental transient events. Aircraft system engineers must understand what happens during an engine compressor stall and inlet unstart so that they can design effective control systems to avoid and/or alleviate the effects of a propulsion pod engine compressor stall and inlet unstart. The objective of the Inlet Unstart Propulsion Airframe Integration test program was to assess the underwing flow field of a High-Speed Civil Transport propulsion system during an engine compressor stall and subsequent inlet unstart. Experimental research testing was conducted in the 10- by 10-Foot Supersonic Wind Tunnel at the NASA Glenn Research Center at Lewis Field. The representative propulsion pod consisted of a two-dimensional, bifurcated inlet mated to a live turbojet engine. The propulsion pod was mounted below a large flat plate that acted as a wing simulator. Because of the plate s long length (nominally 10-ft wide by 18-ft long), realistic boundary layers could form at the inlet cowl plane. Transient instrumentation was used to document the aerodynamic flow-field conditions during an unstart sequence. Acquiring these data was a significant technical challenge because a typical unstart sequence disrupts the local flow field for about only 50 msec. Flow surface information was acquired via static pressure taps installed in the wing simulator, and intrusive pressure probes were used to acquire flow-field information. These data were extensively analyzed to

  15. A smart model of a long-span suspended bridge for wind tunnel tests

    NASA Astrophysics Data System (ADS)

    Cinquemani, S.; Diana, G.; Fossati, L.; Ripamonti, F.

    2015-04-01

    Traditional aeroelastic models rely only on good mechanical design and accurate crafting in order to match the required structural properties. This paper proposes an active regulation of their structural parameters in order to improve accuracy and reliability of wind tunnel tests. Following the design process steps typical of a smart structure, a damping tuning technique allowing to control a specific set of vibration modes is developed and applied on the aeroelastic model of a long-span suspended bridge. Depending on the testing conditions, the structural damping value can be adjusted in a fast, precise and repeatable way in order to highlight the effects of the aerodynamic phenomena of interest. In particular, vortex-induced vibration are taken into consideration, and the response of a bridge section to vortex shedding is assessed. The active parameter regulation allows to widen the pattern of operating conditions in which the model can be tested. The paper discusses the choice of both sensors and actuators to be embedded in the structure and their positioning, as the control algorithm to obtain the desired damping. Experimental results are shown and results are discussed to evaluate the performance of the smart structure in wind dunnel tests.

  16. Testing of the Trim Tab Parametric Model in NASA Langley's Unitary Plan Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Murphy, Kelly J.; Watkins, Anthony N.; Korzun, Ashley M.; Edquist, Karl T.

    2013-01-01

    In support of NASA's Entry, Descent, and Landing technology development efforts, testing of Langley's Trim Tab Parametric Models was conducted in Test Section 2 of NASA Langley's Unitary Plan Wind Tunnel. The objectives of these tests were to generate quantitative aerodynamic data and qualitative surface pressure data for experimental and computational validation and aerodynamic database development. Six component force-and-moment data were measured on 38 unique, blunt body trim tab configurations at Mach numbers of 2.5, 3.5, and 4.5, angles of attack from -4deg to +20deg, and angles of sideslip from 0deg to +8deg. Configuration parameters investigated in this study were forebody shape, tab area, tab cant angle, and tab aspect ratio. Pressure Sensitive Paint was used to provide qualitative surface pressure mapping for a subset of these flow and configuration variables. Over the range of parameters tested, the effects of varying tab area and tab cant angle were found to be much more significant than varying tab aspect ratio relative to key aerodynamic performance requirements. Qualitative surface pressure data supported the integrated aerodynamic data and provided information to aid in future analyses of localized phenomena for trim tab configurations.

  17. Testing of Flexible Ballutes in Hypersonic Wind Tunnels for Planetary Aerocapture

    NASA Technical Reports Server (NTRS)

    Buck, Gregory M.

    2007-01-01

    Studies were conducted for the In-Space Propulsion (ISP) Ultralightweight Ballute Technology Development Program to increase the technical readiness level of inflatable decelerator systems for planetary aerocapture. The present experimental study was conducted to develop the capability for testing lightweight, flexible materials in hypersonic facilities. The primary objectives were to evaluate advanced polymer film materials in a high-temperature, high-speed flow environment and provide experimental data for comparisons with fluid-structure interaction modeling tools. Experimental testing was conducted in the Langley Aerothermodynamics Laboratory 20-Inch Hypersonic CF4 and 31-Inch Mach 10 Air blowdown wind tunnels. Quantitative flexure measurements were made for 60 degree half angle afterbody-attached ballutes, in which polyimide films (1-mil and 3- mil thick) were clamped between a 1/2-inch diameter disk and a base ring (4-inch and 6-inch diameters). Deflection measurements were made using a parallel light silhouette of the film surface through an existing schlieren optical system. The purpose of this paper is to discuss these results as well as free-flying testing techniques being developed for both an afterbody-attached and trailing toroidal ballute configuration to determine dynamic fluid-structural stability. Methods for measuring polymer film temperature were also explored using both temperature sensitive paints (for up to 370 C) and laser-etched thin-film gages.

  18. Testing of Flexible Ballutes in Hypersonic Wind Tunnels for Planetary Aerocapture

    NASA Technical Reports Server (NTRS)

    Buck, Gregory M.

    2006-01-01

    Studies were conducted for the In-Space Propulsion (ISP) Ultralightweight Ballute Technology Development Program to increase the technical readiness level of inflatable decelerator systems for planetary aerocapture. The present experimental study was conducted to develop the capability for testing lightweight, flexible materials in hypersonic facilities. The primary objectives were to evaluate advanced polymer film materials in a high-temperature, high-speed flow environment and provide experimental data for comparisons with fluid-structure interaction modeling tools. Experimental testing was conducted in the Langley Aerothermodynamics Laboratory 20-Inch Hypersonic CF4 and 31-Inch Mach 10 Air blowdown wind tunnels. Quantitative flexure measurements were made for 60 degree half angle afterbody-attached ballutes, in which polyimide films (1-mil and 3-mil thick) were clamped between a 1/2-inch diameter disk and a base ring (4-inch and 6-inch diameters). Deflection measurements were made using a parallel light silhouette of the film surface through an existing schlieren optical system. The purpose of this paper is to discuss these results as well as free-flying testing techniques being developed for both an afterbody-attached and trailing toroidal ballute configuration to determine dynamic fluid-structural stability. Methods for measuring polymer film temperature were also explored using both temperature sensitive paints (for up to 370 C) and laser-etched thin-film gages.

  19. Analysis of the free-tip rotor wind-tunnel test results

    NASA Technical Reports Server (NTRS)

    Stroub, Robert H.

    1985-01-01

    The results from a wind tunnel test of a small scale free-tip rotor are analyzed. The free-tip rotor has blade tips that are free to weathervane into the tip's relative wind, thus producing a more uniform lift around the azimuth. The free-tip assembly, which includes the controller, functioned flawlessly throughout the test. In a test of the free-tip's response after passing through a vertical air jet, the tip pitched freely and in a controlled manner. Analysis of the tip's response characteristics showed the free-tip system's damped natural frequency to be 5.2 per rev. Tip pitch angle responses to the local airstream are presented for an advance-ratio range of 0.1 to 0.397 and for a solidity weighted rotor lift coefficient range of 0.038 to 0.092. Harmonic analysis of the responses showed a dominance by the first harmonic. As a result of the tip being free, forward flight power requirements were reduced by 8% or more. More power reduction was recorded for high thrust conditions. In addition to the power reduction, flatwise blade bending moments were reduced by as much as 30% at the inboard blade stations.

  20. Wind-Tunnel Tests on Airfoil Boundary Layer Control Using a Backward-Opening Slot

    NASA Technical Reports Server (NTRS)

    Bamber, Millard J

    1932-01-01

    This report presents the results of an investigation to determine the effect of boundary layer control on the lift and drag of an airfoil. Boundary layer control was accomplished by means of a backward-opening slot in the upper surface of the hollow airfoil. Air was caused to flow through this slot by a pressure which was maintained inside the airfoil by a blower. Various slot locations, slot openings, and wing pressures were used. The tests were conducted in the 5-foot atmospheric wind tunnel of the Langley Memorial Aeronautical Laboratory. Under the test conditions, the maximum lift coefficient was increased about 96 per cent for one slot arrangement, and the minimum drag coefficient was decreased about 27 per cent for another, both being compared with the results obtained with the unslotted airfoil. It is believed from this investigation that the above effects may be increased by the use of larger slot openings, better slot locations, multiple slots, improved airfoil profiles, and trailing edge flaps.

  1. Wind tunnel performance results of swirl recovery vanes as tested with an advanced high speed propeller

    NASA Technical Reports Server (NTRS)

    Gazzaniga, John A.; Rose, Gayle E.

    1992-01-01

    Tests of swirl recovery vanes designed for use in conjunction with advanced high speed propellers were carried out at the NASA Lewis Research Center. The eight bladed 62.23 cm vanes were tested with a 62.23 cm SR = 7A high speed propeller in the NASA Lewis 2.44 x 1.83 m Supersonic Wind Tunnel for a Mach number range of 0.60 to 0.80. At the design operating condition for cruise of Mach 0.80 at an advance ratio of 3.26, the vane contribution to the total efficiency approached 2 percent. At lower off-design Mach numbers, the vane efficiency is even higher, approaching 4.5 percent for the Mach 0.60 condition. Use of the swirl recovery vanes essentially shifts the peak of the high speed propeller efficiency to a higher operating speed. This allows a greater degree of freedom in the selection of rpm over a wider operating range. Another unique result of the swirl recovery vane configuration is their essentially constant torque split between the propeller and the swirl vanes over a wide range of operating conditions for the design vane angle.

  2. Boeing Smart Rotor Full-scale Wind Tunnel Test Data Report

    NASA Technical Reports Server (NTRS)

    Kottapalli, Sesi; Hagerty, Brandon; Salazar, Denise

    2016-01-01

    A full-scale helicopter smart material actuated rotor technology (SMART) rotor test was conducted in the USAF National Full-Scale Aerodynamics Complex 40- by 80-Foot Wind Tunnel at NASA Ames. The SMART rotor system is a five-bladed MD 902 bearingless rotor with active trailing-edge flaps. The flaps are actuated using piezoelectric actuators. Rotor performance, structural loads, and acoustic data were obtained over a wide range of rotor shaft angles of attack, thrust, and airspeeds. The primary test objective was to acquire unique validation data for the high-performance computing analyses developed under the Defense Advanced Research Project Agency (DARPA) Helicopter Quieting Program (HQP). Other research objectives included quantifying the ability of the on-blade flaps to achieve vibration reduction, rotor smoothing, and performance improvements. This data set of rotor performance and structural loads can be used for analytical and experimental comparison studies with other full-scale rotor systems and for analytical validation of computer simulation models. The purpose of this final data report is to document a comprehensive, highquality data set that includes only data points where the flap was actively controlled and each of the five flaps behaved in a similar manner.

  3. Phase 1 wind tunnel tests of the J-97 powered, external augmentor V/STOL model

    NASA Technical Reports Server (NTRS)

    Garland, D. B.

    1980-01-01

    Test results are presented for a large scale, external augmentor V/STOL model in a 40 ft by 80 ft wind tunnel. The model was powered by a GE J97 engine and featured longitudinal ejectors alongside and external to the fuselage together with an augmentor flap on the low aspect ratio, double-delta wing. A static thrust augmentation ratio of 1.60 was measured for the fuselage augmentor at a nozzle pressure ratio of 3.0 and a nozzle exhaust gas temperature of 700 C. At forward speed the model showed a strong positive lift interference due to the augmentor flap, and a marked absence of negative lift interference due to the fuselage augmentor jet system. The nose-up moment of the fuselage augmentor inlet flow was approximately cancelled by a 60 deg deflection of the augmentor flap. An assessment of the thrust and drag components to allow the prediction of transition performance of aircraft designs based on the present conceptual model was made. Lateral tests showed strong but well ordered effects of power.

  4. Flatback airfoil wind tunnel experiment.

    SciTech Connect

    Mayda, Edward A.; van Dam, C.P.; Chao, David D.; Berg, Dale E.

    2008-04-01

    A computational fluid dynamics study of thick wind turbine section shapes in the test section of the UC Davis wind tunnel at a chord Reynolds number of one million is presented. The goals of this study are to validate standard wind tunnel wall corrections for high solid blockage conditions and to reaffirm the favorable effect of a blunt trailing edge or flatback on the performance characteristics of a representative thick airfoil shape prior to building the wind tunnel models and conducting the experiment. The numerical simulations prove the standard wind tunnel corrections to be largely valid for the proposed test of 40% maximum thickness to chord ratio airfoils at a solid blockage ratio of 10%. Comparison of the computed lift characteristics of a sharp trailing edge baseline airfoil and derived flatback airfoils reaffirms the earlier observed trend of reduced sensitivity to surface contamination with increasing trailing edge thickness.

  5. Numerical Study of the High-Speed Leg of a Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Nayani, Sudheer; Sellers, William L., III; Brynildsen, Scott E.; Everhart, Joel L.

    2015-01-01

    The paper describes the numerical study of the high-speed leg of the NASA Langley 14 by 22-foot Low Speed Wind Tunnel. The high-speed leg consists of the Settling Chamber, Contraction, Test Section, and First Diffuser. Results are shown comparing two different exit boundary conditions and two different methods of determining the surface geometry.

  6. Construction, wind tunnel testing and data analysis for a 1/5 scale ultra-light wing model

    NASA Technical Reports Server (NTRS)

    James, Michael D.; Smith, Howard W.

    1993-01-01

    This report documents the construction, wind tunnel testing, and data analysis of a 1/5 scale ultra-light wing section. Wind tunnel testing provided accurate and meaningful lift, drag, and pitching moment data. This data was processed and graphically presented as follows: C(sub L) vs. gamma; C(sub D) vs. gamma; C(sub M) vs. gamma; and C(sub L) vs. C(sub D). The wing fabric flexure was found to be significant and its possible effects on aerodynamic data was discussed. The fabric flexure is directly related to wing angle of attack and airspeed. Different wing section shapes created by fabric flexure are presented with explanations of the types of pressures that act upon the wing surface. This report provides conclusive aerodynamic data for ultra-light wings.

  7. A Hydrogen Peroxide Hot-Jet Simulator for Wind-Tunnel Tests of Turbojet-Exit Models

    NASA Technical Reports Server (NTRS)

    Runckel, Jack F.; Swihart, John M.

    1959-01-01

    A turbojet-engine-exhaust simulator which utilizes a hydrogen peroxide gas generator has been developed for powered-model testing in wind tunnels with air exchange. Catalytic decomposition of concentrated hydrogen peroxide provides a convenient and easily controlled method of providing a hot jet with characteristics that correspond closely to the jet of a gas turbine engine. The problems associated with simulation of jet exhausts in a transonic wind tunnel which led to the selection of a liquid monopropellant are discussed. The operation of the jet simulator consisting of a thrust balance, gas generator, exit nozzle, and auxiliary control system is described. Static-test data obtained with convergent nozzles are presented and shown to be in good agreement with ideal calculated values.

  8. Space shuttle plume simulation application. Results and math model. [Ames unitary plan wind tunnel test

    NASA Technical Reports Server (NTRS)

    Boyle, W.; Conine, B.

    1978-01-01

    Pressure and gauge wind tunnel data from a transonic test of a 0.02 scale model of the space shuttle launch vehicle was analyzed to define the aerodynamic influence of the main propulsion system and solid rocket booster plumes during the transonic portion of ascent flight. Air was used as a simulant gas to develop the model exhaust plumes. A math model of the plume induced aerodynamic characteristics was developed for a range of Mach numbers to match the forebody aerodynamic math model. The base aerodynamic characteristics are presented in terms of forces and moments versus attitude. Total vehicle base and forebody aerodynamic characteristics are presented in terms of aerodynamic coefficients for Mach number from 0.6 to 1.4 Element and component base and forebody aerodynamic characteristics are presented for Mach numbers of 0.6, 1.05, 1.1, 1.25 and 1.4. The forebody data is available at Mach 1.55. Tolerances for all plume induced aerodynamic characteristics are developed in terms of a math model.

  9. Wind tunnel tests for wind pressure distribution on gable roof buildings.

    PubMed

    Jing, Xiao-kun; Li, Yuan-qi

    2013-01-01

    Gable roof buildings are widely used in industrial buildings. Based on wind tunnel tests with rigid models, wind pressure distributions on gable roof buildings with different aspect ratios were measured simultaneously. Some characteristics of the measured wind pressure field on the surfaces of the models were analyzed, including mean wind pressure, fluctuating wind pressure, peak negative wind pressure, and characteristics of proper orthogonal decomposition results of the measured wind pressure field. The results show that extremely high local suctions often occur in the leading edges of longitudinal wall and windward roof, roof corner, and roof ridge which are the severe damaged locations under strong wind. The aspect ratio of building has a certain effect on the mean wind pressure coefficients, and the effect relates to wind attack angle. Compared with experimental results, the region division of roof corner and roof ridge from AIJ2004 is more reasonable than those from CECS102:2002 and MBMA2006.The contributions of the first several eigenvectors to the overall wind pressure distributions become much bigger. The investigation can offer some basic understanding for estimating wind load distribution on gable roof buildings and facilitate wind-resistant design of cladding components and their connections considering wind load path.

