Altitude Wind Tunnel Control Room at the Aircraft Engine Research Laboratory

NASA Image and Video Library

1944-07-21

Operators in the control room for the Altitude Wind Tunnel at the National Advisory Committee for Aeronautics (NACA) Aircraft Engine Research Laboratory remotely operate a Wright R–3350 engine in the tunnel’s test section. Four of the engines were used to power the B–29 Superfortress, a critical weapon in the Pacific theater during World War II. The wind tunnel, which had been in operation for approximately six months, was the nation’s only wind tunnel capable of testing full-scale engines in simulated altitude conditions. The soundproof control room was used to operate the wind tunnel and control the engine being run in the test section. The operators worked with assistants in the adjacent Exhauster Building and Refrigeration Building to manage the large altitude simulation systems. The operator at the center console controlled the tunnel’s drive fan and operated the engine in the test section. Two sets of pneumatic levers near his right forearm controlled engine fuel flow, speed, and cooling. Panels on the opposite wall, out of view to the left, were used to manage the combustion air, refrigeration, and exhauster systems. The control panel also displayed the master air speed, altitude, and temperature gauges, as well as a plethora of pressure, temperature, and airflow readings from different locations on the engine. The operator to the right monitored the manometer tubes to determine the pressure levels. Despite just being a few feet away from the roaring engine, the control room remained quiet during the tests.

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.

Experimental Investigation of Mars Science Laboratory Entry Vehicle Aeroheating in AEDC Hypervelocity Tunnel 9

NASA Technical Reports Server (NTRS)

Hollis, Brian R.; Collier, Arnold S.

2017-01-01

An experimental investigation of the aeroheating environment of the Mars Science Laboratory entry vehicle was conducted in the Arnold Engineering Development Complex Hypervelocity Wind Tunnel 9. Testing was performed on a 6-in. (0.1524 m) diameter model in the tunnel's Mach 8 and Mach 10 nozzles at free stream Reynolds numbers from 4.1×10*exp 6)/ft to 49×10(exp 6)/ft and from 1.2×10(exp 6)/ft to 19×10(exp 6)/ft, respectively, using pure nitrogen test gas. These conditions spanned the boundary layer flow regimes from completely laminar to fully turbulent flow over the entire forebody. A computational fluid dynamics study was conducted in support of the wind tunnel testing. Laminar and turbulent solutions were generated for all wind tunnel test conditions and comparisons of predicted heating distributions were performed with the data. These comparisons showed agreement for most cases to within the estimated +/-12% experimental uncertainty margin for fully-laminar or fully-turbulent conditions, while transitional heating data were bounded by laminar and turbulent predictions. These results helped to define uncertainty margins on the use of computational tools for vehicle design.

Aerial View of NACA's Lewis Flight Propulsion Research Laboratory

NASA Image and Video Library

1946-05-21

The National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory in Cleveland, Ohio as seen from the west in May 1946. The Cleveland Municipal Airport is located directly behind. The laboratory was built in the early 1940s to resolve problems associated with aircraft engines. The initial campus contained seven principal buildings: the Engine Research Building, hangar, Fuels and Lubricants Building, Administration Building, Engine Propeller Research Building, Altitude Wind Tunnel, and Icing Research Tunnel. These facilities and their associated support structures were located within an area occupying approximately one-third of the NACA’s property. After World War II ended, the NACA began adding new facilities to address different problems associated with the newer, more powerful engines and high speed flight. Between 1946 and 1955, four new world-class test facilities were built: the 8- by 6-Foot Supersonic Wind Tunnel, the Propulsion Systems Laboratory, the Rocket Engine Test Facility, and the 10- by 10-Foot Supersonic Wind Tunnel. These large facilities occupied the remainder of the NACA’s semicircular property. The Lewis laboratory expanded again in the late 1950s and early 1960s as the space program commenced. Lewis purchased additional land in areas adjacent to the original laboratory and acquired a large 9000-acre site located 60 miles to the west in Sandusky, Ohio. The new site became known as Plum Brook Station.

Standardization Tests of NACA No. 1 Wind Tunnel

NASA Technical Reports Server (NTRS)

Reid, Elliott G

1925-01-01

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

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...

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

NASA Technical Reports Server (NTRS)

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

1995-01-01

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

Parachute Testing for Mars Science Laboratory

NASA Image and Video Library

2007-12-20

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

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.

Aeroservoelastic Testing of Free Flying Wind Tunnel Models Part 1: A Sidewall Supported Semispan Model Tested for Gust Load Alleviation and Flutter Suppression

NASA Technical Reports Server (NTRS)

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

2013-01-01

of a two part document. Part 2 is titled: "Aeroservoelastic Testing of Free Flying Wind Tunnel Models, Part 2: A Centerline Supported Fullspan Model Tested for Gust Load Alleviation." A team comprised of the Air Force Research Laboratory (AFRL), Northrop Grumman, Lockheed Martin, and the NASA Langley Research Center conducted three aeroservoelastic wind tunnel tests in the Transonic Dynamics Tunnel to demonstrate active control technologies relevant to large, flexible vehicles. In the first of these three tests, a semispan, aeroelastically scaled, wind tunnel model of a flying wing SensorCraft vehicle was mounted to a force balance to demonstrate gust load alleviation. In the second and third tests, the same wing was mated to a new, multi-degree of freedom, sidewall mount. This mount allowed the half-span model to translate vertically and pitch at the wing root, allowing better simulation of the full span vehicle's rigid body modes. Gust load alleviation (GLA) and Body freedom flutter (BFF) suppression were successfully demonstrated. The rigid body degrees-of-freedom required that the model be flown in the wind tunnel using an active control system. This risky mode of testing necessitated that a model arrestment system be integrated into the new mount. The safe and successful completion of these free flying tests required the development and integration of custom hardware and software. This paper describes the many systems, software, and procedures that were developed as part of this effort.

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.

Results of tests of advanced flexible insulation vortex and flow environments in the North American Aerodynamics Laboratory lowspeed wind tunnel using 0.0405-scale Space Shuttle Orbiter model 16-0 (test OA-309)

NASA Technical Reports Server (NTRS)

Marshall, B. A.; Nichols, M. E.

1984-01-01

An experimental investigation (Test OA-309) was conducted using 0.0405-scale Space Shuttle Orbiter Model 16-0 in the North American Aerodynamics Laboratory 7.75 x 11.00-foot Lowspeed Wind Tunnel. The primary purpose was to locate and study any flow conditions or vortices that might have caused damage to the Advanced Flexible Reusable Surface Insulation (AFRSI) during the Space Transportation System STS-6 mission. A secondary objective was to evaluate vortex generators to be used for Wind Tunnel Test OS-314. Flowfield visualization was obtained by means of smoke, tufts, and oil flow. The test was conducted at Mach numbers between 0.07 and 0.23 and at dynamic pressures between 7 and 35 pounds per square foot. The angle-of-attack range of the model was -5 degrees through 35 degrees at 0 or 2 degrees of sideslip, while roll angle was held constant at zero degrees. The vortex generators were studied at angles of 0, 5, 10, and 15 degrees.

Evaluation of Fatigue Crack Growth and Fracture Properties of Cryogenic Model Materials

NASA Technical Reports Server (NTRS)

Newman, John A.; Forth, Scott C.; Everett, Richard A., Jr.; Newman, James C., Jr.; Kimmel, William M.

2002-01-01

The criteria used to prevent failure of wind-tunnel models and support hardware were revised as part of a project to enhance the capabilities of cryogenic wind tunnel testing at NASA Langley Research Center. Specifically, damage-tolerance fatigue life prediction methods are now required for critical components, and material selection criteria are more general and based on laboratory test data. The suitability of two candidate model alloys (AerMet 100 and C-250 steel) was investigated by obtaining the fatigue crack growth and fracture data required for a damage-tolerance fatigue life analysis. Finally, an example is presented to illustrate the newly implemented damage tolerance analyses required of wind-tunnel model system components.

Large Swing Valve in the 10- by 10-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1956-05-21

A 24-foot diameter swing valve is seen in an open position inside the new 10- by 10-Foot Supersonic Wind Tunnel at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory. The 10- by 10 was the most powerful propulsion wind tunnel in the nation. After over three years of construction the tunnel was ready to conduct its first tests in early 1956. The 10- by 10-foot tunnel 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. This large swinging valve is critical to the operation of the facility. In one position the valve seals off the tunnel exhaust, making the tunnel a closed circuit, which is used for aerodynamic testing of models. In its other position, the valve acts as a seal across the tunnel and leaves the tunnel exhaust open. This arrangement is used when engines are fired. The air going through the tunnel is taken from the atmosphere and returned to the atmosphere after one pass through the tunnel. Engines up to five feet in diameter can be tested in the 10- by 10-foot test section. Air flows up to Mach 3.5 can be fed through the test section by a 250,000-horsepower axial-flow compressor fan. The incoming air must be dehumidified and cooled so that the proper conditions are present for the test. A large air dryer with 1,890 tons of activated alumina soaks up 1.5 tons of water per minute from the air flow. A cooling apparatus equivalent to 250,000 household air conditioners is used to cool the air.

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.

Noise Suppression Addition to the 8- by 6-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1950-08-21

The 8- by 6-Foot Supersonic Wind Tunnel at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory was the largest supersonic wind tunnel in the nation at the time and the only one able to test full-scale engines at supersonic speeds. The 8- by 6 was designed as a non-return and open-throat tunnel. A large compressor created the air flow at one end of the tunnel, squeezed the flow to increase its velocity just before the test section, then reduced the velocity, and expelled it into the atmosphere at the other end of the tunnel. This design worked well for initial aerodynamic testing, but the local community was literally rattled by the noise and vibrations when researchers began running engines in the test section in January 1950. The NACA’s most modern wind tunnel was referred to as “an 87,000-horsepower bugle aimed at the heart of Cleveland.” NACA Lewis responded to the complaints by adding an acoustic housing at the end of the tunnel to dampen the noise. The structure included resonator chambers and a reinforced concrete muffler structure. Modifications continued over the years. A return leg was added, and a second test section, 9 -by 15-foot, was incorporated in the return leg in the 1960s. Since its initial operation in 1948, the 8- by 6-foot tunnel has been aggressively used to support the nation's aeronautics and space programs for the military, industry, and academia.

