Multi-Nozzle Base Flow Model in the 10- by 10-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1964-02-21

Researchers check the setup of a multi-nozzle base flow model in the 10- by 10-Foot Supersonic Wind Tunnel at the National Aeronautics and Space Administration (NASA) Lewis Research Center. NASA researchers were struggling to understand the complex flow phenomena resulting from the use of multiple rocket engines. Robert Wasko and Theodore Cover of the Advanced Development and Evaluation Division’s analysis and operations sections conducted a set of tests in the 10- by 10 tunnel to further understand the flow issues. The Lewis researchers studied four and five-nozzle configurations in the 10- by 10 at simulated altitudes from 60,000 to 200,000 feet. The nozzles were gimbaled during some of the test runs to simulate steering. The flow field for the four-nozzle clusters was surveyed in the center and the lateral areas between the nozzles, whereas the five-nozzle cluster was surveyed in the lateral area only.

Gemini Model in the 10- by 10-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1962-09-21

A researcher at the National Aeronautics and Space Administration (NASA) Lewis Research Center examines a small-scale model of the Gemini capsule in the 10- by 10-Foot Supersonic Wind Tunnel test section. Gemini was added to NASA’s manned space program after its predecessor, Mercury, and its antecedent, Apollo, were already established. Gemini was a transitional mission designed provide the astronauts with practice docking with other spacecraft and withstanding durations in space up to two weeks. The program was officially announced on December 7, 1961, but planning began in mid-1959. It was named Gemini after the zodiac twins because of the spacecraft’s two passenger capacity. The Gemini Program was the first program to start at the new Manned Spacecraft Center in Houston, now the Johnson Space Center. Unlike Mercury and Apollo, Lewis had very little involvement with the Gemini Program. This model was tested in the 10- by 10 tunnel for several weeks in September 1962. Lewis began managing the Agena second-stage rocket program shortly after this photograph was taken. Agenas were used to launch a variety of spacecraft and satellites in the 1960s. They were also used on several Gemini missions to provide targets for the astronauts to practice their rendezvous maneuvers. Gemini had two unmanned and ten manned flights in 1965 and 1966. These yielded the first spacewalks, long-duration space missions, first onboard computer, docking with a second spacecraft, and rendezvous maneuvers.

Pressure distributions obtained on a 0.10-scale model of the Space Shuttle Orbiter's forebody in the Ames Unitary Plan Wind Tunnel

NASA Technical Reports Server (NTRS)

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

1986-01-01

Pressure distribution test data obtained on a 0.10-scale model of the forward fuselage of the Space Shuttle Orbiter are presented without analysis. The tests were completed in the Ames Unitary Wind Tunnel (UPWT). The UPWT tests were conducted in two different test sections operating in the continuous mode, the 8 x 7 feet and 9 x 7 feet test sections. Each test section has its own Mach number range, 1.6 to 2.5 and 2.5 to 3.5 for the 9 x 7 feet and 8 x 7 feet test section, respectively. The test Reynolds number ranged from 1.6 to 2.5 x 10 to the 6th power ft and 0.6 to 2.0 x 10 to the 6th power ft, respectively. The tests were conducted in support of the development of the Shuttle Entry Air Data System (SEADS). In addition to modeling the 20 SEADS orifices, the wind-tunnel model was also instrumented with orifices to match Development Flight Instrumentation (DFI) port locations that existed on the Space Shuttle Columbia (OV-102) during the Orbiter Flight test program. This DFI simulation has provided a means for comparisons between reentry flight pressure data and wind-tunnel and computational data.

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.

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

NASA Technical Reports Server (NTRS)

Gilkey, R. D.

1979-01-01

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

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

Experimental investigation of the subsonic high-altitude operation of the NASA Lewis 10- by 10-foot supersonic wind tunnel

NASA Technical Reports Server (NTRS)

Hughes, Christopher E.; Jeracki, Robert J.

1988-01-01

An experimental investigation was conducted in the NASA Lewis 10- by 10-Foot Supersonic Wind Tunnel during subsonic tunnel operation in the aerodynamic cycle to determine the test section flow characteristics near the Advanced Turboprop Project propeller model plane of rotation. The investigation used an eight-probe pitot static flow survey rake to measure total and static pressures at two locations in the wind tunnel: the test section and the bellmouth section (upstream of the two-dimensional flexible-wall nozzle). A cone angularity probe was used to measure any flow angularity in the test section. The evaluation was conducted at tunnel Mach numbers from 0.10 to 0.35 and at three operating altitudes from 2,000 to 50,000 ft. which correspond to tunnel reference total pressures from 1960 to 245 psfa, respectively. The results of this experimental investigation indicate a total-pressure loss area in the center of the test section and a static-pressure gradient from the test section centerline to the wall. These total and static pressure differences were observed at all tunnel operating altitudes and diminished at lower tunnel velocities. The total-pressure loss area was also found in the bellmouth section, which indicates that the loss mechanism is not the tunnel flexible-wall nozzle. The flow in the test section is essentially axial since very small flow angles were measured. The results also indicate that a correction to the tunnel total and static pressures must be applied in order to determine accurate freestream conditions at the test section centerline.

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.

Aerial View Of The Site From The 40x80 Foot Wind Tunnel At Nasa Ames Research Center.

NASA Image and Video Library

1943-03-12

(03/12/1943) Aerial view of the site from the 40x80 wind tunnel At NASA Ames Research Center. Site includes the 16 foot and 7x10 wind tunnels in the background. Building 200 also under construction. Framing for the drive fans of the 40x80 in scene.

Space Shuttle Model in the 10- by 10-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1975-07-21

Ken Baskin, an engineer from the Facilities and Engineering Branch at the National Aeronautics and Space Administration’s (NASA) Lewis Research Center checks a complete 2.25-scale model of the shuttle in the 10- by 10-Foot Supersonic Wind Tunnel. Baskin’s space shuttle project began in July 1976 during the run-up to the shuttle’s first lift-off scheduled for 1979. The space shuttle was expected to experience multifaceted heating and pressure distributions during the first and second stages of its launch. Rockwell International engineers needed to understand these issues in order to design proper thermal protection. The 10- by 10 tests evaluated the base heating and pressure. The test’s specific objectives were to measure heat transfer and pressure distributions around the orbiter’s external tank and solid rocket booster afterbody caused by rocket exhaust recirculation and impingement, to measure the heat transfer and pressure distributions due to rocket exhaust-induced flow separation, and determine gas recovery temperatures using gas temperature probes and heated model base components. The shuttle model’s main engines and solid rockets were fired during the tests, then just the main engines in an effort to simulate a launch. The researchers conducted 163 runs in the 10- by 10 during the test program.

Test-section noise of the Ames 7 by 10-foot wind tunnel no. 1

NASA Technical Reports Server (NTRS)

Soderman, P. T.

1976-01-01

An investigation was made of the test-section noise levels at various wind speeds in the Ames 7- by 10-Foot Wind Tunnel No. 1. No model was in the test section. Results showed that aerodynamic noise from various struts used to monitor flow conditions in the test section dominated the wind-tunnel background noise over much of the frequency spectrum. A tapered microphone stand with a thin trailing edge generated less noise than did a constant-chord strut with a blunt trailing edge. Noise from small holes in the test-section walls was insignificant.

Overview of X-38 Hypersonic Aerothermodynamic Wind Tunnel Data and Comparison with Numerical Results

NASA Technical Reports Server (NTRS)

Campbell, C.; Caram, J.; Berry, S.; Horvath, T.; Merski, N.; Loomis, M.; Venkatapathy, E.

2004-01-01

A NASA team of engineers has been organized to design a crew return vehicle for returning International Space Station crew members from orbit. The hypersonic aerothermodynamic characteristics of the X-23/X-24A derived X-38 crew return vehicle are being evaluated in various wind tunnels in support of this effort. Aerothermodynamic data from two NASA hypersonic tunnels at Mach 6 and Mach 10 has been obtained with cast ceramic models and a thermographic phosphorus digital imaging system. General windward surface heating features are described based on experimental surface heating images and surface oil flow patterns for the nominal hypersonic aerodynamic orientation. Body flap reattachment heating levels are examined. Computational Fluid Dynamics tools have been applied at the appropriate wind tunnel conditions to make comparisons with this data.

Two-Dimensional Bifurcated Inlet/Engine Tests Completed in 10- by 10-Foot Supersonic Wind Tunnel

NASA Technical Reports Server (NTRS)

Saunders, John D.

1999-01-01

A Two-Dimensional Bifurcated (2DB) Inlet was successfully tested in NASA Lewis Research Center s 10- by 10-Foot Supersonic Wind Tunnel. These tests were the culmination of a collaborative effort between the Boeing Company, General Electric, Pratt & Whitney, and Lewis. Extensive support in-house at Lewis contributed significantly to the progress and accomplishment of this test. The results, which met or exceeded many of the High-Speed Research (HSR) Program goals, were used to revise system studies within the HSR Program. The HSR Program is focused on developing low-noise, low-polluting, high-efficiency supersonic commercial aircraft. A supersonic inlet is an important component of an efficient, low-noise vehicle.

MHz-Rate NO PLIF Imaging in a Mach 10 Hypersonic Wind Tunnel

NASA Technical Reports Server (NTRS)

Jiang, N.; Webster, M.; Lempert, Walter R.; Miller, J. D.; Meyer, T. R.; Danehy, Paul M.

2010-01-01

NO PLIF imaging at repetition rates as high as 1 MHz is demonstrated in the NASA Langley 31 inch Mach 10 hypersonic wind tunnel. Approximately two hundred time correlated image sequences, of between ten and twenty individual frames, were obtained over eight days of wind tunnel testing spanning two entries in March and September of 2009. The majority of the image sequences were obtained from the boundary layer of a 20 flat plate model, in which transition was induced using a variety of cylindrical and triangular shaped protuberances. The high speed image sequences captured a variety of laminar and transitional flow phenomena, ranging from mostly laminar flow, typically at lower Reynolds number and/or in the near wall region of the model, to highly transitional flow in which the temporal evolution and progression of characteristic streak instabilities and/or corkscrew-shaped vortices could be clearly identified. A series of image sequences were also obtained from a 20 compression ramp at a 10 angle of attack in which the temporal dynamics of the characteristic separated flow was captured in a time correlated manner.

Aeroacoustic research in wind tunnels: A status report

NASA Technical Reports Server (NTRS)

Bender, J.; Arndt, R. E. A.

1973-01-01

The increasing attention given to aerodynamically generated noise brings into focus the need for quality experimental research in this area. To meet this need several specialized anechoic wind tunnels have been constructed. In many cases, however, budgetary constraints and the like make it desirable to use conventional wind tunnels for this work. Three basic problems are inherent in conventional facilities: (1) high background noise, (2) strong frequency dependent reverberation effects, and (3) unique instrumentation problems. The known acoustic characteristics of several conventional wind tunnels are evaluated and data obtained in a smaller 4- x 5-foot wind tunnel which is convertible from a closed jet to an open jet mode are presented. The data from these tunnels serve as a guideline for proposed modifications to a 7- x 10-foot wind tunnel. Consideration is given to acoustic treatment in several different portions of the wind tunnel.

GRC-11-02-17-WindTunnel-9x15-001

NASA Image and Video Library

2017-11-02

The Aerosciences Evaluation and Test Capabilities (AETC) Portfolio implemented the Capability Challenge to “Reduce Background Noise Levels for Engine Efficiency Measurements at the NASA Glenn 9x15 Low Speed Wind Tunnel”. The 9x15 Low Speed Wind Tunnel Acoustic Improvements animation documents the acoustic modifications being made to the 9x15 leg of the wind tunnel to reduce background noise levels. A brief history of the 9x15, research testing performed in the wind tunnel, the need to reduce background noise, and the five state of the art acoustic design modifications are documented in the animation. The expected noise reduction is presented audibly and the resulting benefit to NASA is also defined.

9x15 Low Speed Wind Tunnel Improvements Update

NASA Technical Reports Server (NTRS)

Stephens, David

2017-01-01

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 9x15 LSWT was designed for performance testing of VSTOL 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 2018.

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

NASA Technical Reports Server (NTRS)

Craig, J. E.

1981-01-01

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

Vertical/Short Takeoff and Landing Model in the 10- by 10-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1979-05-21

A technician checks a 0.25-scale engine model of a Vought Corporation V-530 engine in the test section of the 10- by 10-Foot Supersonic Wind Tunnel at the National Aeronautics and Space Administration (NASA) Lewis Research Center. Vought created a low-drag tandem-fan Vertical/Short and Takeoff and Landing (V/STOL) engine in the mid-1970s, designated as the V-530. The first fan on the tandem-fan engine was supplied with air through a traditional subsonic inlet, seen on the lower front of the engine. The air was exhausted through the nacelle during normal flight and directed down during takeoffs. The rear fan was supplied by the oval-shaped top inlet during all phases of the flight. The second fan exhausted its air through a rear vectorable nozzle. NASA Lewis and Vought partnered in the late 1970s to collect an array of inlet and nozzle design information on the tandem fan engines for the Navy. Vought created this .25-scale model of the V-530 for extensive testing in Lewis' 10- by 10-foot tunnel. During an early series of tests, the front fan was covered, and a turbofan simulator was used to supply air to the rear fan. The researchers then analyzed the performance of only the front fan inlet. During the final series of tests, the flow from the front fan was used to supply airflow to the rear fan. The researchers studied the inlet's recovery, distortion, and angle-of-attack limits over various flight conditions.

1/50 Scale Model Of The 80x120 Foot Wind Tunnel Model (NFAC) In The Test Section Of The 40x80 Wind Tunnel.

NASA Image and Video Library

1996-06-27

(03/12/1976) 1/50 scale model of the 80x120 foot wind tunnel model (NFAC) in the test section of the 40x80 wind tunnel. Model viewed from the west, mounted on a rotating ground board designed for this test. Ramp leading to ground board includes a generic building placed in front of the 80x120 inlet.

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.

9x15 Low Speed Wind Tunnel Acoustic Improvements

NASA Technical Reports Server (NTRS)

Stark, David; 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 VSTOL aircraft models, but with the addition of the current acoustic treatment in 1986 the tunnel has been used principally for acoustic and performance testing of aircraft propulsions systems. The present document describes an anticipated acoustic upgrade to be completed in 2017.

Technician Works on a Shuttle Model in the 10- by 10-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1977-02-21

A technician prepares a 2.25 percent scale model of the space shuttle for a base heat study in the 10- by 10-Foot Supersonic Wind Tunnel at the National Aeronautics and Space Administration (NASA) Lewis Research Center. This space shuttle project, begun here in July 1976, was aimed at evaluating base heating and pressure prior to the Shuttle’s first lift-off scheduled for 1979. The space shuttle was expected to experience multifaceted heating and pressure distributions during the first and second stages of its launch. Engineers needed to understand these issues in order to design proper thermal protection. The test’s specific objectives were to measure the heat transfer and pressure distributions around the orbiter’s external tank and solid rocket afterbody caused by rocket exhaust recirculation and impingement, to measure the heat transfer and pressure distributions caused by rocket exhaust-induced separation, and determine gas recovery temperatures using gas temperature probes and heated base components. The shuttle model’s main engines and solid rockets were first fired and then just the main engines to simulate a launch during the testing. Lewis researchers conducted 163 runs in the 10- by 10 during the test program.

Field and laboratory comparison of PM10 instruments in high winds

NASA Astrophysics Data System (ADS)

Sharratt, Brenton; Pi, Huawei

2018-06-01

Instruments capable of measuring PM10 (particulate matter ≤10 μm in aerodynamic diameter) concentrations may vary in performance as a result of different technologies utilized in measuring PM10. Therefore, the performance of five instruments capable of measuring PM10 concentrations above eroding soil surfaces was tested during high wind events at field sites in the Columbia Plateau and inside a wind tunnel. Comparisons among the Big Spring Number Eight (BSNE) sampler, DustTrak monitor, E-sampler, High-Volume sampler, and Tapered Element Oscillating Microbalance (TEOM) monitor were made at field sites during nine wind erosion events and inside a wind tunnel at two wind speeds (7 and 12 m s-1) and two ambient PM10 concentrations (2 and 50 mg m-3). PM10 concentrations were similar for the High-Volume sampler and TEOM monitor as well as for the BSNE samplers and DustTrak monitors but higher for the High-Volume sampler and TEOM monitor than the E-sampler during field erosion events. Based upon wind tunnel experiments, the TEOM monitor measured the highest PM10 concentration while the DustTrak monitor typically measured the lowest PM10 concentration as compared with other instruments. In addition, PM10 concentration appeared to lower for all instruments at a wind speed of 12 as compared with 7 m s-1 inside the wind tunnel. Differences in the performance of instruments in measuring PM10 concentration poses risks in comparing PM10 concentration among different instrument types or using multiple instrument types to jointly measure concentrations in the field or laboratory or even the same instrument type subject to different wind speeds.

PM10 emission efficiency for agricultural soils: Comparing a wind tunnel, a dust generator, and the open-air plot

NASA Astrophysics Data System (ADS)

Avecilla, Fernando; Panebianco, Juan E.; Mendez, Mariano J.; Buschiazzo, Daniel E.

2018-06-01

The PM10 emission efficiency of soils has been determined through different methods. Although these methods imply important physical differences, their outputs have never been compared. In the present study the PM10 emission efficiency was determined for soils through a wide range of textures, using three typical methodologies: a rotary-chamber dust generator (EDG), a laboratory wind tunnel on a prepared soil bed, and field measurements on an experimental plot. Statistically significant linear correlation was found (p < 0.05) between the PM10 emission efficiency obtained from the EDG and wind tunnel experiments. A significant linear correlation (p < 0.05) was also found between the PM10 emission efficiency determined both with the wind tunnel and the EDG, and a soil texture index (%sand + %silt)/(%clay + %organic matter) that reflects the effect of texture on the cohesion of the aggregates. Soils with higher sand content showed proportionally less emission efficiency than fine-textured, aggregated soils. This indicated that both methodologies were able to detect similar trends regarding the correlation between the soil texture and the PM10 emission. The trends attributed to soil texture were also verified for two contrasting soils under field conditions. However, differing conditions during the laboratory-scale and the field-scale experiments produced significant differences in the magnitude of the emission efficiency values. The causes of these differences are discussed within the paper. Despite these differences, the results suggest that standardized laboratory and wind tunnel procedures are promissory methods, which could be calibrated in the future to obtain results comparable to field values, essentially through adjusting the simulation time. However, more studies are needed to extrapolate correctly these values to field-scale conditions.

A76-0634. 1/50 Scale Model Of The 80X120 Foot Wind Tunnel Model (Nfac) In The Test Section Of The 40X80 Foot Wind Tunnel.

NASA Image and Video Library

1996-06-27

(03/12/1976) 1/50 scale model of the 80x120 foot wind tunnel model (NFAC) in the test section of the 40x80 foot wind tunnel. Model mounted on a rotating ground board designed for this test, viewed from the west, oriented for North wind.

How Efficient is Vegetation in Reducing Wind Erosion and Emission of Health-Threatening Fine Dust PM10? - A Wind Tunnel Approach

NASA Astrophysics Data System (ADS)

Burri, K.; Graf, F.

2009-04-01

World wide, wind erosion and desertification are most alarming processes of environmental degradation. Not only do they cause tremendous losses of fertile soil, but they also seriously affect human health. Pulmonary tuberculosis (silicosis) is one of the major diseases that have been linked to mineral fine dust (PM10) in the atmosphere. It is widely accepted that the re-establishment of an intact vegetation cover is the most effective measure against wind erosion. However, despite numerous investigations, the mechanisms responsible for the protective effect of vegetation are still not completely understood. Since the phenomenon involves highly variable interactions between soil, plants and atmosphere, it is particularly difficult to quantify the efficiency of vegetation in reducing wind erosion. As an alternative to field investigations, wind tunnel experiments offer the advantage to control specific parameters within this highly complex system. In this study, a series of wind tunnel experiments was performed including measurements of sediment transport and PM10 emission in differently dense grass canopies of Lolium perenne (91, 24, 5 and 0 plants per square meter). The novelty of the present wind tunnel study is the use of living plants instead of artificial imitations or dead plant parts. Although more and more sophisticated imitations of vegetation have been used in recent studies, the behaviour of living plants is likely to differ significantly. Coloured quartz sand was used for visualizing sand erosion and deposition patterns. The vertical profiles of aeolian sediment flux were analysed with a stackable sediment sampler composed of 60 collecting boxes, each with a height of 1 cm. The results of this study confirm that both sediment transport and PM10 emission strongly decrease with increasing plant cover. The protective effect of the plants was found to be linked to characteristic changes in the vertical profile of aeolian sediment flux and to specific spatial

1/50 Scale Model Of The 80X120 Foot Wind Tunnel Model (NFAC) In The Test Section Of The 40X80 Wind Tunnel At Nasa Ames.

NASA Image and Video Library

1976-03-12

(03/12/1976) Overhead view of 1/50 scale model of the 80x120 foot wind tunnel model (NFAC) in the test section of the 40x80 wind tunnel at NASA Ames. Model mounted on a rotating ground board designed for this test.

Comparison of Flight Measured, Predicted and Wind Tunnel Measured Winglet Characteristics on a KC-135 Aircraft

NASA Technical Reports Server (NTRS)

Dodson, R. O., Jr.

1982-01-01

One of the objectives of the KC-135 Winglet Flight Research and Demonstration Program was to obtain experimental flight test data to verify the theoretical and wind tunnel winglet aerodynamic performance prediction methods. Good agreement between analytic, wind tunnel and flight test performance was obtained when the known differences between the tests and analyses were accounted for. The flight test measured fuel mileage improvements for a 0.78 Mach number was 3.1 percent at 8 x 10(5) pounds W/delta and 5.5 percent at 1.05 x 10(6) pounds W/delta. Correcting the flight measured data for surface pressure differences between wind tunnel and flight resulted in a fuel mileage improvement of 4.4 percent at 8 x 10(5) pounds W/delta and 7.2 percent at 1.05 x 10(6) pounds W/delta. The performance improvement obtained was within the wind tunnel test data obtained from two different wind tunnel models. The buffet boundary data obtained for the baseline configuration was in good agreement with previous established data. Buffet data for the 15 deg cant/-4 deg incidence configuration showed a slight improvement, while the 15 deg cant/-2 deg incidence and 0 deg cant/-4 deg incidence data showed a slight deterioration.

On the aero-elastic design of the DTU 10MW wind turbine blade for the LIFES50+ wind tunnel scale model

NASA Astrophysics Data System (ADS)

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

2016-09-01

This paper illustrates the aero-elastic optimal design, the realization and the verification of the wind tunnel scale model blades for the DTU 10 MW wind turbine model, within LIFES50+ project. The aerodynamic design was focused on the minimization of the difference, in terms of thrust coefficient, with respect to the full scale reference. From the Selig low Reynolds database airfoils, the SD7032 was chosen for this purpose and a proper constant section wing was tested at DTU red wind tunnel, providing force and distributed pressure coefficients for the design, in the Reynolds range 30-250 E3 and for different angles of attack. The aero-elastic design algorithm was set to define the optimal spanwise thickness over chord ratio (t/c), the chord length and the twist to match the first flapwise scaled natural frequency. An aluminium mould for the carbon fibre was CNC manufactured based on B-Splines CAD definition of the external geometry. Then the wind tunnel tests at Politecnico di Milano confirmed successful design and manufacturing approaches.

MHz-rate nitric oxide planar laser-induced fluorescence imaging in a Mach 10 hypersonic wind tunnel.

PubMed

Jiang, Naibo; Webster, Matthew; Lempert, Walter R; Miller, Joseph D; Meyer, Terrence R; Ivey, Christopher B; Danehy, Paul M

2011-02-01

Nitric oxide planar laser-induced fluorescence (NO PLIF) imaging at repetition rates as high as 1 MHz is demonstrated in the NASA Langley 31 in. Mach 10 hypersonic wind tunnel. Approximately 200 time-correlated image sequences of between 10 and 20 individual frames were obtained over eight days of wind tunnel testing spanning two entries in March and September of 2009. The image sequences presented were obtained from the boundary layer of a 20° flat plate model, in which transition was induced using a variety of different shaped protuberances, including a cylinder and a triangle. The high-speed image sequences captured a variety of laminar and transitional flow phenomena, ranging from mostly laminar flow, typically at a lower Reynolds number and/or in the near wall region of the model, to highly transitional flow in which the temporal evolution and progression of characteristic streak instabilities and/or corkscrew-shaped vortices could be clearly identified.

High Response Dew Point Measurement System for a Supersonic Wind Tunnel

NASA Technical Reports Server (NTRS)

Blumenthal, Philip Z.

1996-01-01

A new high response on-line measurement system has been developed to continuously display and record the air stream dew point in the NASA Lewis 10 x 10 supersonic wind tunnel. Previous instruments suffered from such problems as very slow response, erratic readings, and high susceptibility to contamination. The system operates over the entire pressure level range of the 10 x 10 SWT, from less than 2 psia to 45 psia, without the need for a vacuum pump to provide sample flow. The system speeds up tunnel testing, provides large savings in tunnel power costs and provides the dew point input for the data-reduction subroutines which calculate test section conditions.

Supersonic Wind Tunnel Capabilities Expanded Into Subsonic Region

NASA Technical Reports Server (NTRS)

Roeder, James W., Jr.

1997-01-01

The operating envelope of the Abe Silverstein 10- by 10-Foot Supersonic Wind Tunnel (10x10 SWT) at the NASA Lewis Research Center was recently expanded to include operation at subsonic test section speeds. This new capability generates test section air speeds ranging from Mach 0.05 to 0.35 (32 to 240 kn). Most of the expansion in air speed range was obtained by running the tunnel's main compressor at much lower speeds than ever before. The compressor drive system, consisting of four large electric motors, was run with only one or two motors energized to obtain the lower compressor speed range. This new capability makes the 10x10 SWT more versatile and gives U.S. researchers an enhanced ability to perform subsonic propulsion and aerodynamic testing.

Evaluation of the NASA Ames no. 1 7 by 10 foot wind tunnel as an acoustic test facility

NASA Technical Reports Server (NTRS)

Wilby, J. F.; Scharton, T. D.

1975-01-01

Measurements were made in the no. 1 7'x10' wind tunnel at NASA Ames Research Center, with the objectives of defining the acoustic characteristics and recommending minimum cost treatments so that the tunnel can be converted into an acoustic research facility. The results indicate that the noise levels in the test section are due to (a) noise generation in the test section, associated with the presence of solid bodies such as the pitot tube, and (b) propagation of acoustic energy from the fan. A criterion for noise levels in the test section is recommended, based on low-noise microphone support systems. Noise control methods required to meet the criterion include removal of hardware items for the test section and diffuser, improved design of microphone supports, and installation of acoustic treatment in the settling chamber and diffuser.

X-33 Aerodynamic and Aeroheating Computations for Wind Tunnel and Flight Conditions

NASA Technical Reports Server (NTRS)

Hollis, Brian R.; Thompson, Richard A.; Murphy, Kelly J.; Nowak, Robert J.; Riley, Christopher J.; Wood, William A.; Alter, Stephen J.; Prabhu, Ramadas K.

1999-01-01

This report provides an overview of hypersonic Computational Fluid Dynamics research conducted at the NASA Langley Research Center to support the Phase II development of the X-33 vehicle. The X-33, which is being developed by Lockheed-Martin in partnership with NASA, is an experimental Single-Stage-to-Orbit demonstrator that is intended to validate critical technologies for a full-scale Reusable Launch Vehicle. As part of the development of the X-33, CFD codes have been used to predict the aerodynamic and aeroheating characteristics of the vehicle. Laminar and turbulent predictions were generated for the X 33 vehicle using two finite- volume, Navier-Stokes solvers. Inviscid solutions were also generated with an Euler code. Computations were performed for Mach numbers of 4.0 to 10.0 at angles-of-attack from 10 deg to 48 deg with body flap deflections of 0, 10 and 20 deg. Comparisons between predictions and wind tunnel aerodynamic and aeroheating data are presented in this paper. Aeroheating and aerodynamic predictions for flight conditions are also presented.

Wind Tunnel Corrections for High Angle of Attack Models,

DTIC Science & Technology

1981-02-01

MAQUETTES EN SOUFFLERIE par X.Vaucheret GERMA * J ACTIVITIES ON WIND TUNNEL CORRECTIONS byH.HoIst A REVIEW OF RESEARCH AT NLR ON WIND TUNNEL...1-10 Ro - ,2.13M106 M - 0.230 ° BALANCE — corrected -T unoorreoted •r r^a—Q * o n ...8217 n * t ?’ A *i o o 1 1 -0.70 -0.65 -0.60 -0.S5 -0.50 -0.45 -0.40 Fig.l 1 Corrected

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

Magnetic Leviation System Design and Implementation for Wind Tunnel Application

NASA Technical Reports Server (NTRS)

Lin, Chin E.; Sheu, Yih-Ran; Jou, Hui-Long

1996-01-01

This paper presents recent work in magnetic suspension wind tunnel development in National Cheng Kung University. In this phase of research, a control-based study is emphasized to implement a robust control system into the experimental system under study. A ten-coil 10 cm x 10 cm magnetic suspension wind tunnel is built using a set of quadrant detectors for six degree of freedom control. To achieve the attitude control of suspended model with different attitudes, a spacial electromagnetic field simulation using OPERA 3D is studied. A successful test for six degree of freedom control is demonstrated in this paper.

NACA Engineer Examines Wind Tunnel Compressor Blades

NASA Image and Video Library

1955-09-21

An engineer examines the main compressor for the 10- by 10-Foot Supersonic Wind Tunnel at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory. The engineers were preparing the new wind tunnel for its initial runs in early 1956. The 10- by 10 was the most powerful propulsion wind tunnel in the nation. The facility was part of Congress’ Unitary Plan Act which coordinated wind tunnel construction at the NACA, Air Force, industry, and universities. The 10- by 10 was the largest of the three NACA tunnels built under the act. The 20-foot diameter eight-stage axial flow compressor, seen in this photograph, could generate air flows up to Mach 2.5 through the test section. The stainless steel compressor had 584 blades ranging from 1.8 to 3.25 feet in length. This main compressor was complemented by a secondary axial flow compressor. Working in tandem the two could generate wind streams up to Mach 3.5. The Cleveland Chamber of Commerce presented NACA Lewis photographer Bill Bowles with a second place award for this photograph in their Business and Professional category. The photograph was published in October 1955 edition of its periodical, The Clevelander, which highlighted local professional photographers. Fellow Lewis photographer Gene Giczy won second place in another category for a photograph of Cleveland Municipal Airport.

Construction of the 40x80 Foot Wind Tunnel at Ames.

NASA Image and Video Library

1943-07-07

Looking South from inside the diffuser of the 40x80 foot wind tunnel at NACA's Ames Research Center. Construction began in late 1941, the mammoth construction task sorely taxing the resources of the new center. Two and a half years later, in dune 1944, the 40 x 80-foot full-scale tunnel went into operation.

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.

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

NASA Technical Reports Server (NTRS)

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

1986-01-01

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

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.

Historical Overview and Recent Improvements at the NASA Glenn Research Center 8x6 9x15 Wind Tunnel Complex

NASA Technical Reports Server (NTRS)

Dussling, Joseph John

2015-01-01

A brief history of the 8x6 Supersonic Wind Tunnel (SWT) and 9x15 Low Speed Wind Tunnel (LSWT) at NASA Glenn Research Center, Cleveland, Ohio is presented along with current capabilities and plans for future upgrades within the facility.

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

NASA Technical Reports Server (NTRS)

Marshall, B. A.; Marroquin, J.

1984-01-01

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

Planar laser-induced fluorescence (PLIF) investigation of hypersonic flowfields in a Mach 10 wind tunnel

NASA Technical Reports Server (NTRS)

Danehy, Paul M.; Wilkes, Jennifer A.; Aderfer, David W.; Jones, Stephen B.; Robbins, Anthony W.; Pantry, Danny P.; Schwartz, Richard J.

2006-01-01

Planar laser-induced fluorescence (PLIF) of nitric oxide (NO) was used to visualize four different hypersonic flowfields in the NASA Langley Research Center 31-Inch Mach 10 Air wind tunnel. The four configurations were: (1) the wake flowfield of a fuselage-only X-33 lifting body, (2) flow over a flat plate containing a rectangular cavity, (3) flow over a 70deg blunted cone with a cylindrical afterbody, formerly studied by an AGARD working group, and (4) an Apollo-geometry entry capsule - relevant to the Crew Exploration Vehicle currently being developed by NASA. In all cases, NO was seeded into the flowfield through tubes inside or attached to the model sting and strut. PLIF was used to visualize the NO in the flowfield. In some cases pure NO was seeded into the flow while in other cases a 5% NO, 95% N2 mix was injected. Several parameters were varied including seeding method and location, seeding mass flow rate, model angle of attack and tunnel stagnation pressure, which varies the unit Reynolds number. The location of the laser sheet was as also varied to provide three dimensional flow information. Virtual Diagnostics Interface (ViDI) technology developed at NASA Langley was used to visualize the data sets in post processing. The measurements demonstrate some of the capabilities of the PLIF method for studying hypersonic flows.