  10. Wind Tunnel Tests for Wind Pressure Distribution on Gable Roof Buildings

    PubMed Central

    2013-01-01

    Gable roof buildings are widely used in industrial buildings. Based on wind tunnel tests with rigid models, wind pressure distributions on gable roof buildings with different aspect ratios were measured simultaneously. Some characteristics of the measured wind pressure field on the surfaces of the models were analyzed, including mean wind pressure, fluctuating wind pressure, peak negative wind pressure, and characteristics of proper orthogonal decomposition results of the measured wind pressure field. The results show that extremely high local suctions often occur in the leading edges of longitudinal wall and windward roof, roof corner, and roof ridge which are the severe damaged locations under strong wind. The aspect ratio of building has a certain effect on the mean wind pressure coefficients, and the effect relates to wind attack angle. Compared with experimental results, the region division of roof corner and roof ridge from AIJ2004 is more reasonable than those from CECS102:2002 and MBMA2006.The contributions of the first several eigenvectors to the overall wind pressure distributions become much bigger. The investigation can offer some basic understanding for estimating wind load distribution on gable roof buildings and facilitate wind-resistant design of cladding components and their connections considering wind load path. PMID:24082851

  11. Vibratory hub load data reduction and analysis from the reverse velocity rotor wind tunnel test, phase 2B

    NASA Technical Reports Server (NTRS)

    Taylor, R. B.

    1976-01-01

    The vibratory hub loads data analysis from the reverse velocity rotor wind tunnel test is reported. Vibratory loads were obtained from the rotating hub balance and also by synthesis of generalized coordinates from the blade flap bending moments. Load trends were defined as a function of speed, rotor thrust and 2 per rev cyclic from each of the data methods. These trends were compared to determine the degree of agreement between each method and provide substantiation for the generalized coordinate approach.

  12. Wind Tunnel Testing of a 120th Scale Large Civil Tilt-Rotor Model in Airplane and Helicopter Modes

    NASA Technical Reports Server (NTRS)

    Theodore, Colin R.; Willink, Gina C.; Russell, Carl R.; Amy, Alexander R.; Pete, Ashley E.

    2014-01-01

    In April 2012 and October 2013, NASA and the U.S. Army jointly conducted a wind tunnel test program examining two notional large tilt rotor designs: NASA's Large Civil Tilt Rotor and the Army's High Efficiency Tilt Rotor. The approximately 6%-scale airframe models (unpowered) were tested without rotors in the U.S. Army 7- by 10-foot wind tunnel at NASA Ames Research Center. Measurements of all six forces and moments acting on the airframe were taken using the wind tunnel scale system. In addition to force and moment measurements, flow visualization using tufts, infrared thermography and oil flow were used to identify flow trajectories, boundary layer transition and areas of flow separation. The purpose of this test was to collect data for the validation of computational fluid dynamics tools, for the development of flight dynamics simulation models, and to validate performance predictions made during conceptual design. This paper focuses on the results for the Large Civil Tilt Rotor model in an airplane mode configuration up to 200 knots of wind tunnel speed. Results are presented with the full airframe model with various wing tip and nacelle configurations, and for a wing-only case also with various wing tip and nacelle configurations. Key results show that the addition of a wing extension outboard of the nacelles produces a significant increase in the lift-to-drag ratio, and interestingly decreases the drag compared to the case where the wing extension is not present. The drag decrease is likely due to complex aerodynamic interactions between the nacelle and wing extension that results in a significant drag benefit.

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

    NASA Technical Reports Server (NTRS)

    Brooks, J. D.

    1976-01-01

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

  14. Comparison of nozzle and afterbody surface pressures from wind tunnel and flight test of the YF-17 aircraft

    NASA Technical Reports Server (NTRS)

    Lucas, E. J.; Fanning, A. E.; Steers, L. I.

    1978-01-01

    Results are reported from the initial phase of an effort to provide an adequate technical capability to accurately predict the full scale, flight vehicle, nozzle-afterbody performance of future aircraft based on partial scale, wind tunnel testing. The primary emphasis of this initial effort is to assess the current capability and identify the cause of limitations on this capability. A direct comparison of surface pressure data is made between the results from an 0.1-scale model wind tunnel investigation and a full-scale flight test program to evaluate the current subscale testing techniques. These data were acquired at Mach numbers 0.6, 0.8, 0.9, 1.2, and 1.5 on four nozzle configurations at various vehicle pitch attitudes. Support system interference increments were also documented during the wind tunnel investigation. In general, the results presented indicate a good agreement in trend and level of the surface pressures when corrective increments are applied for known effects and surface differences between the two articles under investigation.

  15. Entry, Descent, and Landing with Propulsive Deceleration: Supersonic Retropropulsion Wind Tunnel Testing and Shock Phenomena

    NASA Technical Reports Server (NTRS)

    Palaszewski, Bryan

    2013-01-01

    The future exploration of the Solar System will require innovations in transportation and the use of entry, descent, and landing (EDL) systems at many planetary landing sites. The cost of space missions has always been prohibitive, and using the natural planetary and planet's moon atmospheres for entry, and descent can reduce the cost, mass, and complexity of these missions. This paper will describe some of the EDL ideas for planetary entry and survey the overall technologies for EDL that may be attractive for future Solar System missions. Future EDL systems may include an inflatable decelerator for the initial atmospheric entry and an additional supersonic retro-propulsion (SRP) rocket system for the final soft landing. A three engine retro-propulsion configuration with a 2.5 inch diameter sphere-cone aeroshell model was tested in the NASA Glenn 1x1 Supersonic Wind Tunnel (SWT). The testing was conducted to identify potential blockage issues in the tunnel, and visualize the rocket flow and shock interactions during supersonic and hypersonic entry conditions. Earlier experimental testing of a 70 degree Viking-like (sphere-cone) aeroshell was conducted as a baseline for testing of a supersonic retro-propulsion system. This baseline testing defined the flow field around the aeroshell and from this comparative baseline data, retro-propulsion options will be assessed. Images and analyses from the SWT testing with 300- and 500-psia rocket engine chamber pressures are presented here. In addition, special topics of electromagnetic interference with retro-propulsion induced shock waves and retro-propulsion for Earth launched booster recovery are also addressed.

  16. Entry, Descent, and Landing With Propulsive Deceleration: Supersonic Retropropulsion Wind Tunnel Testing

    NASA Technical Reports Server (NTRS)

    Palaszewski, Bryan

    2012-01-01

    The future exploration of the Solar System will require innovations in transportation and the use of entry, descent, and landing (EDL) systems at many planetary landing sites. The cost of space missions has always been prohibitive, and using the natural planetary and planet s moons atmosphere for entry, descent, and landing can reduce the cost, mass, and complexity of these missions. This paper will describe some of the EDL ideas for planetary entry and survey the overall technologies for EDL that may be attractive for future Solar System missions. Future EDL systems may include an inflatable decelerator for the initial atmospheric entry and an additional supersonic retro-propulsion (SRP) rocket system for the final soft landing. As part of those efforts, NASA began to conduct experiments to gather the experimental data to make informed decisions on the "best" EDL options. A model of a three engine retro-propulsion configuration with a 2.5 in. diameter sphere-cone aeroshell model was tested in the NASA Glenn 1- by 1-Foot Supersonic Wind Tunnel (SWT). The testing was conducted to identify potential blockage issues in the tunnel, and visualize the rocket flow and shock interactions during supersonic and hypersonic entry conditions. Earlier experimental testing of a 70 Viking-like (sphere-cone) aeroshell was conducted as a baseline for testing of a supersonic retro-propulsion system. This baseline testing defined the flow field around the aeroshell and from this comparative baseline data, retro-propulsion options will be assessed. Images and analyses from the SWT testing with 300- and 500-psia rocket engine chamber pressures are presented here. The rocket engine flow was simulated with a non-combusting flow of air.

  17. Entry, Descent, and Landing with Propulsive Deceleration: Supersonic Retropropulsion Wind Tunnel Testing and Shock Phenomena

    NASA Technical Reports Server (NTRS)

    Palaszewski, Bryan

    2014-01-01

    The future exploration of the Solar System will require innovations in transportation and the use of entry, descent, and landing (EDL) systems at many planetary landing sites. The cost of space missions has always been prohibitive, and using the natural planetary and planet's moon atmospheres for entry, and descent can reduce the cost, mass, and complexity of these missions. This paper will describe some of the EDL ideas for planetary entry and survey the overall technologies for EDL that may be attractive for future Solar System missions. Future EDL systems may include an inflatable decelerator for the initial atmospheric entry and an additional supersonic retropropulsion (SRP) rocket system for the final soft landing. A three engine retropropulsion configuration with a 2.5 in. diameter sphere-cone aeroshell model was tested in the NASA Glenn Research Center's 1- by 1-ft (1×1) Supersonic Wind Tunnel (SWT). The testing was conducted to identify potential blockage issues in the tunnel, and visualize the rocket flow and shock interactions during supersonic and hypersonic entry conditions. Earlier experimental testing of a 70deg Viking-like (sphere-cone) aeroshell was conducted as a baseline for testing of a SRP system. This baseline testing defined the flow field around the aeroshell and from this comparative baseline data, retropropulsion options will be assessed. Images and analyses from the SWT testing with 300- and 500-psia rocket engine chamber pressures are presented here. In addition, special topics of electromagnetic interference with retropropulsion induced shock waves and retropropulsion for Earth launched booster recovery are also addressed.

  18. Wind-tunnel blockage and actuation systems test of a two-dimensional scramjet inlet unstart model at Mach 6

    NASA Technical Reports Server (NTRS)

    Holland, Scott D.

    1994-01-01

    The present study examines the wind-tunnel blockage and actuation systems effectiveness in starting and forcibly unstarting a two-dimensional scramjet inlet in the NASA Langley 20-Inch Mach 6 Tunnel. The intent of the overall test program is to study (both experimentally and computationally) the dynamics of the inlet unstart; however, prior to the design and fabrication of an expensive, instrumented wind-tunnel model, it was deemed necessary first to examine potential wind-tunnel blockage issues related to model sizing and to examine the adequacy of the actuation systems in accomplishing the start and unstart. The model is equipped with both a moveable cowl and aft plug. Windows in the inlet sidewalls allow limited optical access to the internal shock structure; schlieren video was used to identify inlet start and unstart. A chronology of each actuation sequence is provided in tabular form along with still frames from the schlieren video. A pitot probe monitored the freestream conditions throughout the start/unstart process to determine if there was a blockage effect due to the model start or unstart. Because the purpose of this report is to make the phase I (blockage and actuation systems) data rapidly available to the community, the data is presented largely without analysis of the internal shock interactions or the unstart process. This series of tests indicated that the model was appropriately sized for this facility and identified operability limits required first to allow the inlet to start and second to force the unstart.

  19. The NACA High-Speed Wind Tunnel and Tests of Six Propeller Sections

    NASA Technical Reports Server (NTRS)

    Stack, John

    1934-01-01

    This report gives a description of the high-speed wind tunnel of the National Advisory Committee for Aeronautics. The operation of the tunnel is also described and the method of presenting the data is given. An account of an investigation of the aerodynamic properties of six propeller sections is included.

  20. Space shuttle phase B wind tunnel model and test information. Volume 3: Launch configuration

    NASA Technical Reports Server (NTRS)

    Glynn, J. L.; Poucher, D. E.

    1988-01-01

    Archived wind tunnel data are available for flyback booster or other alternative recoverable configurations as well as reusable orbiters studied during initial development (Phase B) of the Space Shuttle. Considerable wind tunnel data was acquired by the competing contractors and the NASA Centers for an extensive variety of configurations with an array of wing and body planforms. All contractor and NASA wind tunnel data acquired in the Phase B development have been compiled into a data base and are available for application to current winged flyback or recoverable booster aerodynamic studies. The Space Shuttle Phase B Wind Tunnel Database is structured by vehicle component and configuration type. Basic components include booster, orbiter and launch vehicle. Booster configuration types include straight and delta wings, canard, cylindrical, retroglide and twin body. Orbital configuration types include straight and delta wings, lifting body, drop tanks and double delta wings. This is Volume 3 (Part 2) of the report -- Launch Configuration -- which includes booster and orbiter components in various stacked and tandem combinations.

  1. Wind tunnel test of Teledyne Geotech model 1564B cup anemometer

    SciTech Connect

    Parker, M.J.; Addis, R.P.

    1991-04-04

    The Department of Energy (DOE) Environment, Safety and Health Compliance Assessment (Tiger Team) of the Savannah River Site (SRS) questioned the method by which wind speed sensors (cup anemometers) are calibrated by the Environmental Technology Section (ETS). The Tiger Team member was concerned that calibration data was generated by running the wind tunnel to only 26 miles per hour (mph) when speeds exceeding 50 mph are readily obtainable. A wind tunnel experiment was conducted and confirmed the validity of the practice. Wind speeds common to SRS (6 mph) were predicted more accurately by 0--25 mph regression equations than 0--50 mph regression equations. Higher wind speeds were slightly overpredicted by the 0--25 mph regression equations when compared to 0--50 mph regression equations. However, the greater benefit of more accurate lower wind speed predictions accuracy outweight the benefit of slightly better high (extreme) wind speed predictions. Therefore, it is concluded that 0--25 mph regression equations should continue to be utilized by ETS at SRS. During the Department of Energy Tiger Team audit, concerns were raised about the calibration of SRS cup anemometers. Wind speed is measured by ETS with Teledyne Geotech model 1564B cup anemometers, which are calibrated in the ETS wind tunnel. Linear regression lines are fitted to data points of tunnel speed versus anemometer output voltages up to 25 mph. The regression coefficients are then implemented into the data acquisition computer software when an instrument is installed in the field. The concern raised was that since the wind tunnel at SRS is able to generate a maximum wind speed higher than 25 mph, errors may be introduced in not using the full range of the wind tunnel.

  2. Parametric Inlet Tested in Glenn's 10- by 10-Foot Supersonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Slater, John W.; Davis, David O.; Solano, Paul A.

    2005-01-01

    The Parametric Inlet is an innovative concept for the inlet of a gas-turbine propulsion system for supersonic aircraft. The concept approaches the performance of past inlet concepts, but with less mechanical complexity, lower weight, and greater aerodynamic stability and safety. Potential applications include supersonic cruise aircraft and missiles. The Parametric Inlet uses tailored surfaces to turn the incoming supersonic flow inward toward an axis of symmetry. The terminal shock spans the opening of the subsonic diffuser leading to the engine. The external cowl area is smaller, which reduces cowl drag. The use of only external supersonic compression avoids inlet unstart--an unsafe shock instability present in previous inlet designs that use internal supersonic compression. This eliminates the need for complex mechanical systems to control unstart, which reduces weight. The conceptual design was conceived by TechLand Research, Inc. (North Olmsted, OH), which received funding through NASA s Small-Business Innovation Research program. The Boeing Company (Seattle, WA) also participated in the conceptual design. The NASA Glenn Research Center became involved starting with the preliminary design of a model for testing in Glenn s 10- by 10-Foot Supersonic Wind Tunnel (10 10 SWT). The inlet was sized for a speed of Mach 2.35 while matching requirements of an existing cold pipe used in previous inlet tests. The parametric aspects of the model included interchangeable components for different cowl lip, throat slot, and sidewall leading-edge shapes and different vortex generator configurations. Glenn researchers used computational fluid dynamics (CFD) tools for three-dimensional, turbulent flow analysis to further refine the aerodynamic design.

  3. Two-Dimensional Scramjet Inlet Unstart Model: Wind-Tunnel Blockage and Actuation Systems Test

    NASA Technical Reports Server (NTRS)

    Holland, Scott D.

    1994-01-01

    This supplement to NASA TM 109152 shows the Schlieren video (10 min. 52 sec., color, Beta and VHS) of the external flow field and a portion of the internal flow field of a two-dimensional scramjet inlet model in the NASA Langley 20-Inch Mach 6 Tunnel. The intent of the overall test program is to study (both experimentally and computationally) the dynamics of the inlet unstart; this (phase I) effort examines potential wind-tunnel blockage issues related to model sizing and the adequacy of the actuation systems in accomplishing the start and unstart. The model is equipped with both a moveable cowl and aft plug. Windows in the inlet sidewalls allow limited optical access to the internal shock structure. In the video, flow is from right to left, and the inlet is oriented inverted with respect to flight, i.e., with the cowl on top. The plug motion is obvious because the plug is visible in the aft window. The cowl motion, however, is not as obvious because the cowl is hidden from view by the inlet sidewall. The end of the cowl actuator arm, however, becomes visible above the inlet sidewalls between the windows when the cowl is up (see figure 1b of the primary document). The model is injected into the tunnel and observed though several actuation sequences with two plug configurations over a range of unit freestream Reynolds number at a nominal freestream Mach number of 6. The framing rate and shutter speed of the camera were too slow to fully capture the dynamics of the unstart but did prove sufficient to identify inlet start and unstart. This series of tests indicated that the model was appropriately sized for this facility and identified operability limits required first to allow the inlet to start and second to force the unstart.