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.

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...

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)

Turbulent Aeroheating Testing of Mars Science Laboratory Entry Vehicle in Perfect-Gas Nitrogen

NASA Technical Reports Server (NTRS)

Hollis, Brian R.; Collier, Arnold S.

2007-01-01

An experimental investigation of turbulent aeroheating on the Mars Science Laboratory entry vehicle heat shield has been conducted in the Arnold Engineering Development Center Hypervelocity Wind Tunnel No. 9. Testing was performed on a 6-in. (0.1524 m) diameter MSL model in pure N2 gas in the tunnel s Mach 8 and Mach 10 nozzles at free stream Reynolds numbers of 4.1x10(exp 6)/ft to 49x10(exp 6)/ft (1.3x10(exp 7)/m to 16x10(exp 7)/m) and 1.2x10(exp 6)/ft to 19x10(exp 6)/ft (0.39x10(exp 7)/m to 62x10(exp 7)/m), respectively. These conditions were sufficient to span the regime of boundary-layer flow from completely laminar to fully-developed turbulent flow over the entire forebody. A supporting aeroheating test was also conducted in the Langley Research Center 20-Inch Mach 6 Air Tunnel at free stream Reynolds number of 1x10(exp 6)/ft to 7x10(exp 6)/ft (0.36x10(exp 7)/m to 2.2x10(exp 7)/m) in order to help corroborate the Tunnel 9 results. A complementary computational fluid dynamics study was conducted in parallel to the wind tunnel testing. Laminar and turbulent predictions were generated for all wind tunnel test conditions and comparisons were performed with the data for the purpose of helping to define uncertainty margins on predictions for aeroheating environments during entry into the Martian atmosphere. Data from both wind tunnel tests and comparisons with the predictions are presented herein. It was concluded from these comparisons that for perfect-gas conditions, the computational tools could predict fully-laminar or fully-turbulent heating conditions to within 10% of the experimental data

Turbulent Aeroheating Testing of Mars Science Laboratory Entry Vehicle

NASA Technical Reports Server (NTRS)

Hollis, Brian R.; Collier, Arnold S.

2008-01-01

An experimental investigation of turbulent aeroheating on the Mars Science Laboratory entry vehicle heat shield has been conducted in the Arnold Engineering Development Center Hypervelocity Wind Tunnel No. 9. Testing was performed on a 6-in. (0.1524 m) diameter MSL model in pure N2 gas in the tunnel's Mach 8 and Mach 10 nozzles at free stream Reynolds numbers of 4.1 x 10(exp 6)/ft to 49 x 10(exp 6)/ft (1.3 x 10(exp 7)/m to 19 x 10(exp 6/ft) and 1.2 x 10(exp 6)/ft to 19 x 10(exp 6)/ft (0.39 x 10(exp 7)/m to 62 x 10(exp 7)/m), respectively. These conditions were sufficient to span the regime of boundary-layer flow from completely laminar to fully-developed turbulent flow over the entire forebody. A supporting aeroheating test was also conducted in the Langley Research Center 20-Inch Mach 6 Air Tunnel at free stream Reynolds number of 1 x 10(exp 6)/ft to 7 x 10(exp 6)/ft (0.36 x 10(exp 7)/m to 2.2 x 10(exp 7)/m) in order to help corroborate the Tunnel 9 results. A complementary computational fluid dynamics study was conducted in parallel to the wind tunnel testing. Laminar and turbulent predictions were generated for the wind tunnel test conditions and comparisons were performed with the data for the purpose of helping to define uncertainty margins on predictions for aeroheating environments during entry into the Martian atmosphere. Data from both wind tunnel tests and comparisons with the predictions are presented herein. It was concluded from these comparisons that for perfect-gas conditions, the computational tools could predict fully-laminar or fully-turbulent heating conditions to within 12% or better of the experimental data.

Exterior of Flexible Wall at the 10- by 10-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1955-03-21

A mechanic checks the tubing on one of the many jacks which control the nozzle section of the 10- by 10-Foot Supersonic Wind Tunnel at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory. The 10- by 10-foot tunnel, which had its official opening in May 1956, was built under the Congressional 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 10- by 10 wind tunnel can be operated as a closed circuit for aerodynamic tests or as an open circuit for propulsion investigations. The 10-foot tall and 76-foot long stainless steel nozzle section just upstream from the test section can be adjusted to change the speed and composition of the air flow. Hydraulic jacks, seen in this photograph, flex the 1.37-inch thick walls of the tunnel nozzle. The size of the nozzle’s opening controls the velocity of the air through the test section. Seven General Electric motors capable of generating 25,000 horsepower produce the Mach 2.5 and 2.5 airflows. The facility was mostly operated at night due to its large power load requirements.

Preheater in the 10-by 10-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1958-04-21

The 10- by 10-Foot Supersonic Wind Tunnel at the NACA Lewis Flight Propulsion Laboratory was built under the Congressional Unitary Plan Act which coordinated wind tunnel construction at the NACA, Air Force, industry, and universities. The 10- by 10, which began operation in 1956, was the largest of the three NACA tunnels built under the act. Researchers could test engines up to five feet in diameter in the 10- by 10-foot test section. A 250,000-horsepower axial-flow compressor fan can generate airflows up to Mach 3.5 through the test section. The incoming air must be dehumidified and cooled so that the proper conditions are present for the test. A large air dryer with 1,890 tons of activated alumina soaks up 1.5 tons of water per minute from the airflow. A cooling apparatus equivalent to 250,000 household air conditioners is used to cool the air. The air heater is located just upstream from the test section. Natural gas is combusted in the tunnel to increase the air temperature. The system could only be employed when the tunnel was run in its closed-circuit propulsion mode.

Dust emissions of organic soils observed in the field and laboratory

NASA Astrophysics Data System (ADS)

Zobeck, T. M.; Baddock, M. C.; Guo, Z.; Van Pelt, R.; Acosta-Martinez, V.; Tatarko, J.

2011-12-01

According to the U.S. Soil Taxonomy, Histosols (also known as organic soils) are soils that are dominated by organic matter (>20% organic matter) in half or more of the upper 80 cm. These soils, when intensively cropped, are subject to wind erosion resulting in loss in crop productivity and degradation of soil, air, and water quality. Estimating wind erosion on Histosols has been determined by USDA-Natural Resources Conservation Service as a critical need for the Wind Erosion Prediction System (WEPS) model. WEPS has been developed to simulate wind erosion on agricultural land in the US, including soils with organic soil material surfaces. However, additional field measurements are needed to calibrate and validate estimates of wind erosion of organic soils using WEPS. In this study, we used a field portable wind tunnel to generate suspended sediment (dust) from agricultural surfaces for soils with a range of organic contents. The soils were tilled and rolled to provide a consolidated, friable surface. Dust emissions and saltation were measured using an isokinetic vertical slot sampler aspirated by a regulated suction source. Suspended dust was collected on filters of the dust slot sampler and sampled at a frequency of once every six seconds in the suction duct using a GRIMM optical particle size analyzer. In addition, bulk samples of airborne dust were collected using a sampler specifically designed to collect larger dust samples. The larger dust samples were analyzed for physical, chemical, and microbiological properties. In addition, bulk samples of the soils were tested in a laboratory wind tunnel similar to the field wind tunnel and a laboratory dust generator to compare field and laboratory results. For the field wind tunnel study, there were no differences between the highest and lowest organic content soils in terms of their steady state emission rate under an added abrader flux, but the soil with the mid-range of organic matter had less emission by one third. In the laboratory wind tunnel, samples with the same ratio of erodible to non-erodible aggregates as the field soils were abraded and dust emissions were observed with the same sampling system as used in the field wind tunnel. In the dust generator, 5 gm samples < 8 mm diameter of each organic soil were rotated in a 50 cm long tube and the dust generated was observed with the GRIMM during a 20 minute run. Comparisons of the field dust emission rates with the laboratory results will be presented.

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

NASA Technical Reports Server (NTRS)

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

2008-01-01

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

Manometer Boards below the 8- by 6-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1951-02-21

Analysts at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory take data readings from rows of manometers in the basement of the 8- by 6-Foot Supersonic Wind Tunnel. Manometers were mercury-filled glass tubes that indicated different pressure levels in the test section. Manometers look and function very similarly to thermometers. Pressure sensing instruments were installed on the test article inside the wind tunnel or other test facility. Each test could have dozens of such instruments installed and connected to a remotely located manometer tube. The mercury inside the manometer rose and fell with the pressure levels. The dark mercury can be seen at different levels within the tubes. Since the pressure readings were dynamic, it was necessary to note the levels at given points during the test. This was done using both female computers and photography. A camera is seen on a stand to the right in this photograph.

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...

Miniature infrared data acquisition and telemetry system

NASA Technical Reports Server (NTRS)

Stokes, J. H.; Ward, S. M.

1985-01-01

The Miniature Infrared Data Acquisition and Telemetry (MIRDAT) Phase 1 study was performed to determine the technical and commercial feasibility of producing a miniaturized electro-optical telemetry system. This system acquires and transmits experimental data from aircraft scale models for realtime monitoring in wind tunnels. During the Phase 1 study, miniature prototype MIRDAT telemetry devices were constructed, successfully tested in the laboratory and delivered to the user for wind tunnel testing. A search was conducted for commercially available components and advanced hybrid techniques to further miniaturize the system during Phase 2 development. A design specification was generated from laboratory testing, user requirements and discussions with component manufacturers. Finally, a preliminary design of the proposed MIRDAT system was documented for Phase 2 development.

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.

Study and evaluation of ferro-cement for use in wind tunnel construction

NASA Technical Reports Server (NTRS)

Larsen, H. J., Jr. (Compiler)

1972-01-01

The structural suitability and cost effectiveness of ferro-cement for large subsonic wind tunnel structures is investigated. This investigation was carried out in the following four main categories: (1) a state-of-the-art survey into the uses, properties, and costs of ferro-cement; (2) an evaluation of those ferro-cement properties critical to construction of large, subsonic wind tunnels, which have not been adequately established to date; (3) a laboratory testing program to determine preliminary values for those properties; and (4) a study to establish cost factors for ferro-cement as related to a preliminary construction scheme for a nacelle and shroud unit.