Flex-I-Kite flying in 7 x 10 LST (outboard tail model)

NASA Image and Video Library

1959-05-14

On June 26, 1959, then-Langley-research Francis Rogallo examined the Rogallo wing in the 7x10 FT Tunnel. Originally conceived as a means of bringing manned spacecraft to controlled, soft landings, Rogallo's concept was avidly embraced by later generations of hang-gliding enthusiasts. -- Photograph published in Winds of Change, 75th Anniversary NASA publication (page 18), by James Schultz.

Comparison of Ares I-X Wind-Tunnel Derived Buffet Environment with Flight Data

NASA Technical Reports Server (NTRS)

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

2011-01-01

The Ares I-X Flight Test Vehicle (FTV), launched in October 2009, carried with it over 243 buffet verification pressure sensors and was one of the most heavily instrumented launch vehicle flight tests. This flight test represented a unique opportunity for NASA and its partners to compare the wind-tunnel derived buffet environment with that measured during the flight of Ares I-X. It is necessary to define the launch vehicle buffet loads to ensure that structural components and vehicle subsystems possess adequate strength, stress, and fatigue margins when the vehicle structural dynamic response to buffet forcing functions are considered. Ares I-X buffet forcing functions were obtained via wind-tunnel testing of a rigid buffet model (RBM) instrumented with hundreds of unsteady pressure transducers designed to measure the buffet environment across the desired frequency range. This paper discusses the comparison of RBM and FTV buffet environments, including fluctuating pressure coefficient and normalized sectional buffet forcing function root-mean-square magnitudes, frequency content of power-spectral density functions, and force magnitudes of an alternating flow phenomena. Comparison of wind-tunnel model and flight test vehicle buffet environments show very good agreement with root-mean-square magnitudes of buffet forcing functions at the majority of vehicle stations. Spectra proved a challenge to compare because of different wind-tunnel and flight test conditions and data acquisition rates. However, meaningful and promising comparisons of buffet spectra are presented. Lastly, the buffet loads resulting from the transition of subsonic separated flow to supersonic attached flow were significantly over-predicted by wind-tunnel results.

Background Acoustics Levels in the 9x15 Wind Tunnel and Linear Array Testing

NASA Technical Reports Server (NTRS)

Stephens, David

2011-01-01

The background noise level in the 9x15 foot wind tunnel at NASA Glenn has been documented, and the results compare favorably with historical measurements. A study of recessed microphone mounting techniques was also conducted, and a recessed cavity with a micronic wire mesh screen reduces hydrodynamic noise by around 10 dB. A three-microphone signal processing technique can provide additional benefit, rejecting up to 15 dB of noise contamination at some frequencies. The screen and cavity system offers considerable benefit to test efficiency, although there are additional calibration requirements.

North American X-15 model tested in 300MPH Low Speed 7x10 Tunnel

NASA Image and Video Library

1958-09-07

A one-twentieth scale model of the X-15 originally suspended beneath the wing of a B-52 is observed by a scientist of the National Aeronautics and Space Administration (NASA) as it leaves the bomber model in tests to determine the release characteristics and drop motion of the research airplane. Caption: The aerodynamics of air launching the North American X-15 being investigated in the 300MPH Low Speed 7x10 Tunnel, about 1957. Photograph published in Engineer in Charge: A History of the Langley Aeronautical Laboratory, 1917-1958 by James R. Hansen. Page 366. Photograph also published in Sixty Years of Aeronautical Research 1917-1977 By David A. Anderton. A NASA publication. Page 49.

An Assessment of Ares I-X Aeroacoustic Measurements with Comparisons to Pre-Flight Wind Tunnel Test Results

NASA Technical Reports Server (NTRS)

Nance, Donald K.; Reed, Darren K.

2011-01-01

During the recent successful launch of the Ares I-X Flight Test Vehicle, aeroacoustic data was gathered at fifty-seven locations along the vehicle as part of the Developmental Flight Instrumentation. Several of the Ares I-X aeroacoustic measurements were placed to duplicate measurement locations prescribed in pre-flight, sub-scale wind tunnel tests. For these duplicated measurement locations, comparisons have been made between aeroacoustic data gathered during the ascent phase of the Ares I-X flight test and wind tunnel test data. These comparisons have been made at closely matching flight conditions (Mach number and vehicle attitude) in order to preserve a one-to-one relationship between the flight and wind tunnel data. These comparisons and the current wind tunnel to flight scaling methodology are presented and discussed. The implications of using wind tunnel test data scaled under the current methodology to predict conceptual launch vehicle aeroacoustic environments are also discussed.

Supersonic Retropropulsion Experimental Results from the NASA Ames 9- x 7-Foot Supersonic Wind Tunnel

NASA Technical Reports Server (NTRS)

Berry, Scott A.; Rhode, Matthew N.; Edquist, Karl T.

2012-01-01

Supersonic retropropulsion was experimentally examined in the Ames Research Center 9x7-Foot Supersonic Wind Tunnel at Mach 1.8 and 2.4. The experimental model, previously designed for and tested in the Langley Research Center Unitary Plan Wind Tunnel at Mach 2.4, 3.5 and 4.6, was a 5-in diameter 70-deg sphere-cone forebody with a 9.55-in long cylindrical aftbody. The forebody was designed to accommodate up to four 4:1 area ratio nozzles, one on the model centerline and the other three on the half radius spaced 120-deg apart. Surface pressure and flow visualization were the primary measurements, including high-speed data to investigate the dynamics of the interactions between the bow and nozzle shocks. Three blowing configurations were tested with thrust coefficients up to 10 and angles of attack up to 20-deg. Preliminary results and observations from the test are provided

The Design of Low-Turbulence Wind Tunnels

NASA Technical Reports Server (NTRS)

Dryden, Hugh L; Abbott, Ira H

1949-01-01

Within the past 10 years there have been placed in operation in the United States four low-turbulence wind tunnels of moderate cross-sectional area and speed, one at the National Bureau of Standards, two at the NACA Langley Laboratory, and one at the NACA Ames Laboratory. This paper reviews briefly the state of knowledge and those features which make possible the attainment of low turbulence in wind tunnels. Specific applications to two wind tunnels are described.

Second-Generation Large Civil Tiltrotor 7- by 10-Foot Wind Tunnel Test Data Report

NASA Technical Reports Server (NTRS)

Theodore, Colin R.; Russell, Carl R.; Willink, Gina C.; Pete, Ashley E.; Adibi, Sierra A.; Ewert, Adam; Theuns, Lieselotte; Beierle, Connor

2016-01-01

An approximately 6-percent scale model of the NASA Second-Generation Large Civil Tiltrotor (LCTR2) Aircraft was tested in the U.S. Army 7- by 10-Foot Wind Tunnel at NASA Ames Research Center January 4 to April 19, 2012, and September 18 to November 1, 2013. The full model was tested, along with modified versions in order to determine the effects of the wing tip extensions and nacelles; the wing was also tested separately in the various configurations. In both cases, the wing and nacelles used were adopted from the U.S. Army High Efficiency Tilt Rotor (HETR) aircraft, in order to limit the cost of the experiment. The full airframe was tested in high-speed cruise and low-speed hover flight conditions, while the wing was tested only in cruise conditions, with Reynolds numbers ranging from 0 to 1.4 million. In all cases, the external scale system of the wind tunnel was used to collect data. Both models were mounted to the scale using two support struts attached underneath the wing; the full airframe model also used a third strut attached at the tail. The collected data provides insight into the performance of the preliminary design of the LCTR2 and will be used for computational fluid dynamics (CFD) validation and the development of flight dynamics simulation models.

A Wind Tunnel Study of Icing Effects on a Business Jet Airfoil

NASA Technical Reports Server (NTRS)

Addy, Harold E., Jr.; Broeren, Andy P.; Zoeckler, Joesph G.; Lee, Sam

2003-01-01

Aerodynamic wind tunnel tests were conducted to study the effects of various ice accretions on the aerodynamic performance of a 36-inch chord, two-dimensional business jet airfoil. Eight different ice shape configurations were tested. Four were castings made from molds of ice shapes accreted in an icing wind tunnel. Two were made using computationally smoothed tracings of two of the ice shapes accreted in the icing tunnel. These smoothed profiles were then extended in the spanwise direction to form a two-dimensional ice shape. The final two configurations were formed by applying grit to the smoothed ice shapes. The ice shapes resulted in as much as 48% reduction in maximum lift coefficient from that of the clean airfoil. Large increases in drag and changes in pitching moment were also observed. The castings and their corresponding smoothed counterparts yielded similar results. Little change in performance was observed with the addition of grit to the smoothed ice shapes. Changes in the Reynolds number (from 3 x 10(exp 6) to 10.5 x 10(exp 6) and Mach number (from 0.12 to 0.28) did not significantly affect the iced-airfoil performance coefficients.

Wind Tunnel Measurements of Shuttle Orbiter Global Heating with Comparisons to Flight

NASA Technical Reports Server (NTRS)

Berry, Scott A.; Merski, N. Ronald; Blanchard, Robert C.

2002-01-01

An aerothermodynamic database of global heating images was acquired of the Shuttle Orbiter in the NASA Langley Research Center 20-Inch Mach 6 Air Tunnel. These results were obtained for comparison to the global infrared images of the Orbiter in flight from the infrared sensing aeroheating flight experiment (ISAFE). The most recent ISAFE results from STS-103, consisted of port side images, at hypersonic conditions, of the surface features that result from the strake vortex scrubbing along the side of the vehicle. The wind tunnel results were obtained with the phosphor thermography system, which also provides global information and thus is ideally suited for comparison to the global flight results. The aerothermodynamic database includes both windward and port side heating images of the Orbiter for a range of angles of attack (20 to 40 deg), freestream unit Reynolds number (1 x 10(exp 6))/ft to 8 x 10(exp 6)/ft, body flap deflections (0, 5, and 10 deg), speed brake deflections (0 and 45 deg), as well as with boundary layer trips for forced transition to turbulence heating results. Sample global wind tunnel heat transfer images were extrapolated to flight conditions for comparison to Orbiter flight data. A windward laminar case for an angle of attack of 40 deg was extrapolated to Mach 11.6 flight conditions for comparison to STS-2 flight thermocouple results. A portside wind tunnel image for an angle of attack of 25 deg was extrapolated for Mach 5 flight conditions for comparison to STS-103 global surface temperatures. The comparisons showed excellent qualitative agreement, however the extrapolated wind tunnel results over-predicted the flight surface temperatures on the order of 5% on the windward surface and slightly higher on the portside.

Lockheed XFV-1 model in the 40x80 foot wind tunnel at NASA Ames Research Center

NASA Image and Video Library

1952-05-16

Wide shot of 40x 80 wind tunnel settling chamber with Lockheed XFV-1 model. Project engineer Mark Kelly (not shown). Remote controlled model flown in the settling chamber of the 40x80 wind tunnel. Electric motors in the model, controlled the counter-rotating propellers to test vertical takeoff. Test no. 71

Wind Tunnel Results of the B-52B with the X-43A Stack

NASA Technical Reports Server (NTRS)

Davis, Mark C.; Sim, Alexander G.; Rhode, Matthew; Johnson, Kevin D.

2006-01-01

A low-speed wind-tunnel test was performed with a three-percent-scale model of a booster rocket mated to an X-43A research vehicle, a combination referred to as the Hyper-X launch vehicle. The test was conducted both in free-stream air and in the presence of a partial model of the B-52B airplane. The objectives of the test were to obtain force and moment data to generate structural loads affecting the pylon of the B-52B airplane and to determine the aerodynamic influence of the B-52B airplane on the Hyper-X launch vehicle to evaluate launch separation characteristics. The wind-tunnel test was conducted at a low-speed wind tunnel in Hampton, Virginia. All moments and forces reported are based either on the aerodynamic influence of the B-52B airplane or are for the Hyper-X launch vehicle in free-stream air. Overall, the test showed that the B-52B airplane imparts a strong downwash onto the Hyper-X launch vehicle, reducing the net lift of the Hyper-X launch vehicle. Also, pitching and rolling moments are imparted onto the booster and are a strong function of the launch-drop angle of attack.

ARC-2010-ACD10-0020-013

NASA Image and Video Library

2010-01-14

Lawrence Livermore National Labs (LLNL), Navistar and the Department of Energy conduct tests in the NASA Ames National Full-scale Aerodynamic Complex 80x120_foot wind tunnel. The LLNL project is aimed at aerodynamic truck and trailer devices that can reduce fuel consumption at highway speed by 10 percent. Cab being lifted into the tunnel.

ARC-2010-ACD10-0020-023

NASA Image and Video Library

2010-02-03

Lawrence Livermore National Labs (LLNL), Navistar and the Department of Energy conduct tests in the NASA Ames National Full-scale Aerodynamic Complex 80x120_foot wind tunnel. The LLNL project is aimed at aerodynamic truck and trailer devices that can reduce fuel consumption at highway speed by 10 percent. Trailer being lifted into the tunnel.

Analysis of Wind Tunnel Oscillatory Data of the X-31A Aircraft

NASA Technical Reports Server (NTRS)

Smith, Mark S.

1999-01-01

Wind tunnel oscillatory tests in pitch, roll, and yaw were performed on a 19%-scale model of the X-31A aircraft. These tests were used to study the aerodynamic characteristics of the X-31A in response to harmonic oscillations at six frequencies. In-phase and out-of-phase components of the aerodynamic coefficients were obtained over a range of angles of attack from 0 to 90 deg. To account for the effect of frequency on the data, mathematical models with unsteady terms were formulated by use of two different indicial functions. Data from a reduced set of frequencies were used to estimate model parameters, including steady-state static and dynamic stability derivatives. Both models showed good prediction capability and the ability to accurately fit the measured data. Estimated static stability derivatives compared well with those obtained from static wind tunnel tests. The roll and yaw rate derivative estimates were compared with rotary-balanced wind tunnel data and theoretical predictions. The estimates and theoretical predictions were in agreement at small angles of attack. The rotary-balance data showed, in general, acceptable agreement with the steady-state derivative estimates.

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

NASA Technical Reports Server (NTRS)

Nutt, K. W.

1984-01-01

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

High-Speed Schlieren and 10-Hz Kr PLIF for the new AFRL Mach-6 Ludwieg Tube Hypersonic Wind Tunnel

DTIC Science & Technology

2017-11-01

STATES AIR FORCE AFRL-RQ-WP-TP-2017-0167 HIGH -SPEED SCHLIEREN AND 10-HZ KR PLIF FOR THE NEW AFRL MACH 6 LUDWIEG TUBE HYPERSONIC WIND TUNNEL Roger...L. Kimmel and Campbell D. Carter Hypersonic Sciences Branch High Speed Systems Division Joshua D. Pickles and Venkateswaran Narayanaswamy North...Public Affairs Office (PAO) and is available to the general public, including foreign nationals. Copies may be obtained from the Defense Technical

Overview of X-38 Hypersonic Wind Tunnel Data and Comparison with Numerical Results

NASA Technical Reports Server (NTRS)

Campbell, Charles H.; Caram, Jose; Berry, Scott; DiFulvio, Michael; Horvath, Tom

1997-01-01

A NASA team of engineers has been organized to design a crew return vehicle for returning International Space Station crew members from orbit. The hypersonic characteristics of this X-23/X-2&4 derived crew return vehicle (designated X-38) are being evaluated in various wind tunnels in support of this effort. Aerodynamic data has been acquired in three NASA hypersonic facilities at Mach 20, and Mach 6. Computational Fluid Dynamics tools have been applied at the appropriate wind tunnel conditions to make comparisons with portions of this data. Experimental data from the Mach 6 Air and CF4 facilities illustrate a net positive pitching moment increment due to density ratio, as well as increased elevon effectiveness. Chemical nonequilibrium computational fluid dynamics solutions at flight conditions reinforce this conclusion.

14. EXTERIOR VIEW OF OLD TENFOOT WIND TUNNEL (1991). ...

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

14. EXTERIOR VIEW OF OLD TEN-FOOT WIND TUNNEL (1991). - Wright-Patterson Air Force Base, Area B, Buildings 25 & 24,10-foot & 20-foot Wind Tunnel Complex, Northeast side of block bounded by K, G, Third, & Fifth Streets, Dayton, Montgomery County, OH

13. EXTERIOR VIEW OF OLD TENFOOT WIND TUNNEL (1991). ...

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

13. EXTERIOR VIEW OF OLD TEN-FOOT WIND TUNNEL (1991). - Wright-Patterson Air Force Base, Area B, Buildings 25 & 24,10-foot & 20-foot Wind Tunnel Complex, Northeast side of block bounded by K, G, Third, & Fifth Streets, Dayton, Montgomery County, OH

Rudolf Hermann, wind tunnels and aerodynamics

NASA Astrophysics Data System (ADS)

Lundquist, Charles A.; Coleman, Anne M.

2008-04-01

Rudolf Hermann was born on December 15, 1904 in Leipzig, Germany. He studied at the University of Leipzig and at the Aachen Institute of Technology. His involvement with wind tunnels began in 1934 when Professor Carl Wieselsberger engaged him to work at Aachen on the development of a supersonic wind tunnel. On January 6, 1936, Dr. Wernher von Braun visited Dr. Hermann to arrange for use of the Aachen supersonic wind tunnel for Army problems. On April 1, 1937, Dr. Hermann became Director of the Supersonic Wind Tunnel at the Army installation at Peenemunde. Results from the Aachen and Peenemunde wind tunnels were crucial in achieving aerodynamic stability for the A-4 rocket, later designated as the V-2. Plans to build a Mach 10 'hypersonic' wind tunnel facility at Kochel were accelerated after the Allied air raid on Peenemunde on August 17, 1943. Dr. Hermann was director of the new facility. Ignoring destruction orders from Hitler as WWII approached an end in Europe, Dr. Hermann and his associates hid documents and preserved wind tunnel components that were acquired by the advancing American forces. Dr. Hermann became a consultant to the Air Force at its Wright Field in November 1945. In 1951, he was named professor of Aeronautical Engineering at the University of Minnesota. In 1962, Dr. Hermann became the first Director of the Research Institute at the University of Alabama in Huntsville (UAH), a position he held until he retired in 1970.

A Summary of the Experimental Results for a Generic Tractor-Trailer in the Ames Research Center 7- by 10-Foot and 12-Foot Wind Tunnels

NASA Technical Reports Server (NTRS)

Storms, Bruce L.; Satran, Dale R.; Heineck, James T.; Walker, Stephen M.

2006-01-01

Experimental measurements of a generic tractor-trailer were obtained in two wind tunnels at Ames Research Center. After a preliminary study at atmospheric conditions in the 7- by 10-Foot Wind Tunnel, additional testing was conducted at Reynolds numbers corresponding to full-scale highway speeds in the 12-Foot Pressure Wind Tunnel. To facilitate computational modeling, the 1:8-scale geometry, designated the Generic Conventional Model, included a simplified underbody and omitted many small-scale details. The measurements included overall and component forces and moments, static and dynamic surface pressures, and three-component particle image velocimetry. This summary report highlights the effects of numerous drag reduction concepts and provides details of the model installation in both wind tunnels. To provide a basis for comparison, the wind-averaged drag coefficient was tabulated for all configurations tested. Relative to the baseline configuration representative of a modern class-8 tractor-trailer, the most effective concepts were the trailer base flaps and trailer belly box providing a drag-coefficient reduction of 0.0855 and 0.0494, respectively. Trailer side skirts were less effective yielding a drag reduction of 0.0260. The database of this experimental effort is publicly available for further analysis.

9. INTERIOR VIEW OF WIND TUNNEL (1991). WrightPatterson Air ...

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

9. INTERIOR VIEW OF WIND TUNNEL (1991). - Wright-Patterson Air Force Base, Area B, Buildings 25 & 24,10-foot & 20-foot Wind Tunnel Complex, Northeast side of block bounded by K, G, Third, & Fifth Streets, Dayton, Montgomery County, OH

11. INTERIOR VIEW OF WIND TUNNEL (1991). WrightPatterson Air ...

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

11. INTERIOR VIEW OF WIND TUNNEL (1991). - Wright-Patterson Air Force Base, Area B, Buildings 25 & 24,10-foot & 20-foot Wind Tunnel Complex, Northeast side of block bounded by K, G, Third, & Fifth Streets, Dayton, Montgomery County, OH

3. VIEW OF WIND TUNNEL, LOOKING NORTHWEST (1991). WrightPatterson ...

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

3. VIEW OF WIND TUNNEL, LOOKING NORTHWEST (1991). - Wright-Patterson Air Force Base, Area B, Buildings 25 & 24,10-foot & 20-foot Wind Tunnel Complex, Northeast side of block bounded by K, G, Third, & Fifth Streets, Dayton, Montgomery County, OH

Qualitative comparison of calculated turbulence responses with wind-tunnel measurements for a DC-10 derivative wing with an active control system

NASA Technical Reports Server (NTRS)

Perry, B., III

1981-01-01

Comparisons are presented analytically predicted and experimental turbulence responses of a wind tunnel model of a DC-10 derivative wing equipped with an active control system. The active control system was designed for the purpose of flutter suppression, but it had additional benefit of alleviating gust loads (wing bending moment) by about 25%. Comparisions of various wing responses are presented for variations in active control system parameters and tunnel speed. The analytical turbulence responses were obtained using DYLOFLEX, a computer program for dynamic loads analyses of flexible airplanes with active controls. In general, the analytical predictions agreed reasonably well with the experimental data.

Videogrammetric Model Deformation Measurement Technique for Wind Tunnel Applications

NASA Technical Reports Server (NTRS)

Barrows, Danny A.

2006-01-01

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

6. CLOSEUP VIEW OF TENFOOT WIND TUNNEL (1991). WrightPatterson ...

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

6. CLOSE-UP VIEW OF TEN-FOOT WIND TUNNEL (1991). - Wright-Patterson Air Force Base, Area B, Buildings 25 & 24,10-foot & 20-foot Wind Tunnel Complex, Northeast side of block bounded by K, G, Third, & Fifth Streets, Dayton, Montgomery County, OH

Orifice-induced pressure error studies in Langley 7- by 10-foot high-speed tunnel

NASA Technical Reports Server (NTRS)

Plentovich, E. B.; Gloss, B. B.

1986-01-01

For some time it has been known that the presence of a static pressure measuring hole will disturb the local flow field in such a way that the sensed static pressure will be in error. The results of previous studies aimed at studying the error induced by the pressure orifice were for relatively low Reynolds number flows. Because of the advent of high Reynolds number transonic wind tunnels, a study was undertaken to assess the magnitude of this error at high Reynolds numbers than previously published and to study a possible method of eliminating this pressure error. This study was conducted in the Langley 7- by 10-Foot High-Speed Tunnel on a flat plate. The model was tested at Mach numbers from 0.40 to 0.72 and at Reynolds numbers from 7.7 x 1,000,000 to 11 x 1,000,000 per meter (2.3 x 1,000,000 to 3.4 x 1,000,000 per foot), respectively. The results indicated that as orifice size increased, the pressure error also increased but that a porous metal (sintered metal) plug inserted in an orifice could greatly reduce the pressure error induced by the orifice.

The George C. Marshall Space Flight Center's 14 X 14-Inch Trisonic Wind Tunnel: A Historical Perspective

NASA Technical Reports Server (NTRS)

Springer, A.

1994-01-01

A history of the National Aeronautics and Space Administration (NASA) George C. Marshall Space Flight Center's (MSFC) 14 x 14-Inch Trisonic Wind Tunnel is presented. Its early and continuing role in the United States space program is shown through highlights of the tunnel's history and the major programs tested in the tunnel over the past 40 years. The 14-Inch Tunnel has its beginning with the Army in the late 1950's under the Army Ballistic Missile Agency (ABMA). Such programs as the Redstone, Jupiter, Pershing, and early Saturn were tested in the 14-Inch Tunnel in the late 1950's. America's first launch vehicle, the Jupiter C, was designed and developed using the 14-Inch Wind Tunnel. Under NASA, the 14-Inch Wind Tunnel has made large contributions to the Saturn, Space Transportation System, and future launch vehicle programs such as Shuttle-C and the National Launch System. A technical description of the tunnel is presented for background information on the type and capabilities of the 14-Inch Wind Tunnel. The report concludes in stating: the 14-Inch Wind Tunnel as in speed of sound; transonic, at or near the speed of sound the past, will continue to play a large but unseen role in he development of America's space program.

Cryogenic wind tunnel technology. A way to measurement at higher Reynolds numbers

NASA Technical Reports Server (NTRS)

Beck, J. W.

1984-01-01

The goals, design, problems, and value of cryogenic transonic wind tunnels being developed in Europe are discussed. The disadvantages inherent in low-Reynolds-number (Re) wind tunnel simulations of aircraft flight at high Re are reviewed, and the cryogenic tunnel is shown to be the most practical method to achieve high Re. The design proposed for the European Transonic Wind tunnel (ETW) is presented: parameters include cross section. DISPLAY 83A46484/2 = 4 sq m, operating pressure = 5 bar, temperature = 110-120 K, maximum Re = 40 x 10 to the 6th, liquid N2 consumption = 40,000 metric tons/year, and power = 39,5 MW. The smaller Cologne subsonic tunnel being adapted to cryogenic use for preliminary studies is described. Problems of configuration, materials, and liquid N2 evaporation and handling and the research underway to solve them are outlined. The benefits to be gained by the construction of these costly installations are seen more in applied aerodynamics than in basic research in fluid physics. The need for parallel development of both high Re tunnels and computers capable of performing high-Re numerical analysis is stressed.

Qualitative Assessment of the Acoustic Disturbance Environment in the NASA LaRC 20-Inch MACH 6 Wind Tunnel

NASA Technical Reports Server (NTRS)

Horvath, Thomas J.; Berry, Scott A.; Hamilton, H. Harris

2001-01-01

An experimental investigation was conducted on a 5-degree-half-angle cone with a flare in a conventional Mach 6 wind tunnel to examine the effect of facility noise on boundary layer transition. The effect of tunnel noise was inferred by comparing transition onset locations determined from the present test to that previously obtained in a Mach 6 quiet tunnel. Together, the two sets of experiments are believed to represent the first direct comparison of transition onset between a conventional and a quiet hypersonic wind tunnel using a common test model. In the present conventional hypersonic tunnel experiment, adiabatic wall temperatures were measured and heat transfer distributions were inferred on the cone flare model at zero degree angle of attack over a range of length Reynolds numbers (2 x 10(exp 6) to 10 x 10(exp 6)) which resulted in laminar and turbulent flow. Wall-to-total temperature ratio for the transient heating measurements and the adiabatic wall temperature measurements were 0.69 and 0.86, respectively. The cone flare nosetip radius was varied from 0.0001 to 0.125-inch to examine the effects of bluntness on transition onset. At comparable freestream conditions the transition onset Reynolds number obtained on the cone flare model in the conventional "noisy" tunnel was approximately 25% lower than that measured in the low disturbance tunnel.

11. INTERIOR VIEW OF 8FOOT HIGH SPEED WIND TUNNEL. SAME ...

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

11. INTERIOR VIEW OF 8-FOOT HIGH SPEED WIND TUNNEL. SAME CAMERA POSITION AS VA-118-B-10 LOOKING IN THE OPPOSITE DIRECTION. - NASA Langley Research Center, 8-Foot High Speed Wind Tunnel, 641 Thornell Avenue, Hampton, Hampton, VA

The Design of Wind Tunnels and Wind Tunnel Propellers

NASA Technical Reports Server (NTRS)

Warner, Edward P; Norton, F H; Hebbert, C M

1919-01-01

Report discusses the theory of energy losses in wind tunnels, the application of the Drzewiecki theory of propeller design to wind tunnel propellers, and the efficiency and steadiness of flow in model tunnels of various types.

X-38 Experimental Aeroheating at Mach 10

NASA Technical Reports Server (NTRS)

Berry, Scott A.; Horvath, Thomas J.; Weilmuenster, K. James; Alter, Stephan J.; Merski, N. Ronald

2001-01-01

This report provides an update of the hypersonic aerothermodynamic wind tunnel test program conducted at the NASA Langley Research Center in support of the X-38 program. Global surface heat transfer distributions were measured on 0.0177 and 0.0236 scale models of the proposed X-38 configuration at Mach 10 in air. The parametrics that were investigated primarily include freestream unit Reynolds numbers of 0.6 to 2.2 million per foot and body flap deflections of 15, 20, and 25 deg for an angle-of-attack of 40 deg. The model-scale variance was tested to obtain laminar, transitional, and turbulent heating levels on the defected bodyflaps. In addition, a limited investigation of forced boundary layer transition through the use of discrete roughness elements was performed. Comparisons of the present experimental results to computational predictions and previous experimental data were conducted Laminar, transitional, and turbulent heating levels were observed on the deflected body flap, which compared favorably to the computational results and to the predicted heating based on the flight aerothermodynamic database.

Comparison of analytical and experimental performance of a wind-tunnel diffuser section

NASA Technical Reports Server (NTRS)

Shyne, R. J.; Moore, R. D.; Boldman, D. R.

1986-01-01

Wind tunnel diffuser performance is evaluated by comparing experimental data with analytical results predicted by an one-dimensional integration procedure with skin friction coefficient, a two-dimensional interactive boundary layer procedure for analyzing conical diffusers, and a two-dimensional, integral, compressible laminar and turbulent boundary layer code. Pressure, temperature, and velocity data for a 3.25 deg equivalent cone half-angle diffuser (37.3 in., 94.742 cm outlet diameter) was obtained from the one-tenth scale Altitude Wind Tunnel modeling program at the NASA Lewis Research Center. The comparison is performed at Mach numbers of 0.162 (Re = 3.097x19(6)), 0.326 (Re = 6.2737x19(6)), and 0.363 (Re = 7.0129x10(6)). The Reynolds numbers are all based on an inlet diffuser diameter of 32.4 in., 82.296 cm, and reasonable quantitative agreement was obtained between the experimental data and computational codes.

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

NASA Technical Reports Server (NTRS)

Erickson, Gary E.

1991-01-01

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

Aerodynamic characteristics of the 40- by 80/80- by 120-foot wind tunnel at NASA Ames Research Center

NASA Technical Reports Server (NTRS)

Corsiglia, V. R.; Olson, L. E.; Falarski, M. D.

1984-01-01

The design and testing of vane sets and air-exchange inlet for the 40 x 80/80 x 120-ft wind tunnel at NASA Ames are reported. Boundary-layer analysis and 2D and 3D inviscid panel codes are employed in computer models of the system, and a 1/10-scale 2D facility and a 1/50-scale 3D model of the entire wind tunnel are used in experimental testing of the vane sets. The results are presented in graphs, photographs, drawings, and diagrams are discussed. Generally good agreement is found between the predicted and measured performance.

Drive Motor Improved for 8- by 6-Foot Supersonic Wind Tunnel/9- by 15-Foot Low-Speed Wind Tunnel Complex

NASA Technical Reports Server (NTRS)

2005-01-01

An operational change made recently in the drive motor system for the 8- by 6-Foot Supersonic Wind Tunnel (8x6 SWT)/9- by 15-Foot Low-Speed Wind Tunnel (9x15 LSWT) complex resulted in dramatic power savings and expanded operating range. The 8x6 SWT/9x15 LSWT complex offers a unique combination of wind tunnel conditions for both high- and low-speed testing. Prior to the work discussed in this article, the 8- by 6-ft test section offered airflows ranging from Mach 0.36 to 2.0. Subsonic testing was done in the 9-ft high, 15-ft wide test area in the return leg of the facility. The air speed in this test section can range from 0 to 175 mph (Mach 0.23). In the past, we varied the air speed by using a combination of the compressor speed and the position of the tunnel flow-control doors. When very slow speeds were required in the 9x15 LSWT, these large tunnel flow control doors might be very nearly full open, bleeding off large quantities of air, even with the drive system operating at its previous minimum speed of about 510 rpm. Power drawn during this mode of operation varied between 15 and 18 MW/hr, but clearly much of this power was not being used to provide air that would be used for testing in the test section. The air exiting these large doors represented wasted power. Early this year, the facility's tunnel drive system was run on one motor instead of three to see if lower drive speeds could be achieved that would, in turn, result in large power savings because unnecessary air would not be blown out of the flow-control doors unnecessarily. In addition, if the drive could be run slower, then slower speeds would also be possible in the 8x6 SWT test section as an added benefit. Results of the first tests performed early last year showed that in fact the drive, when operating on only one motor, actually reached a steady-state speed of only 337 rpm and drew an amazingly small 6 MW/hr of electrical power. During daytime operation of the drive, this meant that it would be

ARC-2010-ACD10-0020-073

NASA Image and Video Library

2010-02-10

Lawrence Livermore National Labs (LLNL), Navistar and the Department of Energy conduct tests in the NASA Ames National Full-scale Aerodynamic Complex 80x120_foot wind tunnel. The LLNL project is aimed at aerodynamic truck and trailer devices that can reduce fuel consumption at highway speed by 10 percent. Smoke test demo.