  4. Aerodynamic results of wind tunnel tests on a 0.010-scale model (32-QTS) space shuttle integrated vehicle in the AEDC VKF-40-inch supersonic wind tunnel (IA61)

    NASA Technical Reports Server (NTRS)

    Daileda, J. J.

    1976-01-01

    Plotted and tabulated aerodynamic coefficient data from a wind tunnel test of the integrated space shuttle vehicle are presented. The primary test objective was to determine proximity force and moment data for the orbiter/external tank and solid rocket booster (SRB) with and without separation rockets firing for both single and dual booster runs. Data were obtained at three points (t = 0, 1.25, and 2.0 seconds) on the nominal SRB separation trajectory.

  5. Hover test of a full-scale hingeless helicopter rotor: Aeroelastic stability, performance and loads data. [wind tunnel tests

    NASA Technical Reports Server (NTRS)

    Peterson, R. L.; Warmbrodt, W.

    1984-01-01

    A hover test of a full-scale, hingeless rotor system was conducted in the NASA Ames 40- by 80-foot wind tunnel. The rotor was tested on the Ames rotor test apparatus. Rotor aeroelastic stability, performance, and loads at various rotational speeds and thrust coefficients were investigated. The primary objective was to determine the inplane stability characteristics of the rotor system. Rotor inplane damping data were obtained for operation between 350 and 425 rpm (design speed), and for thurst coefficients between 0.0 and 0.12. The rotor was stable for all conditions tested. At constant rotor rotational speed, a minimum inplane dampling level was obtained at a thrust coefficient approximately = 0.02. At constant rotor lift, a minimum in rotor inplane damping was measured at 400 rpm.

  6. Shake test of rotor test apparatus with balance dampers in the 40 by 80 foot wind tunnel

    NASA Technical Reports Server (NTRS)

    Johnson, W.; Biggers, J. G.

    1975-01-01

    A shake test was conducted to determine the dynamic characteristics of a rotor test apparatus on two strut systems with balance dampers in the Ames 40- by 80-ft wind tunnel. The rotor-off hub transfer function (acceleration per unit force as a function of frequency) was measured in the longitudinal and lateral directions, using a combination of broadband and discrete frequency excitation techniques. The dynamic data are summarized for the configurations tested, giving the following properties for each mode identified: the natural frequency, the hub response at resonance, the fixed system damping, the damping ratio, and the modal mass. The complete transfer functions are presented, and the detailed test results are included as an appendix.

  7. NASA Environmentally Responsible Aviation Hybrid Wing Body Flow-Through Nacelle Wind Tunnel CFD

    NASA Technical Reports Server (NTRS)

    Schuh, Michael J.; Garcia, Joseph A.; Carter, Melissa B.; Deere, Karen A.; Tompkins, Daniel M.; Stremel, Paul M.

    2016-01-01

    Wind tunnel tests of a 5.75 scale model of the Boeing Hybrid Wing Body (HWB) configuration were conducted in the NASA Langley Research Center (LaRC) 14x22 and NASA Ames Research Center (ARC) 40x80 low speed wind tunnels as part of the NASA Environmentally Responsible Aviation (ERA) Project. Computational fluid dynamics (CFD) simulations of the flow-through nacelle (FTN) configuration of this model were performed before and after the testing. This paper presents a summary of the experimental and CFD results for the model in the cruise and landing configurations.

  8. NASA Environmentally Responsible Aviation Hybrid Wing Body Flow-Through Nacelle Wind Tunnel CFD

    NASA Technical Reports Server (NTRS)

    Schuh, Michael J.; Garcia, Jospeh A.; Carter, Melissa B.; Deere, Karen A.; Stremel, Paul M.; Tompkins, Daniel M.

    2016-01-01

    Wind tunnel tests of a 5.75% scale model of the Boeing Hybrid Wing Body (HWB) configuration were conducted in the NASA Langley Research Center (LaRC) 14'x22' and NASA Ames Research Center (ARC) 40'x80' low speed wind tunnels as part of the NASA Environmentally Responsible Aviation (ERA) Project. Computational fluid dynamics (CFD) simulations of the flow-through nacelle (FTN) configuration of this model were performed before and after the testing. This paper presents a summary of the experimental and CFD results for the model in the cruise and landing configurations.

  9. Application of infrared thermography to the interpretation of tests in an icing wind tunnel

    NASA Astrophysics Data System (ADS)

    Henry, R.; Guffond, D.

    Wall temperature measurements were obtained in an icing wind tunnel using infrared thermography in order to validate models for the simulation of electrothermal deicers intended for helicopter blades. The measurement procedure involves adjusting the camera for the temperature range considered, determining the directional emissivity of the profile surface and of the ice, and using Mie theory to determine the atmospheric transmission factor. The present results demonstrate the importance of taking the phase change of the ice into account in deicer modeling.

  10. An experimental investigation of a free-tip rotor configuration in a forward flight wind-tunnel test

    NASA Technical Reports Server (NTRS)

    Stroub, R. H.

    1983-01-01

    Results from an experimental evaluation of a free-tip rotor are presented. The evaluation included whirl tests and wind tunnel tests up to advance ratios of 0.4. The free tip extended over the outer 5% of the rotor blade and included a passive mechanical controller whose output characteristics were varied. The controller configuration combined with the free tip aerodynamics resulted in higher power requirements, because the tip's pitch angle was 5 to 10 degrees greater than that of the inboard portion of the blade, and its pitching motion capability was considered to be inhibited by frictional forces. Recommendations are included for design features for a follow-on test.

  11. Tests of the Northrop XSSM-A-3 Missile in the Ames 40- by 80-Foot Wind Tunnel: Wing Modifications

    NASA Technical Reports Server (NTRS)

    Graham, David

    1950-01-01

    Wind-tunnel tests were conducted to determine the longitudinal stability characteristics of a full-scale Northrop XSSM-A-3 missile. Various wing modifications were investigated in an effort to provide a configuration that would maintain longitudinal stability to lift coefficients necessary for landing the missile during flight tests. The results of the tests led to the choice of a wing with an increased leading-edge radius. A short discussion of the results is presented, but no analysis of the data has been made in order to make the data available as soon as possible.

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

    NASA Technical Reports Server (NTRS)

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

    1981-01-01

    Buffet tests of two wings with different leading-edge sweep show that it is feasible to use the standards wing root bending moment technique in a cryogenic wing tunnel. The results for the 65 deg sweep delta wing indicate the importance of matching the reduced frequency parameter in model tests for planforms which are sensitive to reduced frequency parameter if quantitative buffet measurements are required. The unique ability of a pressurized cryogenic wind tunnel to separate the effects of Reynolds number and of aeroelastic distortion by variations in the tunnel stagnation temperature and pressure was demonstrated.

  13. Pressure distribution on the roof of a model low-rise building tested in a boundary layer wind tunnel

    NASA Astrophysics Data System (ADS)

    Goliber, Matthew Robert

    With three of the largest metropolitan areas in the United States along the Gulf coast (Houston, Tampa, and New Orleans), residential populations ever increasing due to the subtropical climate, and insured land value along the coast from Texas to the Florida panhandle greater than $500 billion, hurricane related knowledge is as important now as ever before. This thesis focuses on model low-rise building wind tunnel tests done in connection with full-scale low-rise building tests. Mainly, pressure data collection equipment and methods used in the wind tunnel are compared to pressure data collection equipment and methods used in the field. Although the focus of this report is on the testing of models in the wind tunnel, the low-rise building in the field is located in Pensacola, Florida. It has a wall length of 48 feet, a width of 32 feet, a height of 10 feet, and a gable roof with a pitch of 1:3 and 68 pressure ports strategically placed on the surface of the roof. Built by Forest Products Laboratory (FPL) in 2002, the importance of the test structure has been realized as it has been subjected to numerous hurricanes. In fact, the validity of the field data is so important that the following thesis was necessary. The first model tested in the Bill James Wind Tunnel for this research was a rectangular box. It was through the testing of this box that much of the basic wind tunnel and pressure data collection knowledge was gathered. Knowledge gained from Model 1 tests was as basic as how to: mount pressure tubes on a model, use a pressure transducer, operate the wind tunnel, utilize the pitot tube and reference pressure, and measure wind velocity. Model 1 tests also showed the importance of precise construction to produce precise pressure coefficients. Model 2 was tested in the AABL Wind Tunnel at Iowa State University. This second model was a 22 inch cube which contained a total of 11 rows of pressure ports on its front and top faces. The purpose of Model 2 was to

  14. Aeroservoelastic Testing of Free Flying Wind Tunnel Models Part 1: A Sidewall Supported Semispan Model Tested for Gust Load Alleviation and Flutter Suppression

    NASA Technical Reports Server (NTRS)

    Scott, Robert C.; Vetter, Travis K.; Penning, Kevin B.; Coulson, David A.; Heeg, Jennifer.

    2013-01-01

    of a two part document. Part 2 is titled: "Aeroservoelastic Testing of Free Flying Wind Tunnel Models, Part 2: A Centerline Supported Fullspan Model Tested for Gust Load Alleviation." A team comprised of the Air Force Research Laboratory (AFRL), Northrop Grumman, Lockheed Martin, and the NASA Langley Research Center conducted three aeroservoelastic wind tunnel tests in the Transonic Dynamics Tunnel to demonstrate active control technologies relevant to large, flexible vehicles. In the first of these three tests, a semispan, aeroelastically scaled, wind tunnel model of a flying wing SensorCraft vehicle was mounted to a force balance to demonstrate gust load alleviation. In the second and third tests, the same wing was mated to a new, multi-degree of freedom, sidewall mount. This mount allowed the half-span model to translate vertically and pitch at the wing root, allowing better simulation of the full span vehicle's rigid body modes. Gust load alleviation (GLA) and Body freedom flutter (BFF) suppression were successfully demonstrated. The rigid body degrees-of-freedom required that the model be flown in the wind tunnel using an active control system. This risky mode of testing necessitated that a model arrestment system be integrated into the new mount. The safe and successful completion of these free flying tests required the development and integration of custom hardware and software. This paper describes the many systems, software, and procedures that were developed as part of this effort.

  15. Wind tunnel test of model target thrust reversers for the Pratt and Whitney aircraft JT8D-100 series engines installed on a 727-200 airplane

    NASA Technical Reports Server (NTRS)

    Hambly, D.

    1974-01-01

    The results of a low speed wind tunnel test of 0.046 scale model target thrust reversers installed on a 727-200 model airplane are presented. The full airplane model was mounted on a force balance, except for the nacelles and thrust reversers, which were independently mounted and isolated from it. The installation had the capability of simulating the inlet airflows and of supplying the correct proportions of primary and secondary air to the nozzles. The objectives of the test were to assess the compatibility of the thrust reversers target door design with the engine and airplane. The following measurements were made: hot gas ingestion at the nacelle inlets; model lift, drag, and pitching moment; hot gas impingement on the airplane structure; and qualitative assessment of the rudder effectiveness. The major parameters controlling hot gas ingestion were found to be thrust reverser orientation, engine power setting, and the lip height of the bottom thrust reverser doors on the side nacelles. The thrust reversers tended to increase the model lift, decrease the drag, and decrease the pitching moment.

  16. Orion Multi-Purpose Crew Vehicle (MPCV) Capsule Parachute Assembly System (CPAS) Wake Deficit Wind Tunnel Testing

    NASA Technical Reports Server (NTRS)

    Ross, James C.; Schuster, David M.

    2014-01-01

    During descent after re-entry into the Earth's atmosphere, the Orion CM deploys its drogue parachutes at approximately Mach 0.7. Accurately predicting the dynamic pressure experienced by the drogue parachutes at deployment is critical to properly designing the parachutes. This NASA Engineering and Safety Center assessment was designed to provide a complete set of flowfield measurements on and around an idealized Orion Crew Module shape with the most appropriate wind tunnel simulation of the Orion flight conditions prior to parachute deployment. This document contains the details of testing and the outcome of the assessment.

  17. Investigation of solid plume simulation criteria to produce flight plume effects on multibody configuration in wind tunnel tests

    NASA Technical Reports Server (NTRS)

    Frost, A. L.; Dill, C. C.

    1986-01-01

    An investigation to determine the sensitivity of the space shuttle base and forebody aerodynamics to the size and shape of various solid plume simulators was conducted. Families of cones of varying angle and base diameter, at various axial positions behind a Space Shuttle launch vehicle model, were wind tunnel tested. This parametric evaluation yielded base pressure and force coefficient data which indicated that solid plume simulators are an inexpensive, quick method of approximating the effect of engine exhaust plumes on the base and forebody aerodynamics of future, complex multibody launch vehicles.

  18. A Wind-tunnel Test Technique for Measuring the Dynamic Rotary Stability Derivatives at Subsonic and Supersonic Speeds

    NASA Technical Reports Server (NTRS)

    Beam, Benjamin H

    1956-01-01

    A method is described for measuring the dynamic stability derivatives of a model airplane in a wind tunnel. The characteristic features of this system are that single-degree-of-freedom oscillations were used to obtain combinations of rolling, yawing and pitching motions; that the oscillations were excited and controlled by velocity feedback which permitted operation under conditions unfavorable for more conventional types of oscillatory testing; and that data processing was greatly simplified by using analog computer elements in the strain-gage circuitry. A small number of experimental data are included to illustrate the general scope of results obtainable with this system.

  19. Wind Tunnel Testing of a One-Dimensional Laser Beam Scanning and Laser Sheet Approach to Shock Sensing

    NASA Technical Reports Server (NTRS)

    Tokars, Roger; Adamovsky, Grigory; Anderson, Robert; Hirt, Stefanie; Huang, John; Floyd, Bertram

    2012-01-01

    A 15- by 15-cm supersonic wind tunnel application of a one-dimensional laser beam scanning approach to shock sensing is presented. The measurement system design allowed easy switching between a focused beam and a laser sheet mode for comparison purposes. The scanning results were compared to images from the tunnel Schlieren imaging system. The tests revealed detectable changes in the laser beam in the presence of shocks. The results lend support to the use of the one-dimensional scanning beam approach for detecting and locating shocks in a flow, but some issues must be addressed in regards to noise and other limitations of the system.

  20. Aeroservoelastic Wind-Tunnel Tests of a Free-Flying, Joined-Wing SensorCraft Model for Gust Load Alleviation

    NASA Technical Reports Server (NTRS)

    Scott, Robert C.; Castelluccio, Mark A.; Coulson, David A.; Heeg, Jennifer

    2011-01-01

    A team comprised of the Air Force Research Laboratory (AFRL), Boeing, and the NASA Langley Research Center conducted three aeroservoelastic wind-tunnel tests in the Transonic Dynamics Tunnel to demonstrate active control technologies relevant to large, exible vehicles. In the first of these three tests, a full-span, aeroelastically scaled, wind-tunnel model of a joined-wing SensorCraft vehicle was mounted to a force balance to acquire a basic aerodynamic data set. In the second and third tests, the same wind-tunnel model was mated to a new, two-degree-of-freedom, beam mount. This mount allowed the full-span model to translate vertically and pitch. Trimmed flight at -10% static margin and gust load alleviation were successfully demonstrated. The rigid body degrees of freedom required that the model be own in the wind tunnel using an active control system. This risky mode of testing necessitated that a model arrestment system be integrated into the new mount. The safe and successful completion of these free-flying tests required the development and integration of custom hardware and software. This paper describes the many systems, software, and procedures that were developed as part of this effort. The balance and free ying wind-tunnel tests will be summarized. The design of the trim and gust load alleviation control laws along with the associated results will also be discussed.

  1. Jet V/STOL wind-tunnel simulation and groundplane effects

    NASA Technical Reports Server (NTRS)

    Margason, R. J.

    1982-01-01

    The pretest preparation necessary to define the objectives of an appropriate investigation into the jet V/STOL wind tunnel simulation and ground plane effects were examined. Low speed wind tunnel testing of V/STOL aircraft concepts to determine the aerodynamic propulsion interaction effects during the transition between hover and wingborne flight is a necessary step in the development cycle of this type of aircraft. Powered models are normally used to determine the aerodynamic performance characteristics. Several factors which influence the selection of the model concept and the engine simulator are discussed. Some of the test techniques important for this class of aircraft model are examined. Wind tunnel wall effects important to this aircraft testing with special emphasis on groundplane effects are reviewed.

  2. Comparison of aerodynamic coefficients obtained from theoretical calculations wind tunnel tests and flight tests data reduction for the alpha jet aircraft

    NASA Technical Reports Server (NTRS)

    Guiot, R.; Wunnenberg, H.