Field and Laboratory Responses of Male Leaf Roller Moths, Choristoneura rosaceana and Pandemis pyrusana, to Pheromone Concentrations in an Attracticide Paste Formulation

PubMed Central

Curkovic, Tomislav; Brunner, Jay F.; Landolt, Peter J.

2009-01-01

Male leafroller moths, Choristoneura rosaceana (Harris) (Lepidoptera: Tortricidae) and Pandemis pyrusana (Kearfott), were evaluated for responses to a paste formulation loaded with a range of concentrations of the two species' pheromone blends and evaluated in a laboratory wind tunnel and in the field. Response criteria were flight, flight towards the pheromone source, and contact with the pheromone source for the wind tunnel assays, and capture of moths in traps for the field tests. In the wind tunnel and field, responses of males of both species to the paste generally increased as the pheromone concentration in the paste was increased. There was little response by either species to paste with less than 0.16% pheromone. The relationship between pheromone concentration and response for P. pyrusana was linear and for C. rosaceana was sinusoidal over the range of pheromone concentrations tested. These patterns were seen both in the wind tunnel and in the field. Initial release rates from the paste of (Z)-11-tetradecenyl acetate, the main component of the pheromone blends of both species was 3.6–3.8 ng/h. Inhibitory thresholds for responses were not reached for either species, using pheromone concentrations as high as 16%, in either the wind tunnel or the field. For both species, response of males to rubber septa with one mg pheromone loads was similar to the response to the paste with pheromone at concentrations greater than 3–4%. For C. rosaceana, rates of contact with the paste in the wind tunnel were statistically similar to rates of contact in response to conspecific females, with paste pheromone concentrations above 1.6%. Response rates for males of P. pyrusana were significantly lower to the paste than to conspecific females at all paste pheromone concentrations tested. Overall, the optimum pheromone concentration in the paste for moth attraction to contact was 3.2 % for C. rosaceana and 8% for P. pyrusana. PMID:19619030

Design and validation of a wind tunnel system for odour sampling on liquid area sources.

PubMed

Capelli, L; Sironi, S; Del Rosso, R; Céntola, P

2009-01-01

The aim of this study is to describe the methods adopted for the design and the experimental validation of a wind tunnel, a sampling system suitable for the collection of gaseous samples on passive area sources, which allows to simulate wind action on the surface to be monitored. The first step of the work was the study of the air velocity profiles. The second step of the work consisted in the validation of the sampling system. For this purpose, the odour concentration of some air samples collected by means of the wind tunnel was measured by dynamic olfactometry. The results of the air velocity measurements show that the wind tunnel design features enabled the achievement of a uniform and homogeneous air flow through the hood. Moreover, the laboratory tests showed a very good correspondence between the odour concentration values measured at the wind tunnel outlet and the odour concentration values predicted by the application of a specific volatilization model, based on the Prandtl boundary layer theory. The agreement between experimental and theoretical trends demonstrate that the studied wind tunnel represents a suitable sampling system for the simulation of specific odour emission rates from liquid area sources without outward flow.

Design and calibration of the mixing layer and wind tunnel

NASA Technical Reports Server (NTRS)

Bell, James H.; Mehta, Rabindra D.

1989-01-01

A detailed account of the design, assembly and calibration of a wind tunnel specifically designed for free-shear layer research is contained. The construction of this new facility was motivated by a strong interest in the study of plane mixing layers with varying initial and operating conditions. The Mixing Layer Wind tunnel is located in the Fluid Mechanics Laboratory at NASA Ames Research Center. The tunnel consists of two separate legs which are driven independently by centrifugal blowers connected to variable speed motors. The blower/motor combinations are sized such that one is smaller than the other, giving maximum flow speeds of about 20 and 40 m/s, respectively. The blower speeds can either be set manually or via the Microvax II computer. The two streams are allowed to merge in the test section at the sharp trailing edge of a slowly tapering splitter plate. The test section is 36 cm in the cross-stream direction, 91 cm in the spanwise direction and 366 cm in length. One test section side-wall is slotted for probe access and adjustable so that the streamwise pressure gradient may be controlled. The wind tunnel is also equipped with a computer controlled, three-dimensional traversing system which is used to investigate the flow fields with pressure and hot-wire instrumentation. The wind tunnel calibration results show that the mean flow in the test section is uniform to within plus or minus 0.25 pct and the flow angularity is less than 0.25 deg. The total streamwise free-stream turbulence intensity level is approximately 0.15 pct. Currently the wind tunnel is being used in experiments designed to study the three-dimensional structure of plane mixing layers and wakes.

Development, Analysis and Testing of the High Speed Research Flexible Semispan Model

NASA Technical Reports Server (NTRS)

Schuster, David M.; Spain, Charles V.; Turnock, David L.; Rausch, Russ D.; Hamouda, M-Nabil; Vogler, William A.; Stockwell, Alan E.

1999-01-01

This report presents the work performed by Lockheed Martin Engineering and Sciences (LMES) in support of the High Speed Research (HSR) Flexible Semispan Model (FSM) wind-tunnel test. The test was conducted in order to assess the aerodynamic and aeroelastic character of a flexible high speed civil transport wing. Data was acquired for the purpose of code validation and trend evaluation for this type of wing. The report describes a number of activities in preparing for and conducting the wind-tunnel test. These included coordination of the design and fabrication, development of analytical models, analysis/hardware correlation, performance of laboratory tests, monitoring of model safety issues, and wind-tunnel data acquisition and reduction. Descriptions and relevant evaluations associated with the pretest data are given in sections 1 through 6, followed by pre- and post-test flutter analysis in section 7, and the results of the aerodynamics/loads test in section 8. Finally, section 9 provides some recommendations based on lessons learned throughout the FSM program.

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

NASA Technical Reports Server (NTRS)

2007-01-01

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

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

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

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

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

Development of a quiet supersonic wind tunnel with a cryogenic adaptive nozzle

NASA Technical Reports Server (NTRS)

Wolf, Stephen W. D.

1995-01-01

Low-disturbance or 'quiet' wind tunnels are now considered an essential part of meaningful boundary layer transition research. Advances in Supersonic Laminar Flow Control (SLFC) technology for swept wings depends on a better understanding of the receptivity of the transition phenomena to attachment-line contamination and cross-flows. This need has provided the impetus for building the Laminar Flow Supersonic Wind Tunnel (LFSWT) at NASA-Ames, as part of the NASA High Speed Research Program (HSRP). The LFSWT was designed to provide NASA with an unequaled capability for transition research at low supersonic Mach numbers (<2.5). The following are the objectives in support of the new Fluid Mechanic Laboratory (FML) quiet supersonic wind tunnel: (I) Develop a unique injector drive system using the existing FML indraft compressor; (2) Develop an FML instrumentation capability for quiet supersonic wind tunnel evaluation and transition studies at NASA-Ames; (3) Determine the State of the Art in quiet supersonic wind tunnel design; (4) Build and commission the LFSWT; (5) Make detailed flow quality measurements in the LFSWT; (6) Perform tests of swept wing models in the LFSWT in support of the NASA HSR program; and (7) Provide documentation of research progress.

Wind tunnel pressure distribution tests on a series of biplane wing models Part II : effects of changes in decalage, dihedral, sweepback and overhang

NASA Technical Reports Server (NTRS)

Knight, Montgomery; Noyes, Richard W

1929-01-01

This preliminary report furnishes information on the changes in the forces on each wing of a biplane cellule when the decalage, dihedral, sweepback and overhang are separately varied. The data were obtained from pressure distribution tests made in the Atmospheric Wind Tunnel of the Langley Memorial Aeronautical Laboratory. Since each test was carried up to 90 degree angle of attack, the results may be used in the study of stalled flight and of spinning and in the structural design of biplane wings.

Development of a quiet supersonic wind tunnel with a cryogenic adaptive nozzle

NASA Technical Reports Server (NTRS)

Wolf, Stephen W. D.

1992-01-01

Aspects of the design and construction of the Laminar Flow Supersonic Wind Tunnel at the NASA-Ames Fluid Mechanics Laboratory are discussed. The wind tunnel is to be used as part of the NASA High Speed Research Program (HSRP).

Documentation and archiving of the Space Shuttle wind tunnel test data base. Volume 2: User's Guide to the Archived Data Base

NASA Technical Reports Server (NTRS)

Romere, Paul O.; Brown, Steve Wesley

1995-01-01

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

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

NASA Technical Reports Server (NTRS)

Romere, Paul O.; Brown, Steve Wesley

1995-01-01

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

NACA Technician Cleans a Ramjet in 8- by 6-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1950-04-21

A technician at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory cleans the pitot tube on a 16-inch diameter ramjet in the 8- by 6-Foot Supersonic Wind Tunnel. Pitot tubes are a measurement device used to determine the flow velocity at a specific location in the air stream, not the average velocity of the entire wind stream. NACA Lewis was in the midst of a multi-year program to determine the feasibility of ramjets and design improvements that could be employed for all models. The advantage of the ramjet was its ability to process large volumes of combustion air, resulting in the burning of fuel at the optimal stoichiometric temperatures. This was not possible with turbojets. The higher the Mach number, the more efficient the ramjet operated. The 8- by 6 Supersonic Wind Tunnel had been in operation for just over one year when this photograph was taken. The facility was the NACA’s largest supersonic tunnel and the only facility capable of running an engine at supersonic speeds. The 8- by 6 tunnel was also equipped with a Schlieren camera system that captured the air flow gradient as it passes over the test setup. The ramjet tests in the 8- by 6 tunnel complemented the NACA Lewis investigations using aircraft, the Altitude Wind Tunnel and smaller supersonic tunnels. Researchers studied the ramjet’s performance at different speeds and varying angles -of -attack.