ARC-2010-ACD10-0020-065

NASA Image and Video Library

2010-02-10

Lawrence Livermore National Labs (LLNL), Navistar and the Department of Energy conduct tests in the NASA Ames National Full-scale Aerodynamic Complex 80x120_foot wind tunnel. The LLNL project is aimed at aerodynamic truck and trailer devices that can reduce fuel consumption at highway speed by 10 percent. Smoke test demo.

WIND TUNNEL INVESTIGATION OF THE RESPONSE OF A SONIC ANEMOMETER

EPA Science Inventory

An Applied Technology Inc. (ATI) sonic of the type used by J. C. Kaimal at the Boulder Tower was tested in the large wind tunnel at the U.S. EPA Fluid Modeling Facility. The wind tunnel is approximately 6 ft high, 10 ft wide with a test section bed 60 ft long. The air speed in th...

Wind Tunnel Investigation of Ground Wind Loads for Ares Launch Vehicle

NASA Technical Reports Server (NTRS)

Keller, Donald F.; Ivanco, Thomas G.

2010-01-01

A three year program was conducted at the NASA Langley Research Center (LaRC) Aeroelasticity Branch (AB) and Transonic Dynamics Tunnel (TDT) with the primary objective to acquire scaled steady and dynamic ground-wind loads (GWL) wind-tunnel data for rollout, on-pad stay, and on-pad launch configurations for the Ares I-X Flight Test Vehicle (FTV). The experimental effort was conducted to obtain an understanding of the coupling of aerodynamic and structural characteristics that can result in large sustained wind-induced oscillations (WIO) on such a tall and slender launch vehicle and to generate a unique database for development and evaluation of analytical methods for predicting steady and dynamic GWL, especially those caused by vortex shedding, and resulting in significant WIO. This paper summarizes the wind-tunnel test program that employed two dynamically-aeroelastically scaled GWL models based on the Ares I-X Flight Test Vehicle. The first model tested, the GWL Checkout Model (CM), was a relatively simple model with a secondary objective of restoration and development of processes and methods for design, fabrication, testing, and data analysis of a representative ground wind loads model. In addition, parametric variations in surface roughness, Reynolds number, and protuberances (on/off) were investigated to determine effects on GWL characteristics. The second windtunnel model, the Ares I-X GWL Model, was significantly more complex and representative of the Ares I-X FTV and included the addition of simplified rigid geometrically-scaled models of the Kennedy Space Center (KSC) Mobile Launch Platform (MLP) and Launch Complex 39B primary structures. Steady and dynamic base bending moment as well as model response and steady and unsteady pressure data was acquired during the testing of both models. During wind-tunnel testing of each model, flow conditions (speed and azimuth) where significant WIO occurred, were identified and thoroughly investigated. Scaled data from

An analysis of sound absorbing linings for the interior of the NASA Ames 80 x 120-foot wind tunnel

NASA Technical Reports Server (NTRS)

Wilby, J. F.; White, P. H.

1985-01-01

It is desirable to achieve low frequency sound absorption in the tests section of the NASA Ames 80X120-ft wind tunnel. However, it is difficult to obtain information regarding sound absorption characteristics of potential treatments because of the restrictions placed on the dimensions of the test chambers. In the present case measurements were made in a large enclosure for aircraft ground run-up tests. The normal impedance of the acoustic treatment was measured using two microphones located close to the surface of the treatment. The data showed reasonably good agreement with analytical methods which were then used to design treatments for the wind tunnel test section. A sound-absorbing lining is proposed for the 80X120-ft wind tunnel.

High Reynolds number tests of the CAST 10-2/DOA 2 airfoil in the Langley 0.3-meter transonic cryogenic tunnel, phase 1

NASA Technical Reports Server (NTRS)

Dress, D. A.; Mcguire, P. D.; Stanewsky, E.; Ray, E. J.

1983-01-01

A wind tunnel investigation of an advanced technology airfoil, the CAST 10-2/DOA 2, was conducted in the Langley 0.3 meter Transonic Cryogenic Tunnel (0.3 m TCT). This was the first of a series of tests conducted in a cooperative National Aeronautics and Space Administration (NASA) and the Deutsche Forschungs- und Versuchsanstalt fur Luft- und Raumfahrt e. V. (DFVLR) airfoil research program. Test temperature was varied from 280 K to 100 K to pressures from slightly above 1 to 5.8 atmospheres. Mach number was varied from 0.60 to 0.80, and the Reynolds number (based on airfoil chord) was varied from 4 x 10 to the 8th power to 45 x 10 to the 6th power. This report presents the experimental aerodynamic data obtained for the airfoil and includes descriptions of the airfoil model, the 0.3 m TCT, the test instrumentation, and the testing procedures.

Wind-Tunnel Results of the B-52B with the X-43A Stack

NASA Technical Reports Server (NTRS)

Davis, Mark C.; Sim, Alexander G.; Rhode, Matthew; Johnson, Kevin D., Sr.

2007-01-01

A low-speed wind-tunnel test was performed with a 3%-scale model of a booster rocket mated to an X-43A research vehicle, a combination referred to as the Hyper-X launch vehicle. The test was conducted both in freestream air and in the presence of a partial model of the B-52B airplane. The objectives of the test were to obtain force and moment data to generate structural loads affecting the pylon of the B-52B airplane and to determine the aerodynamic influence of the B-52B on the Hyper-X launch vehicle for evaluating launch separation characteristics. The windtunnel test was conducted at a low-speed wind tunnel in Hampton, Virginia. All moments and forces reported are based either on the aerodynamic influence of the B-52B airplane or are for the Hyper-X launch vehicle in freestream air. Overall, the test showed that the B-52B airplane imparts a strong downwash onto the Hyper-X launch vehicle, reducing the net lift of the Hyper-X launch vehicle. Pitching and rolling moments are also imparted onto the booster and are a strong function of the launch-drop angle of attack.

ARC-2010-ACD10-0020-082

NASA Image and Video Library

2010-02-10

Lawrence Livermore National Labs (LLNL), Navistar and the Department of Energy conduct tests in the NASA Ames National Full-scale Aerodynamic Complex 80x120_foot wind tunnel. The LLNL project is aimed at aerodynamic truck and trailer devices that can reduce fuel consumption at highway speed by 10 percent. Smoke test demo with Ron Schoon, Navistar.

ARC-2010-ACD10-0020-034

NASA Image and Video Library

2010-02-10

Lawrence Livermore National Labs (LLNL), Navistar and the Department of Energy conduct tests in the NASA Ames National Full-scale Aerodynamic Complex 80x120_foot wind tunnel. The LLNL project is aimed at aerodynamic truck and trailer devices that can reduce fuel consumption at highway speed by 10 percent. LLNL's test piece is being installed on truck.

ARC-2010-ACD10-0020-079

NASA Image and Video Library

2010-02-10

Lawrence Livermore National Labs (LLNL), Navistar and the Department of Energy conduct tests in the NASA Ames National Full-scale Aerodynamic Complex 80x120_foot wind tunnel. The LLNL project is aimed at aerodynamic truck and trailer devices that can reduce fuel consumption at highway speed by 10 percent. Smoke test demo with Ron Schoon, Navistar.

Low-speed wind-tunnel results for symmetrical NASA LS(1)-0013 airfoil

NASA Technical Reports Server (NTRS)

Ferris, James C.; Mcghee, Robert J.; Barnwell, Richard W.

1987-01-01

A wind-tunnel test has been conducted in the Langley Low-Turbulence Pressure Tunnel to evaluate the performance of a symmetrical NASA LS(1)-0013 airfoil which is a 13-percent-thick, low-speed airfoil. The airfoil contour was obtained from the thickness distribution of a 13-percent-thick, high-performance airfoil developed for general aviation airplanes. The tests were conducted at Mach numbers from 0.10 tp 0.37 over a Reynolds number range from about 0.6 to 12.0 X 10 to the 6th power. The angle of attack varied from about -8 to 20 degrees. The results indicate that the aerodynamic characteristics of the present airfoil are similar to, but slightly better than, those of the NACA 0012 airfoil.

Forced Boundary-Layer Transition on X-43 (Hyper-X) in NASA LaRC 31-Inch Mach 10 Air Tunnel

NASA Technical Reports Server (NTRS)

Berry, Scott A.; DiFulvio, Michael; Kowalkowski, Matthew K.

2000-01-01

Aeroheating and boundary layer transition characteristics for the X-43 (Hyper-X) configuration have been experimentally examined in the Langley 31-Inch Mach 10 Air Tunnel. Global surface heat transfer distributions, and surface streamline patterns were measured on a 0.333-scale model of the Hyper-X forebody. Parametric variations include angles-of-attack of 0-deg, 2-deg, 3-deg, and 4-deg; Reynolds numbers based on model length of 1.2 to 5.1 million; and inlet cowl door both open and closed. The effects of discrete roughness elements on the forebody boundary layer, which included variations in trip configuration and height, were investigated. This document is intended to serve as a release of preliminary data to the Hyper-X program; analysis is limited to observations of the experimental trends in order to expedite dissemination.

Drive System Enhancement in the NASA Lewis Research Center Supersonic Wind Tunnels

NASA Technical Reports Server (NTRS)

Becks, Edward A.

1998-01-01

An overview of NASA Lewis' Aeropropulsion Wind Tunnel Productivity Improvements was presented at the 19th AIAA Advanced Measurement & Ground Testing Technology Conference. Since that time Lewis has implemented subsonic operation in their 10- by 10-Foot Supersonic Wind Tunnel as had been proven viable in the 8- by 6 and 9- by 15-Foot Wind Tunnel Complex and discussed at the aforementioned conference. In addition, two more years of data have been gathered to help quantify the true productivity increases in these facilities attributable to the drive system and operational improvements. This paper was invited for presentation at the 20th Advanced Measurement and Ground Testing Conference to discuss and quantify the productivity improvements in the 10- by 10 SWT since the implementation of less than full complement motor operation. An update on the increased productivity at the 8- by 6 and 9- by 15-Foot facility due to drive system enhancements will also be presented.

Smart wing wind tunnel test results

NASA Astrophysics Data System (ADS)

Scherer, Lewis B.; Martin, Christopher A.; Appa, Kari; Kudva, Jayanth N.; West, Mark N.

1997-05-01

The use of smart materials technologies can provide unique capabilities in improving aircraft aerodynamic performance. Northrop Grumman built and tested a 16% scale semi-span wind tunnel model of the F/A-18 E/F for the on-going DARPA/WL Smart Materials and Structures-Smart Wing Program. Aerodynamic performance gains to be validated included increase in the lift to drag ratio, increased pitching moment (Cm), increased rolling moment (Cl) and improved pressure distribution. These performance gains were obtained using hingeless, contoured trailing edge control surfaces with embedded shape memory alloy (SMA) wires and spanwise wing twist via a SMA torque tube and are compared to a conventional wind tunnel model with hinged control surfaces. This paper presents an overview of the results from the first wind tunnel test performed at the NASA Langley's 16 ft Transonic Dynamic Tunnel. Among the benefits demonstrated are 8 - 12% increase in rolling moment due to wing twist, a 10 - 15% increase in rolling moment due to contoured aileron, and approximately 8% increase in lift due to contoured flap, and improved pressure distribution due to trailing edge control surface contouring.

A wood-strand material for wind erosion control: effects on total sediment loss, PM10 vertical flux, and PM10 loss.

PubMed

Copeland, N S; Sharratt, B S; Wu, J Q; Foltz, R B; Dooley, J H

2009-01-01

Fugitive dust from eroding land poses risks to environmental quality and human health, and thus, is regulated nationally based on ambient air quality standards for particulate matter with mean aerodynamic diameter < or = 10 microm (PM10) established in the Clean Air Act. Agricultural straw has been widely used for rainfall-induced erosion control; however, its performance for wind erosion mitigation has been less studied, in part because straw is mobile at moderate wind velocities. A wood-based long-strand material has been developed for rainfall-induced erosion control and has shown operational promise for control of wind-induced erosion and dust emissions from disturbed sites. The purpose of this study was to evaluate the efficacy of both agricultural straw and wood-strand materials in controlling wind erosion and fugitive dust emissions under laboratory conditions. Wind tunnel tests were conducted to compare wood strands of several geometries to agricultural wheat straw and bare soil in terms of total sediment loss, PM10 vertical flux, and PM10 loss. Results indicate that the types of wood strands tested are stable at wind speeds of up to 18 m s(-1), while wheat straw is only stable at speeds of up to 6.5 m s(-1). Wood strands reduced total sediment loss and PM10 emissions by 90% as compared to bare soil across the range of wind speeds tested. Wheat straw did not reduce total sediment loss for the range of speeds tested, but did reduce PM10 emissions by 75% compared to a bare soil at wind speeds of up to 11 m s(-1).

The aeolian wind tunnel

NASA Technical Reports Server (NTRS)

Iversen, J. D.

1991-01-01

The aeolian wind tunnel is a special case of a larger subset of the wind tunnel family which is designed to simulate the atmospheric surface layer winds to small scale (a member of this larger subset is usually called an atmospheric boundary layer wind tunnel or environmental wind tunnel). The atmospheric boundary layer wind tunnel is designed to simulate, as closely as possible, the mean velocity and turbulence that occur naturally in the atmospheric boundary layer (defined as the lowest portion of the atmosphere, of the order of 500 m, in which the winds are most greatly affected by surface roughness and topography). The aeolian wind tunnel is used for two purposes: to simulate the physics of the saltation process and to model at small scale the erosional and depositional processes associated with topographic surface features. For purposes of studying aeolian effects on the surface of Mars and Venus as well as on Earth, the aeolian wind tunnel continues to prove to be a useful tool for estimating wind speeds necessary to move small particles on the three planets as well as to determine the effects of topography on the evolution of aeolian features such as wind streaks and dune patterns.

A throat-bypass stability-bleed system using relief valves to increase the transient stability of a mixed-compression inlet. [YF-12 aircraft inlet tests in the Lewis 10 by 10 ft supersonic wind tunnel

NASA Technical Reports Server (NTRS)

Neiner, G. H.; Dustin, M. O.; Cole, G. L.

1979-01-01

A stability-bleed system was installed in a YF-12 flight inlet that was subjected to internal and external airflow disturbances in the NASA Lewis 10 by 10 foot supersonic wind tunnel. The purpose of the system is to allow higher inlet performance while maintaining a substantial tolerance (without unstart) to internal and external disturbances. At Mach numbers of 2.47 and 2.76, the inlet tolerance to decreases in diffuser-exit corrected airflow was increased by approximately 10 percent of the operating-point airflow. The stability-bleed system complemented the terminal-shock-control system of the inlet and did not show interaction problems. For disturbances which caused a combined decrease in Mach number and increase in angle of attack, the system with valves operative kept the inlet started 4 to 28 times longer than with the valves inoperative. Hence, the stability system provides additional time for the inlet control system to react and prevent unstart. This was observed for initial Mach numbers of 2.55 and 2.68. For slow increase in angle of attack at Mach 2.47 and 2.76, the system kept the inlet started beyond the steady-state unstart angle. However, the maximum transient angles of attack without unstart could not be determined because wind-tunnel mechanical-stop limits for angle of attack were reached.

North American Aviation F-100 in the Ames 40x80 Foot Wind Tunnel.

NASA Image and Video Library

1959-02-12

North American F-100-F airplane, equipped with thrust reversers, full scale wind tunnel test. 3/4 front view of F-100-F airplane with North American Aviation thrust reverser. On standard 40x80 struts landing gear down. Mark Kelly, branch chief in photo.

Avrocar Test in Ames 40x80 Foot Wind Tunnel.

NASA Image and Video Library

1961-04-03

Rear view of the Avrocar with tail, mounted on variable height struts. Overhead doors of the wind tunnel test section open. The first Avrocar, S/N 58-7055 (marked AV-7055), after tethered testing, became the "wind tunnel" test model at NASA Ames, where it remained in storage from 1961 until 1966, when it was donated to the National Air and Space Museum, in Suitland, Maryland.

Wind Tunnel Modeling Of Wind Flow Over Complex Terrain

NASA Astrophysics Data System (ADS)

Banks, D.; Cochran, B.

2010-12-01

This presentation will describe the finding of an atmospheric boundary layer (ABL) wind tunnel study conducted as part of the Bolund Experiment. This experiment was sponsored by Risø DTU (National Laboratory for Sustainable Energy, Technical University of Denmark) during the fall of 2009 to enable a blind comparison of various air flow models in an attempt to validate their performance in predicting airflow over complex terrain. Bohlund hill sits 12 m above the water level at the end of a narrow isthmus. The island features a steep escarpment on one side, over which the airflow can be expected to separate. The island was equipped with several anemometer towers, and the approach flow over the water was well characterized. This study was one of only two only physical model studies included in the blind model comparison, the other being a water plume study. The remainder were computational fluid dynamics (CFD) simulations, including both RANS and LES. Physical modeling of air flow over topographical features has been used since the middle of the 20th century, and the methods required are well understood and well documented. Several books have been written describing how to properly perform ABL wind tunnel studies, including ASCE manual of engineering practice 67. Boundary layer wind tunnel tests are the only modelling method deemed acceptable in ASCE 7-10, the most recent edition of the American Society of Civil Engineers standard that provides wind loads for buildings and other structures for buildings codes across the US. Since the 1970’s, most tall structures undergo testing in a boundary layer wind tunnel to accurately determine the wind induced loading. When compared to CFD, the US EPA considers a properly executed wind tunnel study to be equivalent to a CFD model with infinitesimal grid resolution and near infinite memory. One key reason for this widespread acceptance is that properly executed ABL wind tunnel studies will accurately simulate flow separation

Wind tunnel results for a high-speed, natural laminar-flow airfoil designed for general aviation aircraft

NASA Technical Reports Server (NTRS)

Sewall, William G.; Mcghee, Robert J.; Viken, Jeffery K.; Waggoner, Edgar G.; Walker, Betty S.; Millard, Betty F.

1985-01-01

Two dimensional wind tunnel tests were conducted on a high speed natural laminar flow airfoil in both the Langley 6 x 28 inch Transonic Tunnel and the Langley Low Turbulence Pressure Tunnel. The test conditions consisted of Mach numbers ranging from 0.10 to 0.77 and Reynolds numbers ranging from 3 x 1 million to 11 x 1 million. The airfoil was designed for a lift coefficient of 0.20 at a Mach number of 0.70 and Reynolds number of 11 x 1 million. At these conditions, laminar flow would extend back to 50 percent chord of the upper surface and 70 percent chord of the lower surface. Low speed results were also obtained with a 0.20 chord trailing edge split flap deflected 60 deg.

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.

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.

Softwall acoustical characteristics and measurement capabilities of the NASA Lewis 9x15 foot low speed wind tunnel

NASA Technical Reports Server (NTRS)

Rentz, P. E.

1976-01-01

Acoustical characteristics and source directionality measurement capabilities of the wind tunnel in the softwall configuration were evaluated, using aerodynamically clean microphone supports. The radius of measurement was limited by the size of the test section, instead of the 3.0 foot (1 m) limitation of the hardwall test section. The wind-on noise level in the test section was reduced 10 dB. Reflections from the microphone support boom, after absorptive covering, induced measurement errors in the lower frequency bands. Reflections from the diffuser back wall were shown to be significant. Tunnel noise coming up the diffuser was postulated as being responsible, at least partially, for the wind-on noise in the test section and settling chamber. The near field characteristics of finite-sized sources and the theoretical response of a porous strip sensor in the presence of wind are presented.

Kaman K-16 in 40x80 Foot Wind Tunnel at Ames Research Center.

NASA Image and Video Library

1962-09-19

Test No. 175 Kaman K-16 in 40x80 Foot Wind Tunnel at Ames Research Center. Kaman K-16B was an experimental tilt wing aircraft, it used the fuselage of a JRF-5 and was powered by two General Electric YT58-GE-2A engines.

Wind tunnel

NASA Technical Reports Server (NTRS)

Wilson, E. M. (Inventor)

1969-01-01

A supersonic wind wind tunnel is described for testing several air foils mounted in a row. A test section of a wind tunnel contains means for mounting air foil sections in a row, means for rotating each section about an axis so that the angle of attack of each section changes with the other sections, and means for rotating the row with respect to the air stream so that the row forms an oblique angle with the air stream.

Computational design of low aspect ratio wing-winglets for transonic wind-tunnel testing

NASA Technical Reports Server (NTRS)

Kuhlman, John M.; Brown, Christopher K.

1989-01-01

A computational design has been performed for three different low aspect ratio wing planforms fitted with nonplanar winglets; one of the three planforms has been selected to be constructed as a wind tunnel model for testing in the NASA LaRC 7 x 10 High Speed Wind Tunnel. A design point of M = 0.8, CL approx = 0.3 was selected, for wings of aspect ratio equal to 2.2, and leading edge sweep angles of 45 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. This report summarizes the design process and the predicted transonic performance for each configuration.

An evaluation of proposed acoustic treatments for the NASA LaRC 4 x 7 meter wind tunnel

NASA Technical Reports Server (NTRS)

Abrahamson, A. L.

1985-01-01

The NASA LaRC 4 x 7 Meter Wind Tunnel is an existing facility specially designed for powered low speed (V/STOL) testing of large scale fixed wing and rotorcraft models. The enhancement of the facility for scale model acoustic testing is examined. The results are critically reviewed and comparisons are drawn with a similar wind tunnel (the DNW Facility in the Netherlands). Discrepancies observed in the comparison stimulated a theoretical investigation using the acoustic finite element ADAM System, of the ways in which noise propagating around the tunnel circuit radiates into the open test section. The reasons for the discrepancies noted above are clarified and assists in the selection of acoustic treatment options for the facility.

Hyper-X Stage Separation Wind Tunnel Test Program

NASA Technical Reports Server (NTRS)

Woods, W. C.; Holland, S. D.; DiFulvio, M.

2000-01-01

NASA's Hyper-X research program was developed primarily to flight demonstrate a supersonic combustion ramjet engine, fully integrated with a forebody designed to tailor inlet flow conditions and a free expansion nozzle/afterbody to produce positive thrust at design flight conditions. With a point-designed propulsion system, the vehicle must depend upon some other means for boost to its design flight condition. Clean separation from this initial propulsion system stage within less than a second is critical to the success of the flight. This paper discusses the early planning activity, background, and chronology that developed the series of wind tunnel tests to support multi degree of freedom simulation of the separation process. Representative results from each series of tests are presented and issues and concerns during the process and current status will be highlighted.

Hyper-X Stage Separation Wind-Tunnel Test Program

NASA Technical Reports Server (NTRS)

Woods, William C.; Holland, Scott D.; DiFulvio, Michael

2001-01-01

NASA's Hyper-X research program was developed primarily to flight demonstrate a supersonic combustion ramjet engine, fully integrated with a forebody designed to tailor inlet flow conditions and a free expansion nozzle/afterbody to produce positive thrust at design flight conditions. With a point-designed propulsion system the vehicle must depend on some other means for boost to its design flight condition. Clean separation from this initial propulsion system stage within less than a second is critical to the success of the flight. This paper discusses the early planning activity, background, and chronology that developed the series of wind-tunnel tests to support multi-degree-of-freedom simulation of the separation process. Representative results from each series of tests are presented, and issues and concerns during the process and current status are highlighted.

Hyper-X Storage Separation Wind Tunnel Test Program

NASA Technical Reports Server (NTRS)

Woods, William C.; Holland, Scott D.; Difulvio, Michael

2000-01-01

NASA's Hyper-X research program was developed primarily to flight demonstrate a supersonic combustion ramjet engine, fully integrated with a forebody designed to tailor inlet flow, conditions and a free expansion nozzle/afterbody to produce positive thrust at design flight conditions. With a point-designed propulsion system, the vehicle must depend upon some other means for boost to its design flight condition. Clean separation from this initial propulsion system stage within less than a second is critical to the success of the flight. This paper discusses the early planning activity, background, and chronology that developed the series of wind tunnel tests to support multi degree of freedom simulation of the separation process. Representative results from each series of tests are presented and issues and concerns during the process and current status will be highlighted.

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.

Static Aeroelastic Scaling and Analysis of a Sub-Scale Flexible Wing Wind Tunnel Model

NASA Technical Reports Server (NTRS)

Ting, Eric; Lebofsky, Sonia; Nguyen, Nhan; Trinh, Khanh

2014-01-01

This paper presents an approach to the development of a scaled wind tunnel model for static aeroelastic similarity with a full-scale wing model. The full-scale aircraft model is based on the NASA Generic Transport Model (GTM) with flexible wing structures referred to as the Elastically Shaped Aircraft Concept (ESAC). The baseline stiffness of the ESAC wing represents a conventionally stiff wing model. Static aeroelastic scaling is conducted on the stiff wing configuration to develop the wind tunnel model, but additional tailoring is also conducted such that the wind tunnel model achieves a 10% wing tip deflection at the wind tunnel test condition. An aeroelastic scaling procedure and analysis is conducted, and a sub-scale flexible wind tunnel model based on the full-scale's undeformed jig-shape is developed. Optimization of the flexible wind tunnel model's undeflected twist along the span, or pre-twist or wash-out, is then conducted for the design test condition. The resulting wind tunnel model is an aeroelastic model designed for the wind tunnel test condition.

A wind tunnel for measuring selenium volatilization under field-like conditions

DOE Office of Scientific and Technical Information (OSTI.GOV)

Dungan, R.S.; Stork, A.; Frankenberger, W.T. Jr.

2000-04-01

A wind tunnel was developed to measure the loss of volatile selenium (Se) from soil under field-like conditions. The wind tunnel consisted of a volatilization chamber made of Plexiglas (2.4 m long x 0.8 m wide x 1.2 m high), which was set above a stainless steel lysimeter (0.5 m{sup 2} surface area x 0.8 m deep). High air exchange rates (avg. 700 air changes h{sup {minus}1}) were used to simulate field-like environmental conditions inside the wind tunnel. To monitor the fate and transport of the Se, radiolabeled {sup 75}Se as sodium selenite (Na{sub 2}{sup 75}SeO{sub 3}) was incorporated intomore » the top 5 cm of soil. Volatile Se was trapped on activated carbon filters and measured directly using gamma counting. A 135-d bare-soil experiment was carried out, during which 2.7% of the Se was released through volatilization without added C. The average flux rate of gaseous Se was 17 mg m{sup {minus}2} h{sup {minus}1}, with a high of 72 mg m{sup {minus}2} h{sup {minus}1} occurring on Day 6. After 135 d, 59 and 43% of the {sup 75}Se was located in the 0 to 5 and 5 to 10 cm soil layers, respectively. A total of 84.5% of all applied {sup 75}Se was recovered. The purpose of this study was to improve estimates on Se volatilization from seleniferous soils and help close the gap between previous laboratory and field experiments.« less

The self streamlining wind tunnel. [wind tunnel walls

NASA Technical Reports Server (NTRS)

Goodyer, M. J.

1975-01-01

A two dimensional test section in a low speed wind tunnel capable of producing flow conditions free from wall interference is presented. Flexible top and bottom walls, and rigid sidewalls from which models were mounted spanning the tunnel are shown. All walls were unperforated, and the flexible walls were positioned by screw jacks. To eliminate wall interference, the wind tunnel itself supplied the information required in the streamlining process, when run with the model present. Measurements taken at the flexible walls were used by the tunnels computer check wall contours. Suitable adjustments based on streamlining criteria were then suggested by the computer. The streamlining criterion adopted when generating infinite flowfield conditions was a matching of static pressures in the test section at a wall with pressures computed for an imaginary inviscid flowfield passing over the outside of the same wall. Aerodynamic data taken on a cylindrical model operating under high blockage conditions are presented to illustrate the operation of the tunnel in its various modes.

Design and optimization of resistance wire electric heater for hypersonic wind tunnel

NASA Astrophysics Data System (ADS)

Rehman, Khurram; Malik, Afzaal M.; Khan, I. J.; Hassan, Jehangir

2012-06-01

The range of flow velocities of high speed wind tunnels varies from Mach 1.0 to hypersonic order. In order to achieve such high speed flows, a high expansion nozzle is employed in the converging-diverging section of wind tunnel nozzle. The air for flow is compressed and stored in pressure vessels at temperatures close to ambient conditions. The stored air is dried and has minimum amount of moisture level. However, when this air is expanded rapidly, its temperature drops significantly and liquefaction conditions can be encountered. Air at near room temperature will liquefy due to expansion cooling at a flow velocity of more than Mach 4.0 in a wind tunnel test section. Such liquefaction may not only be hazardous to the model under test and wind tunnel structure; it may also affect the test results. In order to avoid liquefaction of air, a pre-heater is employed in between the pressure vessel and the converging-diverging section of a wind tunnel. A number of techniques are being used for heating the flow in high speed wind tunnels. Some of these include the electric arc heating, pebble bed electric heating, pebble bed natural gas fired heater, hydrogen burner heater, and the laser heater mechanisms. The most common are the pebble bed storage type heaters, which are inefficient, contaminating and time consuming. A well designed electrically heating system can be efficient, clean and simple in operation, for accelerating the wind tunnel flow up to Mach 10. This paper presents CFD analysis of electric preheater for different configurations to optimize its design. This analysis has been done using ANSYS 12.1 FLUENT package while geometry and meshing was done in GAMBIT.

Reduction of background noise induced by wind tunnel jet exit vanes

NASA Technical Reports Server (NTRS)

Martin, R. M.; Brooks, T. F.; Hoad, D. R.

1985-01-01

The NASA-Langley 4 x 7 m wind tunnel develops low frequency flow pulsations at certain velocity ranges during open throat mode operation, affecting the aerodynamics of the flow and degrading the resulting model test data. Triangular vanes attached to the trailing edge of flat steel rails, mounted 10 cm from the inside of the jet exit walls, have been used to reduce this effect; attention is presently given to methods used to reduce the inherent noise generation of the vanes while retaining their pulsation reduction features.

Kaman K-16 in 40x80 Foot Wind Tunnel at Ames Research Center.

NASA Image and Video Library

1962-09-19

Test No. 175 Kaman K-16 in 40x80 Foot Wind Tunnel at Ames Research Center. Pictured with two Kaman employees. 3/4 Front view of Airplane. Kaman K-16B was an experimental tilt wing aircraft, it used the fuselage of a JRF-5 and was powered by two General Electric YT58-GE-2A engines.

Kaman K-16 in 40x80 Foot Wind Tunnel at Ames Research Center.

NASA Image and Video Library

1962-09-19

Test No. 175 Kaman K-16 being lowered into the 40x80 foot wind tunnel at NASA's Ames Research Center, viewed from the front. Kaman K-16B was an experimental tilt wing aircraft, it used the fuselage of a JRF-5 and was powered by two General Electric YT58-GE-2A engines.

Enhancement of tunneling magnetoresistance by inserting a diffusion barrier in L10-FePd perpendicular magnetic tunnel junctions

NASA Astrophysics Data System (ADS)

Zhang, De-Lin; Schliep, Karl B.; Wu, Ryan J.; Quarterman, P.; Reifsnyder Hickey, Danielle; Lv, Yang; Chao, Xiaohui; Li, Hongshi; Chen, Jun-Yang; Zhao, Zhengyang; Jamali, Mahdi; Mkhoyan, K. Andre; Wang, Jian-Ping

2018-04-01

We studied the tunnel magnetoresistance (TMR) of L10-FePd perpendicular magnetic tunnel junctions (p-MTJs) with an FePd free layer and an inserted diffusion barrier. The diffusion barriers studied here (Ta and W) were shown to enhance the TMR ratio of the p-MTJs formed using high-temperature annealing, which are necessary for the formation of high quality L10-FePd films and MgO barriers. The L10-FePd p-MTJ stack was developed with an FePd free layer with a stack of FePd/X/Co20Fe60B20, where X is the diffusion barrier, and patterned into micron-sized MTJ pillars. The addition of the diffusion barrier was found to greatly enhance the magneto-transport behavior of the L10-FePd p-MTJ pillars such that those without a diffusion barrier exhibited negligible TMR ratios (<1.0%), whereas those with a Ta (W) diffusion barrier exhibited TMR ratios of 8.0% (7.0%) at room temperature and 35.0% (46.0%) at 10 K after post-annealing at 350 °C. These results indicate that diffusion barriers could play a crucial role in realizing high TMR ratios in bulk p-MTJs such as those based on FePd and Mn-based perpendicular magnetic anisotropy materials for spintronic applications.