    1980-01-01

    The methods by which aerodynamic coefficients are determined and discussed. These include: calculations, wind tunnel experiments and experiments in flight for various prototypes of the Alpha Jet. A comparison of obtained results shows good correlation between expectations and in-flight test results.

  3. The state of the art of conventional flow visualization techniques for wind tunnel testing

    NASA Technical Reports Server (NTRS)

    Settles, G. S.

    1982-01-01

    Conventional wind tunnel flow visualization techniques which consist of surface flow methods, tracers, and optical methods are presented. Different surface flow methods are outlined: (1) liquid films (oil and fluorescent dye and UV lighting, renewable film via porous dispenser in model, volatile carrier fluid, cryogenic colored oil dots, oil film interferometry); (2) reactive surface treatment (reactive gas injection, reversible dye); (3) transition and heat transfer detectors (evaporation, sublimation, liquid crystals, phase change paints, IR thermography); and (4) tufts (fluorescent mini tufts, cryogenic suitability). Other methods are smoke wire techniques, vapor screens, and optical methods.

  4. The 727 airplane side inlet low-speed performance confirmation model test for refanned JT8D engines

    NASA Technical Reports Server (NTRS)

    Schuehle, A. L.

    1974-01-01

    The results of a low-speed wind tunnel test of a 0.3 scale model 727 airplane side inlet for JT8D-100 engines are presented. The objectives of the test were to develop lines for a full-scale flightworthy inlet, to evaluate inlet total pressure recovery and steady-state total pressure distortion, and to obtain model-scale distortion data which can be used in the assessment of the compatibility of the inlet with the JT8D-100 series engines. A secondary objective was to obtain internal/external cowl static pressures for the determination of nacelle loads. Two basic inlet models were tested at static, forward speed, angle-of-attack (inflow angle), and cross-wind conditions. One model was with and one without an acoustic ring. Two modifications to the models were also tested, one with the ring closer to the inlet throat and one with a larger lip. Test measurements consisted of inlet surface static pressure, engine face total pressure, inlet airflow, tunnel total pressure, tunnel total temperature and tunnel velocity. Total pressure traverses were taken directly behind the ring and strut. No dynamic measurements were taken.

  5. Reducing the Effect of Transducer Mount Induced Noise on Aeroacoustic Wind Tunnel Testing Data with a New Transducer Mount Design

    NASA Technical Reports Server (NTRS)

    Herron, Andrew J.; Reed, Darren K.; Nance, Donald K.

    2015-01-01

    Flight vehicle aeroacoustic environments induced during transonic and supersonic flight are usually predicted by subscale wind tunnel testing utilizing high frequency miniature pressure transducers. In order to minimize noise induced by the measurement itself, transducer flush mounting with the model surface is very important. The National Aeronautics and Space Administration (NASA) has accomplished flushness in recent testing campaigns via use of a transducer holder that can be machined and sanded. A single hole in the holder allows the flow medium to interact with the transducer diaphragm. Noise is induced by the resulting cavity however, and is a challenge to remove in post-processing. A new holder design has been developed that minimizes the effects of this transducer mount induced noise (XMIN) by reducing the resonance amplitude or increasing its resonance frequency beyond the range of interest. This paper describes a test conducted at the NASA/George C. Marshall Space Flight Center Trisonic Wind Tunnel intended to verify the effectiveness of this design. The results from this test show that this new transducer holder design does significantly reduce the influence of XMIN on measured fluctuating pressure levels without degrading a transducer's ability to accurately measure the noise external to the model.

  6. Wind-Tunnel Investigation at Low Speed of the Effects of Chordwise Wing Fences and Horizontal-Tail Position on the Static Longitudinal Stability Characteristics of an Airplane Model with a 35 Degree Sweptback Wing

    NASA Technical Reports Server (NTRS)

    Queijo, M J; Jaquet, Byron M; Wolhart, Walter D

    1954-01-01

    Low-speed tests of a model with a wing swept back 35 degrees at the 0.33-chord line and a horizontal tail located well above the extended wing-chord plane indicated static longitudinal instability at moderate angles of attack for all configurations tested. An investigation therefore was made to determine whether the longitudinal stability could be improved by the use of chordwise wing fences, by lowering the horizontal tail, or by a combination of both. The results of the investigation showed that the longitudinal stability characteristics of the model with slats retracted could be improved at moderate angles of attack by placing chordwise wing fences at a spanwise station of about 73 percent of the wing semispan from the plane of symmetry provided the nose of the fence extended slightly beyond or around the wing leading edge.

  7. Full-scale Wind-tunnel and Flight Tests of a Fairchild 22 Airplane Equipped with a Fowler Flap

    NASA Technical Reports Server (NTRS)

    Dearborn, C H; Soule, H A

    1936-01-01

    Full-scale wind-tunnel and flight tests were made of a Fairchild 22 airplane equipped with a Fowler flap to determine the effect of the flap on the performance and control characteristics of the airplane. In the wind-tunnel tests of the airplane with the horizontal tail surfaces removed, the flap was found to increase the maximum lift coefficient from 1.27 to 2.41. In the flight test, the flap was found to decrease the minimum speed from 58.8 to 44.4 miles per hour. The required take-off run to attain an altitude of 50 feet was reduced from 935 feet to 700 feet by the use of the flap, the minimum distance being obtained with five-sixths full deflection. The landing run from a height of 50 feet was reduced one-third. The longitudinal and directional control was adversely affected by the flap, indicating that the design of the tail surfaces is more critical with a flapped than a plain wing.

  8. Full-scale wind-tunnel tests of a small unpowered jet aircraft with a T-tail

    NASA Technical Reports Server (NTRS)

    Soderman, P. T.; Aiken, T. N.

    1971-01-01

    The aerodynamic characteristics of a full scale executive type jet transport aircraft with a T-tail were investigated in a 40 x 80 ft (12.2 by 24.4 meter) wind tunnel (subsonic). Static, longitudinal, and lateral stability, and control characteristics were determined at angles of attack from -2 deg to +42 deg. The aircraft wing had 13 deg of sweep and an aspect ratio of 5.02. The aircraft was tested power off with various wing leading- and trailing-edge high lift devices. The basic configuration was tested with and without such components as engine nacelles, wing tip tanks, and empannage. Hinge-moment data were obtained and downwash angles in the horizontal-tail plane location were calculated. The data were obtained at Reynolds numbers of 4.1 million and 8.7 million based on mean aerodynamic chord. The model had static longitudinal stability through initial stall. Severe tail buffet occurred near the angle of attack for maximum lift. Above initial stall the aircraft had pronounced pitch-up, characteristic of T-tail configurations. A stable trim point was possible at angles of attack between 30 deg and 40 deg (depending on c.g. location and flap setting). Hinge-moment data showed no regions with adverse effects on stick force. Comparisons of wind-tunnel data and flight-test are presented.

  9. National Wind Tunnel Complex (NWTC)

    NASA Technical Reports Server (NTRS)

    1996-01-01

    The National Wind Tunnel Complex (NWTC) Final Report summarizes the work carried out by a unique Government/Industry partnership during the period of June 1994 through May 1996. The objective of this partnership was to plan, design, build and activate 'world class' wind tunnel facilities for the development of future-generation commercial and military aircraft. The basis of this effort was a set of performance goals defined by the National Facilities Study (NFS) Task Group on Aeronautical Research and Development Facilities which established two critical measures of improved wind tunnel performance; namely, higher Reynolds number capability and greater productivity. Initial activities focused upon two high-performance tunnels (low-speed and transonic). This effort was later descoped to a single multipurpose tunnel. Beginning in June 1994, the NWTC Project Office defined specific performance requirements, planned site evaluation activities, performed a series of technical/cost trade studies, and completed preliminary engineering to support a proposed conceptual design. Due to budget uncertainties within the Federal government, the NWTC project office was directed to conduct an orderly closure following the Systems Design Review in March 1996. This report provides a top-level status of the project at that time. Additional details of all work performed have been archived and are available for future reference.

  10. Burning carbon monoxide in the settling chamber of a hotshot wind tunnel for obtaining the CO2 test gas

    NASA Astrophysics Data System (ADS)

    Shumskii, V. V.; Yaroslavtsev, M. I.

    2016-03-01

    A method of formation and heating of CO2 as a test gas in the settling chamber of a hotshot wind tunnel is considered. To form and heat CO2, the chamber is filled with a source gas mixture of CO, O2, and CO2, and after initiation, these substances participate in an exothermic chemical reaction in accordance with the formula CO + 0.5 O2 + xCO2 = (1 + x)CO2. A stoichiometric ratio of the concentrations of carbon monoxide CO and oxygen is used. Variation of the number of moles x of ballast CO2 in the left part of the chemical formula allows changing the temperature of the resultant test gas in a wide range. Experiments in the IT-302M hotshot wind tunnel carried out at ITAM SB RAS have shown that a pressure increase during an isochoric process in the settling chamber due to the joint effect of heat released in the reaction CO + 0.5 O2 and an electric charge provides the completeness of CO combustion almost equal to unity. The time of reaction completion at its initiation by an electric arc is no more than several milliseconds.

  11. Simulation of flow over double-element airfoil and wind tunnel test for use in vertical axis wind turbine

    NASA Astrophysics Data System (ADS)

    Chougule, Prasad; Nielsen, Søren R. K.

    2014-06-01

    Nowadays, small vertical axis wind turbines are receiving more attention due to their suitability in micro-electricity generation. There are few vertical axis wind turbine designs with good power curve. However, the efficiency of power extraction has not been improved. Therefore, an attempt has been made to utilize high lift technology for vertical axis wind turbines in order to improve power efficiency. High lift is obtained by double-element airfoil mainly used in aeroplane wing design. In this current work a low Reynolds number airfoil is selected to design a double-element airfoil blade for use in vertical axis wind turbine to improve the power efficiency. Double-element airfoil blade design consists of a main airfoil and a slat airfoil. Orientation of slat airfoil is a parameter of investigation in this paper and air flow simulation over double-element airfoil. With primary wind tunnel test an orientation parameter for the slat airfoil is initially obtained. Further a computational fluid dynamics (CFD) has been used to obtain the aerodynamic characteristics of double-element airfoil. The CFD simulations were carried out using ANSYS CFX software. It is observed that there is an increase in the lift coefficient by 26% for single-element airfoil at analysed conditions. The CFD simulation results were validated with wind tunnel tests. It is also observe that by selecting proper airfoil configuration and blade sizes an increase in lift coefficient can further be achieved.

  12. Performance and aerodynamic braking of a horizontal-axis wind turbine from small-scale wind tunnel tests

    NASA Technical Reports Server (NTRS)

    Cao, H. V.; Wentz, W. H., Jr.

    1987-01-01

    Wind tunnel tests of three 20" diameter, zero twist, zero pitch wind turbine rotor models were conducted in a 7' x 10' wind tunnel to determine the performance of such rotors with NACA 23024 and NACA 64 sub 3-621 airfoil sections. Aerodynamic braking characteristics of a 38% span, 30% chord, vented aileron configuration were measured on the NACA 23024 rotor. Surface flow patterns were observed using fluorescent mini-tufts attached to the suction side of the rotor blades. Experimental results with and without ailerons are compared to predictions using airfoil section data and a momentum performance code. Results of the performance studies show that the 64 sub 3-621 rotor produces higher peak power than the 23024 rotor for a given rotor speed. Analytical studies, however, indicate that the 23024 should produce higher power. Transition strip experiments show that the 23024 rotor is much more sensitive to roughness than the 64 sub 3-621 rotor. These trends agree with analytical predictions. Results of the aileron test show that this aileron, when deflected, produces a braking torque at all tip speed ratios. In free wheeling coastdowns the rotor blade stopped, then rotated backward at a tip speed ratio of -0.6.

  13. Flow Visualization Techniques in Wind Tunnel Tests of a Full-Scale F/A-18 Aircraft

    NASA Technical Reports Server (NTRS)

    Lanser, Wendy R.; Botha, Gavin J.; James, Kevin D.; Bennett, Mark; Crowder, James P.; Cooper, Don; Olson, Lawrence (Technical Monitor)

    1994-01-01

    The proposed paper presents flow visualization performed during experiments conducted on a full-scale F/A-18 aircraft in the 80- by 120-Foot Wind-Tunnel at NASA Ames Research Center. The purpose of the flow-visualization experiments was to document the forebody and leading edge extension (LEX) vortex interaction along with the wing flow patterns at high angles of attack and low speed high Reynolds number conditions. This investigation used surface pressures in addition to both surface and off-surface flow visualization techniques to examine the flow field on the forebody, canopy, LEXS, and wings. The various techniques used to visualize the flow field were fluorescent tufts, flow cones treated with reflective material, smoke in combination with a laser light sheet, and a video imaging system for three-dimension vortex tracking. The flow visualization experiments were conducted over an angle of attack range from 20 deg to 45 deg and over a sideslip range from -10 deg to 10 deg. The various visualization techniques as well as the pressure distributions were used to understand the flow field structure. The results show regions of attached and separated flow on the forebody, canopy, and wings as well as the vortical flow over the leading-edge extensions. This paper will also present flow visualization comparisons with the F-18 HARV flight vehicle and small-scale oil flows on the F-18.

  14. U.S. aerospace industry opinion of the effect of computer-aided prediction-design technology on future wind-tunnel test requirements for aircraft development programs

    NASA Technical Reports Server (NTRS)

    Treon, S. L.

    1979-01-01

    A survey of the U.S. aerospace industry in late 1977 suggests that there will be an increasing use of computer-aided prediction-design technology (CPD Tech) in the aircraft development process but that, overall, only a modest reduction in wind-tunnel test requirements from the current level is expected in the period through 1995. Opinions were received from key spokesmen in 23 of the 26 solicited major companies or corporate divisions involved in the design and manufacture of nonrotary wing aircraft. Development programs for nine types of aircraft related to test phases and wind-tunnel size and speed range were considered.

  15. Some theoretical considerations of longitudinal stability in power-on flight with special reference to wind-tunnel testing, November 1942

    NASA Technical Reports Server (NTRS)

    Donlan, C. J.

    1976-01-01

    Some problems relating to longitudinal stability in power-on flight are considered. A derivation is included which shows that, under certain conditions, the rate of change of the pitching moment coefficient with lift coefficient as obtained in wind tunnel tests simulating constant power operation is directly proportional to one of the indices of stability commonly associated with flight analysis, (the slope of the curve relating the elevator angle for trim and lift coefficient). The necessity of analyzing power-on wind tunnel data for trim conditions is emphasized, and a method is provided for converting data obtained from constant thrust tests to simulated constant throttle flight conditions.

  16. Design, manufacturing and characterization of aero-elastically scaled wind turbine blades for testing active and passive load alleviation techniques within a ABL wind tunnel

    NASA Astrophysics Data System (ADS)

    Campagnolo, Filippo; Bottasso, Carlo L.; Bettini, Paolo

    2014-06-01

    In the research described in this paper, a scaled wind turbine model featuring individual pitch control (IPC) capabilities, and equipped with aero-elastically scaled blades featuring passive load reduction capabilities (bend-twist coupling, BTC), was constructed to investigate, by means of wind tunnel testing, the load alleviation potential of BTC and its synergy with active load reduction techniques. The paper mainly focus on the design of the aero-elastic blades and their dynamic and static structural characterization. The experimental results highlight that manufactured blades show desired bend-twist coupling behavior and are a first milestone toward their testing in the wind tunnel.

  17. Dynamic Wind-Tunnel Testing of a Sub-Scale Iced Business Jet

    NASA Technical Reports Server (NTRS)

    Lee, Sam; Barnhart, Billy P.; Ratvasky, Thomas P.; Dickes, Edward; Thacker, Michael

    2006-01-01

    The effect of ice accretion on a 1/12-scale complete aircraft model of a business jet was studied in a rotary-balance wind tunnel. Three types of ice accretions were considered: ice protection system failure shape, pre-activation roughness, and runback shapes that form downstream of the thermal ice protection system. The results were compared with those from a 1/12-scale semi-span wing of the same aircraft at similar Reynolds number. The data showed that the full aircraft and the semi-span wing models showed similar characteristics, especially post stall behavior under iced configuration. However, there were also some discrepancies, such as the magnitude in the reductions in the maximum lift coefficient. Most of the ice-induced effects were limited to longitudinal forces. Rotational and forced oscillation studies showed that the effects of ice on lateral forces were relatively minor.

  18. Wind tunnel tests of the dynamic characteristics of the fluidic rudder

    NASA Technical Reports Server (NTRS)

    Belsterling, C. A.

    1976-01-01

    The fourth phase is given of a continuing program to develop the means to stabilize and control aircraft without moving parts or a separate source of power. Previous phases have demonstrated the feasibility of (1) generating adequate control forces on a standard airfoil, (2) controlling those forces with a fluidic amplifier and (3) cascading non-vented fluidic amplifiers operating on ram air supply pressure. The foremost objectives of the fourth phase covered under Part I of this report were to demonstrate a complete force-control system in a wind tunnel environment and to measure its static and dynamic control characteristics. Secondary objectives, covered under Part II, were to evaluate alternate configurations for lift control. The results demonstrate an overall response time of 150 msec, confirming this technology as a viable means for implementing low-cost reliable flight control systems.