Investigation of Condensation/Clustering Effects on Rayleigh Scattering Measurements in a Hypersonic Wind Tunnel

NASA Technical Reports Server (NTRS)

Tyler, Charles

1996-01-01

Rayleigh scattering, a nonintrusive measurement technique for the measurement of density in a hypersonic wind tunnel, is under investigation at Wright Laboratory's Mach 6 wind tunnel. Several adverse effects, i.e., extraneous scatter off walls and windows, hinder Rayleigh scattering measurements. Condensation and clustering of flow constituents also present formidable obstacles. Overcoming some of these difficulties, measurements have been achieved while the Mach 6 test section was pumped down to a vacuum, as well as for actual tunnel operation for various stagnation pressures at fixed stagnation temperatures. Stagnation pressures ranged from 0.69 MPa to 6.9 MPa at fixed stagnation temperatures of 511, 556, and 611 K. Rayleigh scatter results show signal levels much higher than expected for molecular scattering in the wind tunnel. Even with higher than expected signals, scattering measurements have been made in the flowfield of an 8-degree half-angle blunt nose cone with a nose radius of 1.5 cm.

Construction of the Propulsion Systems Laboratory No. 1 and 2

NASA Image and Video Library

1951-01-21

Construction of the Propulsion Systems Laboratory No. 1 and 2 at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory. When it began operation in late 1952, the Propulsion Systems Laboratory was the NACA’s most powerful facility for testing full-scale engines at simulated flight altitudes. The facility contained two altitude simulating test chambers which were a technological combination of the static sea-level test stands and the complex Altitude Wind Tunnel, which recreated actual flight conditions on a larger scale. NACA Lewis began designing the new facility in 1947 as part of a comprehensive plan to improve the altitude testing capabilities across the lab. The exhaust, refrigeration, and combustion air systems from all the major test facilities were linked. In this way, different facilities could be used to complement the capabilities of one another. Propulsion Systems Laboratory construction began in late summer 1949 with the installation of an overhead exhaust pipe connecting the facility to the Altitude Wind Tunnel and Engine Research Building. The large test section pieces arriving in early 1951, when this photograph was taken. The two primary coolers for the altitude exhaust are in place within the framework near the center of the photograph.

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.

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).

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...

Altitude Wind Tunnel at NASA Glenn Research Center: An Interactive History

NASA Technical Reports Server (NTRS)

2008-01-01

When constructed in the Early 1940s, the Altitude Wind Tunnel (AWT) at NASA Glenn Research Center was the nation's only wind tunnel capable of studying full scale engines under realistic flight conditions. It played a significant role in the development of the first U.S. jet engines as well as technologies such as the afterburner and variable-area nozzle. In the late 1950s, the tunnels interior components were removed so that hardware for Project Mercury could be tested in altitude conditions. In 1961, a portion of the tunnel was converted into one of the country's first large vacuum tanks and renamed the Space Power Chamber (SPC). SPC was used extensively throughout the 1960s for the Centaur rocket program. This multimedia piece allows one to interactively learn about the Altitude Wind Tunnel facility. and the research performed there. The piece contains: (1) A chronological history of the AWT from its construction during World War II and the testing of early jet engines, through the Mercury and Centaur programs of the 1960s and up to the final use of the building for the Microwave Systems laboratory. (2) Photographic surveys of the facility in it wind tunnel, vacuum tank and final configurations. (3) Browsable gallery of over 200 captioned photographs and video clips.(4) A nine minute documentary of the AWT produced by NASA in 1961 (5) Links to over 70 reports and publications related to AWT research and the history of the NACA.

Supersonic Research Display for Tour

NASA Image and Video Library

1946-03-21

On March 22, 1946, 250 members of the Institute of Aeronautical Science toured the NACA’s Aircraft Engine Research Laboratory. NACA Chairman Jerome Hunsaker and Secretary John Victory were on hand to brief the attendees in the Administration Building before the visited the lab’s test facilities. At each of the twelve stops, researchers provided brief presentations on their work. Topics included axial flow combustors, materials for turbine blades, engine cooling, icing prevention, and supersonic flight. The laboratory reorganized itself in October 1945 as World War II came to an end to address newly emerging technologies such as the jet engine, rockets, and high-speed flight. While design work began on what would eventually become the 8- by 6-Foot Supersonic Wind Tunnel, NACA Lewis quickly built several small supersonic tunnels. These small facilities utilized the Altitude Wind Tunnel’s massive air handling equipment to generate high-speed airflow. The display seen in this photograph was set up in the building that housed the first of these wind tunnels. Eventually the building would contain 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. The small tunnels were used until the early 1960s to study the aerodynamic characteristics of supersonic inlets and exits.

8- by 6-Foot Supersonic Wind Tunnel's Original Design

NASA Image and Video Library

1949-07-21

Aerial view of the 8- by 6-Foot Supersonic Wind Tunnel in its original configuration at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory. The 8- by 6 was the laboratory’s first large supersonic wind tunnel. It was also the NACA’s most powerful supersonic tunnel, and its first facility capable of running an engine at supersonic speeds. The 8- by 6-foot tunnel has been used to study inlets and exit nozzles, fuel injectors, flameholders, exit nozzles, and controls on ramjet and turbojet propulsion systems. The 8- by 6 was originally an open-throat and non-return tunnel. This meant that the supersonic air flow was blown through the test section and out the other end into the atmosphere. In this photograph, the three drive motors in the structure at the left supplied power to the seven-stage axial-flow compressor in the light-colored structure. The air flow passed through flexible walls which were bent to create the desired speed. The test article was located in the 8- by 6-foot stainless steel test section located inside the steel pressure chamber at the center of this photograph. The tunnel dimensions were then gradually increased to slow the air flow before it exited into the atmosphere. The large two-story building in front of the tunnel was used as office space for the researchers.

Wind Tunnel Pressure Distribution Tests on a Series of Biplane Wing Models. Part III Effects of Charges in Various Combinations of Stagger, Gap, Sweepback, and Decalage

NASA Technical Reports Server (NTRS)

Knight, Montgomery; Noyes, Richard W

1929-01-01

This preliminary report furnishes information on the changes in the forces on each wing of a biplane cellule for various combinations of stagger and gap, stagger and sweepback, stagger and decalage, and gap and decalage. The data were obtained from pressure distribution tests made in the atmospheric wind tunnel of the Langley Memorial Aeronautical Laboratory. Since each test was carried up to 90deg angle of attack, the results may be used in the study of stalled flight and of spinning as well as in the structural design of biplane wings.

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.

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.

40 CFR 53.65 - Test procedure: Loading test.

Code of Federal Regulations, 2013 CFR

2013-07-01

... performing the test in § 53.62 (full wind tunnel test), § 53.63 (wind tunnel inlet aspiration test), or § 53... particle delivery system shall consist of a static chamber or a low velocity wind tunnel having a.... The mean velocity in the test section of the static chamber or wind tunnel shall not exceed 2 km/hr...

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...

Ramjet Testing in the NACA's Altitude Wind Tunnel

NASA Image and Video Library

1946-02-21

A 20-inch diameter ramjet installed in the Altitude Wind Tunnel at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory. The Altitude Wind Tunnel was used in the 1940s to study early ramjet configurations. Ramjets provide a very simple source of propulsion. They are basically a tube which takes in high-velocity air, ignites it, and then expels the expanded airflow at a significantly higher velocity for thrust. Ramjets are extremely efficient and powerful but can only operate at high speeds. Therefore a turbojet or rocket was needed to launch the vehicle. This NACA-designed 20-inch diameter ramjet was installed in the Altitude Wind Tunnel in May 1945. The ramjet was mounted under a section of wing in the 20-foot diameter test section with conditioned airflow ducted directly to the engine. The mechanic in this photograph was installing instrumentation devices that led to the control room. NACA researchers investigated the ramjet’s overall performance at simulated altitudes up to 47,000 feet. Thrust measurements from these runs were studied in conjunction with drag data obtained during small-scale studies in the laboratory’s small supersonic tunnels. An afterburner was attached to the ramjet during the portions of the test program. The researchers found that an increase in altitude caused a reduction in the engine’s horsepower. They also determined the optimal configurations for the flameholders, which provided the engine’s ignition source.

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.

The Brothers Were Wright - An Abridged History of Wind Tunnel Testing at Ames Research Center

NASA Technical Reports Server (NTRS)

Buchholz, Steve

2017-01-01

The Wright Brothers used wind tunnel data to refine their design for the first successful airplane back in 1903. Today, wind tunnels are still in use all over the world gathering data to improve the design of cars, trucks, airplanes, missiles and spacecraft. Ames Research Center is home to many wind tunnels, including the Unitary Plan Wind Tunnel complex. Built in the early 1950s, it is one of the premiere transonic and supersonic testing facilities in the country. Every manned spacecraft has been tested in the wind tunnels at Ames. This is a testing history from past to present.

Altitude Wind Tunnel Control Room

NASA Image and Video Library

1945-05-21

Researchers at the National Advisory Committee for Aeronautics (NACA) Aircraft Engine Research Laboratory monitor a ramjet's performance in the Altitude Wind Tunnel from the control room. The soundproof control room was just a few feet from the tunnel’s 20-foot-diameter test section. In the control room, the operators could control all aspects of the tunnel’s operation, including the air density, temperature, and speed. They also operated the engine or test article in the test section by controlling the angle-of-attack, speed, power, and other parameters. The men in this photograph are monitoring the engine’s thrust and lift. A NACA-designed 20-inch-diameter ramjet was installed in the tunnel in May 1945. Thrust figures from these runs were compared with drag data from tests of scale models in small supersonic tunnels to verify the ramjet’s feasibility. The tunnel was used to analyze the ramjet’s overall performance up to altitudes of 47,000 feet and speeds to Mach 1.84. The researchers found that an increase in altitude caused a reduction in the engine’s horsepower and identified optimal flameholder configurations.