The 1 × 1 m hypersonic wind tunnel Kochel/Tullahoma 1940-1960

NASA Astrophysics Data System (ADS)

Eckardt, Dietrich

2015-03-01

Peenemünde and Cape Canaveral mark cornerstones of space history. Kochel in Southern Germany and Tullahoma in Tennessee, USA also belong in this category. The technically unique Kochel wind tunnel was part of the German long-distance missile development strategy, planned and prepared in secret before the beginning of World War II. A 57 MW closed-circuit wind tunnel facility with 1 × 1 m measuring section was planned for continuous-flow simulation at high Mach numbers Ma 7-10. In the early 1940 s a site beside the Walchensee Power Station at Kochel am See in Upper Bavaria, Germany was chosen to provide the required altitude difference of 200 m for the hydraulic turbine drives. The preparatory activities for the erection of this impressive hypersonic wind tunnel facility were pushed ahead until an enforced temporary pause in September 1944. In early May 1945 US troops occupied the area and, in due course, scientists of General Arnold's Scientific Advisory Group, the `von Kármán team', ordered the transfer to the USA of available equipment, design materials and other paperwork. Here, at the Arnold Engineering Development Center (AEDC) Tullahoma, TN this `Tunnel A' was built to begin operation around 1957. The testing was conducted on the Mach 7 experimental aircraft X-15, space shuttle developments and still secret investigations on unmanned hypersonic vehicles.

Real-time data reduction capabilities at the Langley 7 by 10 foot high speed tunnel

NASA Technical Reports Server (NTRS)

Fox, C. H., Jr.

1980-01-01

The 7 by 10 foot high speed tunnel performs a wide range of tests employing a variety of model installation methods. To support the reduction of static data from this facility, a generalized wind tunnel data reduction program had been developed for use on the Langley central computer complex. The capabilities of a version of this generalized program adapted for real time use on a dedicated on-site computer are discussed. The input specifications, instructions for the console operator, and full descriptions of the algorithms are included.

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.

1988-01-01

A computational design has been performed for three different low aspect ratio wing planforms fitted with nonplanar winglets; one of the three planforms has been selected to be constructed as a wind tunnel model for testing in the NASA LaRC 7 x 10 High Speed Wind Tunnel. A design point of M = 0.8, CL approx = 0.3 was selected, for wings of aspect ratio equal to 2.2, and leading edge sweep angles of 45 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. This report summarizes the design process and the predicted transonic performance for each configuration.

X-33 Rev-F Turbulent Aeroheating Results From Test 6817 in NASA Langley 20-Inch Mach 6 Air Tunnel and Comparisons With Computations

NASA Technical Reports Server (NTRS)

Hollis, Brian R.; Horvath, Thomas J.; Berry, Scott A.

2003-01-01

Measurements and predictions of the X-33 turbulent aeroheating environment have been performed at Mach 6, perfect-gas air conditions. The purpose of this investigation was to compare measured turbulent aeroheating levels on smooth models, models with discrete trips, and models with arrays of bowed panels (which simulate bowed thermal protections system tiles) with each other and with predictions from two Navier-Stokes codes, LAURA and GASP. The wind tunnel testing was conducted at free stream Reynolds numbers based on length of 1.8 x 10(exp 6) to 6.1 x 10(exp 6) on 0.0132 scale X-33 models at a = 40-deg. Turbulent flow was produced by the discrete trips and by the bowed panels at ill but the lowest Reynolds number, but turbulent flow on the smooth model was produced only at the highest Reynolds number. Turbulent aeroheating levels on each of the three model types were measured using global phosphor thermography and were found to agree to within .he estimated uncertainty (plus or minus 15%) of the experiment. Computations were performed at the wind tunnel free stream conditions using both codes. Turbulent aeroheating levels predicted using the LAURA code were generally 5%-10% lower than those from GASP, although both sets of predictions fell within the experimental accuracy of the wind tunnel data.

An Experimental Study of the Ground Transportation System (GTS) Model in the NASA Ames 7- by 10-Ft Wind Tunnel

NASA Technical Reports Server (NTRS)

Storms, Bruce L.; Ross, James C.; Heineck, James T.; Walker, Stephen M.; Driver, David M.; Zilliac, Gregory G.; Bencze, Daniel P. (Technical Monitor)

2001-01-01

The 1/8-scale Ground Transportation System (GTS) model was studied experimentally in the NASA Ames 7- by 10-Ft Wind Tunnel. Designed for validation of computational fluid dynamics (CFD), the GTS model has a simplified geometry with a cab-over-engine design and no tractor-trailer gap. As a further simplification, all measurements of the GTS model were made without wheels. Aerodynamic boattail plates were also tested on the rear of the trailer to provide a simple geometry modification for computation. The experimental measurements include body-axis drag, surface pressures, surface hot-film anemometry, oil-film interferometry, and 3-D particle image velocimetry (PIV). The wind-averaged drag coefficient with and without boattail plates was 0.225 and 0.277, respectively. PIV measurements behind the model reveal a significant reduction in the wake size due to the flow turning provided by the boattail plates. Hot-film measurements on the side of the cab indicate laminar separation with turbulent reattachment within 0.08 trailer width for zero and +/- 10 degrees yaw. Oil film interferometry provided quantitative measurements of skin friction and qualitative oil flow images. A complete set of the experimental data and the surface definition of the model are included on a CD-ROM for further analysis and comparison.

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;

Construction of the Fan Drive Enclosure of the 40x80 Foot Wind Tunnel at Ames.

NASA Image and Video Library

1943-08-26

Concrete frame enclosing the fan drive bents of the 40x80 foot wind tunnel at ames. Once installed, six 40-foot-diameter fans, each powered by a 6000-horsepower electric motor maintained airflow at 230 mph or less (these are still tornado velocities).

Study of the integration of wind tunnel and computational methods for aerodynamic configurations

NASA Technical Reports Server (NTRS)

Browne, Lindsey E.; Ashby, Dale L.

1989-01-01

A study was conducted to determine the effectiveness of using a low-order panel code to estimate wind tunnel wall corrections. The corrections were found by two computations. The first computation included the test model and the surrounding wind tunnel walls, while in the second computation the wind tunnel walls were removed. The difference between the force and moment coefficients obtained by comparing these two cases allowed the determination of the wall corrections. The technique was verified by matching the test-section, wall-pressure signature from a wind tunnel test with the signature predicted by the panel code. To prove the viability of the technique, two cases were considered. The first was a two-dimensional high-lift wing with a flap that was tested in the 7- by 10-foot wind tunnel at NASA Ames Research Center. The second was a 1/32-scale model of the F/A-18 aircraft which was tested in the low-speed wind tunnel at San Diego State University. The panel code used was PMARC (Panel Method Ames Research Center). Results of this study indicate that the proposed wind tunnel wall correction method is comparable to other methods and that it also inherently includes the corrections due to model blockage and wing lift.

GE Fan in Wing VZ-11 VTOL airplane in Ames 40x80 Foot Wind Tunnel.

NASA Image and Video Library

1962-12-27

3/4 front view VZ-11 ground test - variable height struts. Engines of the VZ-11 are a pair of General Electric J85-5 turbojets, mounted in high in the centre fuselage, well away from fan disturbance. Designed in the Ames 40x80 foot wind tunnel.

Sonic Boom Computations for a Mach 1.6 Cruise Low Boom Configuration and Comparisons with Wind Tunnel Data

NASA Technical Reports Server (NTRS)

Elmiligui, Alaa A.; Cliff, Susan E.; Wilcox, Floyd; Nemec, Marian; Bangert, Linda; Aftosmis, Michael J.; Parlette, Edward

2011-01-01

Accurate analysis of sonic boom pressure signatures using computational fluid dynamics techniques remains quite challenging. Although CFD shows accurate predictions of flow around complex configurations, generating grids that can resolve the sonic boom signature far away from the body is a challenge. The test case chosen for this study corresponds to an experimental wind-tunnel test that was conducted to measure the sonic boom pressure signature of a low boom configuration designed by Gulfstream Aerospace Corporation. Two widely used NASA codes, USM3D and AERO, are examined for their ability to accurately capture sonic boom signature. Numerical simulations are conducted for a free-stream Mach number of 1.6, angle of attack of 0.3 and Reynolds number of 3.85x10(exp 6) based on model reference length. Flow around the low boom configuration in free air and inside the Langley Unitary plan wind tunnel are computed. Results from the numerical simulations are compared with wind tunnel data. The effects of viscous and turbulence modeling along with tunnel walls on the computed sonic boom signature are presented and discussed.

Lockheed XFV-1 model in the 40x80 foot Wind Tunnel at NASA Ames Research Center.

NASA Image and Video Library

1952-05-16

Lockheed XFV-1 model. Project engineer Mark Kelly (not shown). Remote controlled model flown in the settling chamber of the 40x80 wind tunnel. Electric motors in the model, controlled the counter-rotating propellers to test vertical takeoff. Test no. 71

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 of...

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 of...

Modifications to the 4x7 meter tunnel for acoustic research: Engineering feasibility study

NASA Technical Reports Server (NTRS)

1986-01-01

The NASA-Langley Research Center 4 x 7 Meter Low Speed Wind Tunnel is currently being used for low speed aerodynamics, V/STOL aerodynamics and, to a limited extent, rotorcraft noise research. The deficiencies of this wind tunnel for both aerodynamics and aeroacoustics research have been recognized for some time. Modifications to the wind tunnel are being made to improve the test section flow quality and to update the model cart systems. A further modification of the 4 x 7 Meter Wind Tunnel to permit rotorcraft model acoustics research has been proposed. As a precursor to the design of the proposed modifications, NASA is conducted both in-house and contracted studies to define the acoustic environment within the wind tunnel and to provide recommendations or the reduction of the wind tunnel background noise to a level acceptable to acoustics researchers. One of these studies by an acoustics consultant, has produced the primary reference documents that define the wind tunnel noise sources and outline recommended solutions.

Investigation in the 7-by-10 Foot Wind Tunnel of Ducts for Cooling Radiators within an Airplane Wing

NASA Technical Reports Server (NTRS)

Harris, Thomas A; Recant, Isidore G

1942-01-01

Report presents the results of an investigation made in the NACA 7 by 10-foot wind tunnel of a large-chord wing model with a duct to house a simulated radiator suitable for a liquid-cooled engine. The duct was expanded to reduce the radiator losses, and the installation of the duct and radiator was made entirely within the wing to reduce form and interference drag. The tests were made using a two-dimensional-flow setup with a full-span duct and radiator. Section aerodynamic characteristics of the basic airfoil are given and also curves showing the characteristics of the various duct-radiator combinations. An expression for efficiency, the primary criterion of merit of any duct, and the effect of the several design parameters of the duct-radiator arrangement are discussed. The problem of throttling is considered and a discussion of the power required for cooling is included.

Two-dimensional wind tunnel

NASA Technical Reports Server (NTRS)

1982-01-01

Information on the Japanese National Aerospace Laboratory two dimensional transonic wind tunnel, completed at the end of 1979 is presented. Its construction is discussed in detail, and the wind tunnel structure, operation, test results, and future plans are presented.

ARC-1969-AAL-5993. Six, 40-Foot-Diameter Fans in the Ames 40x80 Foot Wind Tunnel.

NASA Image and Video Library

1944-06-09

Motor and propeller blades in 40x80ft wind tunnel. Six 40-foot-diameter fans, each powered by a 6000-horsepower electric motor maintained airflow at 230 mph or less (these are still tornado velocities).

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

... 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 Test, and Static Chamber Test Primary Partical Mean Size a (µm) Full Wind Tunnel Test 2 km/hr 24 km/hr Inlet...

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

... 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 Test, and Static Chamber Test Primary Partical Mean Size a (µm) Full Wind Tunnel Test 2 km/hr 24 km/hr Inlet...

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

... 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 Test, and Static Chamber Test Primary Partical Mean Size a (µm) Full Wind Tunnel Test 2 km/hr 24 km/hr Inlet...

Conceptual design for an electron-beam heated hypersonic wind tunnel

DOE Office of Scientific and Technical Information (OSTI.GOV)

Lipinski, R.J.; Kensek, R.P.

1997-07-01

There is a need for hypersonic wind-tunnel testing at about mach 10 and above using natural air and simulating temperatures and pressures which are prototypic of flight at 50 km altitude or below. With traditional wind-tunnel techniques, gas cooling during expansion results in exit temperatures which are too low. Miles, et al., have proposed overcoming this difficulty by heating the air with a laser beam as it expands in the wind-tunnel nozzle. This report discusses an alternative option of using a high-power electron beam to heat the air as it expands. In the e-beam heating concept, the electron beam ismore » injected into the wind-tunnel nozzle near the exit and then is guided upstream toward the nozzle throat by a strong axial magnetic field. The beam deposits most of its power in the dense air near the throat where the expansion rate is greatest. A conceptual design is presented for a large-scale system which achieves Mach 14 for 0.1 seconds with an exit diameter of 2.8 meters. It requires 450 MW of electron beam power (5 MeV at 90 A). The guiding field is 500 G for most of the transport length and increases to 100 kG near the throat to converge the beam to a 1.0-cm diameter. The beam generator is a DC accelerator using a Marx bank (of capacitors) and a diode stack with a hot cathode. 14 refs. 38 figs., 9 tabs.« less

Altitude Wind Tunnel Drive Motor Installation

NASA Image and Video Library

1943-07-21

Construction workers install the drive motor for the Altitude Wind Tunnel (AWT) in the Exhauster Building at the National Advisory Committee for Aeronautics (NACA) Aircraft Engine Research Laboratory. The AWT was capable of operating full-scale engines in air density, speed, and temperature similar to that found at high altitudes. The tunnel could produce wind speeds up to 500 miles per hour through a 20-foot-diameter test section at the standard operating altitude of 30,000 feet. The airflow was created by a large wooden fan near the tunnel’s southeast corner. This photograph shows the installation of the 18,000-horsepower drive motor inside the adjoining Exhauster Building in July 1943. The General Electric motor, whose support frame is seen in this photograph, connected to a drive shaft that extended from the building, through the tunnel shell, and into a 12-bladed, 31-foot-diameter spruce wood fan. Flexible couplings on the shaft allowed for the movement of the shell. The corner of the Exhauster Building was built around the motor after its installation. The General Electric induction motor could produce 10 to 410 revolutions per minute and create wind speeds up to 500 miles per hour, or Mach 0.63, at 30,000 feet. The AWT became operational in January 1944 and tested piston, turbojet and ramjet engines for nearly 20 years.

Test Data Report, Low-Speed Wind Tunnel Drag Test of a 2/5 Scale Lockheed AH-56 Cheyenne Door-Hinge Hub

DTIC Science & Technology

2016-07-01

the U.S. Army 7– by 10–foot Wind Tunnel located at NASA Ames Research Center in Moffett Field, CA. The purpose of the test was to quantify the drag...drag test of a non-rotating 2/5 scale Lockheed AH-56 Cheyenne main rotor hub in the U.S. Army 7– by 10–foot Wind Tunnel located at NASA Ames Research...the U.S. Army 7– by 10–foot wind tunnel at NASA Ames Research Center 5 2.3 Perspective view of the hub mounted with major dimensions and model

Performance of the high speed anechoic wind tunnel at Lyon University

NASA Technical Reports Server (NTRS)

Sunyach, M.; Brunel, B.; Comte-Bellot, G.

1986-01-01

The characteristics of the feed duct, the wind tunnel, and the experiments run in the convergent-divergent anechoic wind tunnel at Lyon University are described. The wind tunnel was designed to eliminate noise from the entrance of air or from flow interactions with the tunnel walls so that noise caused by the flow-test structure interactions can be studied. The channel contains 1 x 1 x 0.2 m glass and metal foil baffles spaced 0.2 m apart. The flow is forced by a 350 kW fan in the primary circuit, and a 110 kW blower in the secondary circuit. The primary circuit features a factor of four throat reductions, followed by a 1.6 reduction before the test section. Upstream and downstream sensors permit monitoring of the anechoic effectiveness of the channel. Other sensors allow modeling of the flow structures in the tunnel. The tunnel was used to examine turbulent boundary layers in flows up to 140 m/sec, tubulence-excited vibrations in walls, and the effects of laminar and turbulent flows on the appearance and locations of noise sources.

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.

Aeroheating (pressure) characteristics on a 0.10-scale version of the vehicle 3 space shuttle configuration (26-OTS) in the Langley Research Center 4-foot wind tunnel (IH4)

NASA Technical Reports Server (NTRS)

Kingsland, R. B.

1976-01-01

Results of wind tunnel tests, conducted at the Langley Research Center Unitary Plan Wind Tunnel, are presented. The model tested was an 0.010-scale version of the Vehicle 3 Space Shuttle Configuration. Pressure measurements were made on the launch configuration, Orbiter alone, external tank alone, and solid rocket booster alone, to provide heat transfer pressure data. The tests were conducted for a Mach number range from 2.36 to 4.6 and Reynolds number range from 1.2 to 5 million per foot. The model was tested at angles of attack from -10 to 20 deg for a sideslip angle range from -5 to +5 deg, and at sideslip angles from -5 to 48 deg for 0 deg angle of attack. Tabulated data are given and photographs of the test configuration are shown.

Dataset from chemical gas sensor array in turbulent wind tunnel.

PubMed

Fonollosa, Jordi; Rodríguez-Luján, Irene; Trincavelli, Marco; Huerta, Ramón

2015-06-01

The dataset includes the acquired time series of a chemical detection platform exposed to different gas conditions in a turbulent wind tunnel. The chemo-sensory elements were sampling directly the environment. In contrast to traditional approaches that include measurement chambers, open sampling systems are sensitive to dispersion mechanisms of gaseous chemical analytes, namely diffusion, turbulence, and advection, making the identification and monitoring of chemical substances more challenging. The sensing platform included 72 metal-oxide gas sensors that were positioned at 6 different locations of the wind tunnel. At each location, 10 distinct chemical gases were released in the wind tunnel, the sensors were evaluated at 5 different operating temperatures, and 3 different wind speeds were generated in the wind tunnel to induce different levels of turbulence. Moreover, each configuration was repeated 20 times, yielding a dataset of 18,000 measurements. The dataset was collected over a period of 16 months. The data is related to "On the performance of gas sensor arrays in open sampling systems using Inhibitory Support Vector Machines", by Vergara et al.[1]. The dataset can be accessed publicly at the UCI repository upon citation of [1]: http://archive.ics.uci.edu/ml/datasets/Gas+sensor+arrays+in+open+sampling+settings.

Wind-tunnel investigation of the flow correction for a model-mounted angle of attack sensor at angles of attack from -10 deg to 110 deg. [Langley 12-foot low speed wind tunnel test

NASA Technical Reports Server (NTRS)

Moul, T. M.

1979-01-01

A preliminary wind tunnel investigation was undertaken to determine the flow correction for a vane angle of attack sensor over an angle of attack range from -10 deg to 110 deg. The sensor was mounted ahead of the wing on a 1/5 scale model of a general aviation airplane. It was shown that the flow correction was substantial, reaching about 15 deg at an angle of attack of 90 deg. The flow correction was found to increase as the sensor was moved closer to the wing or closer to the fuselage. The experimentally determined slope of the flow correction versus the measured angle of attack below the stall angle of attack agreed closely with the slope of flight data from a similar full scale airplane.

Evaluation of the in-flight noise signature of a 32-chute suppressor nozzle: Acoustic data report. [outdoor static and 40 x 80 ft. wind tunnel tests

NASA Technical Reports Server (NTRS)

Moore, M. T.; Doyle, V. L.

1977-01-01

Outdoor static and 40 x 80 FT wind tunnel tests of the J79-15 engine/nacelle system with the conic nozzle and 32-chute exhaust suppressor were conducted to acquire the data necessary to evaluate the simulated in-flight signature of an engine-size 32-chute exhaust nozzle suppressor using the 40 x 80 ft wind tunnel and to study possible engine core noise contamination of the jet signature. The tests are described and and a sampling of the data acquired is presented. Included are aero performance summaries, as-measured and composite 1/3 OBSPL spectra for the 70 ft sideline high and low mics from the outdoor static tests, sideline traverse spectra and internal noise measurements from both the outdoor static and the 40 x 80 ft wind tunnel tests.

Wind tunnel data of the analysis of heat pipe and wind catcher technology for the built environment

PubMed Central

Calautit, John Kaiser; Chaudhry, Hassam Nasarullah; Hughes, Ben Richard

2015-01-01

The data presented in this article were the basis for the study reported in the research articles entitled ‘Climate responsive behaviour heat pipe technology for enhanced passive airside cooling’ by Chaudhry and Hughes [10] which presents the passive airside cooling capability of heat pipes in response to gradually varying external temperatures and related to the research article “CFD and wind tunnel study of the performance of a uni-directional wind catcher with heat transfer devices” by Calautit and Hughes [1] which compares the ventilation performance of a standard roof mounted wind catcher and wind catcher incorporating the heat pipe technology. Here, we detail the wind tunnel test set-up and inflow conditions and the methodologies for the transient heat pipe experiment and analysis of the integration of heat pipes within the control domain of a wind catcher design. PMID:26958604

Wind tunnel data of the analysis of heat pipe and wind catcher technology for the built environment.

PubMed

Calautit, John Kaiser; Chaudhry, Hassam Nasarullah; Hughes, Ben Richard

2015-12-01

The data presented in this article were the basis for the study reported in the research articles entitled 'Climate responsive behaviour heat pipe technology for enhanced passive airside cooling' by Chaudhry and Hughes [10] which presents the passive airside cooling capability of heat pipes in response to gradually varying external temperatures and related to the research article "CFD and wind tunnel study of the performance of a uni-directional wind catcher with heat transfer devices" by Calautit and Hughes [1] which compares the ventilation performance of a standard roof mounted wind catcher and wind catcher incorporating the heat pipe technology. Here, we detail the wind tunnel test set-up and inflow conditions and the methodologies for the transient heat pipe experiment and analysis of the integration of heat pipes within the control domain of a wind catcher design.

X-34 Experimental Aeroheating at Mach 6 and 10

NASA Technical Reports Server (NTRS)

Berry, Scott A.; Horvath, Thomas J.; DiFulvio, Michael; Glass, Christopher; Merski, N. Ronald

1998-01-01

Critical technologies are being developed to support the goals of the NASA Office of Aeronautics and Space Transportation Technology Access to Space initiative for next-generation reusable space transportation systems. From the perspective of aerothermodynamic performance throughout the flight trajectory, the Reusable Launch Vehicle program incorporates conceptual analysis, ground-based testing, and computational fluid dynamics to provide flyable suborbital flight demonstrator vehicles. This report provides an overview of the hypersonic aeroheating wind tunnel test program conducted at the NASA Langley Research Center in support of one of these vehicles, the X-34 small reusable technology demonstrator program. Global surface heat transfer images, surface streamline patterns, and shock shapes were measured on 0.0153- and 0.0183-scale models of proposed X-34 flight vehicles at Mach 6 and 10 in air. The primary parametrics that were investigated include angles-of-attack from 0 to 35 deg. and freestream unit Reynolds numbers from 0.5 to 8 million per foot (which was sufficient to produce laminar, transitional, and turbulent heating data), both with and without control surface deflections. Comparisons of the experimental data to computational predictions are included, along with a discussion of the implications of some of the experimental flow features for the flight vehicle.

A three-dimensional orthogonal laser velocimeter for the NASA Ames 7- by 10-foot wind tunnel

NASA Technical Reports Server (NTRS)

Dunagan, Stephen E.; Cooper, Donald L.

1995-01-01

A three-component dual-beam laser-velocimeter system has been designed, fabricated, and implemented in the 7-by 10-Foot Wind Tunnel at NASA Ames Research Center. The instrument utilizes optical access from both sides and the top of the test section, and is configured for uncoupled orthogonal measurements of the three Cartesian coordinates of velocity. Bragg cell optics are used to provide fringe velocity bias. Modular system design provides great flexibility in the location of sending and receiving optics to adapt to specific experimental requirements. Near-focus Schmidt-Cassegrain optic modules may be positioned for collection of forward or backward scattered light over a large solid angle, and may be clustered to further increase collection solid angle. Multimode fiber optics transmit collected light to the photomultiplier tubes for processing. Counters are used to process the photomultiplier signals and transfer the processed data digitally via buffered interface controller to the host MS-DOS computer. Considerable data reduction and graphical display programming permit on-line control of data acquisition and evaluation of the incoming data. This paper describes this system in detail and presents sample data illustrating the system's capability.

Unitary Plan Wind Tunnel Landmark Dedication and Revitalization

NASA Technical Reports Server (NTRS)

1990-01-01

This video shows construction scenes of unitary plan wind tunnel, aerials, and views of various models, including an MD-II in the 11 ft, an Apollo in the 8x7, Dynasoar in the 8x7, a one inch scale shuttle in the 8x7, and an artist's concept of a 12 ft test section.

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 Test...

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 Test...

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.

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.

Contraction design for small low-speed wind tunnels

NASA Technical Reports Server (NTRS)

Bell, James H.; Mehta, Rabindra D.

1988-01-01

An iterative design procedure was developed for two- or three-dimensional contractions installed on small, low-speed wind tunnels. The procedure consists of first computing the potential flow field and hence the pressure distributions along the walls of a contraction of given size and shape using a three-dimensional numerical panel method. The pressure or velocity distributions are then fed into two-dimensional boundary layer codes to predict the behavior of the boundary layers along the walls. For small, low-speed contractions it is shown that the assumption of a laminar boundary layer originating from stagnation conditions at the contraction entry and remaining laminar throughout passage through the successful designs if justified. This hypothesis was confirmed by comparing the predicted boundary layer data at the contraction exit with measured data in existing wind tunnels. The measured boundary layer momentum thicknesses at the exit of four existing contractions, two of which were 3-D, were found to lie within 10 percent of the predicted values, with the predicted values generally lower. From the contraction wall shapes investigated, the one based on a fifth-order polynomial was selected for installation on a newly designed mixing layer wind tunnel.

Contraction design for small low-speed wind tunnels

NASA Technical Reports Server (NTRS)

Bell, James H.; Mehta, Rabindra D.

1988-01-01

An iterative design procedure was developed for 2- or 3-dimensional contractions installed on small, low speed wind tunnels. The procedure consists of first computing the potential flow field and hence the pressure distributions along the walls of a contraction of given size and shape using a 3-dimensional numerical panel method. The pressure or velocity distributions are then fed into 2-dimensional boundary layer codes to predict the behavior of the boundary layers along the walls. For small, low speed contractions, it is shown that the assumption of a laminar boundary layer originating from stagnation conditions at the contraction entry and remaining laminar throughout passage through the successful designs is justified. This hypothesis was confirmed by comparing the predicted boundary layer data at the contraction exit with measured data in existing wind tunnels. The measured boundary layer momentum thicknesses at the exit of four existing contractions, two of which were 3-D, were found to lie within 10 percent of the predicted values, with the predicted values generally lower. From the contraction wall shapes investigated, the one based on a 5th order polynomial was selected for newly designed mixing wind tunnel installation.

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.

Transverse vorticity measurements in the NASA Ames 80 x 120 wind tunnel boundary layer

NASA Technical Reports Server (NTRS)

Foss, John F.; Bhol, D. G.; Bramkamp, F. D.; Klewicki, J. G.

1994-01-01

The MSU compact four-wire transverse vorticity probe permits omega(sub z)(t) measurements in a nominally 1 sq mm domain. Note that a conventional coordinate system is used with x and y in the streamwise and normal directions respectively. The purpose of this investigation was to acquire time series data in the same access port at the ceiling of the 80 ft x 120 ft wind tunnel (NASA Ames Research Center) as earlier used by the Wallace group from the University of Maryland and to compare the present results with those of the three-component vorticity probe used in that earlier study.

Preparation of polystyrene microspheres for laser velocimetry in wind tunnels

NASA Technical Reports Server (NTRS)

Nichols, Cecil E., Jr.

1987-01-01

Laser Velocimetry (L/V) had made great strides in replacing intrusive devices for wind tunnel flow measurements. The weakness of the L/V has not been the L/V itself, but proper size seeding particles having known drag characteristics. For many Langley Wind Tunnel applications commercial polystyrene latex microspheres suspended in ethanol, injected through a fluid nozzle provides excellent seeding but was not used due to the high cost. This paper provides the instructions, procedures, and formulations for producing polystyrene latex monodisperse microspheres of 0.6, 1.0, 1.7, 2.0, and 2.7 micron diameters. These are presently being used at Langley Research Center as L/V seeding particles.

Wind tunnel measurements of the power output variability and unsteady loading in a micro wind farm model

NASA Astrophysics Data System (ADS)

Bossuyt, Juliaan; Howland, Michael; Meneveau, Charles; Meyers, Johan

2015-11-01

To optimize wind farm layouts for a maximum power output and wind turbine lifetime, mean power output measurements in wind tunnel studies are not sufficient. Instead, detailed temporal information about the power output and unsteady loading from every single wind turbine in the wind farm is needed. A very small porous disc model with a realistic thrust coefficient of 0.75 - 0.85, was designed. The model is instrumented with a strain gage, allowing measurements of the thrust force, incoming velocity and power output with a frequency response up to the natural frequency of the model. This is shown by reproducing the -5/3 spectrum from the incoming flow. Thanks to its small size and compact instrumentation, the model allows wind tunnel studies of large wind turbine arrays with detailed temporal information from every wind turbine. Translating to field conditions with a length-scale ratio of 1:3,000 the frequencies studied from the data reach from 10-4 Hz up to about 6 .10-2 Hz. The model's capabilities are demonstrated with a large wind farm measurement consisting of close to 100 instrumented models. A high correlation is found between the power outputs of stream wise aligned wind turbines, which is in good agreement with results from prior LES simulations. Work supported by ERC (ActiveWindFarms, grant no. 306471) and by NSF (grants CBET-113380 and IIA-1243482, the WINDINSPIRE project).

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.

Heat-transfer test results for a .0275-scale space shuttle external tank with a 10 deg/40 deg double cone-ogive nose in the NASA/AMES 3.5-foot hypersonic wind tunnel (FH14), volume 2

NASA Technical Reports Server (NTRS)

Carroll, H. R.

1977-01-01

A .0275 scale forebody model of the new baseline configuration of the space shuttle external tank vent cap configuration was tested to determine the flow field due to the double cone configuration. The tests were conducted in a 3.5 foot hypersonic wind tunnel at alpha = -5 deg, -4.59 deg, 0 deg, 5 deg, and 10 deg; beta = 0 deg, -3 deg, -5.51 deg, -6 deg, -9 deg, and +6 deg; nominal freestream Reynolds numbers per foot of 1.5 x 1 million, 3.0 x 1 million, and 5.0 x 1 million; and a nominal Mach number of 5. Separation and reattached flow from thermocouple data, shadowgraphs, and oil flows indicate that separation begins about 80% from the tip of the 10 deg cone, then reattaches on the vent cap and produces fully turbulent flow over most of the model forebody. The hardware disturbs the flow over a much larger area than present TPS application has assumed. A correction to the flow disturbance was experimentally suggested from the results of an additional test run.

Correlation of low speed wind tunnel and flight test data for V/STOL aircraft

NASA Technical Reports Server (NTRS)

Cook, W. L.; Hickey, D. H.

1975-01-01

The XV-5B fan-in-wing aircraft and the Y0V-10 RCF rotating cylinder flap aircraft were subjected to wind tunnel tests. These tests were conducted specifically to provide for correlation between wind tunnel and inflight aerodynamics and noise test data. Correlation between aerodynamic and noise data are presented and testing techniques that are related to the accuracy of the data, or that might affect the correlations, are discussed.

X-33 Computational Aeroheating/Aerodynamic Predictions and Comparisons With Experimental Data

NASA Technical Reports Server (NTRS)

Hollis, Brian R.; Thompson, Richard A.; Berry, Scott A.; Horvath, Thomas J.; Murphy, Kelly J.; Nowak, Robert J.; Alter, Stephen J.

2003-01-01

This report details a computational fluid dynamics study conducted in support of the phase II development of the X-33 vehicle. Aerodynamic and aeroheating predictions were generated for the X-33 vehicle at both flight and wind-tunnel test conditions using two finite-volume, Navier-Stokes solvers. Aerodynamic computations were performed at Mach 6 and Mach 10 wind-tunnel conditions for angles of attack from 10 to 50 with body-flap deflections of 0 to 20. Additional aerodynamic computations were performed over a parametric range of free-stream conditions at Mach numbers of 4 to 10 and angles of attack from 10 to 50. Laminar and turbulent wind-tunnel aeroheating computations were performed at Mach 6 for angles of attack of 20 to 40 with body-flap deflections of 0 to 20. Aeroheating computations were performed at four flight conditions with Mach numbers of 6.6 to 8.9 and angles of attack of 10 to 40. Surface heating and pressure distributions, surface streamlines, flow field information, and aerodynamic coefficients from these computations are presented, and comparisons are made with wind-tunnel data.