  19. A computer-controlled, on-board data acquisition system for wind-tunnel testing

    NASA Technical Reports Server (NTRS)

    Finger, H. J.; Cambra, J. M.

    1974-01-01

    A computer-controlled data acquisition system has been developed for the 40x80-foot wind tunnel at Ames Research Center. The system, consisting of several small onboard units installed in the model and a data-managing, data-displaying ground station, is capable of sampling up to 256 channels of raw data at a total sample rate of 128,000 samples/sec. Complete signal conditioning is contained within the on-board units. The sampling sequence and channel gain selection is completely random and under total control of the ground station. Outputs include a bar-graph display, digital-to-analog converters, and digital interface to the tunnel's central computer, an SEL 840MP. The system can be run stand-alone or under the control of the SEL 840MP.

  20. Aerodynamic Parameters of High Performance Aircraft Estimated from Wind Tunnel and Flight Test Data

    NASA Technical Reports Server (NTRS)

    Klein, Vladislav; Murphy, Patrick C.

    1998-01-01

    A concept of system identification applied to high performance aircraft is introduced followed by a discussion on the identification methodology. Special emphasis is given to model postulation using time invariant and time dependent aerodynamic parameters, model structure determination and parameter estimation using ordinary least squares an mixed estimation methods, At the same time problems of data collinearity detection and its assessment are discussed. These parts of methodology are demonstrated in examples using flight data of the X-29A and X-31A aircraft. In the third example wind tunnel oscillatory data of the F-16XL model are used. A strong dependence of these data on frequency led to the development of models with unsteady aerodynamic terms in the form of indicial functions. The paper is completed by concluding remarks.

  1. Aerodynamic Parameters of High Performance Aircraft Estimated from Wind Tunnel and Flight Test Data

    NASA Technical Reports Server (NTRS)

    Klein, Vladislav; Murphy, Patrick C.

    1999-01-01

    A concept of system identification applied to high performance aircraft is introduced followed by a discussion on the identification methodology. Special emphasis is given to model postulation using time invariant and time dependent aerodynamic parameters, model structure determination and parameter estimation using ordinary least squares and mixed estimation methods. At the same time problems of data collinearity detection and its assessment are discussed. These parts of methodology are demonstrated in examples using flight data of the X-29A and X-31A aircraft. In the third example wind tunnel oscillatory data of the F-16XL model are used. A strong dependence of these data on frequency led to the development of models with unsteady aerodynamic terms in the form of indicial functions. The paper is completed by concluding remarks.

  2. Small scale noise and wind tunnel tests of upper surface blowing nozzle flap concepts. Volume 1. Aerodynamic test results

    NASA Technical Reports Server (NTRS)

    Renselaer, D. J.; Nishida, R. S.; Wilkin, C. A.

    1975-01-01

    The results and analyses of aerodynamic and acoustic studies conducted on the small scale noise and wind tunnel tests of upper surface blowing nozzle flap concepts are presented. Various types of nozzle flap concepts were tested. These are an upper surface blowing concept with a multiple slot arrangement with seven slots (seven slotted nozzle), an upper surface blowing type with a large nozzle exit at approximately mid-chord location in conjunction with a powered trailing edge flap with multiple slots (split flow or partially slotted nozzle). In addition, aerodynamic tests were continued on a similar multi-slotted nozzle flap, but with 14 slots. All three types of nozzle flap concepts tested appear to be about equal in overall aerodynamic performance but with the split flow nozzle somewhat better than the other two nozzle flaps in the landing approach mode. All nozzle flaps can be deflected to a large angle to increase drag without significant loss in lift. The nozzle flap concepts appear to be viable aerodynamic drag modulation devices for landing.

  3. Acoustical modeling study of the open test section of the NASA Langley V/STOL wind tunnel

    NASA Technical Reports Server (NTRS)

    Ver, I. L.; Andersen, D. W.; Bliss, D. B.

    1975-01-01

    An acoustic model study was carried out to identify effective sound absorbing treatment of strategically located surfaces in an open wind tunnel test section. Also an aerodynamic study done concurrently, sought to find measures to control low frequency jet pulsations which occur when the tunnel is operated in its open test section configuration. The acoustical modeling study indicated that lining of the raised ceiling and the test section floor immediately below it, results in a substantial improvement. The aerodynamic model study indicated that: (1) the low frequency jet pulsations are most likely caused or maintained by coupling of aerodynamic and aeroacoustic phenomena in the closed tunnel circuit, (2) replacing the hard collector cowl with a geometrically identical but porous fiber metal surface of 100 mks rayls flow resistance does not result in any noticable reduction of the test section noise caused by the impingement of the turbulent flow on the cowl.

  4. Evaluation of four subcritical response methods for on-line prediction of flutter onset in wind-tunnel tests

    NASA Technical Reports Server (NTRS)

    Ruhlin, C. L.; Watson, J. J.; Ricketts, R. H.; Doggett, R. V., Jr.

    1982-01-01

    Four subcritical response methods were evaluated for on-line use in transonic wind-tunnel tests where the flutter model is excited solely by airstream turbulence. The methods were: randomdec, power-spectral-density, peak-hold, and cross-spectrum. Subcritical response data were obtained during tests in the Langley Transonic Dynamics Tunnel of a cantilevered flutter model wing. The test procedure was to maintain a constant Mach number and increase the dynamic pressure (q) in incremental steps. The four methods provided damping trends by which the flutter mode could be tracked and extrapolated to a flutter-onset q. A hard flutter point was obtained at M = 0.82. The peak-mold and cross-spectrum methods gave reliable results and could be most readily used for on-line testing.

  5. Capabilities of wind tunnels with two-adaptive walls to minimize boundary interference in 3-D model testing

    NASA Technical Reports Server (NTRS)

    Rebstock, Rainer; Lee, Edwin E., Jr.

    1989-01-01

    An initial wind tunnel test was made to validate a new wall adaptation method for 3-D models in test sections with two adaptive walls. First part of the adaptation strategy is an on-line assessment of wall interference at the model position. The wall induced blockage was very small at all test conditions. Lift interference occurred at higher angles of attack with the walls set aerodynamically straight. The adaptation of the top and bottom tunnel walls is aimed at achieving a correctable flow condition. The blockage was virtually zero throughout the wing planform after the wall adjustment. The lift curve measured with the walls adapted agreed very well with interference free data for Mach 0.7, regardless of the vertical position of the wing in the test section. The 2-D wall adaptation can significantly improve the correctability of 3-D model data. Nevertheless, residual spanwise variations of wall interference are inevitable.

  6. Development of pneumatic test techniques for subsonic high-lift and in-ground-effect wind tunnel investigations

    NASA Technical Reports Server (NTRS)

    Englar, Robert J.

    1994-01-01

    Wind tunnel evaluations of two-dimensional high-lift airfoils and of vehicles operating in ground effect near the tunnel floor require special test facilities and procedures. These are needed to avoid errors caused by proximity to the walls and interference from the wall boundary layers. Pneumatic test techniques and facilities were developed for GTRI aerodynamic research tunnels and calibrated to verify that these wall effects had been removed. The modified facilities were then employed to evaluate the aerodynamic characteristics of blown very-high-lift airfoils and of racing hydroplanes operating in ground effect at various levels above the floor. The pneumatic facilities, techniques and calibrations are discussed and typical aerodynamic data recorded both with and without the test-section blowing systems are presented.

  7. Configuration design studies and wind tunnel tests of an energy efficient transport with a high-aspect-ratio supercritical wing

    NASA Technical Reports Server (NTRS)

    Henne, P. A.; Dahlin, J. A.; Peavey, C. C.; Gerren, D. S.

    1982-01-01

    The results of design studies and wind tunnel tests of high aspect ratio supercritical wings suitable for a medium range, narrow body transport aircraft flying near M=0.80 were presented. The basic characteristics of the wing design were derived from system studies of advanced transport aircraft where detailed structural and aerodynamic tradeoffs were used to determine the most optimum design from the standpoint of fuel usage and direct operating cost. These basic characteristics included wing area, aspect ratio, average thickness, and sweep. The detailed wing design was accomplished through application of previous test results and advanced computational transonic flow procedures. In addition to the basic wing/body development, considerable attention was directed to nacelle/plyon location effects, horizontal tail effects, and boundary layer transition effects. Results of these tests showed that the basic cruise performance objectives were met or exceeded.

  8. Experimental Determination of Jet Boundary Corrections for Airfoil Tests in Four Open Wind Tunnel Jets of Different Shapes

    NASA Technical Reports Server (NTRS)

    Knight, Montgomery; Harris, Thomas A

    1931-01-01

    This experimental investigation was conducted primarily for the purpose of obtaining a method of correcting to free air conditions the results of airfoil force tests in four open wind tunnel jets of different shapes. Tests were also made to determine whether the jet boundaries had any appreciable effect on the pitching moments of a complete airplane model. Satisfactory corrections for the effect of the boundaries of the various jets were obtained for all the airfoils tested, the span of the largest being 0.75 of the jet width. The corrections for angle of attack were, in general, larger than those for drag. The boundaries had no appreciable effect on the pitching moments of either the airfoils or the complete airplane model. Increasing turbulence appeared to increase the minimum drag and maximum lift and to decrease the pitching moment.

  9. 40 CFR 205.54-1 - Low speed sound emission test procedures.

    Code of Federal Regulations, 2013 CFR

    2013-07-01

    ... 40 Protection of Environment 26 2013-07-01 2013-07-01 false Low speed sound emission test....54-1 Low speed sound emission test procedures. (a) Instrumentation. The following instrumentation... checked annually to verify that its output has not changed. (3) An engine-speed tachometer which...

  10. 40 CFR 205.54-1 - Low speed sound emission test procedures.

    Code of Federal Regulations, 2010 CFR

    2010-07-01

    ... 40 Protection of Environment 24 2010-07-01 2010-07-01 false Low speed sound emission test....54-1 Low speed sound emission test procedures. (a) Instrumentation. The following instrumentation... checked annually to verify that its output has not changed. (3) An engine-speed tachometer which...

  11. 40 CFR 205.54-1 - Low speed sound emission test procedures.

    Code of Federal Regulations, 2012 CFR

    2012-07-01

    ... 40 Protection of Environment 26 2012-07-01 2011-07-01 true Low speed sound emission test....54-1 Low speed sound emission test procedures. (a) Instrumentation. The following instrumentation... checked annually to verify that its output has not changed. (3) An engine-speed tachometer which...

  12. 40 CFR 205.54-1 - Low speed sound emission test procedures.

    Code of Federal Regulations, 2014 CFR

    2014-07-01

    ... 40 Protection of Environment 25 2014-07-01 2014-07-01 false Low speed sound emission test....54-1 Low speed sound emission test procedures. (a) Instrumentation. The following instrumentation... checked annually to verify that its output has not changed. (3) An engine-speed tachometer which...

  13. Wind tunnel test results for the direction controlled antitank DCAT missile at Mach numbers from 0.64 to 2.50

    NASA Technical Reports Server (NTRS)

    Martin, T. A.; Spring, D. J.

    1973-01-01

    Wind tunnel test results are presented to show aerodynamic characteristics over the Mach number range of 0.64 to 2.50 of the DCAT missile. Data are presented showing the interference created by the rear mounted reaction control system. Two candidate fins were installed on the model during tests: a flat folding fin and a curved wrap around fin.

  14. Exploratory testing of supersonic ribbon parachutes in the NASA Ames 9-ft by 7-ft wind tunnel

    SciTech Connect

    Pepper, W.B.; Buffington, R.J.; Peterson, C.W.

    1986-01-01

    Sandia National Laboratories conducted a test of ribbon parachutes in the NASA Ames Research Center 9-ft by 7-ft Supersonic Wind Tunnel. The Ribbon parachutes were 15 inches in diameter with Kevlar-29 suspension lines and nylon ribbons. Both 20-degree conical parachutes and hemisflo parachutes were tested behind two forebody shapes with several forebody diameters. Drag data and color Schlieren photographs were obtained over a Mach number range of 1.5 to 2.5 and a dynamic pressure range of 200 to 400 lbs/ft/sup 2/. Correlations of supersonic parachute drag with Mach number, forebody shape and diameter, canopy porosity, suspension line length, and inflated canopy diameter and stability are presented.

  15. Icing testing in the large Modane wind-tunnel on full-scale and reduced scale models

    NASA Technical Reports Server (NTRS)

    Charpin, F.; Fasso, G.

    1979-01-01

    Icing tests on full scale models of parts of aircraft (wings, tailplanes, radome) equipped with actual de-icing systems were carried out in the large Modane wind tunnel of ONERA. For studying icing on the Concorde, it was necessary to use a 1/6 scale half model. The equations governing the relevant parameter ratios to obtain reasonably good similitude water catching and ice accretion are recalled. Despite the inherent limitations of this particular kind of testing, i.e., the impossibility of duplicating both the Mach and Reynolds conditions for the main flow pattern, it is possible to obtain on a reduced scale model a reasonably good representation of icing cloud catching and of the shape of resulting ice accretion.

  16. Exploratory study to induce fan noise in the test section of the NASA Langley full-scale wind tunnel

    NASA Technical Reports Server (NTRS)

    Ver, I. L.; Hayden, R. E.; Myles, M. M.; Murray, B. E.

    1975-01-01

    Measures to reduce the intensity of fan noise in the NASA Langley 30 ft x 60 ft subsonic wind tunnel were sought. Measurements were first performed to document existing aerodynamic and acoustic conditions. The purpose of these experiments was to (1) obtain the transfer function between the sound power output of the fan and the sound pressure on the test platform, (2) evaluate the sound attenuation around the tunnel circuit, (3) measure simultaneously the flow profile and the turbulence spectrum of the inflow to the fan and the noise on the test platform, and (4) perform flow observations and identify secondary noise sources. Subsequently, these data were used to predict (1) the relative contribution of the major aerodynamic parameters to total fan noise and (2) the effect of placing a dissipative silencer in the collector duct upstream of the fan. Promising noise control measures were identified and recommendations were made on how to evaluate them.

  17. Performance and loads data from a wind tunnel test of a full-scale, coaxial, hingeless rotor helicopter

    NASA Technical Reports Server (NTRS)

    Felker, F. F., III

    1981-01-01

    A full-scale XH-59A advancing blade concept helicopter was tested in Ames Research Center's 40 by 80 foot wind tunnel. The helicopter was tested with the rotor on and off, rotor hub fairings on and off, interrotor shaft fairing on and off, rotor instrumentation module on and off, and auxiliary propulsion thrust on and off. An advance ratio range of 0.25 and 0.45 with the rotor on and from 60 to 180 knots with the rotor off was investigated. Data on aerodynamic forces and moments, rotor loads, rotor control positions and vibration for the XH-59A as well as the aerodynamic performance of the isolated rotor are presented.

  18. Preliminary Computational Study for Future Tests in the NASA Ames 9 foot' x 7 foot Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Pearl, Jason M.; Carter, Melissa B.; Elmiligui, Alaa A.; WInski, Courtney S.; Nayani, Sudheer N.

    2016-01-01

    The NASA Advanced Air Vehicles Program, Commercial Supersonics Technology Project seeks to advance tools and techniques to make over-land supersonic flight feasible. In this study, preliminary computational results are presented for future tests in the NASA Ames 9 foot x 7 foot supersonic wind tunnel to be conducted in early 2016. Shock-plume interactions and their effect on pressure signature are examined for six model geometries. Near- field pressure signatures are assessed using the CFD code USM3D to model the proposed test geometries in free-air. Additionally, results obtained using the commercial grid generation software Pointwise Reigistered Trademark are compared to results using VGRID, the NASA Langley Research Center in-house mesh generation program.

  19. 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.

  20. 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.

  1. Reducing the Effect of Transducer Mount Induced Noise on Aeroacoustic Wind Tunnel Testing Data with a New Transducer Mount Design

    NASA Technical Reports Server (NTRS)

    Herron, A. J.; Reed, D. K.; Nance, D. K.