Neural Network Modeling of UH-60A Pilot Vibration

NASA Technical Reports Server (NTRS)

Kottapalli, Sesi

2003-01-01

Full-scale flight-test pilot floor vibration is modeled using neural networks and full-scale wind tunnel test data for low speed level flight conditions. Neural network connections between the wind tunnel test data and the tlxee flight test pilot vibration components (vertical, lateral, and longitudinal) are studied. Two full-scale UH-60A Black Hawk databases are used. The first database is the NASMArmy UH-60A Airloads Program flight test database. The second database is the UH-60A rotor-only wind tunnel database that was acquired in the NASA Ames SO- by 120- Foot Wind Tunnel with the Large Rotor Test Apparatus (LRTA). Using neural networks, the flight-test pilot vibration is modeled using the wind tunnel rotating system hub accelerations, and separately, using the hub loads. The results show that the wind tunnel rotating system hub accelerations and the operating parameters can represent the flight test pilot vibration. The six components of the wind tunnel N/rev balance-system hub loads and the operating parameters can also represent the flight test pilot vibration. The present neural network connections can significandy increase the value of wind tunnel testing.

Development of a quiet supersonic wind tunnel with a cryogenic adaptive nozzle

NASA Technical Reports Server (NTRS)

Wolf, Stephen W. D.

1993-01-01

The main objective of this work is to develop an interim Quiet (low-disturbance) supersonic wind tunnel for the NASA-Ames Fluid Mechanics Laboratory (FML). The main emphasis is to bring on-line a full-scale Mach 1.6 tunnel as rapidly as possible to impact the NASA High Speed Research Program (HSRP). The development of a cryogenic adaptive nozzle and other sophisticated features of the tunnel will now happen later, after the full scale wind tunnel is in operation. The work under this contract for the period of this report can be summarized as follows: provide aerodynamic design requirements for the NASA-Ames Fluid Mechanics Laboratory (FML) Laminar Flow Supersonic Wind Tunnel (LFSWT); research design parameters for a unique Mach 1.6 drive system for the LFSWT using an 1/8th-scale Proof-of-Concept (PoC) supersonic wind tunnel; carry out boundary layer transition studies in PoC to aid the design of critical components of the LFSWT; appraise the State of the Art in quiet supersonic wind tunnel design; and help develop a supersonic research capability within the FML particularly in the areas of high speed transition measurements and schlieren techniques. The body of this annual report summarizes the work of the Principal Investigator.

Engineer Measures Ice Formation on an Instrument Antenna Model

NASA Image and Video Library

1945-05-21

A National Advisory Committee for Aeronautics (NACA) researcher measures the ice thickness on a landing antenna model in the Icing Research Tunnel at the Aircraft Engine Research Laboratory. NACA design engineers added the Icing Research Tunnel to the original layout of the new Aircraft Engine Research Laboratory to take advantage of the massive refrigeration system being built for the Altitude Wind Tunnel. The Icing Research Tunnel was built to study the formation of ice on aircraft surfaces and methods of preventing or eradicating that ice. Ice buildup adds extra weight, effects aerodynamics, and sometimes blocks air flow through engines. The Icing Research Tunnel is a closed-loop atmospheric wind tunnel with a 6- by 9-foot test section. Carrier Corporation refrigeration equipment reduced the internal air temperature to -45 degrees F and a spray bar system injected water droplets into the air stream. The 24-foot diameter drive fan, seen in this photograph, created air flows velocities up to 400 miles per hour. The Icing Research Tunnel began testing in June of 1944. Early testing, seen in this photograph, studied ice accumulation on propellers and antenna of a military aircraft. The Icing Research Tunnel’s designers, however, struggled to develop a realistic spray system since they did not have access to data on the size of naturally occurring water droplets. The system would have to generate small droplets, distribute them uniformly throughout the airstream, and resist freezing and blockage. For five years a variety of different designs were painstakingly developed and tested before the system was perfected.

The use of wind tunnel facilities to estimate hydrodynamic data

NASA Astrophysics Data System (ADS)

Hoffmann, Kristoffer; Tophøj Rasmussen, Johannes; Hansen, Svend Ole; Reiso, Marit; Isaksen, Bjørn; Egeberg Aasland, Tale

2016-03-01

Experimental laboratory testing of vortex-induced structural oscillations in flowing water is an expensive and time-consuming procedure, and the testing of high Reynolds number flow regimes is complicated due to the requirement of either a large-scale or high-speed facility. In most cases, Reynolds number scaling effects are unavoidable, and these uncertainties have to be accounted for, usually by means of empirical rules-of-thumb. Instead of performing traditional hydrodynamic measurements, wind tunnel testing in an appropriately designed experimental setup may provide an alternative and much simpler and cheaper framework for estimating the structural behavior under water current and wave loading. Furthermore, the fluid velocities that can be obtained in a wind tunnel are substantially higher than in a water testing facility, thus decreasing the uncertainty from scaling effects. In a series of measurements, wind tunnel testing has been used to investigate the static response characteristics of a circular and a rectangular section model. Motivated by the wish to estimate the vortex-induced in-line vibration characteristics of a neutrally buoyant submerged marine structure, additional measurements on extremely lightweight, helium-filled circular section models were conducted in a dynamic setup. During the experiment campaign, the mass of the model was varied in order to investigate how the mass ratio influences the vibration amplitude. The results show good agreement with both aerodynamic and hydrodynamic experimental results documented in the literature.

Design and laboratory testing of a new flow-through directional passive air sampler for ambient particulate matter.

PubMed

Lin, Chun; Solera Garcia, Maria Angeles; Timmis, Roger; Jones, Kevin C

2011-03-01

A new type of directional passive air sampler (DPAS) is described for collecting particulate matter (PM) in ambient air. The prototype sampler has a non-rotating circular sampling tray that is divided into covered angular channels, whose ends are open to winds from sectors covering the surrounding 360°. Wind-blown PM from different directions enters relevant wind-facing channels, and is retained there in collecting pools containing various sampling media. Information on source direction and type can be obtained by examining the distribution of PM between channels. Wind tunnel tests show that external wind velocities are at least halved over an extended area of the collecting pools, encouraging PM to settle from the air stream. Internal and external wind velocities are well-correlated over an external velocity range of 2.0-10.0 m s⁻¹, which suggests it may be possible to relate collected amounts of PM simply to ambient concentrations and wind velocities. Measurements of internal wind velocities in different channels show that velocities decrease from the upwind channel round to the downwind channel, so that the sampler effectively resolves wind directions. Computational fluid dynamics (CFD) analyses were performed on a computer-generated model of the sampler for a range of external wind velocities; the results of these analyses were consistent with those from the wind tunnel. Further wind tunnel tests were undertaken using different artificial particulates in order to assess the collection performance of the sampler in practice. These tests confirmed that the sampler can resolve the directions of sources, by collecting particulates preferentially in source-facing channels.

Immersion and contact freezing experiments in the Mainz wind tunnel laboratory

NASA Astrophysics Data System (ADS)

Eppers, Oliver; Mayer, Amelie; Diehl, Karoline; Mitra, Subir; Borrmann, Stephan; Szakáll, Miklós

2016-04-01

Immersion and contact freezing are of outmost important ice nucleation processes in mixed phase clouds. Experimental studies are carried out in the Mainz vertical wind tunnel laboratory in order to characterize these nucleation processes for different ice nucleating particles (INP), such as for mineral dust or biological particles. Immersion freezing is investigated in our laboratory with two different experimental techniques, both attaining contact-free levitation of liquid droplets and cooling of the surrounding air down to about -25 °C. In an acoustic levitator placed in the cold room of our laboratory, drops with diameters of 2 mm are investigated. In the vertical air stream of the wind tunnel droplets with diameter of 700 micron are freely floated at their terminal velocities, simulating the flow conditions of the free atmosphere. Furthermore, the wind tunnel offers a unique platform for contact freezing experiments. Supercooled water droplets are floated in the vertical air stream at their terminal velocities and INP are injected into the tunnel air stream upstream of them. As soon as INP collides with the supercooled droplet the contact freezing is initiated. The first results of immersion and contact freezing experiments with cellulose particles both in the acoustic levitator and in the wind tunnel will be presented. Cellulose is considered as typical INP of biological origin and a macrotracer for plant debris. Nucleating properties of cellulose will be provided, mainly focusing on the temperature, INP concentration, and specific surface area dependences of the freezing processes. Direct comparison between the different experimental techniques (acoustic levitator and wind tunnel), as well as between nucleation modes (immersion and contact freezing) will be presented. The work is carried out within the framework of the German research unit INUIT.

Design, construction and commissioning of the Braunschweig Icing Wind Tunnel

NASA Astrophysics Data System (ADS)

Bansmer, Stephan E.; Baumert, Arne; Sattler, Stephan; Knop, Inken; Leroy, Delphine; Schwarzenboeck, Alfons; Jurkat-Witschas, Tina; Voigt, Christiane; Pervier, Hugo; Esposito, Biagio

2018-06-01

Beyond its physical importance in both fundamental and climate research, atmospheric icing is considered as a severe operational condition in many engineering applications like aviation, electrical power transmission and wind-energy production. To reproduce such icing conditions in a laboratory environment, icing wind tunnels are frequently used. In this paper, a comprehensive overview on the design, construction and commissioning of the Braunschweig Icing Wind Tunnel is given. The tunnel features a test section of 0.5 m × 0.5 m with peak velocities of up to 40 m s-1. The static air temperature ranges from -25 to +30 °C. Supercooled droplet icing with liquid water contents up to 3 g m-3 can be reproduced. The unique aspect of this facility is the combination of an icing tunnel with a cloud chamber system for making ice particles. These ice particles are more realistic in shape and density than those usually used for mixed phase and ice crystal icing experiments. Ice water contents up to 20 g m-3 can be generated. We further show how current state-of-the-art measurement techniques for particle sizing are performed on ice particles. The data are compared to those of in-flight measurements in mesoscale convective cloud systems in tropical regions. Finally, some applications of the icing wind tunnel are presented.

ARC-2009-ACD09-0053-005

NASA Image and Video Library

2009-04-08

Mars Science Laboratory (MSL) parachuste test in the NASA Ames 80x120ft Subsonic Wind Tunnel at Moffett Field, Calif. (TR #22 - Phase 6) is the largest ever built to fly on an extraterrestrail mission. This image shows the qualification-test parachute beginning to open a few seconds after it was launched from a mortar into an 80-mile-per-hour (36-meter-per-second) wind.

Development of an impulsive noise source to study the acoustic reflection characteristics of hard-walled wind tunnels

NASA Technical Reports Server (NTRS)

Salikuddin, M.; Burrin, R. H.; Ahuja, K. K.; Bartel, H. W.