Evaluation of a novel wind tunnel for the measurement of the kinetics of odour emissions from piggery effluent.

PubMed

Sohn, J H; Smith, R; Yoong, E; Hudson, N; Kim, T I

2004-01-01

A novel laboratory wind tunnel, with the capability to control factors such as air flow-rate, was developed to measure the kinetics of odour emissions from liquid effluent. The tunnel allows the emission of odours and other volatiles under an atmospheric transport system similar to ambient conditions. Sensors for wind speed, temperature and humidity were installed and calibrated. To calibrate the wind tunnel, trials were performed to determine the gas recovery efficiency under different air flow-rates (ranging from 0.001 to 0.028m3/s) and gas supply rates (ranging from 2.5 to 10.0 L/min) using a standard CO gas mixture. The results have shown gas recovery efficiencies ranging from 61.7 to 106.8%, while the average result from the trials was 81.14%. From statistical analysis, it was observed that the highest, most reliable gas recovery efficiency of the tunnel was 88.9%. The values of air flow-rate and gas supply rate corresponding to the highest gas recovery efficiency were 0.028 m3/s and 10.0 L/min respectively. This study suggested that the wind tunnel would provide precise estimates of odour emission rate. However, the wind tunnel needs to be calibrated to compensate for errors caused by different air flow-rates.

10. Photocopy of drawing (original in possession of Naval Surface ...

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

10. Photocopy of drawing (original in possession of Naval Surface Warfare Center Carderock Division, Bethesda, MD) SUPERSONIC WIND TUNNEL BUILDING, GROUND FLOOR PLAN, 1947 - Naval Surface Warfare Center, Supersonic Wind Tunnel Building, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD

Wind-Tunnel Tests of Seven Static-Pressure Probes at Transonic Speeds

NASA Technical Reports Server (NTRS)

Capone, Francis J.

1961-01-01

Wind-tunnel tests have been conducted to determine the errors of 3 seven static-pressure probes mounted very close to the nose of a body of revolution simulating a missile forebody. The tests were conducted at Mach numbers from 0.80 to 1.08 and at angles of attack from -1.7 deg to 8.4 deg. The test Reynolds number per foot varied from 3.35 x 10(exp 6) to 4.05 x 10(exp 6). For three 4-vane, gimbaled probes, the static-pressure errors remained constant throughout the test angle-of-attack range for all Mach numbers except 1.02. For two single-vane, self-rotating probes having two orifices at +/-37.5 deg. from the plane of symmetry on the lower surface of the probe body, the static-pressure error varied as much as 1.5 percent of free-stream static pressure through the test angle-of- attack range for all Mach numbers. For two fixed, cone-cylinder probes of short length and large diameter, the static-pressure error varied over the test angle-of-attack range at constant Mach numbers as much as 8 to 10 percent of free-stream static pressure.

Development of a Test to Evaluate Aerothermal Response of Materials to Hypersonic Flow Using a Scramjet Wind Tunnel (Postprint)

DTIC Science & Technology

2010-05-01

SCRAMJET WIND TUNNEL (POSTPRINT) 5a. CONTRACT NUMBER FA8650-10-D-5226-0002 5b. GRANT NUMBER 5c. PROGRAM ELEMENT NUMBER 62102F 6. AUTHOR(S...prototype scramjet engine as a wind tunnel . A sample holder was designed using combustion fluid dynamics results as inputs into structural models. The...Z39-18 Development of a Test to Evaluate Aerothermal Response of Materials to Hypersonic Flow Using a Scramjet Wind Tunnel Triplicane A

F-111B in Ames 40x80 Foot Wind Tunnel.

NASA Image and Video Library

1969-02-06

Installation Photos, 3/4 front view from below. F-111B in Ames 40x80 Foot Wind Tunnel. The General Dynamics/Grumman F-111B was a long-range carrier-based interceptor aircraft that was planned to be a follow-on to the F-4 Phantom II. The F-111B was developed in the 1960s by General Dynamics in conjunction with Grumman for the United States Navy (USN) as part of the joint Tactical Fighter Experimental (TFX) with the United States Air Force (USAF) to produce a common fighter for the services that could perform a variety of missions. It incorporated innovations such as variable-geometry wings, afterburning turbofan engines, and a long-range radar and missile weapons system.

10. Photocopy of photograph (original photograph in the collection of ...

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

10. Photocopy of photograph (original photograph in the collection of Naval Surface Warfare Center Carderock Division, Bethesda, MD) VIEW SOUTHWEST, SUBSONIC WIND TUNNEL BUILDING, 1943 - Naval Surface Warfare Center, Subsonic Wind Tunnel Building, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD

Results of winglet development studies for DC-10 derivatives

NASA Technical Reports Server (NTRS)

Shollenberger, C. A.; Humphreys, J. W.; Heiberger, F. S.; Pearson, R. M.

1983-01-01

The results of investigations into the application of winglets to the DC-10 aircraft are presented. The DC-10 winglet configuration was developed and its cruise performance determined in a previous investigation. This study included high speed and low speed wind tunnel tests to evaluate aerodynamic characteristics, and a subsonic flutter wind tunnel test with accompanying analysis and evaluation of results. Additionally, a configuration integration study employed the results of the wind tunnel studies to determine the overall impact of the installation of winglets on the DC-10 aircraft. Conclusions derived from the high speed and low speed tests indicate that the winglets had no significant effects on the DC-10 stability characteristics or high speed buffet. It was determined that winglets had a minimal effect on aircraft lift characteristics and improved the low speed aircraft drag under high lift conditions. The winglets affected the DC-10 flutter characteristics by reducing the flutter speed of the basic critical mode and introducing a new critical mode involving outer wing torsion and longitudinal bending. The overall impact of winglets was determined to be of sufficient benefit to merit flight evaluation.

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

DOE Office of Scientific and Technical Information (OSTI.GOV)

Macha, J.M.; Buffington, R.J.; Henfling, 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.

Water tunnel flow visualization and wind tunnel data analysis of the F/A-18. [leading edge extension vortex effects

NASA Technical Reports Server (NTRS)

Erickson, G. E.

1982-01-01

Six degree of freedom studies were utilized to extract a band of yawing and rolling moment coefficients from the F/A-18 aircraft flight records. These were compared with 0.06 scale model data obtained in a 16T wind tunnel facility. The results, indicate the flight test yawing moment data exhibit an improvement over the wind tunnel data to near neutral stability and a significant reduction in lateral stability (again to anear neutral level). These data are consistent with the flight test results since the motion was characterized by a relatively slo departure. Flight tests repeated the slow yaw departure at M 0.3. Only 0.16 scale model wind tunnel data showed levels of lateral stability similar to the flight test results. Accordingly, geometric modifications were investigated on the 0.16 scale model in the 30x60 foot wind tunnel to improve high angle of attack lateral stability.

Automated Boundary Conditions for Wind Tunnel Simulations

NASA Technical Reports Server (NTRS)

Carlson, Jan-Renee

2018-01-01

Computational fluid dynamic (CFD) simulations of models tested in wind tunnels require a high level of fidelity and accuracy particularly for the purposes of CFD validation efforts. Considerable effort is required to ensure the proper characterization of both the physical geometry of the wind tunnel and recreating the correct flow conditions inside the wind tunnel. The typical trial-and-error effort used for determining the boundary condition values for a particular tunnel configuration are time and computer resource intensive. This paper describes a method for calculating and updating the back pressure boundary condition in wind tunnel simulations by using a proportional-integral-derivative controller. The controller methodology and equations are discussed, and simulations using the controller to set a tunnel Mach number in the NASA Langley 14- by 22-Foot Subsonic Tunnel are demonstrated.

The Langley Wind Tunnel Enterprise

NASA Technical Reports Server (NTRS)

Paulson, John W., Jr.; Kumar, Ajay; Kegelman, Jerome T.

1998-01-01

After 4 years of existence, the Langley WTE is alive and growing. Significant improvements in the operation of wind tunnels have been demonstrated and substantial further improvements are expected when we are able to truly address and integrate all the processes affecting the wind tunnel testing cycle.

Identification of the sources of PM10 in a subway tunnel using positive matrix factorization.

PubMed

Park, Duckshin; Lee, Taejeong; Hwang, Doyeon; Jung, Wonseok; Lee, Yongil; Cho, KiChul; Kim, Dongsool; Lees, Kiyoung

2014-12-01

The level of particulate matter of less than 10 μm diameter (PM10) at subway platforms can be significantly reduced by installing a platform screen-door system. However, both workers and passengers might be exposed to higher PM10 levels while the cars are within the tunnel because it is a more confined environment. This study determined the PM10 levels in a subway tunnel, and identified the sources of PM10 using elemental analysis and receptor modeling. Forty-four PM10 samples were collected in the tunnel between the Gireum and Mia stations on Line 4 in metropolitan Seoul and analyzed using inductively coupled plasma-atomic emission spectrometry and ion chromatography. The major PM10 sources were identified using positive matrix factorization (PMF). The average PM10 concentration in the tunnels was 200.8 ± 22.0 μg/m3. Elemental analysis indicated that the PM10 consisted of 40.4% inorganic species, 9.1% anions, 4.9% cations, and 45.6% other materials. Iron was the most abundant element, with an average concentration of 72.5 ± 10.4 μg/m3. The PM10 sources characterized by PMF included rail, wheel, and brake wear (59.6%), soil combustion (17.0%), secondary aerosols (10.0%), electric cable wear (8.1%), and soil and road dust (5.4%). Internal sources comprising rail, wheel, brake, and electric cable wear made the greatest contribution to the PM10 (67.7%) in tunnel air. Implications: With installation of a platform screen door, PM10 levels in subway tunnels were higher than those on platforms. Tunnel PM10 levels exceeded 150 µg/m3 of the Korean standard for subway platform. Elemental analysis of PM10 in a tunnel showed that Fe was the most abundant element. Five PM10 sources in tunnel were identified by positive matrix factorization. Railroad-related sources contributed 68% of PM10 in the subway tunnel.

Wind tunnel tests of the GA(W)-2 airfoil with 20% aileron, 25% slotted flap, 30% Fowler flap and 10% slot-lip spoiler

NASA Technical Reports Server (NTRS)

Wentz, W. H., Jr.

1977-01-01

Two dimensional wind tunnel tests were conducted for the GA(W)-2 airfoil section with: 20% aileron, 25% slotted flap; 30% Fowler flap, and 10% slot-lip spoiler. All tests were conducted at a Reynolds number of 2,200,000 and a Mach Number of 0.13. In addition to force measurements, tuft studies were conducted for the slotted and Fowler flap configurations. Aileron and spoiler hinge moments were obtained by integration of surface pressure measurements. Tests results show that a value of 3.82 was obtained with 30% Fowler flap. Aileron control effectiveness and hinge moments were similar to other airfoils. The slot-lip spoiler provided powerful, positive roll control at all flap settings.

Design and Characterization of the UTIAS Anechoic Wind Tunnel

NASA Astrophysics Data System (ADS)

Chow, Derrick H. F.

The anechoic open-jet wind tunnel facility at the University of Toronto Institute for Aerospace Studies was updated and characterized to meet the needs of current and future aeroacoustic experiments. The wind tunnel inlet was resurfaced and flow-conditioning screens were redesigned to improve the freestream turbulence intensity to below 0.4% in the test section. The circular nozzle was replaced with a square secondary contraction that increased the maximum test section velocity to 75 m/s and improved flow uniformity to over 99% across a usable cross-sectional area of 500 mm x 500 mm. Acoustic baffles were installed in front of the wind tunnel inlet and foam wedges were installed in the anechoic chamber. The overall background sound pressure levels in the chamber were improved by 8-18 db over the range of operational freestream velocities. The anechoic chamber cut-off frequency is 170 Hz and the reverberation time for a 60 dB sound power decay is 0.032 s.

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.

A three degree of freedom manipulator used for store separation wind tunnel test

NASA Astrophysics Data System (ADS)

Wei, R.; Che, B.-H.; Sun, C.-B.; Zhang, J.; Lu, Y.-Q.

2018-06-01

A three degree of freedom manipulator is presented, which is used for store separation wind tunnel test. It is a kind of mechatronics product, have small volume and large moment of torque. The paper researched the design principle of wind tunnel test equipment, also introduced the transmission principle design, physical design, control system design, drive element selection calculation and verification, dynamics computation and static structural computation of the manipulator. To satisfy the design principle of wind tunnel test equipment, some optimization design are made include optimizes the structure of drive element and cable, fairing configuration, overall dimension so that to make the device more suitable for the wind tunnel test. Some tests are made to verify the parameters of the manipulator. The results show that the device improves the load from 100 Nm to 250 Nm, control accuracy from 0.1°to 0.05°in pitch and yaw, also improves load from 10 Nm to 20 Nm, control accuracy from 0.1°to 0.05°in roll.

Wind Tunnel Investigation of a Balloon as Decelerator at Mach Numbers from 1.47 to 2.50

NASA Technical Reports Server (NTRS)

McShera, John T.; Keyes, J. Wayne

1961-01-01

A wind-tunnel investigation was conducted to study the characteristics of a towed spherical balloon as a drag device at Mach numbers from 1.47 to 2.50, Reynolds numbers from 0.36 x 10(exp 6) to 1.0 x 10(exp 6) , and angles of attack from -15 to 15 degrees. Tow-cable length was approximately 24 inches from asymmetric body to cone on the upstream side of the balloon. As the tow cable was lengthened the balloon reached a point in the test section where wall-reflected shocks intersected the balloon and caused severe oscillations. As a result, the tow cable broke and the inflatable balloon model was destroyed. Further tests used a model rigid plastic sphere 6.75 inches in diameter. Tow cable length was approximately 24 inches from asymmetric body to the upstream side of the sphere.

Automation of Some Operations of a Wind Tunnel Using Artificial Neural Networks

NASA Technical Reports Server (NTRS)

Decker, Arthur J.; Buggele, Alvin E.

1996-01-01

Artificial neural networks were used successfully to sequence operations in a small, recently modernized, supersonic wind tunnel at NASA-Lewis Research Center. The neural nets generated correct estimates of shadowgraph patterns, pressure sensor readings and mach numbers for conditions occurring shortly after startup and extending to fully developed flow. Artificial neural networks were trained and tested for estimating: sensor readings from shadowgraph patterns, shadowgraph patterns from shadowgraph patterns and sensor readings from sensor readings. The 3.81 by 10 in. (0.0968 by 0.254 m) tunnel was operated with its mach 2.0 nozzle, and shadowgraph was recorded near the nozzle exit. These results support the thesis that artificial neural networks can be combined with current workstation technology to automate wind tunnel operations.

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.

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.

Flight determined lift and drag characteristics of an F-8 airplane modified with a supercritical wing with comparison to wind-tunnel results

NASA Technical Reports Server (NTRS)

Pyle, J. S.; Steers, L. L.

1975-01-01

Flight measurements obtained with a TF-8A airplane modified with a supercritical wing are presented for altitudes from 7.6 kilometers (25,000 feet) to 13.7 kilometers (45,000 feet), Mach numbers from 0.6 to 1.2, and Reynolds numbers from 0.8 x 10 to the 7th power to 2.3 x 10 to the 7th power. Flight results for the airplane with and without area-rule fuselage fairings are compared. The techniques used to determine the lift and drag characteristics of the airplane are discussed. Flight data are compared with wind-tunnel model results, where applicable.

Calibration of the Naval Postgraduate School 3.5 X 5.0 Academic Wind Tunnel

DTIC Science & Technology

1990-09-01

design of the wind tunnel great care is taken to ensure undisturbed, uniform flow through the test section. Even so, there will exist some disturbances...the longitudinal pressure gradient will determine if there is flow leakage in the test section doors. The information obtained also makes possible an...turbulence calibrations were performed. At the completion of these measurements it was determined that the flow quality could be improved by wind tunnel

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.

A Photogrammetric System for Model Attitude Measurement in Hypersonic Wind Tunnels

NASA Technical Reports Server (NTRS)

Jones, Thomas W.; Lunsford, Charles B.

2007-01-01

A series of wind tunnel tests have been conducted to evaluate a multi-camera videogrammetric system designed to measure model attitude in hypersonic facilities. The technique utilizes processed video data and photogrammetric principles for point tracking to compute model position including pitch, roll and yaw. A discussion of the constraints encountered during the design, and a review of the measurement results obtained from the NASA Langley Research Center (LaRC) 31-Inch Mach 10 tunnel are presented.

Wind tunnel model and method

NASA Technical Reports Server (NTRS)

Jackson, C. M., Jr.; Summerfield, D. G. (Inventor)

1974-01-01

The design and development of a wind tunnel model equipped with pressure measuring devices are discussed. The pressure measuring orifices are integrally constructed in the wind tunnel model and do not contribute to distortions of the aerodynamic surface. The construction of a typical model is described and a drawing of the device is included.

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.

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.

The X-43A Flight Research Program: Lessons Learned on the Road to Mach 10

NASA Technical Reports Server (NTRS)

Peebles, Curtis

2007-01-01

During an aerospace engineer's undergraduate studies, he or she will attend classes in aerodynamics, thermodynamics, structures, stability and control, dynamics, design, propulsion, and computer science, along with the related courses in mathematics, physics, statistics, and chemistry required to understand the material. Upon graduation, the new engineer will have acquired a basic knowledge of how to build an aerospace vehicle. What only comes through experience, however, is the understanding of the inevitable imperfect process through which an aerospace vehicle is built. This is the adventure of turning a basic concept into functional hardware. Engineers working on a project must often deal with ambiguous situations. They are routinely asked by management to provide risk assessments of a project, yet even after careful analysis uncertainties remain. The project must be accomplished within finite limits of time and money. The question an engineer answers is whether the solution to potential problem is worth the cost and schedule delay, or if the solution might actually be worse than the problem it is meant to solve. Review protocols are established to ensure that an unknown has not been overlooked. But these cannot protect against an unknown unknown. Examples of these situations can be found in the history of the X-43A Hyper-X (Hypersonic Experiment) program. In this NASA project, a supersonic combustion ramjet (scramjet) engine was flight tested on a subscale vehicle. The X-43A Hyper-X Research Vehicle (HXRV) was launched from a B-52B mothership, then boosted to the test speed by a modified Pegasus rocket first stage, called the Hyper-X Launch Vehicle (HXLV). Once at the proper speed and altitude, the X-43A separated from the booster, stabilized itself, and then the engine test began. Although wind-tunnel scramjet engine tests had begun in the late 1950s, before the Hyper-X program there had never been an actual in-flight test of such an engine integrated with an

Measurement of Model Noise in a Hard-Wall Wind Tunnel

NASA Technical Reports Server (NTRS)

Soderman, Paul T.

2006-01-01

Identification, analysis, and control of fluid-mechanically-generated sound from models of aircraft and automobiles in special low-noise, semi-anechoic wind tunnels are an important research endeavor. Such studies can also be done in aerodynamic wind tunnels that have hard walls if phased microphone arrays are used to focus on the noise-source regions and reject unwanted reflections or background noise. Although it may be difficult to simulate the total flyover or drive-by noise in a closed wind tunnel, individual noise sources can be isolated and analyzed. An acoustic and aerodynamic study was made of a 7-percent-scale aircraft model in a NASA Ames 7-by-10-ft (about 2-by-3-m) wind tunnel for the purpose of identifying and attenuating airframe noise sources. Simulated landing, takeoff, and approach configurations were evaluated at Mach 0.26. Using a phased microphone array mounted in the ceiling over the inverted model, various noise sources in the high-lift system, landing gear, fins, and miscellaneous other components were located and compared for sound level and frequency at one flyover location. Numerous noise-alleviation devices and modifications of the model were evaluated. Simultaneously with acoustic measurements, aerodynamic forces were recorded to document aircraft conditions and any performance changes caused by geometric modifications. Most modern microphone-array systems function in the frequency domain in the sense that spectra of the microphone outputs are computed, then operations are performed on the matrices of microphone-signal cross-spectra. The entire acoustic field at one station in such a system is acquired quickly and interrogated during postprocessing. Beam-forming algorithms are employed to scan a plane near the model surface and locate noise sources while rejecting most background noise and spurious reflections. In the case of the system used in this study, previous studies in the wind tunnel have identified noise sources up to 19 d

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.

Wind tunnel studies of Martian aeolian processes

NASA Technical Reports Server (NTRS)

Greeley, R.; Iversen, J. D.; Pollack, J. B.; Udovich, N.; White, B.

1973-01-01

Preliminary results are reported of an investigation which involves wind tunnel simulations, geologic field studies, theoretical model studies, and analyses of Mariner 9 imagery. Threshold speed experiments were conducted for particles ranging in specific gravity from 1.3 to 11.35 and diameter from 10.2 micron to 1290 micron to verify and better define Bagnold's (1941) expressions for grain movement, particularly for low particle Reynolds numbers and to study the effects of aerodynamic lift and surface roughness. Wind tunnel simulations were conducted to determine the flow field over raised rim craters and associated zones of deposition and erosion. A horseshoe vortex forms around the crater, resulting in two axial velocity maxima in the lee of the crater which cause a zone of preferential erosion in the wake of the crater. Reverse flow direction occurs on the floor of the crater. The result is a distinct pattern of erosion and deposition which is similar to some martian craters and which indicates that some dark zones around Martian craters are erosional and some light zones are depositional.

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.

Flowfield measurements in the NASA Lewis Research Center 9- by 15-foot low-speed wind tunnel

NASA Technical Reports Server (NTRS)

Hughes, Christopher E.

1989-01-01

An experimental investigation was conducted in the NASA Lewis 9- by 15-Foot Low-Speed Wind Tunnel to determine the flow characteristics in the test section during wind tunnel operation. In the investigation, a 20-probe horizontally-mounted Pitot-static flow survey rake was used to obtain cross-sectional total and static pressure surveys at four axial locations in the test section. At each axial location, the cross-sectional flowfield surveys were made by repositioning the Pitot-static flow survey rake vertically. In addition, a calibration of the new wind tunnel rake instrumentation, used to determine the wind tunnel operating conditions, was performed. Boundary laser surveys were made at three axial locations in the test section. The investigation was conducted at tunnel Mach numbers 0.20, 0.15, 0.10, and 0.05. The test section profile results from the investigation indicate that fairly uniform total pressure profiles (outside the test section boundary layer) and fairly uniform static pressure and Mach number profiles (away from the test section walls and downstream of the test section entrance) exist throughout in the wind tunnel test section.

Flight and wind-tunnel correlation of boundary-layer transition on the AEDC transition cone

NASA Technical Reports Server (NTRS)

Fisher, D. L.; Dougherty, N. S., Jr.

1982-01-01

Transition and fluctuating surface pressure data were acquired on a 10 deg included angle cone, using the same instrumentation and technique over a wide range of Mach and Reynolds numbers in 23 wind tunnels and in flight. Transition was detected with a traversing pitot-pressure probe in contact with the surface. The surface pressure fluctuations were measured with microphones set flush in the cone surface. Good correlation of end of transition Reynolds number RE(T) was obtained between data from the lower disturbance wind tunnels and flight up to a boundary layer edge Mach number, M(e) = 1.2. Above M(e) = 1.2, however, this correlation deteriorates, with the flight Re(T) being 25 to 30% higher than the wind tunnel Re(T) at M(e) = 1.6. The end of transition Reynolds number correlated within + or - 20% with the surface pressure fluctuations, according to the equation used. Broad peaks in the power spectral density distributions indicated that Tollmien-Schlichting waves were the probable cause of transition in flight and in some of the wind tunnels.

Semi-span wind tunnel testing without conventional peniche

NASA Astrophysics Data System (ADS)

Skinner, S. N.; Zare-Behtash, H.

2017-12-01

Low-speed wind tunnel tests of a flexible wing semi-span model have been implemented in the 9× 7 ft de Havilland wind tunnel at the University of Glasgow. The main objective of this investigation is to quantify the effect of removing the traditional peniche boundary layer spacer utilised in this type of testing. Removal of the peniche results in a stand-off gap between the wind tunnel wall and the model's symmetry plane. This offers the advantage of preventing the development of a horseshoe vortex in front of the model, at the peniche/wall juncture. The formation of the horseshoe vortex is known to influence the flow structures around the entire model and thus alters the model's aerodynamic behaviours. To determine the influence of the stand-off gap, several gap heights have been tested for a range of angles of attack at Re=1.5× 10^6, based on the wing mean aerodynamic chord (MAC). Force platform data have been used to evaluate aerodynamic coefficients, and how they vary with stand-off heights. Stereoscopic Particle Imaging Velocimetry (sPIV) was used to examine the interaction between the tunnel boundary layer and model's respective stand-off gap. In addition, clay and tuft surface visualisation enhanced the understanding of how local flow structures over the length of the fuselage vary with stand-off height and angle of attack. The presented results show that a stand-off gap of four-to-five times the displacement thickness of the tunnel wall boundary layer is capable of achieving a flow field around the model fuselage that is representative of what would be expected for an equivalent full-span model in free-air—this cannot be achieved with the application of a peniche.

The X-ray-emitting trail of the nearby pulsar PSR1929 + 10

NASA Technical Reports Server (NTRS)

Wang, Q. D.; Li, Zhi-Yun; Begelman, Mitchell C.

1993-01-01

The paper reports the detection by the Rosat satellite of a nebula associated with the nearby pulsar PSR1929 + 10, which is of a type different from the pulsar-wind nebulae produced by direct interaction of the relativistic wind from older pulsars with the interstellar medium (ISM) (Blandorf et al., 1973). The PSR1929 + 10 nebula appears as a linear diffuse X-ray feature in the direction opposite to the pulsar's proper motion, with the pulsar wind confined by the ram-pressure arising from the high velocity of the pulsar through the ISM. This results in a trail of relativistic electrons with enhanced emissions of synchrotron radiation.

An experimental investigation of three dimensional low speed minimum interference wind tunnel for high lift wings

NASA Technical Reports Server (NTRS)

Shindo, S.; Joppa, R. G.

1980-01-01

As a means to achieve a minimum interference correction wind tunnel, a partially actively controlled test section was experimentally examined. A jet flapped wing with 0.91 m (36 in) span and R = 4.05 was used as a model to create moderately high lift coefficients. The partially controlled test section was simulated using an insert, a rectangular box 0.96 x 1.44 m (3.14 x 4.71 ft) open on both ends in the direction of the tunnel air flow, placed in the University of Washington Aeronautical Laboratories (UWAL) 2.44 x 3.66 m (8 x 12 ft) wind tunnel. A tail located three chords behind the wing was used to measure the downwash at the tail region. The experimental data indicates that, within the range of momentum coefficient examined, it appears to be unnecessary to actively control all four sides of the test section walls in order to achieve the near interference free flow field environment in a small wind tunnel. The remaining wall interference can be satisfactorily corrected by the vortex lattice method.

A tilting wind tunnel for fire behavior studies

Treesearch

David R. Weise

1994-01-01

The combined effects of wind velocity and slope on wildland fire behavior can be studied in the laboratory using a tilting wind tunnel. The tilting wind tunnel requires a commercially available fan to induce wind and can be positioned to simulate heading and backing fires spreading up and down slope. The tunnel is portable and can be disassembled for transport using a...

Thermal and Pressure Characterization of a Wind Tunnel Force Balance Using the Single Vector System. Experimental Design and Analysis Approach to Model Pressure and Temperature Effects in Hypersonic Wind Tunnel Research

NASA Technical Reports Server (NTRS)

Lynn, Keith C.; Commo, Sean A.; Johnson, Thomas H.; Parker, Peter A,

2011-01-01

Wind tunnel research at NASA Langley Research Center s 31-inch Mach 10 hypersonic facility utilized a 5-component force balance, which provided a pressurized flow-thru capability to the test article. The goal of the research was to determine the interaction effects between the free-stream flow and the exit flow from the reaction control system on the Mars Science Laboratory aeroshell during planetary entry. In the wind tunnel, the balance was exposed to aerodynamic forces and moments, steady-state and transient thermal gradients, and various internal balance cavity pressures. Historically, these effects on force measurement accuracy have not been fully characterized due to limitations in the calibration apparatus. A statistically designed experiment was developed to adequately characterize the behavior of the balance over the expected wind tunnel operating ranges (forces/moments, temperatures, and pressures). The experimental design was based on a Taylor-series expansion in the seven factors for the mathematical models. Model inversion was required to calculate the aerodynamic forces and moments as a function of the strain-gage readings. Details regarding transducer on-board compensation techniques, experimental design development, mathematical modeling, and wind tunnel data reduction are included in this paper.

Videometric Applications in Wind Tunnels

NASA Technical Reports Server (NTRS)

Burner, A. W.; Radeztsky, R. H.; Liu, Tian-Shu

1997-01-01

Videometric measurements in wind tunnels can be very challenging due to the limited optical access, model dynamics, optical path variability during testing, large range of temperature and pressure, hostile environment, and the requirements for high productivity and large amounts of data on a daily basis. Other complications for wind tunnel testing include the model support mechanism and stringent surface finish requirements for the models in order to maintain aerodynamic fidelity. For these reasons nontraditional photogrammetric techniques and procedures sometimes must be employed. In this paper several such applications are discussed for wind tunnels which include test conditions with Mach number from low speed to hypersonic, pressures from less than an atmosphere to nearly seven atmospheres, and temperatures from cryogenic to above room temperature. Several of the wind tunnel facilities are continuous flow while one is a short duration blowdown facility. Videometric techniques and calibration procedures developed to measure angle of attack, the change in wing twist and bending induced by aerodynamic load, and the effects of varying model injection rates are described. Some advantages and disadvantages of these techniques are given and comparisons are made with non-optical and more traditional video photogrammetric techniques.

Review of Aeronautical Wind Tunnel Facilities

NASA Technical Reports Server (NTRS)

1988-01-01

The nation's aeronautical wind tunnel facilities constitute a valuable technological resource and make a significant contribution to the global supremacy of U.S. aircraft, both civil and military. At the request of NASA, the National Research Council's Aeronautics and Space Engineering Board organized a commitee to review the state of repair, adequacy, and future needs of major aeronautical wind tunnel facilities in meeting national goals. The comittee identified three main areas where actions are needed to sustain the capability of NASA's aeronautical wind tunnel facilities to support the national aeronautical research and development activities: tunnel maintenance and upgrading, productivity enhancement, and accommodation of new requirements (particularly in hypersonics). Each of these areas are addressed and the committee recommendations for appropriate actions presented.

Base pressure and heat transfer tests of the 0.0225-scale space shuttle plume simulation model (19-OTS) in yawed flight conditions in the NASA-Lewis 10x10-foot supersonic wind tunnel (test IH83)

NASA Technical Reports Server (NTRS)

Foust, J. W.

1979-01-01

Wind tunnel tests were performed to determine pressures, heat transfer rates, and gas recovery temperatures in the base region of a rocket firing model of the space shuttle integrated vehicle during simulated yawed flight conditions. First and second stage flight of the space shuttle were simulated by firing the main engines in conjunction with the SRB rocket motors or only the SSME's into the continuous tunnel airstream. For the correct rocket plume environment, the simulated altitude pressures were halved to maintain the rocket chamber/altitude pressure ratio. Tunnel freestream Mach numbers from 2.2 to 3.5 were simulated over an altitude range of 60 to 130 thousand feet with varying angle of attack, yaw angle, nozzle gimbal angle and SRB chamber pressure. Gas recovery temperature data derived from nine gas temperature probe runs are presented. The model configuration, instrumentation, test procedures, and data reduction are described.

Aeronautical Wind Tunnels, Europe and Asia

DTIC Science & Technology

2006-02-01

User Fees Contact Information Dr. Surjatin Wiriadidjaja, UPT-LAGG, BPP Teknologi, Puspiptek, Serpong, Tangerang 15310, Indonesia. Tel: (62) 21 756...of the tunnel, FFA T1500 Transonic Wind Tunnel Circuit (Sweden) manufactured by The Swedish Defense Research Agency (FOI). 2.4 m Transonic Wind

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.

Scramjet Tests in a Shock Tunnel at Flight Mach 7, 10, and 15 Conditions

NASA Technical Reports Server (NTRS)

Rogers, R. C.; Shih, A. T.; Tsai, C.-Y.; Foelsche, R. O.