    2015-01-01

    Characterization of launch vehicle unsteady aerodynamics is a field best studied through experimentation, which is often carried out in the form of large scale wind tunnel testing. Measurement of the fluctuating pressures induced by the boundary layer noise is customarily made with miniature pressure transducers installed into a model of the vehicle of interest. Literature shows that noise level increases between two to five decibels (dB referenced to 20 micropascal) can be induced when the transducer surface is not mounted perfectly flush with the model outer surface. To reduce this artificially induced noise, special transducer holders have been used for aeroacoustic wind tunnel testing by NASA. This holder is a sleeve into which the transducer fits, with a cap that allows it to be mounted in a recessed hole in the model. A single hole in the holder allows the transport of the tunnel medium so the transducer can discriminate the fluctuating pressure due to the turbulent boundary layer noise. The holder is first dry fitted into the model and any difference in height between the holder and the model surface can be sanded flush. The holder is then removed from the model, the transducer glued inside the holder, and the holder replaced in the model, secured also with glue, thus eliminating the problem of noise level increases due to lack of flushness. In order to work with this holder design, special transducers have been ordered with their standard screen removed and the diaphragm moved as close to the top of the casing as possible to minimize any cavity volume. Although this greatly reduces induced noise due to the transducers being out of flush, the holders can also induce a cavity resonance that is usually at a very high frequency. This noise is termed transducer mount induced noise (XMIN). The peak of the mode can vary with the cavity depth, boundary layer noise that can excite the mode, tunnel flow medium, and the build of the transducers. Because the boundary

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

    NASA Astrophysics Data System (ADS)

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

    2015-02-01

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

  3. Developing, mechanizing and testing of a digital active flutter suppression system for a modified B-52 wind-tunnel model

    NASA Technical Reports Server (NTRS)

    Matthew, J. R.

    1980-01-01

    A digital flutter suppression system was developed and mechanized for a significantly modified version of the 1/30-scale B-52E aeroelastic wind tunnel model. A model configuration was identified that produced symmetric and antisymmetric flutter modes that occur at 2873N/sq m (60 psf) dynamic pressure with violent onset. The flutter suppression system, using one trailing edge control surface and the accelerometers on each wing, extended the flutter dynamic pressure of the model beyond the design limit of 4788N/sq m (100 psf). The hardware and software required to implement the flutter suppression system were designed and mechanized using digital computers in a fail-operate configuration. The model equipped with the system was tested in the Transonic Dynamics Tunnel at NASA Langley Research Center and results showed the flutter dynamic pressure of the model was extended beyond 4884N/sq m (102 psf).

  4. Tests of N.A.C.A. airfoils in the variable density wind tunnel Series 44 and 64

    NASA Technical Reports Server (NTRS)

    Jacobs, Eastman N; Pinkerton, Robert M

    1931-01-01

    This note is one of a series covering an investigation of a number of related airfoils. It presents the results obtained from tests in the N.A.C.A. Variable Density Wind Tunnel of two groups of six airfoils each. One group, the 44 series, has a maximum mean camber of 4 percent of the chord at a position 0.4 of the chord behind the leading edge and the other group, the 64 series, has a maximum mean camber of 6 percent of the chord at the same position. The members within each group differ only in maximum thickness, the maximum thickness/chord ratios being: 0.06, 0.09, 0.12, 0.15, 0.18, and 0.21. The results are analyzed with a view to indicating the variation of the aerodynamic characteristics with profile thickness for airfoils having a certain mean camber line form.

  5. Tests of N.A.C.A. airfoils in the variable-density wind tunnel Series 43 and 63

    NASA Technical Reports Server (NTRS)

    Jacobs, Eastman N; Pinkerton, Robert M

    1931-01-01

    This note is one of a series covering an investigation of a family of related airfoils. It gives in preliminary form the results obtained from tests in the N.A.C.A. Variable-Density Wind Tunnel of two groups of six airfoils each. One group, the 43 series, has a maximum mean camber of 4 per cent of the chord at a position 0.3 of the chord from the leading edge; the other group, the 63 series, has a maximum mean camber of 6 per cent of the chord at the same position. The members within each group differ only in maximum thickness, the maximum thickness/chord ratios being:0.06, 0.09, 0.12, 0.15, 0.18, and 0.21. The results are analyzed with a view to indicating the variation of the aerodynamic characteristics with profile thickness for airfoils having a certain mean camber line.

  6. A wind tunnel database using RIM

    NASA Technical Reports Server (NTRS)

    Wray, W. O., Jr.

    1984-01-01

    Engineering data base development which has become increasingly widespread to industry with the availability of data management systems is examined. A large data base was developed for wind tunnel data and related model test information, using RIM as the data base manager. The arrangement of the wind tunnel data into the proper schema for the most efficient database utilization is discussed. The FORTRAN interface program of RIM is used extensively in the loading phases of the data base and by the users. Several examples to illustrate how the Wind Tunnel Data base might be searched for specific data items and test information using RIM are presented.

  7. Wind tunnel pressurization and recovery system

    NASA Technical Reports Server (NTRS)

    Pejack, Edwin R.; Meick, Joseph; Ahmad, Adnan; Lateh, Nordin; Sadeq, Omar

    1988-01-01

    The high density, low toxicity characteristics of refrigerant-12 (dichlorofluoromethane) make it an ideal gas for wind tunnel testing. Present limitations on R-12 emissions, set to slow the rate of ozone deterioration, pose a difficult problem in recovery and handling of large quantities of R-12. This preliminary design is a possible solution to the problem of R-12 handling in wind tunnel testing. The design incorporates cold temperature condensation with secondary purification of the R-12/air mixture by adsorption. Also discussed is the use of Freon-22 as a suitable refrigerant for the 12 foot wind tunnel.

  8. Analysis of wind tunnel test results for a 9.39-per cent scale model of a VSTOL fighter/attack aircraft. Volume 2: Evaluation of prediction methodologies

    NASA Technical Reports Server (NTRS)

    Lummus, J. R.; Joyce, G. T.; Omalley, C. D.

    1980-01-01

    An evaluation of current prediction methodologies to estimate the aerodynamic uncertainties identified for the E205 configuration is presented. This evaluation was accomplished by comparing predicted and wind tunnel test data in three major categories: untrimmed longitudinal aerodynamics; trimmed longitudinal aerodynamics; and lateral-directional aerodynamic characteristics.

  9. 40 CFR 53.62 - Test procedure: Full wind tunnel test.

    Code of Federal Regulations, 2012 CFR

    2012-07-01

    ... PROGRAMS (CONTINUED) AMBIENT AIR MONITORING REFERENCE AND EQUIVALENT METHODS Procedures for Testing... penetration curves are then each numerically integrated with three idealized ambient particle size... test sampler, or portion thereof, as installed in the test section for testing, blocks no more than...

  10. 40 CFR 53.62 - Test procedure: Full wind tunnel test.

    Code of Federal Regulations, 2014 CFR

    2014-07-01

    ... PROGRAMS (CONTINUED) AMBIENT AIR MONITORING REFERENCE AND EQUIVALENT METHODS Procedures for Testing... penetration curves are then each numerically integrated with three idealized ambient particle size... test sampler, or portion thereof, as installed in the test section for testing, blocks no more than...

  11. 40 CFR 53.62 - Test procedure: Full wind tunnel test.

    Code of Federal Regulations, 2011 CFR

    2011-07-01

    ... PROGRAMS (CONTINUED) AMBIENT AIR MONITORING REFERENCE AND EQUIVALENT METHODS Procedures for Testing... penetration curves are then each numerically integrated with three idealized ambient particle size... test sampler, or portion thereof, as installed in the test section for testing, blocks no more than...

  12. 40 CFR 53.42 - Generation of test atmospheres for wind tunnel tests.

    Code of Federal Regulations, 2012 CFR

    2012-07-01

    ... 40 Protection of Environment 6 2012-07-01 2012-07-01 false Generation of test atmospheres for wind... Testing Performance Characteristics of Methods for PM10 § 53.42 Generation of test atmospheres for wind... of multiplets (doublets and triplets) in a test particle atmosphere shall not exceed 10 percent....

  13. Estimation of the Unsteady Aerodynamic Load on Space Shuttle External Tank Protuberances from a Component Wind Tunnel Test

    NASA Technical Reports Server (NTRS)

    Panda, Jayatana; Martin, Fred W.; Sutliff, Daniel L.

    2008-01-01

    At the wake of the Columbia (STS-107) accident it was decided to remove the Protuberance Aerodynamic Load (PAL) Ramp that was originally intended to protect various protuberances outside of the Space Shuttle External Tank from high buffet load induced by cross-flows at transonic speed. In order to establish the buffet load without the PAL ramp, a wind tunnel test was conducted where segments of the protuberances were instrumented with dynamic pressure transducers; and power-spectra of sectional lift and drag forces at various span-wise locations between two adjacent support brackets were measured under different cross flow angles, Mach number and other conditions. Additionally, frequency-dependent spatial correlations between the sectional forces were also established. The sectional forces were then adjusted by the correlation length to establish span-averaged spectra of normal and lateral forces that can be suitably "added" to various other unsteady forces encountered by the protuberance. This paper describes the methodology used for calculating the correlation-adjusted power spectrum of the buffet load. A second part of the paper describes wind-tunnel results on the difference in the buffet load on the protuberances with and without the PAL ramp. In general when the ramp height is the same as that of the protuberance height, such as that found on the liquid Oxygen part of the tank, the ramp is found to cause significant reduction of the unsteady aerodynamic load. However, on the liquid Hydrogen part of the tank, where the Oxygen feed-line is far larger in diameter than the height of the PAL ramp, little protection is found to be available to all but the Cable Tray.

  14. Wind tunnel buffet load measuring technique

    NASA Technical Reports Server (NTRS)

    Chang, C. S.; Ellison, A. M.

    1972-01-01

    Indirect force measurement technique estimates unsteady forces acting on elastic model during wind tunnel tests. Measurement of forces is practically insensitive to errors in aeroelastic scaling between model and full-scale structure, simplifying design, fabrication and dynamic calibration.

  15. On a new type of wind tunnel

    NASA Technical Reports Server (NTRS)

    Munk, Max

    1921-01-01

    Discussed here is a new type of wind tunnel, its advantages, the difficulties attendant upon its use, and the special methods required for its operation. The main difference between the new type of wind tunnel and the ones now in operation is the use of a different fluid. The idea is to diminish the effect of viscosity If air is compressed, it becomes a fluid with new properties - a fluid that is best suited for reliable and exact tests on models. When air is compressed, its density increases, but its viscosity does not. It is argued that the increase of pressure greatly increases the range and value of wind tunnel tests. Reynolds number, deductions from the Reynolds law, the causes of errors that result in differences between tests on models and actual flights, and the dimensions of a compressed air wind tunnel are covered.

  16. Mass spectrometric gas composition measurements associated with jet interaction tests in a high-enthalpy wind tunnel

    NASA Technical Reports Server (NTRS)

    Lewis, B. W.; Brown, K. G.; Wood, G. M., Jr.; Puster, R. L.; Paulin, P. A.; Fishel, C. E.; Ellerbe, D. A.

    1986-01-01

    Knowledge of test gas composition is important in wind-tunnel experiments measuring aerothermodynamic interactions. This paper describes measurements made by sampling the top of the test section during runs of the Langley 7-Inch High-Temperature Tunnel. The tests were conducted to determine the mixing of gas injected from a flat-plate model into a combustion-heated hypervelocity test stream and to monitor the CO2 produced in the combustion. The Mass Spectrometric (MS) measurements yield the mole fraction of N2 or He and CO2 reaching the sample inlets. The data obtained for several tunnel run conditions are related to the pressures measured in the tunnel test section and at the MS ionizer inlet. The apparent distributions of injected gas species and tunnel gas (CO2) are discussed relative to the sampling techniques. The measurements provided significant real-time data for the distribution of injected gases in the test section. The jet N2 diffused readily from the test stream, but the jet He was mostly entrained. The amounts of CO2 and Ar diffusing upward in the test section for several run conditions indicated the variability of the combustion-gas test-stream composition.

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

    NASA Astrophysics Data System (ADS)

    Marroquin, J.; Lemoine, P.

    1992-10-01

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

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

    NASA Astrophysics Data System (ADS)

    Marroquin, J.; Lemoine, P.

    1992-10-01

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

  19. 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.

  20. 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.

  1. An Investigation of the Low-speed Stability and Control Characteristics of Swept-forward and Swept-back Wing in the Ames 40- by 80-foot Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Mccormack, Gerald M; Stevens, Victor I , Jr

    1947-01-01

    An investigation has been made at large scale of the characteristics of highly swept wings. Data were obtained at several angles of sideslip on wings having angles of sweep of plus or minus 45 degrees, plus or minus 30 degrees, and 0 degrees. The airfoil sections of the wings varied from approximately NACA 0015 at the root to NACA 23009 at the tip. Each wing was investigated with flaps under flection, partial-span split flaps deflected 60 degrees, full-span split flaps defected 60 degrees and split-flap-type ailerons deflected plus or minus 15 degrees. Values of maximum lift were obtained at Reynolds numbers raging from 5.7 to 9.2 times 10 to the 6th power. In this report the summarized results are compared with the predictions made by use of the simplified theory for the effect of sweep and with existing small-scale data. The basic wind-tunnel results from which these summary data were taken are included in an appendix. The primary problems accompanying the use of weep as revealed by this investigation are the loss in maximum lift, the high effective dihedral, and the sharp reduction in lateral-control effectiveness. In general, simple theory enables good predictions to be made of the gross effects of sweep but further refinements are necessary to obtain the accuracy required for design purposes. In cases where comparisons can be made, the indications are that, as sweep increases, scale effects diminish and large-scale results approach small-scale results.

  2. Performance and test section flow characteristics of the National Full-Scale Aerodynamics Complex 80- by 120-Foot Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Zell, Peter T.

    1993-01-01

    Results from the performance and test section flow calibration of the 80- by 120-Foot Wind Tunnel are presented. Measurements indicating the 80- by 120-ft test section flow quality were obtained throughout the tunnel operational envelope and for atmospheric wind speeds up to approximately 20 knots. Tunnel performance characteristics and a dynamic pressure system calibration were also documented during the process of mapping the test section flow field. Experimental results indicate that the test section flow quality is relatively insensitive to dynamic pressure and the level of atmospheric winds experienced during the calibration. The dynamic pressure variation in the test section is within +/-75 percent of the average. The axial turbulence intensity is less than 0.5 percent up to the maximum test section speed of 100 knots, and the vertical and lateral flow angle variations are within +/-5 deg and +/-7 deg, respectively. Atmospheric winds were found to affect the pressure distribution in the test section only at high ratios of wind speed to test section speed.

  3. External aerodynamics of heavy ground vehicles: Computations and wind tunnel testing

    NASA Astrophysics Data System (ADS)

    Bayraktar, Ilhan

    Aerodynamic characteristics of a ground vehicle affect vehicle operation in many ways. Aerodynamic drag, lift and side forces have influence on fuel efficiency, vehicle top speed and acceleration performance. In addition, engine cooling, air conditioning, wind noise, visibility, stability and crosswind sensitivity are some other tasks for vehicle aerodynamics. All of these areas benefit from drag reduction and changing the lift force in favor of the operating conditions. This can be achieved by optimization of external body geometry and flow modification devices. Considering the latter, a thorough understanding of the airflow is a prerequisite. The present study aims to simulate the external flow field around a ground vehicle using a computational method. The model and the method are selected to be three dimensional and time-dependent. The Reynolds-averaged Navier Stokes equations are solved using a finite volume method. The Renormalization Group (RNG) k-epsilon model was elected for closure of the turbulent quantities. Initially, the aerodynamics of a generic bluff body is studied computationally and experimentally to demonstrate a number of relevant issues including the validation of the computational method. Experimental study was conducted at the Langley Full Scale Wind Tunnel using pressure probes and force measurement equipment. Experiments and computations are conducted on several geometric configurations. Results are compared in an attempt to validate the computational model for ground vehicle aerodynamics. Then, the external aerodynamics of a heavy truck is simulated using the validated computational fluid dynamics method, and the external flow is presented using computer visualization. Finally, to help the estimation of the error due to two commonly practiced engineering simplifications, a parametric study on the tires and the moving ground effect are conducted on full-scale tractor-trailer configuration. Force and pressure coefficients and velocity

  4. Large-Scale Wind Turbine Testing in the NASA 24.4m (80) by 36.6m(120) Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Zell, Peter T.; Imprexia, Cliff (Technical Monitor)

    2000-01-01

    The 80- by 120-Foot Wind Tunnel at NASA Ames Research Center in California provides a unique capability to test large-scale wind turbines under controlled conditions. This special capability is now available for domestic and foreign entities wishing to test large-scale wind turbines. The presentation will focus on facility capabilities to perform wind turbine tests and typical research objectives for this type of testing.

  5. Propulsion and airframe aerodynamic interactions of supersonic V/STOL configurations. Volume 2: Wind tunnel test force and moment data report

    NASA Technical Reports Server (NTRS)

    Zilz, D. E.

    1985-01-01

    A wind tunnel model of a supersonic V/STOL fighter configuration has been tested to measure the aerodynamic interaction effects which can result from geometrically close-coupled propulsion system/airframe components. The approach was to configure the model to represent two different test techniques. One was a conventional test technique composed of two test modes. In the Flow-Through mode, absolute configuration aerodynamics are measured, including inlet/airframe interactions. In the Jet-Effects mode, incremental nozzle/airframe interactions are measured. The other test technique is a propulsion simulator approach, where a sub-scale, externally powered engine is mounted in the model. This allows proper measurement of inlet/airframe and nozzle/airframe interactions simultaneously. This is Volume 2 of 2: Wind Tunnel Test Force and Moment Data Report.