1986-01-01

Two impulsive sound sources, one using multiple acoustic drivers and the other using a spark discharge were developed to study the acoustic reflection characteristics of hard-walled wind tunnels, and the results of laboratory tests are presented. The analysis indicates that though the intensity of the pulse generated by the spark source was higher than that obtained from the acoustic source, the number of averages needed for a particular test may require an unacceptibly long tunnel-run time due to the low spark generation repeat rate because of capacitor charging time. The additional hardware problems associated with the longevity of electrodes and electrode holders in sustaining the impact of repetitive spark discharges, show the multidriver acoustic source to be more suitable for this application.

Post-test data report for the space shuttle full-scale AFRSI sequence of environments test (OS-305-1 to -5) in the NASA/Ames Research Center 11x11-foot wind tunnel

NASA Technical Reports Server (NTRS)

Marshall, B. A.

1984-01-01

The Advanced Flexible Reusable Surface Insulation (AFRSI) test article was wind tunnel tested. The AFRSI was exposed to a simulated ascent airloads environment and data was obtained which could be used to support the AFRSI certification program. The AFRSI sequence of environments also included radiant heating (1500 degrees Fahrenheit) and wind/rain environments. The test article was wind/rain conditioned before each wind tunnel entry and was thermally conditioned after each wind tunnel entry. The AFRSI failed and the test was aborted before reaching the ascent environment. The AFRSI test article sequentially exposed to 50 wind/rain and 49 simulated entry thermal missions, as well as four wind tunnel entries equivalent to 40 ascent missions.

A Modified Bagnold-Type Wind Tunnel for Laboratory Use

ERIC Educational Resources Information Center

Logan, Alan

1975-01-01

Using the basic Bagnold design, a relatively inexpensive suction-type wind tunnel can be constructed for laboratory demonstration of sand-grain movement, ripple development, and other eolian processes. Its simple design provides no workshop problems and it can be made for a total cost in materials of approximately $225. (Author/CP)

Comparisons of Mixed-Phase Icing Cloud Simulations with Experiments Conducted at the NASA Propulsion Systems Laboratory

NASA Technical Reports Server (NTRS)

Bartkus, Tadas; Tsao, Jen-Ching; Struk, Peter

2017-01-01

This paper builds on previous work that compares numerical simulations of mixed-phase icing clouds with experimental data. The model couples the thermal interaction between ice particles and water droplets of the icing cloud with the flowing air of an icing wind tunnel for simulation of NASA Glenn Research Centers (GRC) Propulsion Systems Laboratory (PSL). Measurements were taken during the Fundamentals of Ice Crystal Icing Physics Tests at the PSL tunnel in March 2016. The tests simulated ice-crystal and mixed-phase icing that relate to ice accretions within turbofan engines.

AFRL Ludwieg Tube Initial Performance

DTIC Science & Technology

2017-11-01

Nov 2016 14. ABSTRACT The Air Force Research Laboratory has developed and constructed a Ludwieg tube wind tunnel for hypersonic experimental ...Ludwieg tube wind tunnel for hypersonic experimental research. This wind tunnel is now operational and its initial performance has been...opening and the first expansion wave reflection was 100 ms, as expected. About 80 ms of quasi -steady pressure was obtained after the valve-opening

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

Code of Federal Regulations, 2013 CFR

2013-07-01

... tunnel tests. 53.42 Section 53.42 Protection of Environment ENVIRONMENTAL PROTECTION AGENCY (CONTINUED... particle delivery system shall consist of a blower system and a wind tunnel having a test section of... particles delivered to the test section of the wind tunnel shall be established using the operating...

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

Code of Federal Regulations, 2012 CFR

2012-07-01

... tunnel tests. 53.42 Section 53.42 Protection of Environment ENVIRONMENTAL PROTECTION AGENCY (CONTINUED... particle delivery system shall consist of a blower system and a wind tunnel having a test section of... particles delivered to the test section of the wind tunnel shall be established using the operating...

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

Code of Federal Regulations, 2014 CFR

2014-07-01

... tunnel tests. 53.42 Section 53.42 Protection of Environment ENVIRONMENTAL PROTECTION AGENCY (CONTINUED... particle delivery system shall consist of a blower system and a wind tunnel having a test section of... particles delivered to the test section of the wind tunnel shall be established using the operating...

Some new tests at the Gottingen laboratory

NASA Technical Reports Server (NTRS)

1921-01-01

The tests at the Gottingen laboratory included: friction tests on a surface treated with omelette, verification tests on the M.V.A. 356 wing, and comparative tests of wing no. 36 at the Eiffel laboratory. The examination of all these experiments leads to the belief that, at large incidences, the speeds registered by the suction manometer of the testing chamber of the Eiffel laboratory wind tunnel are, owing to pressure drop, greater than the actual speeds. Therefore, the values of k(sub x) and k(sub y) measured at the Eiffel laboratory at large incidences are too low.

NASA Ames Sonic Boom Testing

NASA Technical Reports Server (NTRS)

Durston, Donald A.; Kmak, Francis J.

2009-01-01

Multiple sonic boom wind tunnel models were tested in the NASA Ames Research Center 9-by 7-Foot Supersonic Wind Tunnel to reestablish related test techniques in this facility. The goal of the testing was to acquire higher fidelity sonic boom signatures with instrumentation that is significantly more sensitive than that used during previous wind tunnel entries and to compare old and new data from established models. Another objective was to perform tunnel-to-tunnel comparisons of data from a Gulfstream sonic boom model tested at the NASA Langley Research Center 4-foot by 4-foot Unitary Plan Wind Tunnel.

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.

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.

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.

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

NASA Technical Reports Server (NTRS)

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

1982-01-01

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

Procedures and requirements for testing in the Langley Research Center unitary plan wind tunnel

NASA Technical Reports Server (NTRS)

Wassum, Donald L.; Hyman, Curtis E., Jr.

1988-01-01

Information is presented to assist those interested in conducting wind-tunnel testing within the Langley Unitary Plan Wind Tunnel. Procedures, requirements, forms and examples necessary for tunnel entry are included.

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

NASA Technical Reports Server (NTRS)

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

2002-01-01

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

Attraction of the cutaneous leishmaniasis vector Nyssomyia neivai (Diptera: Psychodidae) to host odour components in a wind tunnel

PubMed Central

2012-01-01

Background Laboratory studies of host-seeking olfactory behaviour in sandflies have largely been restricted to the American visceral leishmaniasis vector Lutzomyia longipalpis. In comparison, almost nothing is known about the chemical ecology of related species, which transmit American cutaneous leishmaniasis (ACL), due in part to difficulties in raising these insects in the laboratory. Understanding how ACL vectors locate their hosts will be essential to developing new vector control strategies to combat this debilitating disease. Methods This study examined host-odour seeking behaviour of the ACL vector Nyssomyia neivai (Pinto) (=Lutzomyia neivai) using a wind tunnel olfactometer. The primary aim was to determine whether field-collected female N. neivai would respond to host odours in the laboratory, thereby eliminating the need to maintain colonies of these insects for behavioural experiments. Responses to two key host odour components, 1-octen-3-ol and lactic acid, and a commercially-available mosquito lure (BG-Lure™) were assessed and compared relative to an air control. We also tested whether trials could be conducted outside of the normal evening activity period of N. neivai without impacting on fly behaviour, and whether the same flies could be used to assess baseline responses to air without affecting responses to octenol, thereby reducing the number of flies required for experiments. Results Octenol was found to both activate host-seeking behaviour and attract female N. neivai in the wind tunnel, while lactic acid elicited weaker responses of activation and attractiveness under identical conditions. The BG-Lure did not activate or attract N. neivai under test conditions. Further experiments showed that sandfly behaviour in the wind tunnel was not affected by time of day, such that experiments need not be restricted to nocturnal hours. Moreover, using the same flies to measure both baseline responses to air and attraction to test compounds did not affect odour-seeking behaviour. Conclusions The results of this study demonstrate that N. neivai taken from the field are suitable for use in laboratory olfactometer experiments. It is hoped this work will facilitate further research into chemical ecology of this species, and other ACL vectors. PMID:23009099

Attraction of the cutaneous leishmaniasis vector Nyssomyia neivai (Diptera: Psychodidae) to host odour components in a wind tunnel.

PubMed

Pinto, Mara C; Bray, Daniel P; Eiras, Alvaro E; Carvalheira, Henrique P; Puertas, Camila P

2012-09-25

Laboratory studies of host-seeking olfactory behaviour in sandflies have largely been restricted to the American visceral leishmaniasis vector Lutzomyia longipalpis. In comparison, almost nothing is known about the chemical ecology of related species, which transmit American cutaneous leishmaniasis (ACL), due in part to difficulties in raising these insects in the laboratory. Understanding how ACL vectors locate their hosts will be essential to developing new vector control strategies to combat this debilitating disease. This study examined host-odour seeking behaviour of the ACL vector Nyssomyia neivai (Pinto) (=Lutzomyia neivai) using a wind tunnel olfactometer. The primary aim was to determine whether field-collected female N. neivai would respond to host odours in the laboratory, thereby eliminating the need to maintain colonies of these insects for behavioural experiments. Responses to two key host odour components, 1-octen-3-ol and lactic acid, and a commercially-available mosquito lure (BG-Lure™) were assessed and compared relative to an air control. We also tested whether trials could be conducted outside of the normal evening activity period of N. neivai without impacting on fly behaviour, and whether the same flies could be used to assess baseline responses to air without affecting responses to octenol, thereby reducing the number of flies required for experiments. Octenol was found to both activate host-seeking behaviour and attract female N. neivai in the wind tunnel, while lactic acid elicited weaker responses of activation and attractiveness under identical conditions. The BG-Lure did not activate or attract N. neivai under test conditions. Further experiments showed that sandfly behaviour in the wind tunnel was not affected by time of day, such that experiments need not be restricted to nocturnal hours. Moreover, using the same flies to measure both baseline responses to air and attraction to test compounds did not affect odour-seeking behaviour. The results of this study demonstrate that N. neivai taken from the field are suitable for use in laboratory olfactometer experiments. It is hoped this work will facilitate further research into chemical ecology of this species, and other ACL vectors.