2001-01-01

Tests of the Hyper-X scramjet engine flowpath have been conducted in the HYPULSE shock tunnel at conditions duplicating the stagnation enthalpy at flight Mach 7, 10, and 15. For the tests at Mach 7 and 10 HYPULSE was operated as a reflected-shock tunnel; at the Mach 15 condition, HYPULSE was operated as a shock-expansion tunnel. The test conditions matched the stagnation enthalpy of a scramjet engine on an aerospace vehicle accelerating through the atmosphere along a 1000 psf dynamic pressure trajectory. Test parameter variation included fuel equivalence ratios from lean (0.8) to rich (1.5+); fuel composition from pure hydrogen to mixtures of 2% and 5% silane in hydrogen by volume; and inflow pressure and Mach number made by changing the scramjet model mounting angle in the HYPULSE test chamber. Data sources were wall pressures and heat flux distributions and schlieren and fuel plume imaging in the combustor/nozzle sections. Data are presented for calibration of the facility nozzles and the scramjet engine model. Comparisons of pressure distributions and flowpath streamtube performance estimates are made for the three Mach numbers tested.

Wind tunnel wall interference

NASA Technical Reports Server (NTRS)

Newman, Perry A.; Mineck, Raymond E.; Barnwell, Richard W.; Kemp, William B., Jr.

1986-01-01

About a decade ago, interest in alleviating wind tunnel wall interference was renewed by advances in computational aerodynamics, concepts of adaptive test section walls, and plans for high Reynolds number transonic test facilities. Selection of NASA Langley cryogenic concept for the National Transonic Facility (NTF) tended to focus the renewed wall interference efforts. A brief overview and current status of some Langley sponsored transonic wind tunnel wall interference research are presented. Included are continuing efforts in basic wall flow studies, wall interference assessment/correction procedures, and adaptive wall technology.

A century of wind tunnels since Eiffel

NASA Astrophysics Data System (ADS)

Chanetz, Bruno

2017-08-01

Fly higher, faster, preserve the life of test pilots and passengers, many challenges faced by man since the dawn of the twentieth century, with aviation pioneers. Contemporary of the first aerial exploits, wind tunnels, artificially recreating conditions encountered during the flight, have powerfully contributed to the progress of aeronautics. But the use of wind tunnels is not limited to aviation. The research for better performance, coupled with concern for energy saving, encourages manufacturers of ground vehicles to perform aerodynamic tests. Buildings and bridge structures are also concerned. This article deals principally with the wind tunnels built at ONERA during the last century. Somme wind tunnels outside ONERA, even outside France, are also evocated when their characteristics do not exist at ONERA.

An Experimental Device for Generating High Frequency Perturbations in Supersonic Wind Tunnels

NASA Technical Reports Server (NTRS)

Melcher, Kevin J.; Ibrahim, Mounir B.

1996-01-01

This paper describes the analytical study of a device that has been proposed as a mechanism for generating gust-like perturbations in supersonic wind tunnels. The device is envisioned as a means to experimentally validate dynamic models and control systems designed for high-speed inlets. The proposed gust generator is composed of two flat trapezoidal plates that modify the properties of the flow ingested by the inlet. One plate may be oscillated to generate small perturbations in the flow. The other plate is held stationary to maintain a constant angle-of-attack. Using an idealized approach, design equations and performance maps for the new device were developed from the compressible flow relations. A two-dimensional CFD code was used to confirm the correctness of these results. The idealized approach was then used to design and evaluate a new gust generator for a 3.05-meter by 3.05-meter (10-foot by 10-foot) supersonic wind tunnel.

Wind-Tunnel Investigations of Blunt-Body Drag Reduction Using Forebody Surface Roughness

NASA Technical Reports Server (NTRS)

Whitmore, Stephen A.; Sprague, Stephanie; Naughton, Jonathan W.; Curry, Robert E. (Technical Monitor)

2001-01-01

This paper presents results of wind-tunnel tests that demonstrate a novel drag reduction technique for blunt-based vehicles. For these tests, the forebody roughness of a blunt-based model was modified using micomachined surface overlays. As forebody roughness increases, boundary layer at the model aft thickens and reduces the shearing effect of external flow on the separated flow behind the base region, resulting in reduced base drag. For vehicle configurations with large base drag, existing data predict that a small increment in forebody friction drag will result in a relatively large decrease in base drag. If the added increment in forebody skin drag is optimized with respect to base drag, reducing the total drag of the configuration is possible. The wind-tunnel tests results conclusively demonstrate the existence of a forebody dragbase drag optimal point. The data demonstrate that the base drag coefficient corresponding to the drag minimum lies between 0.225 and 0.275, referenced to the base area. Most importantly, the data show a drag reduction of approximately 15% when the drag optimum is reached. When this drag reduction is scaled to the X-33 base area, drag savings approaching 45,000 N (10,000 lbf) can be realized.

Wind-tunnel experiments of scalar transport in aligned and staggered wind farms

NASA Astrophysics Data System (ADS)

Zhang, W.; Markfort, C. D.; Porté-Agel, F.

2012-04-01

Wind energy is the fastest growing renewable energy worldwide, and it is expected that many more large-scale wind farms will be built and will cover a significant portion of land and ocean surfaces. By extracting kinetic energy from the atmospheric boundary layer, wind farms may affect the exchange/transport of momentum, heat and moisture between the atmosphere and land surface. To ensure the long-term sustainability of wind energy, it is important to understand the influence of large-scale wind farms on land-atmosphere interaction. Knowledge of this impact will also be useful to improve parameterizations of wind farms in numerical prediction tools, such as large-scale weather models and large-eddy simulation. Here, we present wind-tunnel measurements of the surface scalar (heat) flux from model wind farms, consisting of more than 10 rows of wind turbines, in a turbulent boundary layer with a surface heat source. Spatially distributed surface heat flux was obtained in idealized aligned and staggered wind farm layouts, having the same turbine distribution density. Measurements, using surface-mounted heat flux sensors, were taken at the 11th out of 12 rows of wind turbines, where the mean flow achieves a quasi-equilibrium state. In the aligned farm, there exist two distinct regions of increased and decreased surface heat flux on either side of turbine columns. The regions are correlated with coherent wake rotation in the turbine-array. On the upwelling side there is decreased flux, while on the downwelling side cool air moves towards the surface causing increased flux. For the staggered farm, the surface heat flux exhibits a relatively uniform distribution and an overall reduction with respect to the boundary layer flow, except in the vicinity of the turbine tower. This observation is also supported by near-surface temperature and turbulent heat flux measured using a customized x-wire/cold-wire. The overall surface heat flux, relative to that of the boundary layer

DARPA/ARFL/NASA Smart Wing second wind tunnel test results

NASA Astrophysics Data System (ADS)

Scherer, Lewis B.; Martin, Christopher A.; West, Mark N.; Florance, Jennifer P.; Wieseman, Carol D.; Burner, Alpheus W.; Fleming, Gary A.

1999-07-01

To quantify the benefits of smart materials and structures adaptive wing technology. Northrop Grumman Corp. built and tested two 16 percent scale wind tunnel models of a fighter/attach aircraft under the DARPA/AFRL/NASA Smart Materials and Structures Development - Smart Wing Phase 1. Performance gains quantified included increased pitching moment, increased rolling moment and improved pressure distribution. The benefits were obtained for hingeless, contoured trailing edge control surfaces with embedded shape memory alloy wires and spanwise wing twist effected by SMA torque tube mechanism, compared to convention 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 16 ft Transonic Dynamic Tunnel in June 1998. Successful results obtained were: 1) 5 degrees of spanwise twist and 8-12 percent increase in rolling moment utilizing a single SMA torque tube, 2) 12 degrees of deflection, and 10 percent increase in rolling moment due to hingeless, contoured aileron, and 3) demonstration of optical techniques for measuring spanwise twist and deflected shape.

Real-Time Frequency Response Estimation Using Joined-Wing SensorCraft Aeroelastic Wind-Tunnel Data

NASA Technical Reports Server (NTRS)

Grauer, Jared A; Heeg, Jennifer; Morelli, Eugene A

2012-01-01

A new method is presented for estimating frequency responses and their uncertainties from wind-tunnel data in real time. The method uses orthogonal phase-optimized multi- sine excitation inputs and a recursive Fourier transform with a least-squares estimator. The method was first demonstrated with an F-16 nonlinear flight simulation and results showed that accurate short period frequency responses were obtained within 10 seconds. The method was then applied to wind-tunnel data from a previous aeroelastic test of the Joined- Wing SensorCraft. Frequency responses describing bending strains from simultaneous control surface excitations were estimated in a time-efficient manner.

The Denis-gruson Six-component Wind-tunnel Balance

NASA Technical Reports Server (NTRS)

1935-01-01

The 6.C.1 balance is the first fully automatic balance assuring a continuous and simultaneous record of the aerodynamic characteristics of an airfoil in a wind tunnel. Because of the rapidity of the measurements a complete polar (six components) requires only about three minutes of wind, that is to say, of motive power, which is of interest for wind tunnels with high efficiency factors and may lead to the economical design of large size wind tunnels.

Jet-Boundary Corrections for Reflection-Plane Models in Rectangular Wind Tunnels

DTIC Science & Technology

1943-01-01

clock- wiso) located at a distance d above the tunnel center line and at distances equal to VIand —vI from the reflection wall. Tho single trailing vortex...neglected. The angle-.xkmnge is usually small, les than %O. M & fw-refleciicur plane models in 7-by 10-foof cIosedrectangular wind fu-meIs 04

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.

Hardwall acoustical characteristics and measurement capabilities of the NASA Lewis 9 x 15 foot low speed wind tunnel

NASA Technical Reports Server (NTRS)

Rentz, P. E.

1976-01-01

Experimental evaluations of the acoustical characteristics and source sound power and directionality measurement capabilities of the NASA Lewis 9 x 15 foot low speed wind tunnel in the untreated or hardwall configuration were performed. The results indicate that source sound power estimates can be made using only settling chamber sound pressure measurements. The accuracy of these estimates, expressed as one standard deviation, can be improved from + or - 4 db to + or - 1 db if sound pressure measurements in the preparation room and diffuser are also used and source directivity information is utilized. A simple procedure is presented. Acceptably accurate measurements of source direct field acoustic radiation were found to be limited by the test section reverberant characteristics to 3.0 feet for omni-directional and highly directional sources. Wind-on noise measurements in the test section, settling chamber and preparation room were found to depend on the sixth power of tunnel velocity. The levels were compared with various analytic models. Results are presented and discussed.

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.

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)

NASA ERA Integrated CFD for Wind Tunnel Testing of Hybrid Wing-Body Configuration

NASA Technical Reports Server (NTRS)

Garcia, Joseph A.; Melton, John E.; Schuh, Michael; James, Kevin D.; Long, Kurtis R.; Vicroy, Dan D.; Deere, Karen A.; Luckring, James M.; Carter, Melissa B.; Flamm, Jeffrey D.;

NASA ERA Integrated CFD for Wind Tunnel Testing of Hybrid Wing-Body Configuration

NASA Technical Reports Server (NTRS)

Garcia, Joseph A.; Melton, John E.; Schuh, Michael; James, Kevin D.; Long, Kurt R.; Vicroy, Dan D.; Deere, Karen A.; Luckring, James M.; Carter, Melissa B.; Flamm, Jeffrey D.;

Evaporation Rates of Chemical Warfare Agents Using 5-CM Wind Tunnels I. CASARM Sulfur Mustard (HD) from Glass

DTIC Science & Technology

2008-10-01

Blank CONTENTS 1. INTRODUCTION 9 2. EXPERIMENTAL PROCEDURES 9 2.1 Wind Tunnel 9 2.2 Agent 10 2.3 Gas Chromatography /Mass Spectrometry Detection 10...protective equipment. 2.3 Gas Chromatography /Mass Spectrometry Detection (GC/MSD) The GC/MSD analysis of the thermal desorption tubes was performed on a...coupled to thermal desorption tubes that were analyzed using gas chromatography /mass spectrometry detection (GC/MSD). Differences between the tunnels

Wind Tunnel to Atmospheric Mapping for Static Aeroelastic Scaling

NASA Technical Reports Server (NTRS)

Heeg, Jennifer; Spain, Charles V.; Rivera, J. A.

2004-01-01

Wind tunnel to Atmospheric Mapping (WAM) is a methodology for scaling and testing a static aeroelastic wind tunnel model. The WAM procedure employs scaling laws to define a wind tunnel model and wind tunnel test points such that the static aeroelastic flight test data and wind tunnel data will be correlated throughout the test envelopes. This methodology extends the notion that a single test condition - combination of Mach number and dynamic pressure - can be matched by wind tunnel data. The primary requirements for affecting this extension are matching flight Mach numbers, maintaining a constant dynamic pressure scale factor and setting the dynamic pressure scale factor in accordance with the stiffness scale factor. The scaling is enabled by capabilities of the NASA Langley Transonic Dynamics Tunnel (TDT) and by relaxation of scaling requirements present in the dynamic problem that are not critical to the static aeroelastic problem. The methodology is exercised in two example scaling problems: an arbitrarily scaled wing and a practical application to the scaling of the Active Aeroelastic Wing flight vehicle for testing in the TDT.

Fast Laser Holographic Interferometry For Wind Tunnels

NASA Technical Reports Server (NTRS)

Lee, George

1989-01-01

Proposed system makes holographic interferograms quickly in wind tunnels. Holograms reveal two-dimensional flows around airfoils and provide information on distributions of pressure, structures of wake and boundary layers, and density contours of flow fields. Holograms form quickly in thermoplastic plates in wind tunnel. Plates rigid and left in place so neither vibrations nor photgraphic-development process degrades accuracy of holograms. System processes and analyzes images quickly. Semiautomatic micro-computer-based desktop image-processing unit now undergoing development moves easily to wind tunnel, and its speed and memory adequate for flows about airfoils.

Fan array wind tunnel: a multifunctional, complex environmental flow manipulator

NASA Astrophysics Data System (ADS)

Dougherty, Christopher; Veismann, Marcel; Gharib, Morteza

2017-11-01

The recent emergence of small unmanned aerial vehicles (UAVs) has reshaped the aerospace testing environment. Traditional closed-loop wind tunnels are not particularly suited nor easily retrofit to take advantage of these coordinated, controls-based rotorcraft. As such, a highly configurable, novel wind tunnel aimed at addressing the unmet technical challenges associated with single or formation flight performance of autonomous drone systems is presented. The open-loop fan array wind tunnel features 1296 individually controllable DC fans arranged in a 2.88m x 2.88m array. The fan array can operate with and without a tunnel enclosure and is able to rotate between horizontal and vertical testing configurations. In addition to standard variable speed uniform flow, the fan array can generate both unsteady and shear flows. Through the aid of smaller side fan array units, vortex flows are also possible. Conceptual design, fabrication, and validation of the tunnel performance will be presented, including theoretical and computational predictions of flow speed and turbulence intensity. Validation of these parameters is accomplished through standard pitot-static and hot-wire techniques. Particle image velocimetry (PIV) of various complex flows will also be shown. This material is based upon work supported by the Center for Autonomous Systems and Technologies (CAST) at the Graduate Aerospace Laboratories of the California Institute of Technology (GALCIT).

Wind-Tunnel Investigation of a Balloon as a Towed Decelerator at Mach Numbers from 1.47 to 2.50

NASA Technical Reports Server (NTRS)

McShera, John T.; Keyes, J. Wayne

1961-01-01

A wind-tunnel investigation has been conducted to study the characteristics of a towed spherical balloon as a drag device at Mach numbers from 1.47 to 2.50, Reynolds numbers from 0.36 x 10(exp 6) to 1.0 x 10(exp 6) , and angles of attack from -15 to 15 deg. Towed spherical balloons were found to be stable at supersonic speeds. The drag coefficient of the balloon is reduced by the presence of a tow cable and a further reduction occurs with the addition of a payload. The balloon inflation pressure required to maintain an almost spherical shape is about equal to the free-stream dynamic pressure. Measured pressure and temperature distribution around the balloon alone were in fair agreement with predicted values. There was a pronounced decrease in the pressure coefficients on the balloon when attached to a tow cable behind a payload.

Wind Tunnel Results of the Aerodynamic Performance of a 1/8-Scale Model of a Twin-Engine Transport with Multi-Element Wing

NASA Technical Reports Server (NTRS)

Laflin, Brenda E. Gile; Applin, Zachary T.; Jones, Kenneth M.

1997-01-01

A wind tunnel investigation was performed in the 14- by 22-Foot Subsonic Tunnel on a pressure instrumented 1/8-scale twin-engine subsonic transport to better understand the flow physics on a multi-element wing section. The wing consisted of a part-span, triple-slotted trailing edge flap, inboard leading-edge Krueger flap and an outboard leading-edge slat. The model was instrumented with flush pressure ports at the fuselage centerline and seven spanwise wing locations. The model was tested in cruise, take-off and landing configurations at dynamic pressures and Mach numbers from 10 lbf/ft(exp 2) to 50 lbf/ft(exp 2) and 0.08 to 0.17, respectively. This resulted in corresponding Reynolds numbers of 0.8 x 10(exp 5) to 1.8 x 10(exp 6). Pressure data were collected using electronically scanned pressure devices and force and moment data were collected with a six component strain gauge balance. Results are presented for various control surface deflections over an angle-of-attack range from -4 degrees to 16 degrees and sideslip angle range from -10 degrees to 10 degrees. Longitudinal and lateral directional aerodynamic data are presented as well as chordwise pressure distributions at the seven spanwise wing locations and the fuselage centerline.

Comparison of Space Shuttle Orbiter low-speed static stability and control derivatives obtained from wind-tunnel and approach and landing flight tests

NASA Technical Reports Server (NTRS)

Freeman, D. C., Jr.; Spencer, B., Jr.

1980-01-01

Tests were conducted in the 8 foot transonic pressure tunnel to obtain wind tunnel data for comparison with static stability and control parameters measured on the space shuttle orbiter approach and landing flight tests. The longitudinal stability, elevon effectiveness, lateral directional stability, and aileron effectiveness derivatives were determined from the wind tunnel data and compared with the flight test results. The comparison covers a range of angles of attack from approximately 2 deg to 10 deg at subsonic Mach numbers of 0.41 to 0.56. In general the wind tunnel results agreed well with the flight test results, indicating the wind tunnel data is applicable to the design of entry vehicles for subsonic speeds over the angle of attack range studied.

Wind tunnel interference factors for high-lift wings in closed wind tunnels. Ph.D. Thesis - Princeton Univ.

NASA Technical Reports Server (NTRS)

Joppa, R. G.

1973-01-01

A problem associated with the wind tunnel testing of very slow flying aircraft is the correction of observed pitching moments to free air conditions. The most significant effects of such corrections are to be found at moderate downwash angles typical of the landing approach. The wind tunnel walls induce interference velocities at the tail different from those induced at the wing, and these induced velocities also alter the trajectory of the trailing vortex system. The relocated vortex system induces different velocities at the tail from those experienced in free air. The effect of the relocated vortex and the walls is to cause important changes in the measured pitching moments in the wind tunnel.

Experimental hingeless rotor characteristics at low advance ratio with thrust. [wind tunnel tests of rotary wing operating at moderate to high lift

NASA Technical Reports Server (NTRS)

London, R. J.; Watts, G. A.; Sissingh, G. J.

1973-01-01

An experimental investigation to determine the dynamic characteristics of a hingeless rotor operating at moderate to high lift was conducted on a small scale, 7.5-foot diameter, four-bladed hingeless rotor model in a 7 x 10-foot wind tunnel. The primary objective of this research program was the empirical determination of the rotor steady-state and frequency responses to swashplate and body excitations. Collective pitch was set from 0 to 20 degrees, with the setting at a particular advance ratio limited by the cyclic pitch available for hub moment trim. Advance ratio varied from 0.00 to 0.36 for blades with nondimensional first-flap frequencies at 1.15, 1.28 and 1.33 times the rotor rotation frequency. Several conditions were run with the rotor operating in the transition regime. Rotor response at high lift is shown to be generally nonlinear in this region. As a secondary objective an experimental investigation of the rotor response to 4/revolution swashplate excitations at advance ratios of 0.2 to 0.85 and at a nondimensional, first-flap modal frequency of 1.34 was also conducted, using the 7 x 10-foot wind tunnel. It is shown that 4/revolution swashplate inputs are a method for substantially reducing rotor-induced, shafttransmitted vibratory forces.

Low Pressure Seeder Development for PIV in Large Scale Open Loop Wind Tunnels

NASA Astrophysics Data System (ADS)

Schmit, Ryan

2010-11-01

A low pressure seeding techniques have been developed for Particle Image Velocimetry (PIV) in large scale wind tunnel facilities was performed at the Subsonic Aerodynamic Research Laboratory (SARL) facility at Wright-Patterson Air Force Base. The SARL facility is an open loop tunnel with a 7 by 10 foot octagonal test section that has 56% optical access and the Mach number varies from 0.2 to 0.5. A low pressure seeder sprayer was designed and tested in the inlet of the wind tunnel. The seeder sprayer was designed to produce an even and uniform distribution of seed while reducing the seeders influence in the test section. ViCount Compact 5000 using Smoke Oil 180 was using as the seeding material. The results show that this low pressure seeder does produce streaky seeding but excellent PIV images are produced.

Survey Of Wind Tunnels At Langley Research Center

NASA Technical Reports Server (NTRS)

Bower, Robert E.

1989-01-01

Report presented at AIAA 14th Aerodynamic Testing Conference on current capabilities and planned improvements at NASA Langley Research Center's major wind tunnels. Focuses on 14 major tunnels, 8 unique in world, 3 unique in country. Covers Langley Spin Tunnel. Includes new National Transonic Facility (NTF). Also surveys Langley Unitary Plan Wind Tunnel (UPWT). Addresses resurgence of inexpensive simple-to-operate research tunnels. Predicts no shortage of tools for aerospace researcher and engineer in next decade or two.

Design of Rail Instrumentation for Wind Tunnel Sonic Boom Measurements and Computational-Experimental Comparisons

NASA Technical Reports Server (NTRS)

Cliff, Susan E.; Elmiligui, A.; Aftosmis, M.; Morgenstern, J.; Durston, D.; Thomas, S.

2012-01-01

An innovative pressure rail concept for wind tunnel sonic boom testing of modern aircraft configurations with very low overpressures was designed with an adjoint-based solution-adapted Cartesian grid method. The computational method requires accurate free-air calculations of a test article as well as solutions modeling the influence of rail and tunnel walls. Specialized grids for accurate Euler and Navier-Stokes sonic boom computations were used on several test articles including complete aircraft models with flow-through nacelles. The computed pressure signatures are compared with recent results from the NASA 9- x 7-foot Supersonic Wind Tunnel using the advanced rail design.

Propulsion simulator for magnetically-suspended wind tunnel models

NASA Technical Reports Server (NTRS)

Joshi, P. B.; Malonson, M. R.; Sacco, G. P.; Goldey, C. L.; Garbutt, Keith; Goodyer, M.

1992-01-01

In order to demonstrate the measurement of aerodynamic forces/moments, including the effects of exhaust jets in Magnetic Suspension and Balance System (MSBS) wind tunnels, two propulsion simulator models were developed at Physical Sciences Inc. (PSI). Both the small-scale model (1 in. diameter X 8 in. long) and the large-scale model (2.5 in. diameter X 15 in. long) employed compressed, liquefied carbon dioxide as a propellant. The small-scale simulator, made from a highly magnetizable iron alloy, was demonstrated in the 7 in. MSBS wind tunnel at the University of Southampton. It developed a maximum thrust of approximate 1.3 lbf with a 0.098 in. diameter nozzle and 0.7 lbf with a 0.295 in. diameter nozzle. The Southampton MSBS was able to control the simulator at angles-of attack up to 20 deg. The large-scale simulator was demonstrated to operate in both a steady-state and a pulse mode via a miniaturized solinoid valve. It developed a stable and repeatable thrust of 2.75 lbf over a period of 4s and a nozzle pressure ratio (NPR) of 5.

Flight-measured X-24A lifting body control surface hinge moments and correlation with wind tunnel predictions

NASA Technical Reports Server (NTRS)

Tang, M. H.; Pearson, G. P. E.

1973-01-01

Control-surface hinge-moment measurements obtained in the X-24A lifting body flight-test program are compared with results from wind-tunnel tests. The effects of variations in angle of attack, angle of sideslip, rudder bias, rudder deflection, upper-flap deflection, lower-flap deflection, Mach number, and rocket-engine operation on the control-surface hinge moments are presented. In-flight motion pictures of tufts attached to the inboard side of the right fin and the rudder and upper-flap surfaces are discussed.

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.

Cryogenic wind tunnels: Unique capabilities for the aerodynamicist

NASA Technical Reports Server (NTRS)

Hall, R. M.

1976-01-01

The cryogenic wind-tunnel concept as a practical means for improving ground simulation of transonic flight conditions. The Langley 1/3-meter transonic cryogenic tunnel is operational, and the design of a cryogenic National Transonic Facility is undertaken. A review of some of the unique capabilities of cryogenic wind tunnels is presented. In particular, the advantages of having independent control of tunnel Mach number, total pressure, and total temperature are highlighted. This separate control over the three tunnel parameters will open new frontiers in Mach number, Reynolds number, aeroelastic, and model-tunnel interaction studies.

Boundary Condition Study for the Juncture Flow Experiment in the NASA Langley 14x22-Foot Subsonic Wind Tunnel

NASA Technical Reports Server (NTRS)

Rumsey, C. L.; Carlson, J.-R.; Hannon, J. A.; Jenkins, L. N.; Bartram, S. M.; Pulliam, T. H.; Lee, H. C.

2017-01-01

Because future wind tunnel tests associated with the NASA Juncture Flow project are being designed for the purpose of CFD validation, considerable effort is going into the characterization of the wind tunnel boundary conditions, particularly at inflow. This is important not only because wind tunnel flowfield nonuniformities can play a role in integrated testing uncertainties, but also because the better the boundary conditions are known, the better CFD can accurately represent the experiment. This paper describes recent investigative wind tunnel tests involving two methods to measure and characterize the oncoming flow in the NASA Langley 14- by 22-Foot Subsonic Tunnel. The features of each method, as well as some of their pros and cons, are highlighted. Boundary conditions and modeling tactics currently used by CFD for empty-tunnel simulations are also described, and some results using three different CFD codes are shown. Preliminary CFD parametric studies associated with the Juncture Flow model are summarized, to determine sensitivities of the flow near the wing-body juncture region of the model to a variety of modeling decisions.

Reduction of Background Noise in the NASA Ames 40- by 80-Foot Wind Tunnel

NASA Technical Reports Server (NTRS)

Jaeger, Stephen M.; Allen, Christopher S.; Soderman, Paul T.; Olson, Larry E. (Technical Monitor)

1995-01-01

Background noise in both open-jet and closed wind tunnels adversely affects the signal-to-noise ratio of acoustic measurements. To measure the noise of increasingly quieter aircraft models, the background noise will have to be reduced by physical means or through signal processing. In a closed wind tunnel, such as the NASA Ames 40- by 80- Foot Wind Tunnel, the principle background noise sources can be classified as: (1) fan drive noise; (2) microphone self-noise; (3) aerodynamically induced noise from test-dependent hardware such as model struts and junctions; and (4) noise from the test section walls and vane set. This paper describes the steps taken to minimize the influence of each of these background noise sources in the 40 x 80.

The F2 wind tunnel at Fauga-Mauzac

NASA Technical Reports Server (NTRS)

Afchain, D.; Broussaud, P.; Frugier, M.; Rancarani, G.

1984-01-01

Details on the French subsonic wind-tunnel F2 that becomes operational on July 1983 are presented. Some of the requirements were: (1) installation of models on any wall of the facility, (2) good observation points due to transparent walls, (3) smooth flow, (4) a laser velocimeter, and (5) easy access and handling. The characteristics include a nonpressurized return circuit, dimensions of 5 x 1.4 x 1.8 m, maximum velocity of 100 m/s and a variable speed fan of 683 kW.

Calibration and test capabilities of the Langley 7- by 10- foot high speed tunnel

NASA Technical Reports Server (NTRS)

Fox, C. H., Jr.; Huffman, J. K.

1977-01-01

The results of a new subsonic calibration of the Langley 7 by 10 foot high speed tunnel with the test section in a solid wall configuration are presented. A description of the test capabilities of the 7 by 10 foot high speed tunnel is also given.

Operational flow visualization techniques in the Langley Unitary Plan Wind Tunnel

NASA Technical Reports Server (NTRS)

Corlett, W. A.

1982-01-01

The unitary plan wind tunnel (UPWT) uses in daily operation are shown. New ideas for improving the quality of established flow visualization methods are developed and programs on promising new flow visualization techniques are pursued. The unitary plan wind tunnel is a supersonic facility, referred to as a production facility, although the majority of tests are inhouse basic research investigations. The facility has two 4 ft. by 4 ft. test sections which span a Mach range from 1.5 to 4.6. The cost of operation is about $10 per minute. Problems are the time required for a flow visualization test setup and investigation costs and the ability to obtain consistently repeatable results. Examples of sublimation, vapor screen, oil flow, minitufts, schlieren, and shadowgraphs taken in UPWT are presented. All tests in UPWT employ one or more of the flow visualization techniques.

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

NASA Technical Reports Server (NTRS)

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

1985-01-01

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

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.

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.

An isentropic compression-heated Ludweig tube transient wind tunnel

NASA Technical Reports Server (NTRS)

Magari, Patrick J.; Lagraff, John E.

1991-01-01

Theoretical development and experimental results show that the Ludweig tube with isentropic heating (LICH) transient wind tunnel described is a viable means of producing flow conditions that are suitable for a variety of experimental investigations. A complete analysis of the wave dynamics of the pump tube compression process is presented. The LICH tube operating conditions are very steady and run times are greater than those of other types of transient facilities such as shock tubes and gas tunnels. This facility is well suited for producing flow conditions that are dynamically similar to those found in a gas turbine, i.e., transonic Mach number, gas-to-wall temperature ratios of about 1.5, and Reynolds numbers greater than 10 to the 6th.

Wind Tunnel Tests Conducted to Develop an Icing Flight Simulator

NASA Technical Reports Server (NTRS)

Ratvasky, Thomas P.

2001-01-01

As part of NASA's Aviation Safety Program goals to reduce aviation accidents due to icing, NASA Glenn Research Center is leading a flight simulator development activity to improve pilot training for the adverse flying characteristics due to icing. Developing flight simulators that incorporate the aerodynamic effects of icing will provide a critical element in pilot training programs by giving pilots a pre-exposure of icing-related hazards, such as ice-contaminated roll upset or tailplane stall. Integrating these effects into training flight simulators will provide an accurate representation of scenarios to develop pilot skills in unusual attitudes and loss-of-control events that may result from airframe icing. In order to achieve a high level of fidelity in the flight simulation, a series of wind tunnel tests have been conducted on a 6.5-percent-scale Twin Otter aircraft model. These wind tunnel tests were conducted at the Wichita State University 7- by 10-ft wind tunnel and Bihrle Applied Research's Large Amplitude Multiple Purpose Facility in Neuburg, Germany. The Twin Otter model was tested without ice (baseline), and with two ice configurations: 1) Ice on the horizontal tail only; 2) Ice on the wing, horizontal tail, and vertical tail. These wind tunnel tests resulted in data bases of aerodynamic forces and moments as functions of angle of attack; sideslip; control surface deflections; forced oscillations in the pitch, roll, and yaw axes; and various rotational speeds. A limited amount of wing and tail surface pressure data were also measured for comparison with data taken at Wichita State and with flight data. The data bases from these tests will be the foundation for a PC-based Icing Flight Simulator to be delivered to Glenn in fiscal year 2001.

L10-MnGa based magnetic tunnel junction for high magnetic field sensor

NASA Astrophysics Data System (ADS)

Zhao, X. P.; Lu, J.; Mao, S. W.; Yu, Z. F.; Wang, H. L.; Wang, X. L.; Wei, D. H.; Zhao, J. H.

2017-07-01

We report on the investigation of the magnetic tunnel junction structure designed for high magnetic field sensors with a perpendicularly magnetized L10-MnGa reference layer and an in-plane magnetized Fe sensing layer. A large linear tunneling magnetoresistance ratio up to 27.4% and huge dynamic range up to 5600 Oe have been observed at 300 K, with a low nonlinearity of 0.23% in the optimized magnetic tunnel junction (MTJ). The field response of tunneling magnetoresistance is discussed to explain the field sensing properties in the dynamic range. These results indicate that L10-MnGa based orthogonal MTJ is a promising candidate for a high performance magnetic field sensor with a large dynamic range, high endurance and low power consumption.