  6. Method for Standardizing Sonic-Boom Model Pressure Signatures Measured at Several Wind-Tunnel Facilities

    NASA Technical Reports Server (NTRS)

    Mack, Robert J.

    2007-01-01

    Low-boom model pressure signatures are often measured at two or more wind-tunnel facilities. Preliminary measurements are made at small separation distances in a wind tunnel close at hand, and a second set of pressure signatures is measured at larger separation distances in a wind-tunnel facility with a larger test section. In this report, a method for correcting and standardizing the wind-tunnel-measured pressure signatures obtained in different wind tunnel facilities is presented and discussed.

  7. Preliminary wing model tests in the variable density wind tunnel of the National Advisory Committee for Aeronautics

    NASA Technical Reports Server (NTRS)

    Munk, Max M

    1926-01-01

    This report contains the results of a series of tests with three wing models. By changing the section of one of the models and painting the surface of another, the number of models tested was increased to five. The tests were made in order to obtain some general information on the air forces on wing sections at a high Reynolds number and in particular to make sure that the Reynolds number is really the important factor, and not other things like the roughness of the surface and the sharpness of the trailing edge. The few tests described in this report seem to indicate that the air forces at a high Reynolds number are not equivalent to respective air forces at a low Reynolds number (as in an ordinary atmospheric wind tunnel). The drag appears smaller at a high Reynolds number and the maximum lift is increased in some cases. The roughness of the surface and the sharpness of the trailing edge do not materially change the results, so that we feel confident that tests with systematic series of different wing sections will bring consistent results, important and highly useful to the designer.

  8. Acoustic Quality of the 40- by 80- Foot Wind Tunnel Test Section After Installation of a Deep Acoustic Lining

    NASA Technical Reports Server (NTRS)

    Soderman, Paul T.; Jaeger, Stephen M.; Hayes, Julie A.; Allen, Christopher S.

    2002-01-01

    A recessed, 42-inch deep acoustic lining has been designed and installed in the 40- by 80- Foot Wind Tunnel (40x80) test section to greatly improve the acoustic quality of the facility. This report describes the test section acoustic performance as determined by a detailed static calibration-all data were acquired without wind. Global measurements of sound decay from steady noise sources showed that the facility is suitable for acoustic studies of jet noise or similar randomly generated sound. The wall sound absorption, size of the facility, and averaging effects of wide band random noise all tend to minimize interference effects from wall reflections. The decay of white noise with distance was close to free field above 250 Hz. However, tonal sound data from propellers and fans, for example, will have an error band to be described that is caused by the sensitivity of tones to even weak interference. That error band could be minimized by use of directional instruments such as phased microphone arrays. Above 10 kHz, air absorption began to dominate the sound field in the large test section, reflections became weaker, and the test section tended toward an anechoic environment as frequency increased.

  9. Parametric Data from a Wind Tunnel Test on a Rocket-Based Combined-Cycle Engine Inlet

    NASA Technical Reports Server (NTRS)

    Fernandez, Rene; Trefny, Charles J.; Thomas, Scott R.; Bulman, Mel J.

    2001-01-01

    A 40-percent scale model of the inlet to a rocket-based combined-cycle (RBCC) engine was tested in the NASA Glenn Research Center 1- by 1-Foot Supersonic Wind Tunnel (SWT). The full-scale RBCC engine is scheduled for test in the Hypersonic Tunnel Facility (HTF) at NASA Glenn's Plum Brook Station at Mach 5 and 6. This engine will incorporate the configuration of this inlet model which achieved the best performance during the present experiment. The inlet test was conducted at Mach numbers of 4.0, 5.0, 5.5, and 6.0. The fixed-geometry inlet consists of an 8 deg.. forebody compression plate, boundary layer diverter, and two compressive struts located within 2 parallel sidewalls. These struts extend through the inlet, dividing the flowpath into three channels. Test parameters investigated included strut geometry, boundary layer ingestion, and Reynolds number (Re). Inlet axial pressure distributions and cross-sectional Pitot-pressure surveys at the base of the struts were measured at varying back-pressures. Inlet performance and starting data are presented. The inlet chosen for the RBCC engine self-started at all Mach numbers from 4 to 6. Pitot-pressure contours showed large flow nonuniformity on the body-side of the inlet. The inlet provided adequate pressure recovery and flow quality for the RBCC cycle even with the flow separation.

  10. Subsonic wind-tunnel tests of a trailing-cone device for calibrating aircraft static pressure systems

    NASA Technical Reports Server (NTRS)

    Jordan, F. L., Jr.; Ritchie, V. S.

    1973-01-01

    A trailing-cone device for calibrating aircraft static-pressure systems was tested in a transonic wind tunnel to investigate the pressure-sensing characteristics of the device including effects of several configuration changes. The tests were conducted at Mach numbers from 0.30 to 0.95 with Reynolds numbers from (0.9 x one million to 4.1 x one million per foot). The results of these tests indicated that the pressures sensed by the device changed slightly but consistently as the distance between the device pressure orifices and cone was varied from 4 to 10 cone diameters. Differences between such device-indicated pressures and free-stream static pressure were small, however, and corresponded to Mach number differences of less than 0.001 for device configurations with pressure orifices located 5 or 6 cone diameters ahead of the cone. Differences between device-indicated and free-stream static pressures were not greatly influenced by a protection skid at the downstream end of the pressure tube of the device nor by a 2-to-1 change in test Reynolds number.

  11. Investigation of correlation between full-scale and fifth-scale wind tunnel tests of a Bell helicopter Textron Model 222

    NASA Technical Reports Server (NTRS)

    Squires, P. K.

    1982-01-01

    Reasons for lack of correlation between data from a fifth-scale wind tunnel test of the Bell Helicopter Textron Model 222 and a full-scale test of the model 222 prototype in the NASA Ames 40-by 80-foot tunnel were investigated. This investigation centered around a carefully designed fifth-scale wind tunnel test of an accurately contoured model of the Model 222 prototype mounted on a replica of the full-scale mounting system. The improvement in correlation for drag characteristics in pitch and yaw with the fifth-scale model mounted on the replica system is shown. Interference between the model and mounting system was identified as a significant effect and was concluded to be a primary cause of the lack of correlation in the earlier tests.

  12. Wind Tunnel Aeroacoustic Tests of Six Airfoils for Use on Small Wind Turbines; Period of Performance: August 23, 2002 through March 31, 2004

    SciTech Connect

    Oerlemans, S.

    2004-08-01

    The U.S. Department of Energy, working through the National Renewable Energy Laboratory, is engaged in a comprehensive research effort to improve our understanding of wind turbine aeroacoustics. Quiet wind turbines are an inducement to widespread deployment, so the goal of NREL's aeroacoustic research is to develop tools that the U.S. wind industry can use in developing and deploying highly efficient, quiet wind turbines at low wind speed sites. NREL's National Wind Technology Center is implementing a multifaceted approach that includes wind tunnel tests, field tests, and theoretical analyses in direct support of low wind speed turbine development by its industry partners. To that end, wind tunnel aerodynamic tests and aeroacoustic tests have been performed on six airfoils that are candidates for use on small wind turbines. Results are documented in this report.

  13. Wind sensitivity studies of a non-return wind tunnel with a 216- by 432-mm (8.5- by 17.0-inches) test section, phase 2

    NASA Technical Reports Server (NTRS)

    Eckert, W. T.; Mort, K. W.; Piazza, J. E.

    1973-01-01

    The refinement of inlet and exit treatments were studied which would minimize the effect of external wind on the test-section flow quality of a nonreturn wind tunnel. The investigation was conducted in the Ames Research Center 40- by 80-foot Wind Tunnel which served as the wind source. Several inlets and two exits were tested at wind directions ranging from 0 to 180 degrees and at wind-to-test-section velocity ratios from zero to somewhat greater than one. For the best inlet configuration the flow quality was good, with a velocity deviation in each of the three component directions generally less. The loss in total pressure due to the inlet treatment was low: about 0.035 of the test-section dynamic pressure for the no-wind case.

  14. Wind-tunnel tests on a 3-dimensional fixed-geometry scramjet inlet at M = 2.30 to 4.60

    NASA Technical Reports Server (NTRS)

    Mueller, J. N.; Trexler, C. A.; Souders, S. W.

    1977-01-01

    Wind-tunnel tests were conducted on a baseline scramjet inlet model having fixed geometry and swept leading edges at M = 2.30, 2.96, 3.95, and 4.60 in the Langley unitary plan wind tunnel. The unit Reynolds number of the tests was held constant at 6.56 million per meter (2 million per foot). The objectives of the tests were to establish inlet performance and starting characteristics in the lower Mach number range of operation (less than M = 5). Surface pressures obtained on the inlet components are presented, along with the results of the internal flow surveys made at the throat and capture stations of the inlet. Contour plots of the inlet-flow-field parameters such as Mach numbers, pressure recovery, flow capture, local static and total pressure ratios at the survey stations are shown for the test Mach numbers.

  15. Wind-Tunnel Tests of a 1/5-Scale Semispan Model of the Republic XF-12 Horizontal Tail Surface

    NASA Technical Reports Server (NTRS)

    Denaci, H. G.

    1945-01-01

    Wind-tunnel tests of a 1/5-scale semispan model of the Republic XF-12 horizontal tail surface equipped with an internally balanced elevator were conducted in the 6- by 6-foot test section of the Langley stability tunnel. The tests included measurements of the aerodynamic characteristics of the horizontal tail with and without a beveled trailing edge and also included measurements of the tab characteristics. The variation of the aerodynamic characteristics with boundary-layer conditions and leakage in the internal-balance chambers, measurements of the boundary-layer displacement thickness near the elevator hinge axis, and pressure distributions at the mean geometric chord were also obtained. The results showed that the hinge-moment characteristics of the elevator were critical to boundary-layer conditions and internal-balance leakage. Increasing the boundary-layer displacement thickness by use of roughness strips reduced the rate of change of elevator hinge moments with tab deflection by about 20 percent. The present horizontal tail appears to be unsatisfactory for longitudinal stability with power on, however, an increase in horizontal-tail lift effectiveness should correct this difficulty. The maneuvering stick force per unit acceleration will be extremely critical to minor variations of the elevator hinge moments if the elevator is linked directly to the stick.

  16. Performance tests for the NASA Ames Research Center 20 cm x 40 cm oscillating flow wind tunnel

    NASA Technical Reports Server (NTRS)

    Cook, W. J.; Giddings, T. A.

    1984-01-01

    An evaluation is presented of initial tests conducted to assess the performance of the NASA Ames 20 cm x 40 cm oscillating flow wind tunnel. The features of the tunnel are described and two aspects of tunnel operation are discussed. The first is an assessment of the steady mainstream and boundary layer flows and the second deals with oscillating mainstream and boundary layer flows. Experimental results indicate that in steady flow the test section mainstream velocity is uniform in the flow direction and in cross section. The freestream turbulence intensity is about 0.2 percent. With minor exceptions the steady turbulent boundary layer generated on the top wall of the test section exhibits the characteristics of a zero pressure gradient turbulent boundary layer generated on a flat plate. The tunnel was designed to generate sinusoidal oscillating mainstream flows. Experiments confirm that the tunnel produces sinusoidal mainstream velocity variations for the range of frequencies (up to 15 Hz). The results of this study demonstrate that the tunnel essentially produces the flows that it was designed to produce.

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

    NASA Technical Reports Server (NTRS)

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

    1974-01-01

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

  18. Acoustic Data Processing and Transient Signal Analysis for the Hybrid Wing Body 14- by 22-Foot Subsonic Wind Tunnel Test

    NASA Technical Reports Server (NTRS)

    Bahr, Christopher J.; Brooks, Thomas F.; Humphreys, William M.; Spalt, Taylor B.; Stead, Daniel J.

    2014-01-01

    An advanced vehicle concept, the HWB N2A-EXTE aircraft design, was tested in NASA Langley's 14- by 22-Foot Subsonic Wind Tunnel to study its acoustic characteristics for var- ious propulsion system installation and airframe con gurations. A signi cant upgrade to existing data processing systems was implemented, with a focus on portability and a re- duction in turnaround time. These requirements were met by updating codes originally written for a cluster environment and transferring them to a local workstation while en- abling GPU computing. Post-test, additional processing of the time series was required to remove transient hydrodynamic gusts from some of the microphone time series. A novel automated procedure was developed to analyze and reject contaminated blocks of data, under the assumption that the desired acoustic signal of interest was a band-limited sta- tionary random process, and of lower variance than the hydrodynamic contamination. The procedure is shown to successfully identify and remove contaminated blocks of data and retain the desired acoustic signal. Additional corrections to the data, mainly background subtraction, shear layer refraction calculations, atmospheric attenuation and microphone directivity corrections, were all necessary for initial analysis and noise assessments. These were implemented for the post-processing of spectral data, and are shown to behave as expected.

  19. The optimum hypersonic wind tunnel

    NASA Technical Reports Server (NTRS)

    Trimmer, L. L.; Cary, A., Jr.; Voisinet, R. L. P.

    1986-01-01

    The capabilities of existing hypersonic wind tunnels in the U.S. are assessed to form a basis for recommendations for a new, costly facility which would provide data for modeling the hypervelocity aerodynamics envisioned for the new generation of aerospace vehicles now undergoing early studies. Attention is given to the regimes, both entry and aerodynamic, which the new vehicles will encounter, and the shortcomings of data generated for the Orbiter before flight are discussed. The features of foreign-gas, impulse, aeroballistic range, arc-heated and combustion-heated facilities are examined, noting that in any hypersonic wind tunnel the flow must be preheated to prevent liquefaction upon expansion in the test channel. The limitations of the existing facilities and the identification of the regimes which must be studied lead to a description of the characteristics of an optimum hypersonic wind tunnel, including the operations and productivity, the instrumentation, the nozzle design and the flow quality. Three different design approaches are described, each costing at least $100 million to achieve workability.

  20. A vegetation modeling concept for Building and Environmental Aerodynamics wind tunnel tests and its application in pollutant dispersion studies.

    PubMed

    Gromke, Christof

    2011-01-01

    A new vegetation modeling concept for Building and Environmental Aerodynamics wind tunnel investigations was developed. The modeling concept is based on fluid dynamical similarity aspects and allows the small-scale modeling of various kinds of vegetation, e.g. field crops, shrubs, hedges, single trees and forest stands. The applicability of the modeling concept was validated in wind tunnel pollutant dispersion studies. Avenue trees in urban street canyons were modeled and their implications on traffic pollutant dispersion were investigated. The dispersion experiments proved the modeling concept to be practicable for wind tunnel studies and suggested to provide reliable concentration results. Unfavorable effects of trees on pollutant dispersion and natural ventilation in street canyons were revealed. Increased traffic pollutant concentrations were found in comparison to the tree-free reference case.

  1. Flexibility in flight behaviour of barn swallows (Hirundo rustica) and house martins (Delichon urbica) tested in a wind tunnel.

    PubMed

    Bruderer, L; Liechti, F; Bilo, D

    2001-04-01

    The flight behaviour of barn swallows (Hirundo rustica) and house martins (Delichon urbica) was tested in a wind tunnel at 15 combinations of flight angles and speeds. In contrast to that of most other small passerines, the intermittent flight of hirundines rarely consists of regular patterns of flapping and rest phases. To vary mechanical power output, both species used intermittent flight, controlling the number of single, pulse-like wingbeats per unit time. House martins in descent tended to concentrate their wingbeats into bursts and performed true gliding flight during rest phases. Barn swallows mainly performed partial bounds during brief interruptions of upstrokes, which they progressively prolonged with decreasing flight angle. Thus, identification of distinct flapping phases to calculate wingbeat frequencies was not feasible. Instead, an effective wingbeat frequency for flight intervals of 20 s, including partial bounds, was introduced. The effective wingbeat frequencies of house martins (N=3) ranged from 2 to 10.5 s(-1), those of barn swallows (N=4) from 2.5 to 8.5 s(-1). In both hirundine species, effective wingbeat frequency was found to decrease almost linearly with decreasing flight angle. With changes in air speed, wingbeat frequency varied according to a U-shaped curve, suggesting a minimum power speed of roughly 9 m s(-1). The duration of the down- and upstrokes varied systematically depending on flight angle and air speed.

  2. Flow visualization of leading-edge vortex enhancement by spanwise blowing. [swept wings - wind tunnel stability tests

    NASA Technical Reports Server (NTRS)

    Erickson, G. E.; Campbell, J. F.

    1975-01-01

    Flow visualization studies were conducted in a small pilot wind tunnel to determine qualitative effects of blowing a discrete jet essentially parallel to the leading edge of a 45 deg-swept trapezoidal wing featuring leading- and trailing-edge flaps. Test parameters included wing angle-of-attack, jet momentum coefficient, leading- and trailing-edge flap deflections, and nozzle chordwise displacement. Results of this study indicate that blowing from a reflection plane over the wing enhances the leading-edge vortex and delays vortex bursting to higher angles-of-attack and greater span distances. Increased blowing rates decrease vortex size, growth rate, and vertical displacement above the wing surface at a given span station and also extend the spanwise effectiveness of lateral blowing. Deflection of a leading-edge flap delays the beneficial effects of spanwise blowing to higher angles-of-attack. Nozzle chordwise locations investigated for the wing with and without leading-edge flap deflection appear equally effective in enhancing the separated leading-edge flow.