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

NASA Technical Reports Server (NTRS)

Kayten, Gerald G

1945-01-01

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

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.

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, vortex shedding, and local turbulence intensity and wind shear values. To achieve accurate results, attention must of course be paid to issues such as ensuring Reynolds number independence, avoiding blockage issues, and properly matching the velocity power spectrum, but once this is done, the laws of fluid mechanics take care of the rest. There will not be an overproduction of turbulent kinetic energy at the top of escarpments, or unacceptable dissipation of inlet turbulence levels. Modern atmospheric boundary layer wind tunnels are also often used to provide validation data for evaluating the performance of CFD model in complex flow environments. Present day computers have further increased the quality and quantity of data that can be economically obtained in a timely manner, for example through wind speed measurement using a computer controlled 3-D measurement positioning system Given this accuracy and widespread acceptance, it is perhaps surprising that ours was the only wind tunnel model in the Bolund blind experiment, an indication of how seldom physical modelling is used when estimating terrain effect for wind farms. In demonstrating how the Bolund test was modeled, this presentation will provide background on wind tunnel testing, including the governing scaling parameters. And we’ll see how our results compared to the full scale tests.

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.

Transition Analysis for the Mars Science Laboratory Entry Vehicle

NASA Technical Reports Server (NTRS)

Chang, Chau-Lyan; Choudhari, Meelan M.; Hollis, Brian R.; Li, Fei

2009-01-01

Viscous Laminar-turbulent transition plays an important role in the design of the Mars Science Laboratory (MSL) entry vehicle. The lift-to-drag ratio required for the precision landing trajectory will be achieved via an angle of attack equal to 16 degrees. At this relatively high angle of attack, the boundary layer flow near the leeward meridian is expected to transition early in the trajectory, resulting in substantially increased heating loads. This paper presents stability calculations and transition correlations for a series of wind tunnel models of the MSL vehicle. Experimentally measured transition onset locations are used to correlate with the N-factor calculations for various wind tunnel conditions. Due to relatively low post-shock Mach numbers near the edge of the boundary layer, the dominant instability waves are found to be of the first mode type. The N-factor values correlating with measured transition onset at selected test points from the Mach 6 conventional facility experiments fall between 3.5 and 4.5 and apparently vary linearly with the wind tunnel unit Reynolds number, indicating strong receptivity effect. The small transition N value is consistent with previous correlations for second-mode dominant transition in the same wind tunnel facility. Stability calculations for stationary and traveling crossflow instability waves in selected configurations indicate that an N value of 4 and 6, respectively, correlates reasonably well with transition onset discerned from one experimentally measured thermographic image.

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.

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

NASA Technical Reports Server (NTRS)

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

2013-01-01

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

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.

NACA Computer at the Lewis Flight Propulsion Laboratory

NASA Image and Video Library

1951-02-21

A female computer at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory with a slide rule and Friden adding machine to make computations. The computer staff was introduced during World War II to relieve short-handed research engineers of some of the tedious computational work. The Computing Section was staffed by “computers,” young female employees, who often worked overnight when most of the tests were run. The computers obtained test data from the manometers and other instruments, made the initial computations, and plotted the data graphically. Researchers then analyzed the data and summarized the findings in a report or made modifications and ran the test again. There were over 400 female employees at the laboratory in 1944, including 100 computers. The use of computers was originally planned only for the duration of the war. The system was so successful that it was extended into the 1960s. The computers and analysts were located in the Altitude Wind Tunnel Shop and Office Building office wing during the 1940s and transferred to the new 8- by 6-Foot Supersonic Wind Tunnel in 1948.

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

14. View north of Tropic wind tunnel and frontal view ...

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

14. View north of Tropic wind tunnel and frontal view of main fan (typical). - Natick Research & Development Laboratories, Climatic Chambers Building, U.S. Army Natick Research, Development & Engineering Center (NRDEC), Natick, Middlesex County, MA

7. Detail view west of Arctic Chamber wind tunnel shell ...

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

7. Detail view west of Arctic Chamber wind tunnel shell (typical) in east elevation. - Natick Research & Development Laboratories, Climatic Chambers Building, U.S. Army Natick Research, Development & Engineering Center (NRDEC), Natick, Middlesex County, MA

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.

Increased Mach Number Capability for the NASA Glenn 10x10 Supersonic Wind Tunnel

NASA Technical Reports Server (NTRS)

Slater, John; Saunders, John

2014-01-01

Computational simulations and wind tunnel testing were conducted to explore the operation of the Abe Silverstein Supersonic Wind Tunnel at the NASA Glenn Research Center at test section Mach numbers above the current limit of Mach 3.5. An increased Mach number would enhance the capability for testing of supersonic and hypersonic propulsion systems. The focus of the explorations was on understanding the flow within the second throat of the tunnel, which is downstream of the test section and is where the supersonic flow decelerates to subsonic flow. Methods of computational fluid dynamics (CFD) were applied to provide details of the shock boundary layer structure and to estimate losses in total pressure. The CFD simulations indicated that the tunnel could be operated up to Mach 4.0 if the minimum width of the second throat was made smaller than that used for previous operation of the tunnel. Wind tunnel testing was able to confirm such operation of the tunnel at Mach 3.6 and 3.7 before a hydraulic failure caused a stop to the testing. CFD simulations performed after the wind tunnel testing showed good agreement with test data consisting of static pressures along the ceiling of the second throat. The CFD analyses showed increased shockwave boundary layer interactions, which was also observed as increased unsteadiness of dynamic pressures collected in the wind tunnel testing.

Increased Mach Number Capability for the NASA Glenn 10x10 Supersonic Wind Tunnel

NASA Technical Reports Server (NTRS)

Slater, J. W.; Saunders, J. D.

2015-01-01

Computational simulations and wind tunnel testing were conducted to explore the operation of the Abe Silverstein Supersonic Wind Tunnel at the NASA Glenn Research Center at test section Mach numbers above the current limit of Mach 3.5. An increased Mach number would enhance the capability for testing of supersonic and hypersonic propulsion systems. The focus of the explorations was on understanding the flow within the second throat of the tunnel, which is downstream of the test section and is where the supersonic flow decelerates to subsonic flow. Methods of computational fluid dynamics (CFD) were applied to provide details of the shock boundary layer structure and to estimate losses in total pressure. The CFD simulations indicated that the tunnel could be operated up to Mach 4.0 if the minimum width of the second throat was made smaller than that used for previous operation of the tunnel. Wind tunnel testing was able to confirm such operation of the tunnel at Mach 3.6 and 3.7 before a hydraulic failure caused a stop to the testing. CFD simulations performed after the wind tunnel testing showed good agreement with test data consisting of static pressures along the ceiling of the second throat. The CFD analyses showed increased shockwave boundary layer interactions, which was also observed as increased unsteadiness of dynamic pressures collected in the wind tunnel testing.

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.

Subsonic Wind Tunnel Tests of the FBTV Configuration in Proximity of the B-52

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

Priebe, R.W.

1966-12-01

Wind tunnel tests were conducted on a .075 scale Sandia FBTV store model in an 8-foot transonic wind tunnel during December `66. These tests were performed to obtain longitudinal and lateral stability characteristics.

40 CFR 53.43 - Test procedures.

Code of Federal Regulations, 2012 CFR

2012-07-01

... of the test section of the wind tunnel. The mean wind speed in the test section must be within ±10... wind tunnel and allow the particle concentration to stabilize. (vi) Install an array of five or more evenly spaced isokinetic samplers in the sampling zone (see § 53.42(d)) of the wind tunnel. Collect...

40 CFR 53.43 - Test procedures.

Code of Federal Regulations, 2014 CFR

2014-07-01

...-sectional area of the test section of the wind tunnel. The mean wind speed in the test section must be... into the wind tunnel and allow the particle concentration to stabilize. (vi) Install an array of five or more evenly spaced isokinetic samplers in the sampling zone (see § 53.42(d)) of the wind tunnel...

40 CFR 53.43 - Test procedures.

Code of Federal Regulations, 2013 CFR

2013-07-01

...-sectional area of the test section of the wind tunnel. The mean wind speed in the test section must be... into the wind tunnel and allow the particle concentration to stabilize. (vi) Install an array of five or more evenly spaced isokinetic samplers in the sampling zone (see § 53.42(d)) of the wind tunnel...

Reduction of Helicopter BVI Noise, Vibration, and Power Consumption Through Individual Blade Control

NASA Technical Reports Server (NTRS)

Jacklin, Stephen A.; Blaas, Achim; Teves, Dietrich; Kube, Roland; Warmbrodt, William (Technical Monitor)

1994-01-01

A wind tunnel test was conducted with a full-scale BO 105 helicopter rotor to evaluate the potential of open-loop individual blade control (IBC) to improve rotor performance, to reduce blade vortex interaction (BVI) noise, and to alleviate helicopter vibrations. The wind tunnel test was an international collaborative effort between NASA/U.S. Army AFDD, ZF Luftfahrttechnik, Eurocopter Deutschland, and the German Aerospace Laboratory (DLR) and was conducted under the auspices of the U.S./German MOU on Rotorcraft Aeromechanics. In this test the normal blade pitch links of the rotor were replaced by servo-actuators so that the pitch of each blade could be controlled independently of the other blades. The specially designed servoactuators and IBC control system were designed and manufactured by ZF Luftfahrttechnik, GmbH. The wind tunnel test was conducted in the 40- by 80-Foot Wind Tunnel at the NASA Ames Research Center. An extensive amount of measurement information was acquired for each IBC data point. These data include rotor performance, static and dynamic hub forces and moments, rotor loads, control loads, inboard and outboard blade pitch motion, and BVI noise data. The data indicated very significant (80 percent) simultaneous reductions in both BVI noise and hub vibrations could be obtained using multi-harmonic input at the critical descent (terminal approach) condition. The data also showed that performance improvements of up to 7 percent could be obtained using 2P input at high-speed forward flight conditions.