A wind tunnel study on the effects of complex topography on wind turbine performance

NASA Astrophysics Data System (ADS)

Howard, Kevin; Hu, Stephen; Chamorro, Leonardo; Guala, Michele

2012-11-01

A set of wind tunnel experiments were conducted to study the response of a wind turbine under flow conditions typically observed at the wind farm scale, in complex terrain. A scale model wind turbine was placed in a fully developed turbulent boundary layer flow obtained in the SAFL Wind Tunnel. Experiments focused on the performance of a turbine model, under the effects induced by a second upwind turbine or a by three-dimensional, sinusoidal hill, peaking at the turbine hub height. High frequency measurements of fluctuating streamwise and wall normal velocities were obtained with a X-wire anemometer simultaneously with the rotor angular velocity and the turbine(s) voltage output. Velocity measurements in the wake of the first turbine and of the hill were used to determine the inflow conditions for the downwind test turbine. Turbine performance was inferred by the mean and fluctuating voltage statistics. Specific experiments were devoted to relate the mean voltage to the mean hub velocity, and the fluctuating voltage to the unsteadiness in the rotor kinematics induced by the perturbed (hill or turbine) or unperturbed (boundary layer) large scales of the incoming turbulent flow. Results show that the voltage signal can be used to assess turbine performance in complex flows.

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.

Wind Tunnel Complex at the Aircraft Engine Research Laboratory

NASA Image and Video Library

1945-09-21

This aerial photograph shows the entire original wind tunnel complex at the National Advisory Committee for Aeronautics (NACA) Aircraft Engine Research Laboratory. The large Altitude Wind Tunnel (AWT) at the center of the photograph dominates the area. The Icing Research Tunnel to the right was incorporated into the lab’s design to take advantage of the AWT’s powerful infrastructure. The laboratory’s first supersonic wind tunnel was added to this complex just prior to this September 1945 photograph. The AWT was the nation’s only wind tunnel capable of studying full-scale engines in simulated flight conditions. The AWT’s test section and control room were within the two-story building near the top of the photograph. The exhauster equipment used to thin the airflow and the drive motor for the fan were in the building to the right of the tunnel. The unique refrigeration equipment was housed in the structure to the left of the tunnel. The Icing Research Tunnel was an atmospheric tunnel that used the AWT’s refrigeration equipment to simulate freezing rain inside its test section. A spray bar system inside the tunnel was originally used to create the droplets. The 18- by 18-inch supersonic wind tunnel was built in the summer of 1945 to take advantage of the AWT’s powerful exhaust system. It was the lab’s first supersonic tunnel and could reach Mach 1.91. Eventually the building would house three small supersonic tunnels, referred to as the “stack tunnels” because of the vertical alignment. The two other tunnels were added to this structure in 1949 and 1951.

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.

Comparative wind tunnel test at high Reynolds numbers of NACA 64 621 airfoils with two aileron configurations

NASA Technical Reports Server (NTRS)

Gregorek, G. M.

1995-01-01

An experimental program to measure the aerodynamic characteristics of the NACA 64-621 airfoil when equipped with plain ailerons of 0.38 chord and 0.30 chord and with 0.38 chord balanced aileron has been conducted in the pressurized O.S.U. 6 x 12 ft High Reynolds Number Wind Tunnel. Surface pressures were measured and integrated to yield lift and pressure drag coefficients for angles of attack from -3 to +42 deg and for selected aileron deflections from 0 to -90 deg at nominal Mach and Reynolds numbers of 0.25 and 5 x 10(exp 6). When resolved into thrust coefficient for wind turbine aerodynamic control applications, the data indicated the anticipated decrease in thrust coefficient with negative aileron deflection at low angles of attack; however, as angle of attack increased, thrust coefficients eventually became positive. All aileron configurations, even at -90 deg deflections showed this trend. Hinge moments for each configuration complete the data set.

Screens Would Protect Wind-Tunnel Fan Blades

NASA Technical Reports Server (NTRS)

Farmer, Moses G.

1992-01-01

Butterfly screen installed in wind tunnel between test section and fan blades to prevent debris from reaching fan blades if model structure fails. Protective screens deployed manually or automatically. Concept beneficial anywhere wind tunnels employed. Also useful in areas outside of aerospace industry, such as in airflow design of automobiles and other vehicles.

A new low-turbulence wind tunnel for animal and small vehicle flight experiments

NASA Astrophysics Data System (ADS)

Quinn, Daniel B.; Watts, Anthony; Nagle, Tony; Lentink, David

2017-03-01

Our understanding of animal flight benefits greatly from specialized wind tunnels designed for flying animals. Existing facilities can simulate laminar flow during straight, ascending and descending flight, as well as at different altitudes. However, the atmosphere in which animals fly is even more complex. Flow can be laminar and quiet at high altitudes but highly turbulent near the ground, and gusts can rapidly change wind speed. To study flight in both laminar and turbulent environments, a multi-purpose wind tunnel for studying animal and small vehicle flight was built at Stanford University. The tunnel is closed-circuit and can produce airspeeds up to 50 m s-1 in a rectangular test section that is 1.0 m wide, 0.82 m tall and 1.73 m long. Seamless honeycomb and screens in the airline together with a carefully designed contraction reduce centreline turbulence intensities to less than or equal to 0.030% at all operating speeds. A large diameter fan and specialized acoustic treatment allow the tunnel to operate at low noise levels of 76.4 dB at 20 m s-1. To simulate high turbulence, an active turbulence grid can increase turbulence intensities up to 45%. Finally, an open jet configuration enables stereo high-speed fluoroscopy for studying musculoskeletal control in turbulent flow.

Wind-Tunnel Development of Ailerons for the Curtiss XP-60 Airplanem Special Report

NASA Technical Reports Server (NTRS)

Rogallo, F. M.; Lowry, John G.

1942-01-01

An investigation was made in the LWAL 7- by 10-foot tunnel of internally balanced, sealed ailerons for the Curtiss XP-60 airplane. Ailerons with tabs and. with various amounts of balance were tested. Stick forces were estimated for several aileron arrangements including an arrangement recommended for the airplane. Flight tests of the recommended arrangement are discussed briefly in an appendix, The results of the wind-tunnel and flight tests indicate that the ailerons of large or fast airplanes may be satisfactorily balanced by the method developed.

Real-time simulator for helicopter rotor wind-tunnel operations

NASA Technical Reports Server (NTRS)

Talbot, P. D.; Peterson, R. L.; Graham, D. R.

1986-01-01

This paper describes the elements and operation of a simulator that is being used to train operators of the Rotor Test Apparatus (RTA) in the large-scale 40- by 80-Foot Wind Tunnel at Ames Research Center. The simulator, named TUTOR (for Tunnel Utilization Trainer with Operating Rotor) duplicates the controls of the rotor and its dynamic behavior, as well as the wind-tunnel controls. The simulation software uses a preexisting blade-element model of a four-bladed rotor with flapping and lead-lag degrees of freedom. Equations were developed for all hardware and controls of the RTA and of the wind tunnel that are normally required to perform a wind-tunnel test of a helicopter rotor. The simulator hardware consists of consoles designed to have the same appearance and functions as those in the control room of the 40- by 80-Foot Wind Tunnel, allowing input from three operators who normally establish the required operating conditions during a test run. Normal operating procedures can be practiced, as well as simulated emergencies such as rotor power failure.

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.

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

NASA Technical Reports Server (NTRS)

DeLoach, Richard

2014-01-01

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

Aerodynamic performance of a low-speed wind tunnel.

PubMed

Frechen, F-B; Frey, M; Wett, M; Löser, C

2004-01-01

The determination of the odour mass flow emitted from a source is a very important step and forms the basis for all subsequent considerations and calculations. Wastewater treatment plants, as well as waste treatment facilities, consist of different kinds of odour sources. Unfortunately, most of the sources are passive sources, where no outward air flow-rate can be measured, but where odorants are obviously emitted. Thus, a type of sampling is required that allows to measure the emitted odour flow-rate (OFR). To achieve this, different methods are in use worldwide. Besides indirect methods, such as micrometeorological atmospheric dispersion models, which have not been used in Germany (in other countries due to different problems, direct methods are also used). Direct measurements include hood methods, commonly divided into static flux chambers, dynamic flux chambers and wind tunnels. The wind tunnel that we have been operating in principle since 1983 is different from all subsequent presented wind tunnels, in that we operate it at a considerably lower wind speed than the others. To describe the behaviour of this wind tunnel, measurement of the flow pattern in this low-speed tunnel are under way, and some initial results are presented here.

A large volume 2000 MPA air source for the radiatively driven hypersonic wind tunnel

DOE Office of Scientific and Technical Information (OSTI.GOV)

Constantino, M

1999-07-14

An ultra-high pressure air source for a hypersonic wind tunnel for fluid dynamics and combustion physics and chemistry research and development must provide a 10 kg/s pure air flow for more than 1 s at a specific enthalpy of more than 3000 kJ/kg. The nominal operating pressure and temperature condition for the air source is 2000 MPa and 900 K. A radial array of variable radial support intensifiers connected to an axial manifold provides an arbitrarily large total high pressure volume. This configuration also provides solutions to cross bore stress concentrations and the decrease in material strength with temperature. [hypersonic,more » high pressure, air, wind tunnel, ground testing]« less

On Blockage Corrections for Two-dimensional Wind Tunnel Tests Using the Wall-pressure Signature Method

NASA Technical Reports Server (NTRS)

Allmaras, S. R.

1986-01-01

The Wall-Pressure Signature Method for correcting low-speed wind tunnel data to free-air conditions has been revised and improved for two-dimensional tests of bluff bodies. The method uses experimentally measured tunnel wall pressures to approximately reconstruct the flow field about the body with potential sources and sinks. With the use of these sources and sinks, the measured drag and tunnel dynamic pressure are corrected for blockage effects. Good agreement is obtained with simpler methods for cases in which the blockage corrections were about 10% of the nominal drag values.

Influences of wind and precipitation on different-sized particulate matter concentrations (PM2.5, PM10, PM2.5-10)

NASA Astrophysics Data System (ADS)

Zhang, Boen; Jiao, Limin; Xu, Gang; Zhao, Suli; Tang, Xin; Zhou, Yue; Gong, Chen

2018-06-01

Though it is recognized that meteorology has a great impact on the diffusion, accumulation and transport of air pollutants, few studies have investigated the impacts on different-sized particulate matter concentrations. We conducted a systematic comparative analysis and used the framework of generalized additive models (GAMs) to explore the influences of critical meteorological parameters, wind and precipitation, on PM2.5, PM10 and PM2.5-10 concentrations in Wuhan during 2013-2016. Overall, results showed that the impacts of wind and precipitation on different-sized PM concentrations are significantly different. The fine PM concentrations decreased gradually with the increase of wind speed, while coarse PM concentrations would increase due to dust resuspension under strong wind. Wind direction exerts limited influence on coarse PM concentrations. Wind speed was linearly correlated with log-transformed PM2.5 concentrations, but nonlinearly correlated with log-transformed PM10 and PM2.5-10 concentrations. We also found the PM2.5 and PM2.5-10 concentrations decreased by nearly 60 and 15% when the wind speed was up to 6 m/s, respectively, indicating a stronger negative impact of wind-speed on fine PM than coarse PM. The scavenging efficiency of precipitation on PM2.5-10 was over twice as high as on PM2.5. Our findings may help to understand the impacts of meteorology on different PM concentrations as well as discriminate and forecast variation in particulate matter concentrations.

An Optical and Infrared Time-domain Study of the Supergiant Fast X-Ray Transient Candidate IC 10 X-2

NASA Astrophysics Data System (ADS)

Kwan, Stephanie; Lau, Ryan M.; Jencson, Jacob; Kasliwal, Mansi M.; Boyer, Martha L.; Ofek, Eran; Masci, Frank; Laher, Russ

2018-03-01

We present an optical and infrared (IR) study of IC 10 X-2, a high-mass X-ray binary in the galaxy IC 10. Previous optical and X-ray studies suggest that X-2 is a Supergiant Fast X-ray Transient: a large-amplitude (factor of ∼100), short-duration (hours to weeks) X-ray outburst on 2010 May 21. We analyze R- and g-band light curves of X-2 from the intermediate Palomar Transient Factory taken between 2013 July 15 and 2017 February 14 that show high-amplitude (≳1 mag), short-duration (≲8 days) flares and dips (≳0.5 mag). Near-IR spectroscopy of X-2 from Palomar/TripleSpec show He I, Paschen-γ, and Paschen-β emission lines with similar shapes and amplitudes as those of luminous blue variables (LBVs) and LBV candidates (LBVc). Mid-IR colors and magnitudes from Spitzer/Infrared Array Camera photometry of X-2 resemble those of known LBV/LBVcs. We suggest that the stellar companion in X-2 is an LBV/LBVc and discuss possible origins of the optical flares. Dips in the optical light curve are indicative of eclipses from optically thick clumps formed in the winds of the stellar counterpart. Given the constraints on the flare duration (0.02–0.8 days) and the time between flares (15.1 ± 7.8 days), we estimate the clump volume filling factor in the stellar winds, f V , to be 0.01< {f}V< 0.71, which overlaps with values measured from massive star winds. In X-2, we interpret the origin of the optical flares as the accretion of clumps formed in the winds of an LBV/LBVc onto the compact object.

Effects of wind velocity and slope on fire behavior

Treesearch

D.R. Weise; G.S. Biging

1994-01-01

Effects of wind velocity and slope on fire spread rate and flame length were examined. Fuel beds of vertical sticks (13.97 cm x 0.455 cm x 0.1 10 cm) and coarse excelsior were burned in an open-topped tilting wind tunnel. Mean fuel moisture content of sticks and excelsior was 11% and 12%, respectively. Mean surface area to volume ratio was 23 cm-! Five slopes (negative...

A multiple-fan active control wind tunnel for outdoor wind speed and direction simulation

NASA Astrophysics Data System (ADS)

Wang, Jia-Ying; Meng, Qing-Hao; Luo, Bing; Zeng, Ming

2018-03-01

This article presents a new type of active controlled multiple-fan wind tunnel. The wind tunnel consists of swivel plates and arrays of direct current fans, and the rotation speed of each fan and the shaft angle of each swivel plate can be controlled independently for simulating different kinds of outdoor wind fields. To measure the similarity between the simulated wind field and the outdoor wind field, wind speed and direction time series of two kinds of wind fields are recorded by nine two-dimensional ultrasonic anemometers, and then statistical properties of the wind signals in different time scales are analyzed based on the empirical mode decomposition. In addition, the complexity of wind speed and direction time series is also investigated using multiscale entropy and multivariate multiscale entropy. Results suggest that the simulated wind field in the multiple-fan wind tunnel has a high degree of similarity with the outdoor wind field.

A continuously weighing, high frequency sand trap: Wind tunnel and field evaluations

NASA Astrophysics Data System (ADS)

Yang, Fan; Yang, XingHua; Huo, Wen; Ali, Mamtimin; Zheng, XinQian; Zhou, ChengLong; He, Qing

2017-09-01

A new continuously weighing, high frequency sand trap (CWHF) has been designed. Its sampling efficiency is evaluated in a wind tunnel and the potential of the new trap has been demonstrated in field trials. The newly designed sand trap allows fully automated and high frequency measurement of sediment fluxes over extensive periods. We show that it can capture the variations and structures of wind-driven sand transport processes and horizontal sediment flux, and reveal the relationships between sand transport and meteorological parameters. Its maximum sampling frequency can reach 10 Hz. Wind tunnel tests indicated that the sampling efficiency of the CWHF sand trap varies between 39.2 to 64.3%, with an average of 52.5%. It achieved a maximum sampling efficiency of 64.3% at a wind speed of 10 m s- 1. This is largely achieved by the inclusion of a vent hole which leads to a higher sampling efficiency than that of a step-like sand trap at high wind speeds. In field experiments, we show a good agreement between the mass of sediment from the CWHF sand trap, the wind speed at 2 m and the number of saltating particles at 5 cm above the ground surface. According to analysis of the horizontal sediment flux at four heights from the CWHF sand trap (25, 35, 50, and 100 cm), the vertical distribution of the horizontal sediment flux up to a height of 100 cm above the sand surface follows an exponential function. Our field experiments show that the new instrument can capture more detailed information on sediment transport with much reduced labor requirement. Therefore, it has great potential for application in wind-blown sand monitoring and process studies.

A Correlation Between Flight-Determined Derivatives and Wind-Tunnel Data for the X-24B Research Aircraft

NASA Technical Reports Server (NTRS)

Sim, Alex G.

1976-01-01

Longitudinal and lateral-directional estimates of the aerodynamic derivatives of the X-24B research aircraft were obtained from flight data by using a modified maximum likelihooa estimation method. Data were obtained over a Mach number range from 0.35 to 1.72 and over an angle of attack range from 3.5deg to 15.7deg. Data are presented for a subsonic and a transonic configuration. The flight derivatives were generally consistent and documented the aircraft well. The correlation between the flight data and wind-tunnel predictions is presented and discussed.

A Correlation Between Flight-Determined Derivatives and Wind-Tunnel Data for the X-24B Research Aircraft

NASA Technical Reports Server (NTRS)

Sim, Alex G.

1997-01-01

Longitudinal and lateral-directional estimates of the aerodynamic derivatives of the X-24B research aircraft were obtained from flight data by using a modified maximum likelihood estimation method. Data were obtained over a Mach number range from 0.35 to 1.72 and over an angle of attack range from 3.5 deg. to 15.7 deg. Data are presented for a subsonic and transonic configuration. The flight derivatives were generally consistent and documented the aircraft well. The correlation between the flight data and wind-tunnel predictions is presented and discussed.

Comparison of conventional and adaptive wall wind tunnel results with regard to Reynolds number effects

NASA Technical Reports Server (NTRS)

Stanewsky, E.; Freimuth, P.

1989-01-01

A comparison of results from conventional and adaptive wall wind tunnels with regard to Reynolds number effects was carried out. The special objective of this comparison was to confirm or reject earlier conclusions, soley based on conventional wind tunnel results, concerning the influence of viscous effects on the characteristics of partially open wind tunnel walls, hence wall interference. The following postulations could be confirmed: (1) certain classes of supercritical airfoils exhibit a non-linear increase in lift which is, at least in part, related to viscous-inviscid interactions on the airfoil. This non-linear lift characteristic can erroneously be suppressed by sidewall interference effects in addition to being affected by changes in Reynolds number. Adaptive walls seem to relieve the influence of sidewall interference; (2) the degree of (horizontal) wall interference effects can be significantly affected by changes in Reynolds number, thus appearing as true Reynolds number effects; (3) perforated wall characteristics seem much more susceptible to viscous changes than the characteristics of slotted walls; here, blockage interference may be most severely influenced by viscous changes; and (4) real Reynolds number effects are present on the CAST 10-2/DOA 2 airfoil; they were shown to be appreciable also by the adaptive wall wind tunnel tests.

8. VIEW SOUTHWEST, INTERIOR VIEW, WIND TUNNEL 139 Naval ...

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

8. VIEW SOUTHWEST, INTERIOR VIEW, WIND TUNNEL 139 - Naval Surface Warfare Center, Subsonic Wind Tunnel Building, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD

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... extracts an ambient aerosol at elevated wind speeds. This wind tunnel test uses a single-sized, liquid... this subpart (under the heading of “wind tunnel inlet aspiration test”). The candidate sampler must...

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... extracts an ambient aerosol at elevated wind speeds. This wind tunnel test uses a single-sized, liquid... this subpart (under the heading of “wind tunnel inlet aspiration test”). The candidate sampler must...

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... extracts an ambient aerosol at elevated wind speeds. This wind tunnel test uses a single-sized, liquid... this subpart (under the heading of “wind tunnel inlet aspiration test”). The candidate sampler must...

Wind tunnel technology for the development of future commercial aircraft

NASA Technical Reports Server (NTRS)

Szodruch, J.

1986-01-01

Requirements for new technologies in the area of civil aircraft design are mainly related to the high cost involved in the purchase of modern, fuel saving aircraft. A second important factor is the long term rise in the price of fuel. The demonstration of the benefits of new technologies, as far as these are related to aerodynamics, will,for the foreseeable future, still be based on wind tunnel measurements. Theoretical computation methods are very successfully used in design work, wing optimization, and an estimation of the Reynolds number effect. However, wind tunnel tests are still needed to verify the feasibility of the considered concepts. Along with other costs, the cost for the wind tunnel tests needed for the development of an aircraft is steadily increasing. The present investigation is concerned with the effect of numerical aerodynamics and civil aircraft technology on the development of wind tunnels. Attention is given to the requirements for the wind tunnel, investigative methods, measurement technology, models, and the relation between wind tunnel experiments and theoretical methods.

Wall Interference in Two-Dimensional Wind Tunnels

NASA Technical Reports Server (NTRS)

Kemp, William B., Jr.

1986-01-01

Viscosity and tunnel-wall constraints introduced via boundary conditions. TWINTN4 computer program developed to implement method of posttest assessment of wall interference in two-dimensional wind tunnels. Offers two methods for combining sidewall boundary-layer effects with upper and lower wall interference. In sequential procedure, Sewall method used to define flow free of sidewall effects, then assessed for upper and lower wall effects. In unified procedure, wind-tunnel flow equations altered to incorporate effects from all four walls at once. Program written in FORTRAN IV for batch execution.

Aeroacoustic Characterization of the NASA Ames Experimental Aero-Physics Branch 32- by 48-Inch Subsonic Wind Tunnel with a 24-Element Phased Microphone Array

NASA Technical Reports Server (NTRS)

Costanza, Bryan T.; Horne, William C.; Schery, S. D.; Babb, Alex T.

2011-01-01

The Aero-Physics Branch at NASA Ames Research Center utilizes a 32- by 48-inch subsonic wind tunnel for aerodynamics research. The feasibility of acquiring acoustic measurements with a phased microphone array was recently explored. Acoustic characterization of the wind tunnel was carried out with a floor-mounted 24-element array and two ceiling-mounted speakers. The minimum speaker level for accurate level measurement was evaluated for various tunnel speeds up to a Mach number of 0.15 and streamwise speaker locations. A variety of post-processing procedures, including conventional beamforming and deconvolutional processing such as TIDY, were used. The speaker measurements, with and without flow, were used to compare actual versus simulated in-flow speaker calibrations. Data for wind-off speaker sound and wind-on tunnel background noise were found valuable for predicting sound levels for which the speakers were detectable when the wind was on. Speaker sources were detectable 2 - 10 dB below the peak background noise level with conventional data processing. The effectiveness of background noise cross-spectral matrix subtraction was assessed and found to improve the detectability of test sound sources by approximately 10 dB over a wide frequency range.

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

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

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

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

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

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

McDonnell Model XV-1 Convertiplane in the Ames 40x80 Foot Wind Tunnel.

NASA Image and Video Library

1954-05-17

Foreword, front view of McDonnell Model XV-1 Convertiplane in the Ames 40x80 Foot Wind Tunnel. The McDonnell XV-1 was an experimental compound gyroplane developed for a joint research program between the United States Air Force and the United States Army to explore technologies to develop an aircraft that could take off and land like a helicopter but fly at faster airspeeds, similar to a conventional airplane. The XV-1 would reach a speed of 200 mph (322 km/h), faster than any previous rotorcraft, but the program was terminated due to the tip-jet noise and complexity of the technology which gave only a modest gain in performance.

Plasma Wind Tunnel Testing of Electron Transpiration Cooling Concept

DTIC Science & Technology

2017-02-28

AFRL-AFOSR-UK-TR-2017-0012 Plasma Wind Tunnel Testing of Electron Transpiration Cooling Concept Olivier Chazot INSTITUT VON KARMAN DE DYNAMIQUE DES...28-02-2017 2. REPORT TYPE Final 3. DATES COVERED (From - To) 01 Dec 2015 to 30 Nov 2016 4. TITLE AND SUBTITLE Plasma Wind Tunnel Testing of Electron ...Aeronautics and Aerospace Department B-1640 Rhode Saint Genèse Belgium Internal Ref: ARR 1605 February 2017 Plasma Wind Tunnel Testing of Electron

Wind Tunnel Management and Resource Optimization: A Systems Modeling Approach

NASA Technical Reports Server (NTRS)

Jacobs, Derya, A.; Aasen, Curtis A.

2000-01-01

Time, money, and, personnel are becoming increasingly scarce resources within government agencies due to a reduction in funding and the desire to demonstrate responsible economic efficiency. The ability of an organization to plan and schedule resources effectively can provide the necessary leverage to improve productivity, provide continuous support to all projects, and insure flexibility in a rapidly changing environment. Without adequate internal controls the organization is forced to rely on external support, waste precious resources, and risk an inefficient response to change. Management systems must be developed and applied that strive to maximize the utility of existing resources in order to achieve the goal of "faster, cheaper, better". An area of concern within NASA Langley Research Center was the scheduling, planning, and resource management of the Wind Tunnel Enterprise operations. Nine wind tunnels make up the Enterprise. Prior to this research, these wind tunnel groups did not employ a rigorous or standardized management planning system. In addition, each wind tunnel unit operated from a position of autonomy, with little coordination of clients, resources, or project control. For operating and planning purposes, each wind tunnel operating unit must balance inputs from a variety of sources. Although each unit is managed by individual Facility Operations groups, other stakeholders influence wind tunnel operations. These groups include, for example, the various researchers and clients who use the facility, the Facility System Engineering Division (FSED) tasked with wind tunnel repair and upgrade, the Langley Research Center (LaRC) Fabrication (FAB) group which fabricates repair parts and provides test model upkeep, the NASA and LARC Strategic Plans, and unscheduled use of the facilities by important clients. Expanding these influences horizontally through nine wind tunnel operations and vertically along the NASA management structure greatly increases the

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.

Turning Vanes inside the Altitude Wind Tunnel

NASA Image and Video Library

1944-02-21

Men stand in front of turning vanes inside the Altitude Wind Tunnel (AWT) at the National Advisory Committee for Aeronautics (NACA) Aircraft Engine Research Laboratory. The AWT was the only wind tunnel capable of testing full-size aircraft engines in simulated altitude conditions. A large wooden drive fan, located on the other side of these vanes, created wind speeds up to 500 miles per hour. The drive shaft connected the fan to the induction motor located in an adjacent building. Turning vanes were located in each corner of the rectangular tunnel to straighten the airflow and direct it around the corners. This set of vanes was located in the 31-foot-diameter southeast corner of the tunnel. These elliptical panels consisted of 36 to 42 vertical vanes that were supported by three horizontal supports. The individual vanes were 2.5 feet long and half-moon shaped. The panel of vanes was affixed to the curved corner rings of the tunnel. Each set of turning vanes had a moveable vane in the middle of the lower level for personnel access. Each set of vanes took weeks to assemble before they were installed during the summer of 1943. This publicity photograph was taken just weeks after the tunnel became operational in February 1944.

Wind tunnel test of musi VI bridge

NASA Astrophysics Data System (ADS)

Permata, Robby; Andika, Matza Gusto; Syariefatunnisa, Risdhiawan, Eri; Hermawan, Budi; Noordiana, Indra

2017-11-01

Musi VI Bridge is planned to cross the Musi River in Palembang City, South Sumatera Province, Indonesia. The main span is a steel arch type with 200 m length and side span length is 75 m. Finite element analysis results showed that the bridge has frequency ratio for torsional and heaving mode (torsional frequency/heaving frequency)=1.14. This close to unity value rises concern about aerodynamic behaviour and stability of the bridge deck under wind loading. Sectional static and free vibration wind tunnel test were performed to clarify this phenomena in B2TA3 facility in Serpong, Indonesia. The test followed the draft of Guide of Wind Tunnel Test for Bridges developed by Indonesian Ministry of Public Works. Results from wind tunnel testing show that the bridge is safe from flutter instability and no coupled motion vibration observed. Therefore, low value of frequency ratio has no effect to aerodynamic behaviour of the bridge deck. Vortex-induced vibration in heaving mode occurred in relatively low wind velocity with permissible maximum amplitude value.

Construction of the 8- by 6-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1948-06-21

The 8- by 6-Foot Supersonic Wind Tunnel at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory was the nation’s largest supersonic facility when it began operation in April 1949. The emergence of new propulsion technologies such as turbojets, ramjets, and rockets during World War II forced the NACA and the aircraft industry to develop new research tools. In late 1945 the NACA began design work for new large supersonic wind tunnels at its three laboratories. The result was the 4- by 4-Foot Supersonic Wind Tunnel at Langley Memorial Aeronautical Laboratory, 6- by 6-foot supersonic wind tunnel at Ames Aeronautical Laboratory, and the largest facility, the 8- by 6-Foot Supersonic Wind Tunnel in Cleveland. The two former tunnels were to study aerodynamics, while the 8- by 6 facility was designed for supersonic propulsion. The 8- by 6-Foot Supersonic Wind Tunnel was used to study propulsion systems, including inlets and exit nozzles, combustion fuel injectors, flame holders, exit nozzles, and controls on ramjet and turbojet engines. Flexible sidewalls alter the tunnel’s nozzle shape to vary the Mach number during operation. A seven-stage axial compressor, driven by three electric motors that yield a total of 87,000 horsepower, generates air speeds from Mach 0.36 to 2.0. A section of the tunnel is seen being erected in this photograph.

A new low-turbulence wind tunnel for animal and small vehicle flight experiments

PubMed Central

Watts, Anthony; Nagle, Tony; Lentink, David

2017-01-01

Our understanding of animal flight benefits greatly from specialized wind tunnels designed for flying animals. Existing facilities can simulate laminar flow during straight, ascending and descending flight, as well as at different altitudes. However, the atmosphere in which animals fly is even more complex. Flow can be laminar and quiet at high altitudes but highly turbulent near the ground, and gusts can rapidly change wind speed. To study flight in both laminar and turbulent environments, a multi-purpose wind tunnel for studying animal and small vehicle flight was built at Stanford University. The tunnel is closed-circuit and can produce airspeeds up to 50 m s−1 in a rectangular test section that is 1.0 m wide, 0.82 m tall and 1.73 m long. Seamless honeycomb and screens in the airline together with a carefully designed contraction reduce centreline turbulence intensities to less than or equal to 0.030% at all operating speeds. A large diameter fan and specialized acoustic treatment allow the tunnel to operate at low noise levels of 76.4 dB at 20 m s−1. To simulate high turbulence, an active turbulence grid can increase turbulence intensities up to 45%. Finally, an open jet configuration enables stereo high-speed fluoroscopy for studying musculoskeletal control in turbulent flow. PMID:28405384

Characterization of a hypersonic quiet wind tunnel nozzle

NASA Astrophysics Data System (ADS)

Sweeney, Cameron J.

The Boeing/AFOSR Mach-6 Quiet Tunnel at Purdue University has been able to achieve low-disturbance flows at high Reynolds numbers for approximately ten years. The flow in the nozzle was last characterized in 2010. However, researchers have noted that the performance of the nozzle has changed in the intervening years. Understand ing the tunnel characteristics is critical for the hypersonic boundary-layer transition research performed at the facility and any change in performance could have signif icant effects on research performed at the facility. Pitot probe measurements were made using Kulite and PCB pressure transducers to quantify the performance changes since characterization was last performed. Aspects of the nozzle that were investi gated include the radial uniformity of the flow, the effects that time and stagnation pressure have on the flow, and the Reynolds number limits of low-disturbance flows. Measurements showed that freestream noise levels are consistently around 0.01% to 0.02% for the majority of the quiet flow core, with quiet flow now achievable for Reynolds numbers up to Re = 13.0x10 6/m. Additionally, while pitot probes are a widely used measurement technique for quantifying freestream disturbances, pitot probes are not without drawbacks. In order to provide a more complete methodology for freestream noise measurement other researchers have started experimenting with alternate geometries, such as cones. Using a newly designed 30° half-angle cone model, measurements were performed to quantify the freestream noise in the BAM6QT and compare the performance with another hypersonic wind tunnel. Also, measurements were made with three newly designed pitot sleeves to study the effects of probe geometry on freestream noise measurements. The results were compared to recent DNS calculations.

A wind tunnel study of aeolian sediment transport response to unsteady winds

NASA Astrophysics Data System (ADS)

Li, Bailiang; McKenna Neuman, Cheryl

2014-06-01

Although moderate attention has been paid to the response of the aeolian mass transport rate to wind gusts, it is still unclear how the particle size and volumetric concentration affect this relation. Very little is known about the response time of the particle speed, and specifically, how the sensor scale and elevation affect measurements of this variable. The present study addresses this knowledge gap through a series of wind tunnel experiments in which a gusty wind was generated by programming the fan motor to adjust to a randomly selected rpm every 10 s. Beds consisting of either medium or coarse sand were investigated through synchronous, co-located measurements of the local wind speed and particle speed/count rate obtained via a customized laser Doppler anemometry (LDA) system. The vertically integrated sand transport rate (Q) and the wind speed in the freestream were quantified using a passive sand trap and pitot tube, respectively. The results of the experiments indicate that the response of the aeolian transport system to wind gusts is generally faster in terms of the particle speed than the mass transport rate, while the degree of correlation is found to vary with the sensor elevation, as well as with the particle size and volumetric concentration. In essence, the coupling within the transport system is demonstrated to be strongly scale dependent.