  3. An investigation of drag reduction for tractor trailer vehicles with air deflector and boattail. [wind tunnel tests

    NASA Technical Reports Server (NTRS)

    Muirhead, V. U.

    1981-01-01

    A wind tunnel investigation was conducted to determine the influence of several physical variables on the aerodynamic drag of a trailer model. The physical variables included: a cab mounted wind deflector, boattail on trailer, flow vanes on trailer front, forced transition on trailer, and decreased gap between tractor and trailer. Tests were conducted at yaw angles (relative wind angles) of 0, 5, 10, 20, and 30 degrees and Reynolds numbers of 3.58 x 10 to the 5th power 6.12 x 10 to the 5th power based upon the equivalent diameter of the vehicles. The wind deflector on top of the cab produced a calculated reduction in fuel consumption of about 5 percent of the aerodynamic portion of the fuel budget for a wind speed of 15.3 km/hr (9.5 mph) over a wind angle range of 0 deg to 180 deg and for a vehicle speed of 88.5 km/hr (55 mph). The boattail produced a calculated 7 percent to 8 percent reduction in fuel consumption under the same conditions. The decrease in gap reduced the calculated fuel consumption by about 5 percent of the aerodynamic portion of the fuel budget.

  4. Two-Dimensional Bifurcated Inlet Variable Cowl Lip Test Completed in 10- by 10-Foot Supersonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Hoffman, T. R.

    2000-01-01

    Researchers at the NASA Glenn Research Center at Lewis Field successfully tested a variable cowl lip inlet at simulated takeoff conditions in Glenn s 10- by 10-Foot Supersonic Wind Tunnel (10x10 SWT) as part of the High-Speed Research Program. The test was a follow-on to the Two-Dimensional Bifurcated (2DB) Inlet/Engine test. At the takeoff condition for a High-Speed Civil Transport aircraft, the inlet must provide adequate airflow to the engine with an acceptable distortion level and high-pressure recovery. The test was conducted to study the effectiveness of installing two rotating lips on the 2DB Inlet cowls to increase mass flow rate and eliminate or reduce boundary layer flow separation near the lips. Hardware was mounted vertically in the test section so that it extended through the tunnel ceiling and that the 2DB Inlet was exposed to the atmosphere above the test section. The tunnel was configured in the aerodynamic mode, and exhausters were used to pump down the tunnel to vacuum levels and to provide a maximum flow rate of approximately 58 lb/sec. The test determined the (1) maximum flow in the 2DB Inlet for each variable cowl lip, (2) distortion level and pressure recovery for each lip configuration, (3) boundary layer conditions near variable lips inside the 2DB Inlet, (4) effects of a wing structure adjacent to the 2DB Inlet, and (5) effects of different 2DB Inlet exit configurations. It also employed flow visualization to generate enough qualitative data on variable lips to optimize the variable lip concept. This test was a collaborative effort between the Boeing Company and Glenn. Extensive inhouse support at Glenn contributed significantly to the progress and accomplishment of this test.

  5. Wind Tunnel Database Development using Modern Experiment Design and Multivariate Orthogonal Functions

    NASA Technical Reports Server (NTRS)

    Morelli, Eugene A.; DeLoach, Richard

    2003-01-01

    A wind tunnel experiment for characterizing the aerodynamic and propulsion forces and moments acting on a research model airplane is described. The model airplane called the Free-flying Airplane for Sub-scale Experimental Research (FASER), is a modified off-the-shelf radio-controlled model airplane, with 7 ft wingspan, a tractor propeller driven by an electric motor, and aerobatic capability. FASER was tested in the NASA Langley 12-foot Low-Speed Wind Tunnel, using a combination of traditional sweeps and modern experiment design. Power level was included as an independent variable in the wind tunnel test, to allow characterization of power effects on aerodynamic forces and moments. A modeling technique that employs multivariate orthogonal functions was used to develop accurate analytic models for the aerodynamic and propulsion force and moment coefficient dependencies from the wind tunnel data. Efficient methods for generating orthogonal modeling functions, expanding the orthogonal modeling functions in terms of ordinary polynomial functions, and analytical orthogonal blocking were developed and discussed. The resulting models comprise a set of smooth, differentiable functions for the non-dimensional aerodynamic force and moment coefficients in terms of ordinary polynomials in the independent variables, suitable for nonlinear aircraft simulation.

  6. Reducing the Effect of Transducer Mount Induced Noise (XMIN) on Aeroacoustic Wind Tunnel Testing Data with a New Transducer Mount Design

    NASA Technical Reports Server (NTRS)

    Herron, Andrew J.; Reed, Darren K.; Nance, Donald K.

    2015-01-01

    Characterization of flight vehicle unsteady aerodynamics is often studied via large scale wind tunnel testing. Boundary layer noise is measured by miniature pressure transducers installed in a model. Noise levels (2-5 dB ref. 20 µPa) can be induced when transducer is mounted out of flush with model outer surface. This effect must be minimized to accurately determine aerodynamically induced acoustic environments.

  7. 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.

  8. Full scale wind tunnel investigation of a bearingless main helicopter rotor. [Ames 40 by 80 foot wind tunnel test using the BO-105 helicopter

    NASA Technical Reports Server (NTRS)

    1980-01-01

    A stability test program was conducted to determine the effects of airspeed, collective pitch, rotor speed and shaft angle on stability and loads at speeds beyond that attained in the BMR/BO-105 flight test program. Loads and performance data were gathered at forward speeds up to 165 knots. The effect of cyclic pitch perturbations on rotor response was investigated at simulated level flight conditions. Two configuration variations were tested for their effect on stability. One variable was the control system stiffness. An axially softer pitch link was installed in place of the standard BO-105 pitch link. The second variation was the addition of elastomeric damper strips to increase the structural damping. The BMR was stable at all conditions tested. At fixed collective pitch, shaft angle and rotor speed, damping generally increased between hover and 60 knots, remained relatively constant from 60 to 90 knots, then decreased above 90 knots. Analytical predictions are in good agreement with test data up to 90 knots, but the trend of decreasing damping above 90 knots is contrary to the theory.

  9. Phase 2 and 3 wind tunnel tests of the J-97 powered, external augmentor V/STOL model. [conducted in Ames 40- by 80-foot wind tunnel

    NASA Technical Reports Server (NTRS)

    Garland, D. B.

    1980-01-01

    Modifications were made to the model to improve longitudinal acceleration capability during transition from hovering to wing borne flight. A rearward deflection of the fuselage augmentor thrust vector is shown to be beneficial in this regard. Other agmentor modifications were tested, notably the removal of both endplates, which improved acceleration performance at the higher transition speeds. The model tests again demonstrated minimal interference of the fuselage augmentor on aerodynamic lift. A flapped canard surface also shows negligible influence on the performance of the wing and of the fuselage augmentor.

  10. Flow direction measurement criteria and techniques planned for the 40- by 80-/80- x 120-foot wind tunnel integrated systems tests

    NASA Technical Reports Server (NTRS)

    Zell, P. T.; Hoffmann, J.; Sandlin, D. R.

    1985-01-01

    A study was performed in order to develop the criteria for the selection of flow direction indicators for use in the Integrated Systems Tests (ISTs) of the 40 by 80/80 by 120 Foot Wind Tunnel System. The problems, requirements, and limitations of flow direction measurement in the wind tunnel were investigated. The locations and types of flow direction measurements planned in the facility were discussed. A review of current methods of flow direction measurement was made and the most suitable technique for each location was chosen. A flow direction vane for each location was chosen. A flow direction vane that employs a Hall Effect Transducer was then developed and evaluated for application during the ISTs.

  11. Wind tunnel test of the 0.019 scale space shuttle integrated vehicle (model 14-OTS) in the CALSPAN 8-foot transonic wind tunnel (IA36), volume 2

    NASA Technical Reports Server (NTRS)

    Hardin, R. B.; Burrows, R. R.

    1975-01-01

    A test is presented which was performed to determine the effect of cold jet gas plumes generated from main propulsion system and solid rocket motor nozzles on: (1) six-component force and moment data, (2) wing static pressures, (3) wing hinge moment, (4) elevon hinge moment, (5) rudder hinge moment, and (6) orbiter MPS nozzle pressure loads. The effects of rudder deflection, nozzle gimbal angle, and plume size were also obtained.

  12. Eliminating Wind Tunnel Flow Breakdown

    NASA Technical Reports Server (NTRS)

    Hackett, J. E.

    1983-01-01

    Undesirable vortexes near floor in small wind tunnels suppressed by simple device that alters flow pattern there. Air is injected along floor and interacts with backflow from wind-tunnel model. Results in smoother, more correct air-flow and to more-reliable wind-tunnel data.

  13. Quiet wind tunnel

    NASA Technical Reports Server (NTRS)

    Howard, P. W.; Schutzenhofer, L. A.

    1978-01-01

    Simple and inexpensive technique suppresses background noise generated by pores in wind tunnel wall lining and makes aerodynamic data more accurate and reliable. Porous walls are covered with wire-mesh screen. Screen offers smoother surface to airflow and damps vortexes and resonance caused by wall perforations; yet it provides enough open area for perforations to cancel shock waves generated by model.

  14. Assessment of analytical and experimental techniques utilized in conducting plume technology tests 575 and 593. [exhaust flow simulation (wind tunnel tests) of scale model Space Shuttle Orbiter

    NASA Technical Reports Server (NTRS)

    Baker, L. R.; Sulyma, P. R.; Tevepaugh, J. A.; Penny, M. M.

    1976-01-01

    Since exhaust plumes affect vehicle base environment (pressure and heat loads) and the orbiter vehicle aerodynamic control surface effectiveness, an intensive program involving detailed analytical and experimental investigations of the exhaust plume/vehicle interaction was undertaken as a pertinent part of the overall space shuttle development program. The program, called the Plume Technology program, has as its objective the determination of the criteria for simulating rocket engine (in particular, space shuttle propulsion system) plume-induced aerodynamic effects in a wind tunnel environment. The comprehensive experimental program was conducted using test facilities at NASA's Marshall Space Flight Center and Ames Research Center. A post-test examination of some of the experimental results obtained from NASA-MSFC's 14 x 14-inch trisonic wind tunnel is presented. A description is given of the test facility, simulant gas supply system, nozzle hardware, test procedure and test matrix. Analysis of exhaust plume flow fields and comparison of analytical and experimental exhaust plume data are presented.

  15. Overview of the 1989 Wind Tunnel Calibration Workshop

    NASA Technical Reports Server (NTRS)

    Henderson, Arthur, Jr.; Mckinney, L. Wayne

    1993-01-01

    An overview of the 1989 Wind Tunnel Calibration Workshop held at NASA LaRC in Hampton, VA on 19-20 Apr. 1989 is presented. The purpose of the Workshop was to explore wind tunnel calibration requirements as they relate to test quality and data accuracy, with the ultimate goal of developing wind tunnel calibration requirements for the major NASA wind tunnels at ARC, LaRC, and LeRC. The two sessions addressed the following topics: (1) what constitutes a properly calibrated wind tunnel; and (2) the status of calibration of NASA's major wind tunnels. The most significant contributions to the stated goals are highlighted, and the consensus of the Workshop's conclusions and recommendations regarding formulation and implementation of that goal are presented.

  16. Jet V/STOL wind-tunnel simulation and groundplane effects

    NASA Technical Reports Server (NTRS)

    Margason, R. J.

    1981-01-01

    Low-speed wind-tunnel testing of V/STOL aircraft concepts to determine the aerodynamic-propulsion interaction effects during the transition between hover and wingborne flight is a necessary step in the development cycle of this type of aircraft. Powered models are normally used to determine the aerodynamic performance characteristics. This paper examines some of the pretest preparation necessary to define the objectives of an appropriate investigation. Several factors which influence the selection of the model concept and the engine simulator are discussed. In addition, some of the test techniques important for this class of aircraft model are examined. Finally, the paper reviews some of the wind-tunnel wall effects important to this type of aircraft testing with special emphasis on groundplane effects.

  17. Wind-tunnel measurements of wing-canard interference and a comparison with various theories

    NASA Technical Reports Server (NTRS)

    Feistel, T. W.; Corsiglia, V. R.; Levin, D. B.

    1981-01-01

    Wind-tunnel tests and analyses of the aerodynamics of wing-canard combinations for low speed applications are presented. Systematic tests are conducted in a 7 x 10 wind tunnel to explore various combinations of wing-canard vertical and horizontal positioning. The goals of the tests are (1) to investigate potential improved stalling characteristics over conventional tail-aft configurations, (2) to investigate the existence of a lift coefficient advantage, and (3) to determine induced drag levels. The measurements obtained are compared with calculations made using the Prandtl-Munk theory, and with a vortex-lattice panel code. Results indicate that the panel code gives excellent results for lift and induced drag at moderate lift coefficient, whereas Prandtl-Munk theory gives conservative results for induced drag. The application is a light transport aircraft used for short-haul operations.

  18. Acoustic Modifications of the Ames 40x80 Foot Wind Tunnel and Test Techniques for High-Speed Research Model Testing

    NASA Technical Reports Server (NTRS)

    Soderman, Paul T.; Olson, Larry (Technical Monitor)

    1995-01-01

    The NFAC 40- by 80- Foot Wind Tunnel at Ames is being refurbished with a new, deep acoustic lining in the test section which will make the facility nearly anechoic over a large frequency range. The modification history, key elements, and schedule will be discussed. Design features and expected performance gains will be described. Background noise reductions will be summarized. Improvements in aeroacoustic research techniques have been developed and used recently at NFAC on several wind tunnel tests of High Speed Research models. Research on quiet inflow microphones and struts will be described. The Acoustic Survey Apparatus in the 40x80 will be illustrated. A special intensity probe was tested for source localization. Multi-channel, high speed digital data acquisition is now used for acoustics. And most important, phased microphone arrays have been developed and tested which have proven to be very powerful for source identification and increased signal-to-noise ratio. Use of these tools for the HEAT model will be illustrated. In addition, an acoustically absorbent symmetry plane was built to satisfy the HEAT semispan aerodynamic and acoustic requirements. Acoustic performance of that symmetry plane will be shown.

  19. Wind tunnel tests of the 0.010-scale space shuttle integrated vehicle (model 52-QT) in the NASA/Ames 3.5-foot hypersonic wind tunnel (IA18)

    NASA Technical Reports Server (NTRS)

    Esparza, V.; Chee, E.; Stone, J.; Mellenthin, J. A.

    1975-01-01

    Experimental aerodynamic investigations were conducted in the NASA/Ames Research Center 3.5-foot hypersonic wind tunnel on an 0.010-scale model of the space shuttle integrated vehicle consisting of an orbiter and external tank. The basic hypersonic stability characteristics of the orbiter attached rigidly to the external tank and the basic hypersonic stability characteristics of external tank alone simulating RTLS abort conditions were evaluated. The integrated vehicle was tested at angles of attack from- 8 deg through +30 deg and angles of sideslip of- 8 deg through +8 deg at fixed angles of attack of -4 deg, 0 deg, and +4 deg. A maximum angle of attack range of +15 deg through +40 deg was obtained for this configuration, at Mach number 7.3, for one run only. External tank alone testing was conducted at angles of attack from +8 deg through -30 deg and angles of sideslip of -8 deg at fixed angles of attack of -4 deg, 0 deg and +4 deg. Six-component force data and static base pressures were recorded during the test.

  20. Wind tunnel test of the 0.015-scale Rockwell International space shuttle vehicle orbiter in the Ames 6 by 6 foot supersonic wind tunnel. [to determine longitudinal and lateral-directional characteristics

    NASA Technical Reports Server (NTRS)

    Milam, M. D.; Dziubala, T. J.

    1973-01-01

    Experimental investigations were performed in a 6- by 6-Foot Supersonic wind tunnel on a 0.015-scale model of the Rockwell International space shuttle vehicle (SSV) 2A orbiter. The purpose of the test was to investigate the longitudinal and lateral-directional characteristics of the vehicle. In addition, hinge moments were measured on the rudder and elevons. Buffet onset was investigated using wing trailing edge pressures and a strain gauge instrumented panel mounted in the wing. The model was tested through a Mach range from 0.6 to 2.0 at a constant unit Reynolds number of 2.5 million. Pitch runs were made at angles of attack from minus 2 deg to +26 deg with beta = 0 deg and 5 deg; yaw runs were made in the range from minus 5 deg to 10 deg of sideslip at angles of attack of 0 deg and 10 deg. Static pressures were measured at the fuselage base and the trailing edges of the wing and rudder. Boundary layer transition was fixed for some runs using distributed roughness strips.