Programmable controller system for wind tunnel diversion vanes

NASA Technical Reports Server (NTRS)

King, R. F.

1982-01-01

A programmable controller (PC) system automatic sequence control, which acts as a supervisory controller for the servos, selects the proper drives, and automatically sequences the vanes, was developed for use in a subsonic wind tunnel. Tunnel modifications include a new second test section (80 ft x 100 ft with a maximum air speed capability of 110 knots) and an increase in maximum velocity flow from 200 knots to 300 knots. A completely automatic sequence control is necessary in order to allow intricate motion of the 14 triangularly arranged vanes which can be as large as 70 ft high x 35 ft wide and which require precise acceleration and deceleration control. Rate servos on each drive aid in this control, and servo cost was minimized by using four silicon controlled rectifier controllers to control the 20 dc drives. The PC has a programming capacity which facilitated the implementation of extensive logic design. A series of diagrams sequencing the vanes and a block diagram of the system are included.

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.

Experimental investigation of a 10-percent-thick helicopter rotor airfoil section designed with a viscous transonic analysis code

NASA Technical Reports Server (NTRS)

Noonan, K. W.

1981-01-01

An investigation was conducted in the Langley 6- by 28-Inch Transonic Tunnel to determine the two dimensional aerodynamic characteristics of a 10-percent-thick helicopter rotor airfoil at Mach numbers from 0.33 to 0.87 and respective Reynolds numbers from 4.9 x 10 to the 6th to 9.8 x 10 to the 6th. This airfoil, designated the RC-10(N)-1, was also investigated at Reynolds numbers from 3.0 x 10 to the 6th to 7.3 x 10 to the 6th at respective Mach numbers of 0.33 to 0.83 for comparison wit the SC 1095 (with tab) airfoil. The RC-10(N)-1 airfoil was designed by the use of a viscous transonic analysis code. The results of the investigation indicate that the RC-10(N)-1 airfoil met all the design goals. At a Reynolds number of about 9.4 x 10 to the 6th the drag divergence Mach number at zero normal-force coefficient was 0.815 with a corresponding pitching-moment coefficient of zero. The drag divergence Mach number at a normal-force coefficient of 0.9 and a Reynolds number of about 8.0 x 10 to the 6th was 0.61. The drag divergence Mach number of this new airfoil was higher than that of the SC 1095 airfoil at normal-force coefficients above 0.3. Measurements in the same wind tunnel at comparable Reynolds numbers indicated that the maximum normal-force coefficient of the RC-10(N)-1 airfoil was higher than that of the NACA 0012 airfoil for Mach numbers above about 0.35 and was about the same as that of the SC 1095 airfoil for Mach numbers up to 0.5.

NACA Transonic Wind-tunnel Test Sections

NASA Technical Reports Server (NTRS)

Wright, Ray H; Ward, Vernon G

1955-01-01

Report presents an approximate subsonic theory for the solid-blockage interference in circular wind tunnels with walls slotted in the direction of flow. This theory indicated the possibility of obtaining zero blockage interference. Tests in a circular slotted tunnel based on the theory confirmed the theoretical predictions.

Temperature dependence of a superconducting tunnel junction x-ray detector

NASA Astrophysics Data System (ADS)

Hiller, Lawrence J.; Labov, Simon E.; Mears, Carl A.; Barfknecht, Andrew T.; Frank, Matthias A.; Netel, Harrie; Lindeman, Mark A.

1995-09-01

Superconducting tunnel junctions can be used as part of a high-resolution, energy-dispersive x- ray detector. The energy of the absorbed x ray is used to break superconducting electron pairs, producing on the order of 10(superscript 6) excitations, called quasiparticles. The number of quasiparticles produced is proportional to the energy of the absorbed x ray. When a bias voltage is maintained across the barrier, these quasiparticles produce a net tunneling current. Either the peak tunneling current or the total tunneled charge may be measured to determine the energy of the absorbed x ray. The tunneling rate, and therefore the signal, is enhanced by the use of a quasiparticle trap near the tunnel barrier. The trapping efficiency is improved by decreasing the energy gap, though this reduces the maximum temperature at which the device may operate. In our niobium/aluminum configuration, we can very the energy gap in the trapping layer by varying its thickness. This paper examines the performance of two devices with 50 nm aluminum traps at temperatures ranging from 100 mK to 700 mK. We found that this device has a very good energy resolution of about 12 eV FWHM at 1 keV. This energy resolution is independent of temperature for much of this temperature range.

Study of optical techniques for the Ames unitary wind tunnel, part 7

NASA Technical Reports Server (NTRS)

Lee, George

1993-01-01

A summary of optical techniques for the Ames Unitary Plan wind tunnels are discussed. Six optical techniques were studied: Schlieren, light sheet and laser vapor screen, angle of attack, model deformation, infrared imagery, and digital image processing. The study includes surveys and reviews of wind tunnel optical techniques, some conceptual designs, and recommendations for use of optical methods in the Ames Unitary Plan wind tunnels. Particular emphasis was placed on searching for systems developed for wind tunnel use and on commercial systems which could be readily adapted for wind tunnels. This final report is to summarize the major results and recommendations.

NASA Ames Laminar Flow Supersonic Wind Tunnel (LFSWT) Tests of a 10 deg Cone at Mach 1.6

NASA Technical Reports Server (NTRS)

Wolf, Stephen W. D.; Laub, James A.

1997-01-01

This work is part of the ongoing qualification of the NASA Ames Laminar Flow Supersonic Wind Tunnel (LFSWT) as a low-disturbance (quiet) facility suitable for transition research. A 10 deg cone was tested over a range of unit Reynolds numbers (Re = 2.8 to 3.8 million per foot (9.2 to 12.5 million per meter)) and angles of incidence (O deg to 10 deg) at Mach 1.6. The location of boundary layer transition along the cone was measured primarily from surface temperature distributions, with oil flow interferometry and Schlieren flow visualization providing confirmation measurements. With the LFSWT in its normal quiet operating mode, no transition was detected on the cone in the test core, over the Reynolds number range tested at zero incidence and yaw. Increasing the pressure disturbance levels in the LFSWT test section by a factor of five caused transition onset on the cone within the test core, at zero incidence and yaw. When operating the LFSWT in its normal quiet mode, transition could only be detected in the test core when high angles of incidence (greater than 5 deg) for cones were set. Transition due to elevated pressure disturbances (Tollmien-Schlichting) and surface trips produced a skin temperature rise of order 4 F (2.2 C). Transition due to cross flows on the leeward side of the cone at incidence produced a smaller initial temperature rise of only order 2.5 F (1.4 C), which indicates a slower transition process. We can conclude that these cone tests add further proof that the LFSWT test core is normally low-disturbance (pressure fluctuations greater than 0.1%), as found by associated direct flow quality measurements discussed in this report. Furthermore, in a quiet test environment, the skin temperature rise is sensitive to the type of dominant instability causing transition. The testing of a cone in the LFSWT provides an excellent experiment for the development of advanced transition detection techniques.

Measurement Techniques for Flow Diagnostic in ITAM Impulse Wind Tunnels

DTIC Science & Technology

2010-04-01

time of wind - tunnel operation, so that oscillations caused by initial shock loads could decay and a comparatively long time period with constant flow...Flow Diagnostic in ITAM Impulse Wind Tunnels 7 - 4 RTO-EN-AVT-186 A strain-gauge pressure probe is an elastic element (membrane) in a sealed...Diagnostic in ITAM Impulse Wind Tunnels RTO-EN-AVT-186 7 - 5 probes are individually calibrated. Piezoelectric pressure gauges are based

10. ENTRY STAIRWELL TO CABLE TUNNEL, ABOUT THREE QUARTERS THE ...

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

10. ENTRY STAIRWELL TO CABLE TUNNEL, ABOUT THREE QUARTERS THE DISTANCE TO THE SLED LAUNCHING PAD FROM THE FIRING CONTROL BLOCKHOUSE 0545. Looking west northwest. - Edwards Air Force Base, South Base Sled Track, Firing Control Blockhouse, South of Sled Track at east end, Lancaster, Los Angeles County, CA

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.

Drag Corrections in High-Speed Wind Tunnels

NASA Technical Reports Server (NTRS)

Ludwieg, H.

1947-01-01

In the vicinity of a body in a wind tunnel the displacement effect of the wake, due to the finite dimensions of the stream, produces a pressure gradient which evokes a change of drag. In incompressible flow this change of drag is so small, in general, that one does not have to take it into account in wind-tunnel measurements; however, in compressible flow it beoomes considerably larger, so that a correction factor is necessary for measured values. Correction factors for a closed tunnel and an open jet with circular cross sections are calculated and compared with the drag - corrections already bown for high-speed tunnnels.

Development of Dynamic Flow Field Pressure Probes Suitable for Use in Large Scale Supersonic Wind Tunnels

NASA Technical Reports Server (NTRS)

Porro, A. Robert

2000-01-01

A series of dynamic flow field pressure probes were developed for use in large-scale supersonic wind tunnels at NASA Glenn Research Center. These flow field probes include pitot, static, and five-hole conical pressure probes that are capable of capturing fast acting flow field pressure transients that occur on a millisecond time scale. The pitot and static probes can be used to determine local Mach number time histories during a transient event. The five-hole conical pressure probes are used primarily to determine local flow angularity, but can also determine local Mach number. These probes were designed, developed, and tested at the NASA Glenn Research Center. They were also used in a NASA Glenn 10-by 10-Foot Supersonic Wind Tunnel (SWT) test program where they successfully acquired flow field pressure data in the vicinity of a propulsion system during an engine compressor staff and inlet unstart transient event. Details of the design, development, and subsequent use of these probes are discussed in this report.

An inventory of aeronautical ground research facilities. Volume 1: Wind tunnels

NASA Technical Reports Server (NTRS)

Pirrello, C. J.; Hardin, R. D.; Heckart, M. V.; Brown, K. R.

1971-01-01

A survey of wind tunnel research facilities in the United States is presented. The inventory includes all subsonic, transonic, and hypersonic wind tunnels operated by governmental and private organizations. Each wind tunnel is described with respect to size, mechanical operation, construction, testing capabilities, and operating costs. Facility performance data are presented in charts and tables.

Mitigation of wind tunnel wall interactions in subsonic cavity flows

DOE Office of Scientific and Technical Information (OSTI.GOV)

Wagner, Justin L.; Casper, Katya Marie; Beresh, Steven J.

In this study, the flow over an open aircraft bay is often represented in a wind tunnel with a cavity. In flight, this flow is unconfined, though in experiments, the cavity is surrounded by wind tunnel walls. If untreated, wind tunnel wall effects can lead to significant distortions of cavity acoustics in subsonic flows. To understand and mitigate these cavity–tunnel interactions, a parametric approach was taken for flow over an L/D = 7 cavity at Mach numbers 0.6–0.8. With solid tunnel walls, a dominant cavity tone was observed, likely due to an interaction with a tunnel duct mode. Furthermore, anmore » acoustic liner opposite the cavity decreased the amplitude of the dominant mode and its harmonics, a result observed by previous researchers. Acoustic dampeners were also placed in the tunnel sidewalls, which further decreased the dominant mode amplitudes and peak amplitudes associated with nonlinear interactions between cavity modes. This then indicates that cavity resonance can be altered by tunnel sidewalls and that spanwise coupling should be addressed when conducting subsonic cavity experiments. Though mechanisms for dominant modes and nonlinear interactions likely exist in unconfined cavity flows, these effects can be amplified by the wind tunnel walls.« less

Mitigation of wind tunnel wall interactions in subsonic cavity flows

DOE PAGES

Wagner, Justin L.; Casper, Katya Marie; Beresh, Steven J.; ...

2015-03-06

In this study, the flow over an open aircraft bay is often represented in a wind tunnel with a cavity. In flight, this flow is unconfined, though in experiments, the cavity is surrounded by wind tunnel walls. If untreated, wind tunnel wall effects can lead to significant distortions of cavity acoustics in subsonic flows. To understand and mitigate these cavity–tunnel interactions, a parametric approach was taken for flow over an L/D = 7 cavity at Mach numbers 0.6–0.8. With solid tunnel walls, a dominant cavity tone was observed, likely due to an interaction with a tunnel duct mode. Furthermore, anmore » acoustic liner opposite the cavity decreased the amplitude of the dominant mode and its harmonics, a result observed by previous researchers. Acoustic dampeners were also placed in the tunnel sidewalls, which further decreased the dominant mode amplitudes and peak amplitudes associated with nonlinear interactions between cavity modes. This then indicates that cavity resonance can be altered by tunnel sidewalls and that spanwise coupling should be addressed when conducting subsonic cavity experiments. Though mechanisms for dominant modes and nonlinear interactions likely exist in unconfined cavity flows, these effects can be amplified by the wind tunnel walls.« less

Tunnel magnetoresistance in ultrathin L10 MnGa/MgO perpendicular magnetic tunnel junctions

NASA Astrophysics Data System (ADS)

Suzuki, K. Z.; Miura, Y.; Ranjbar, R.; Sugihara, A.; Mizukami, S.

2018-06-01

L10 MnGa is one of the interesting magnetic alloys for spin-transfer-torque based applications because such alloys have high perpendicular magnetic anisotropy, small magnetization, and low Gilbert damping. Magnetic tunnel junctions (MTJs) with ultrathin MnGa electrodes have recently been demonstrated using the room temperature growth technique of MnGa on paramagnetic B2-ordered CoGa templates, which exhibited a small TMR ratio of  ∼3%. To obtain a higher TMR ratio, we systematically investigated the annealing dependence of the TMR ratio with MTJs with 1–5 nm thick MnGa electrodes in this study. The TMR ratios were 2%–3% without annealing, which were the same as those reported previously, and the TMR ratios reached their maximum values of 6%–8% at an annealing temperature of approximately 250 °C for the MTJs with 2–5 nm MnGa electrodes. The TMR ratio increased to approximately 25% at 10 K for those MTJs. These TMR ratios were slightly higher than those reported in MTJs with 30 nm-thick MnGa electrodes. The annealing temperature at which TMR showed the maximum value tended to decrease with decreasing MnGa thickness, and this low annealing endurance may be attributed to the atomic mixing between MnGa and barrier/buffer layers. The TMR ratio was discussed in terms of both coherent tunneling based on first principles calculations with different element terminations at the interface and incoherent tunneling.

ARC-2010-ACD10-0029-027

NASA Image and Video Library

2010-02-16

Lawrence Livermore National Laboratories media Day for their LLNL project aimed at aerodynamic truck and trailer devices. Tests are being preformed in the Ames Full-Scale Aerodynamic Complex 80x120 foot wind tunnel. Gabriel and Sharon Lozano.

The results of a wind tunnel investigation of a model rotor with a free tip

NASA Technical Reports Server (NTRS)

Stroub, Robert H.; Young, Larry A.

1985-01-01

The results of a wind-tunnel test of the free tip rotor are presented. The free tip extended over the outer 10% of the rotor blade and included a simple, passive controller mechanism. Wind-tunnel test hardware is described. The free-tip assembly, which includes the controller, functioned flawlessly throughout the test. The tip pitched freely and responded to airflow perturbation in a sharp, quick, and stable manner. Tip pitch-angle responses are presented for an advance ratio range of 0.1 to 0.397 and for a thrust coefficient range of 0.038 to 0.092. The free tip reduced power requirements, loads going into the control system, and some flatwise blade-bending moments. Chordwise loads were not reduced by the free tip.

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

Wind Tunnel Study of Turbulent Flow Structure in the Convective Boundary Layer Capped by a Temperature Inversion.

NASA Astrophysics Data System (ADS)

Fedorovich, Evgeni; Kaiser, Rolf; Rau, Matthias; Plate, Erich

1996-05-01

Experiments on simulating the atmospheric convective boundary layer (CBL), capped by a temperature inversion and affected by surface shear, were carried out in the thermally stratified wind tunnel of the Institute of Hydrology and Water Resources, University of Karlsruhe. The tunnel is of the closed-circuit type, with a test section 10 m long, 1.5 m wide, and 1.5 m high. The return section of the tunnel is subdivided into 10 layers, each driven by its own fan and heating system. By this means, velocity and temperature profiles can be preshaped at the inlet of the test section, which allows for the reproduction of developed CBL over comparatively short fetches. The bottom heating is controlled to produce the constant heat flux through the floor of the test section. The flow velocity components in the tunnel are measured with a laser Doppler system; for temperature measurements, the resistance-wire technique is employed.A quasi-stationary, horizontally evolving CBL was reproduced in the tunnel, with convective Richardson numbers RiT and RiN up to 10 and 20, respectively, and the shear/buoyancy dynamic ratio u(/w( in the range of 0.2-0.5. Within the employed modeling approach, means and other statistics of the flow were calculated by temporal averaging. Deardorff mixed-layer scaling was used as a framework for processing and interpreting the experimental results. The comparison of the wind tunnel data with results of atmospheric, water tank, and numerical studies of the CBL shows the crucial dependence of the turbulence statistics in the upper part of the layer on the parameters of entrainment, as well as the modification of the CBL turbulence regime by the surface shear.

A directional microphone array for acoustic studies of wind tunnel models

NASA Technical Reports Server (NTRS)

Soderman, P. T.; Noble, S. C.

1974-01-01

An end-fire microphone array that utilizes a digital time delay system has been designed and evaluated for measuring noise in wind tunnels. The directional response of both a four- and eight-element linear array of microphones has enabled substantial rejection of background noise and reverberations in the NASA Ames 40- by 80-foot wind tunnel. In addition, it is estimated that four- and eight-element arrays reject 6 and 9 dB, respectively, of microphone wind noise, as compared with a conventional omnidirectional microphone with nose cone. Array response to two types of jet engine models in the wind tunnel is presented. Comparisons of array response to loudspeakers in the wind tunnel and in free field are made.

Installation of the Douglas XSB2D-1 in the Test Section of the 40x80 Foot Wind Tunnel at Ames.

NASA Image and Video Library

1944-06-12

Test section of the Ames 40 x 80 foot wind tunnel with the overhead doors open. XSB2D-1 airplane being lowered onto the struts by the overhead crane. Mechanics and engineers on orchard ladders aligning the model with ball sockets on the struts. The Douglas BTD Destroyer was an American dive/ torpedo bomber developed for the United States Navy during World War II.

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

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

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

Aeolian transport of biota with dust: A wind tunnel experiment

NASA Astrophysics Data System (ADS)

Rivas, J. A., Jr.; Gill, T. E.; Van Pelt, R. S.; Walsh, E.

2015-12-01

Ephemeral wetlands are ideal sources for dust emission, as well as repositories for dormant stages of aquatic invertebrates. An important component of invertebrate dispersal and colonization to new areas is the ability to be entrained into the atmosphere. Aquatic invertebrate eggs fall within the size of dust and sand grains (30-600μm), are less dense and aerodynamically shaped. We have shown previously that aquatic invertebrates can be dispersed long distances in dust storms but the extent of transport of taxa based on diapausing egg size/morphology has not been investigated. Here, we control the wind erosion process in a wind tunnel to test entrainment of diapausing stages of brine shrimp, clam shrimp, tadpole shrimp, fairy shrimp, Daphnia, and the rotifers Brachionus plicatilis and B. calyciflorus into the air by saltation. Diapausing eggs were mixed with sterilized wind-erodible soil. The soil/egg mixture was moistened with distilled water and air dried to form a crust. Dust was generated in a wind tunnel by releasing sand grains that act as saltator material similar to wind-entrained natural sands. Maximum wind velocity was 10m/s and entrained particles were sampled through an isokinetic horizontal intake opening. Aeolian sediment was collected from three points in the system; transfer section for coarse sediment, the pan subtending a settling chamber for finer saltation-sized sediment, and two paper filters for suspension-sized sediment. Samples were then passed through 250 and 350 μm sieves to remove abrader sand and rehydrated with various sterile media depending on the type of organism. We retrieved viable brine, fairy, and tadpole shrimp, ostracods, Daphnia, and diapausing eggs of the rotifers after hydration. This experiment demonstrates that resting stages of many invertebrates can be wind-eroded due to size and egg morphology and remain viable under controlled conditions mimicking dust emission.

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.

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.

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.

Detailed Uncertainty Analysis for Ares I Ascent Aerodynamics Wind Tunnel Database

NASA Technical Reports Server (NTRS)

Hemsch, Michael J.; Hanke, Jeremy L.; Walker, Eric L.; Houlden, Heather P.

2008-01-01

A detailed uncertainty analysis for the Ares I ascent aero 6-DOF wind tunnel database is described. While the database itself is determined using only the test results for the latest configuration, the data used for the uncertainty analysis comes from four tests on two different configurations at the Boeing Polysonic Wind Tunnel in St. Louis and the Unitary Plan Wind Tunnel at NASA Langley Research Center. Four major error sources are considered: (1) systematic errors from the balance calibration curve fits and model + balance installation, (2) run-to-run repeatability, (3) boundary-layer transition fixing, and (4) tunnel-to-tunnel reproducibility.

Turbulence Intensity at Inlet of 80- by 120-Foot Wind Tunnel Caused by Upwind Blockage

NASA Technical Reports Server (NTRS)

Salazar, Denise; Yuricich, Jillian

2014-01-01

In order to estimate the magnitude of turbulence in the National Full-Scale Aerodynamics Complex (NFAC) 80- by 120-Foot Wind Tunnel (80 x 120) caused by buildings located upwind from the 80 x 120 inlet, a 150th-scale study was performed that utilized a nominal two-dimensional blockage placed ahead of the inlet. The distance of the blockage ahead of the inlet was varied. This report describes velocity measurements made in the plane of the 80 x 120 model inlet for the case of zero ambient (atmospheric) wind.

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.

6. VIEW OF FIVEFOOT WIND TUNNEL WITH AIR STRAIGHTENER AND ...

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

6. VIEW OF FIVE-FOOT WIND TUNNEL WITH AIR STRAIGHTENER AND OPERATOR STATION IN FOREGROUND (1991). - Wright-Patterson Air Force Base, Area B, Building No. 19, Five-Foot Wind Tunnel, Dayton, Montgomery County, OH

2. VIEW SOUTH OF WIND TUNNEL 138 AND COOLING SYSTEM ...

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

2. VIEW SOUTH OF WIND TUNNEL 138 AND COOLING SYSTEM 140, NORTH ELEVATION - Naval Surface Warfare Center, Subsonic Wind Tunnel Building, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD

The Variable Density Wind Tunnel of the National Advisory Committee for Aeronautics

NASA Technical Reports Server (NTRS)

Munk, Max M; Miller, Elton W

1926-01-01

This report contains an exact description of the new wind tunnel of the National Advisory Committee for Aeronautics. This is the first american type wind tunnel. It differs from ordinary wind tunnels by its being surrounded by a strong steel shell, 35 feet long and 15 feet in diameter. A compressor system is provided to fill this shell - and hence the entire wind tunnel - with air compressed to a density up to 25 times the ordinary atmospheric density. It is demonstrated in the report that the increase of the air density makes up for a corresponding decrease in the scale of the model. Hence such american type wind tunnel is free from scale effect. The report is illustrated by many drawings and photographs. All construction details are described, and many dimensions given. The method of conducting tests is also described and some preliminary results given in the report. So far, the tests have confirmed the chief feature of this wind tunnel - absence of scale effect.

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

NASA Technical Reports Server (NTRS)

Erickson, Gary E.; Inenaga, Andrew S.

1994-01-01

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

Results of the AFRSI rewaterproofing systems screening test in the NASA/Ames Research Center (ARC) 2 x 2-foot transonic wind tunnel

NASA Technical Reports Server (NTRS)

Marroquin, J.; Kingsland, R. B.

1985-01-01

An experimental investigation was conducted in the NASA/Ames Research Center 2x2-foot Transonic Wind Tunnel to evaluate two AFRSI rewaterproofing systems and to investigate films as a means of reducing blanket joint distortion. The wind tunnel wall slot configuration influenced on the flow field over the test panel was investigated; primarily using oil flow data, and resulted in a closed slot configuration to provide a satisfactory screening environment flow field for the test. Sixteen AFRSI test panels, configured to represent the test system or film, were subjected to this screening environment (a flow field of separated and reattached flow at a freestream Mach numnber of 0.65 and q = 650 or 900 psf). Each condition was held until damage to the test article was observed or 55 minutes if no damage was incurred. All objectives related to AFRSI rewaterproofing and to the use of films to stiffen the blanket fibers were achieved.

Aeroelastic instability stoppers for wind tunnel models

NASA Technical Reports Server (NTRS)

Doggett, R. V., Jr.; Ricketts, R. H. (Inventor)

1981-01-01

A mechanism for diverting the flow in a wind tunnel from the wing of a tested model is described. The wing is mounted on the wall of a tunnel. A diverter plate is pivotally mounted on the tunnel wall ahead of the model. An actuator fixed to the tunnel is pivotably connected to the diverter plate, by plunger. When the model is about to become unstable during the test the actuator moves the diverter plate from the tunnel wall to divert maintaining stable model conditions. The diverter plate is then retracted to enable normal flow.

0.4 Percent Scale Space Launch System Wind Tunnel Test

NASA Image and Video Library

2011-11-15

0.4 Percent Scale Space Launch System Wind Tunnel Test 0.4 Percent Scale SLS model installed in the NASA Langley Research Center Unitary Plan Wind Tunnel Test Section 1 for aerodynamic force and movement testing.

Structural integrity of wind tunnel wooden fan blades

NASA Technical Reports Server (NTRS)

Young, Clarence P., Jr.; Wingate, Robert T.; Rooker, James R.; Mort, Kenneth W.; Zager, Harold E.

1991-01-01

Information is presented which was compiled by the NASA Inter-Center Committee on Structural Integrity of Wooden Fan Blades and is intended for use as a guide in design, fabrication, evaluation, and assurance of fan systems using wooden blades. A risk assessment approach for existing NASA wind tunnels with wooden fan blades is provided. Also, state of the art information is provided for wooden fan blade design, drive system considerations, inspection and monitoring methods, and fan blade repair. Proposed research and development activities are discussed, and recommendations are provided which are aimed at future wooden fan blade design activities and safely maintaining existing NASA wind tunnel fan blades. Information is presented that will be of value to wooden fan blade designers, fabricators, inspectors, and wind tunnel operations personnel.

Calibration of a Direct Detection Doppler Wind Lidar System using a Wind Tunnel

NASA Astrophysics Data System (ADS)

Rees, David

2012-07-01

As a critical stage of a Project to develop an airborne Direct-Detection Doppler Wind Lidar System, it was possible to exploit a Wind Tunnel of the VZLU, Prague, Czech Republic for a comprehensive series of tests against calibrated Air Speed generated by the Wind Tunnel. The initial results from these test sequences will be presented. The rms wind speed errors were of order 0.25 m/sec - very satisfactory for this class of Doppler Wind Lidar measurements. The next stage of this Project will exploit a more highly-developed laser and detection system for measurements of wind shear, wake vortex and other potentially hazardous meteorological phenomena at Airports. Following the end of this Project, key parts of the instrumentation will be used for routine ground-based Doppler Wind Lidar measurements of the troposphere and stratosphere.

7. VIEW WEST OF SCALE ROOM IN FULLSCALE WIND TUNNEL; ...

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

7. VIEW WEST OF SCALE ROOM IN FULL-SCALE WIND TUNNEL; SCALES ARE USED TO MEASURE FORCES ACTING ON MODEL AIRCRAFT SUSPENDED ABOVE. - NASA Langley Research Center, Full-Scale Wind Tunnel, 224 Hunting Avenue, Hampton, Hampton, VA

Modeling cumulus clouds in a two-phase wind tunnel

NASA Astrophysics Data System (ADS)

Bordás, R.; Thévenin, D.

2009-04-01

Experiments in wind-tunnels concerning meteorological flows are not very frequent in the literature. However, they are indispensable for a well-controlled and accurate investigation of turbulence-droplet interactions at the micro-scale. Of course it is impossible to reproduce perfectly the turbulent properties of clouds in a comparatively small wind-tunnel. The enormous length scales that are predominant in nature (integral length scale of typically 100 meters) lead to very high Reynolds numbers, roughly 107 calculated with the cloud dimensions or 104 as Taylor Reynolds number Reλ. Nevertheless, it is not necessary to reproduce exactly the whole turbulence spectrum to investigate the issue of rain formation in cumulus clouds. Only those scales and turbulence properties should be reproduced in the wind tunnel, which are physically important for the droplet population. In this work the key properties of cumulus clouds will be identified and implemented in a two-phase wind tunnel, allowing reproducible and accurate measurements. These properties are in particular the droplet number density, the turbulent kinetic energy and its dissipation rate. It is demonstrated by means of non-intrusive optical measurement techniques that the flow velocity, droplet number density, and key turbulence properties have been matched and are in the right order of magnitude. In this manner wind-tunnel investigations become possible and deliver realistic information concerning the interaction between droplets and turbulence in cumulus clouds.

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, liquid...

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.

A Vision in Aeronautics: The K-12 Wind Tunnel Project

NASA Technical Reports Server (NTRS)

1997-01-01

A Vision in Aeronautics, a project within the NASA Lewis Research Center's Information Infrastructure Technologies and Applications (IITA) K-12 Program, employs small-scale, subsonic wind tunnels to inspire students to explore the world of aeronautics and computers. Recently, two educational K-12 wind tunnels were built in the Cleveland area. During the 1995-1996 school year, preliminary testing occurred in both tunnels.

Sound propagation from a simple source in a wind tunnel

NASA Technical Reports Server (NTRS)

Cole, J. E., III

1975-01-01

The nature of the acoustic field of a simple source in a wind tunnel under flow conditions was examined theoretically and experimentally. The motivation of the study was to establish aspects of the theoretical framework for interpreting acoustic data taken (in wind) tunnels using in wind microphones. Three distinct investigations were performed and are described in detail.

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

Pioneering Russian wind tunnels and first experimental investigations, 1871-1915

NASA Astrophysics Data System (ADS)

Gorbushin, A. R.

2017-11-01

A review of foreign and Russian sources is given mentioning the pioneering wind tunnels built in Russia at the turn of 19th and 20th centuries. The first wind tunnel in Russia was constructed by V.A. Pashkevich at the Mikhailovsky Artillery Academy in St. Petersburg in 1871. In total from 1871 through 1915, 18 wind tunnels were constructed in Russia: 11 in Moscow, 5 in St. Petersburg and 2 in Kaluga. An overview of the pioneering Russian wind tunnels built by V.A. Pashkevich, K.E. Tsiolkovsky, prof. N.E. Zhukovsky, D.P. Ryabushinsky and prof. K.P. Boklevsky is given. Schemes, photographs, formulas, description of the research and test results taken from the original papers published by the wind tunnel designers are given. Photographs from the N.E. Zhukovsky Scientific and Memorial Museum and the Archive of the Russian Academy of Sciences are used in the article. Methods of flow visualization and results of their application are presented. The Russian scientists and researchers' contribution to the development of techniques and methods of aerodynamic experiment is shown, including one of the most important aspects - the wall interference problem.

Improvement of Flow Quality in NAL Chofu Mach 10 Nozzle

NASA Technical Reports Server (NTRS)

Lacey, John; Inoue, Yasutoshi; Higashida, Akio; Inoue, Manabu; Ishizaka, Kouichi; Korte, John J.

2002-01-01

As a result of CFD analysis and remachining of the nozzle, the flow quality of the Mach 10 Hypersonic Wind Tunnel at NAL Chofu, Japan was improved. The subsequent test results validated the CFD analytical predictions by NASA and MHL.

Hot-bench simulation of the active flexible wing wind-tunnel model

NASA Technical Reports Server (NTRS)

Buttrill, Carey S.; Houck, Jacob A.

1990-01-01

Two simulations, one batch and one real-time, of an aeroelastically-scaled wind-tunnel model were developed. The wind-tunnel model was a full-span, free-to-roll model of an advanced fighter concept. The batch simulation was used to generate and verify the real-time simulation and to test candidate control laws prior to implementation. The real-time simulation supported hot-bench testing of a digital controller, which was developed to actively control the elastic deformation of the wind-tunnel model. Time scaling was required for hot-bench testing. The wind-tunnel model, the mathematical models for the simulations, the techniques employed to reduce the hot-bench time-scale factors, and the verification procedures are described.

Wind tunnel test of Teledyne Geotech model 1564B cup anemometer

DOE Office of Scientific and Technical Information (OSTI.GOV)

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 mphmore » 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.« less

Wind tunnel test of Teledyne Geotech model 1564B cup anemometer

NASA Astrophysics Data System (ADS)

Parker, M. J.; Addis, R. P.

1991-04-01

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 outweigh 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.