Altitude Wind Tunnel Operating at Night

NASA Image and Video Library

1945-04-21

The Altitude Wind Tunnel (AWT) during one of its overnight runs at the National Advisory Committee for Aeronautics (NACA) Aircraft Engine Research Laboratory in Cleveland, Ohio. The AWT was run during night hours so that its massive power loads were handled when regional electric demands were lowest. At the time the AWT was among the most complex wind tunnels ever designed. In order to simulate conditions at high altitudes, NACA engineers designed innovative new systems that required tremendous amounts of electricity. The NACA had an agreement with the local electric company that it would run its larger facilities overnight when local demand was at its lowest. In return the utility discounted its rates for the NACA during those hours. The AWT could produce wind speeds up to 500 miles per hour through its 20-foot-diameter test section at the standard operating altitude of 30,000 feet. The airflow was created by a large fan that was driven by an 18,000-horsepower General Electric induction motor. The altitude simulation was accomplished by large exhauster and refrigeration systems. The cold temperatures were created by 14 Carrier compressors and the thin atmosphere by four 1750-horsepower exhausters. The first and second shifts usually set up and broke down the test articles, while the third shift ran the actual tests. Engineers would often have to work all day, then operate the tunnel overnight, and analyze the data the next day. The night crew usually briefed the dayshift on the tests during morning staff meetings.

The steady-state flow quality in a model of a non-return wind tunnel

NASA Technical Reports Server (NTRS)

Mort, K. W.; Eckert, W. T.; Kelly, M. W.

1972-01-01

The structural cost of non-return wind tunnels is significantly less than that of the more conventional closed-circuit wind tunnels. However, because of the effects of external winds, the flow quality of non-return wind tunnels is an area of concern at the low test speeds required for V/STOL testing. The flow quality required at these low speeds is discussed and alternatives to the traditional manner of specifying the flow quality requirements in terms of dynamic pressure and angularity are suggested. The development of a non-return wind tunnel configuration which has good flow quality at low as well as at high test speeds is described.

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.

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.

Multi-angular Flame Measurements and Analysis in a Supersonic Wind Tunnel Using Fiber-Based Endoscopes

DTIC Science & Technology

2016-09-14

angular Flame Measurements and Analysis in a Supersonic Wind Tunnel Using Fiber-Based Endoscopes This paper reports new measurements and analysis made in...the Research Cell 19 super- sonic wind -tunnel facility housed at the Air Force Research Laboratory. The measure- ments include planar chemiluminescence...ASME for publication in the JOURNAL OF ENGINEERING FOR GAS TURBINES AND POWER. Manuscript received July 14, 2015; final manuscript received July 30

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

NASA Technical Reports Server (NTRS)

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

2001-01-01

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

Flight effects on noise generated by the JT8D-17 engine in a quiet nacelle and a conventional nacelle as measured in the NASA-Ames 40- by 80-foot wind tunnel

NASA Technical Reports Server (NTRS)

Strout, F. G.

1976-01-01

A JT8D-17 turbofan engine was tested in the NASA-Ames 40- by 80-foot wind tunnel to determine flight effects on jet and fan noise. Baseline, quiet nacelle with 20-lobe ejector/suppressor, and internal mixer configurations were tested over a range of engine power settings and tunnel velocities. Flight effects derived from the 40- by 80-foot wind tunnel test are compared with 727/JT8D flight test data and with model data obtained in a smaller wind tunnel. Procedures are defined for measuring noise data in a wind tunnel relatively near the sources and analyzing the results to obtain far-field flight effects. Wind tunnel and 727 flight test noise results compare favorably for both the baseline and quiet nacelle configurations. Two reports are provided, including a comprehensive version with extensive test results and analysis and the subject summary version that emphasizes data analysis and program finding.

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.

The Resistance of Spheres in Wind Tunnels and In Air

NASA Technical Reports Server (NTRS)

Bacon, D L; Reid, E G

1924-01-01

To supplement the standardization tests now in progress at several laboratories, a broad investigation of the resistance of spheres in wind tunnels and free air has been carried out by the National Advisory Committee for Aeronautics. The subject has been classed in aerodynamic research, and in consequence there is available a great mass of data from previous investigations. This material was given careful consideration in laying out the research, and explanation of practically all the disagreement between former experiments has resulted. A satisfactory confirmation of Reynolds law has been accomplished, the effect of means of support determined, the range of experiment greatly extended by work in the new variable density wind tunnel, and the effects of turbulence investigated by work in the tunnels and by towing and dropping tests in free air. It is concluded that the erratic nature of most of the previous work is due to support interference and differing turbulence conditions. While the question of support has been investigated thoroughly, a systematic and comprehensive study of the effects of scale and quality of turbulence will be necessary to complete the problem, as this phase was given only general treatment.

Development of a quiet supersonic wind tunnel with a cryogenic adaptive nozzle

NASA Technical Reports Server (NTRS)

Wolf, Stephen D.

1991-01-01

The main objectives of this work is to demonstrate the potential of a cryogenic adaptive nozzle to generate quiet (low disturbance) supersonic flow. A drive system was researched for the Fluid Mechanics Laboratory (FML) Laminar Flow Supersonic Wind Tunnel (LFSWT) using a pilot tunnel. A supportive effort for ongoing Proof of Concept (PoC) research leading to the design of critical components of the LFSWT was maintained. The state-of-the-art in quiet supersonic wind tunnel design was investigated. A supersonic research capability was developed within the FML.

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

NASA Technical Reports Server (NTRS)

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

2000-01-01

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

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.

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.

Around Marshall

NASA Image and Video Library

1984-01-01

An engineer at the Marshall Space Flight Center (MSFC) observes a model of the Space Shuttle Orbiter being tested in the MSFC's 14x14-Inch Trisonic Wind Tunnel. The 14-Inch Wind Tunnel is a trisonic wind tunnel. This means it is capable of running subsonic, below the speed of sound; transonic, at or near the speed of sound (Mach 1,760 miles per hour at sea level); or supersonic, greater than Mach 1 up to Mach 5. It is an intermittent blowdown tunnel that operates by high pressure air flowing from storage to either vacuum or atmospheric conditions. The MSFC 14x14-Inch Trisonic Wind Tunnel has been an integral part of the development of the United States space program Rocket and launch vehicles from the Jupiter-C in 1958, through the Saturn family up to the current Space Shuttle and beyond have been tested in this Wind Tunnel. MSFC's 14x14-Inch Trisonic Wind Tunnel, as with most other wind tunnels, is named after the size of the test section. The 14-Inch Wind Tunnel, as in the past, will continue to play a large but unseen role in the development of America's space program.

A laser-based ice shape profilometer for use in icing wind tunnels

NASA Technical Reports Server (NTRS)

Hovenac, Edward A.; Vargas, Mario

1995-01-01

A laser-based profilometer was developed to measure the thickness and shape of ice accretions on the leading edge of airfoils and other models in icing wind tunnels. The instrument is a hand held device that is connected to a desk top computer with a 10 meter cable. It projects a laser line onto an ice shape and used solid state cameras to detect the light scattered by the ice. The instrument corrects the image for camera angle distortions, displays an outline of the ice shape on the computer screen, saves the data on a disk, and can print a full scale drawing of the ice shape. The profilometer has undergone extensive testing in the laboratory and in the NASA Lewis Icing Research Tunnel. Results of the tests show very good agreement between profilometer measurements and known simulated ice shapes and fair agreement between profilometer measurements and hand tracing techniques.

De-icing of the altitude wind tunnel turning vanes by electro-magnetic impulse

NASA Technical Reports Server (NTRS)

Zumwalt, G. W.; Ross, R.

1986-01-01

The Altitude Wind Tunnel at the NASA-Lewis facility is being proposed for a refurbishment and moderization. Two major changes are: (1) the increasing of the test section Mach number to 0.90, and (2) the addition of spray nozzles to provide simulation of flight in icing clouds. Features to be retained are the simulation of atmospheric temperature and pressure to 50,000 foot altitude and provision for full-scale aircraft engine operation by the exhausting of the aircraft combustion gases and ingestion of air to replace that used in combustion. The first change required a re-design of the turning vanes in the two corners downstream of the test section due to the higher Mach number at the corners. The second change threatens the operation of the turning vanes by the expected ice build-up, particulary on the first-corner vanes. De-icing by heat has two drawbacks: (1) an extremely large amount of heat is required, and (2) the melted ice would tend to collect as ice on some other surfaces in the tunnel, namely, the tunnel propellers and the cooling coils. An alternate de-icing method had been under development for three years under NASA-Lewis grants to the Wichita State University. This report describes the electro-impulse de-icing (EIDI) method and the testing work done to assess its applicability to wind tunnel turning vane de-icing. Tests were conducted in the structural dynamics laboratory and in the NASA Icing Research Tunnel. Good ice protection was achieved at lower power consumption and at a wide range of tunnel operations conditions. Recommendations for design and construction of the system for this application of the EIDI method are given.

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

NASA Technical Reports Server (NTRS)

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

1984-01-01

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

Evaluation of New Technologies for Protection of Military Personnel from Filth and Biting Flies

DTIC Science & Technology

2007-10-01

Page 1. INTRODUCTION 3 2. RESEACH ACCOMPLISHMENTS FOR YEAR 3: 5 A. Students 5 B. Most Recent Research Results 6 I. Wind Tunnel for Testing ...were released per arena, and morbidity (knockdown) was recorded at 2-5 h post -treatment and mortality was recorded at 24, 48, and 72 h. Field Test ...Keuls test ) 42 Figure 3-1. Laboratory and field experimental design elements. A). Laboratory arena constructed of PVC pipe. Cords

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

NASA Technical Reports Server (NTRS)

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

1995-01-01

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

Aerospace Technology: Technical Data and Information on Foreign Test Facilities

DTIC Science & Technology

1990-06-22

effects of an airflow on various active models (nozzles or rotors ) or pas- sive models (airfoils). It is specially dedicated to acoustic testing driven by...Tunnel Figure V.3: Aerospatiale Rotor Test Bench and 99 Microphones Installed Inside Test Chamber of the CEPRA 19 Anechoic Wind Tunnel Figure V.4...Figure V.26: Ground Effect Test on Airbus A320 Model in 127 Test Section of the ONERA S1MA Wind Tunnel Figure V.27: ONERA S3Ch Transonic Wind Tunnel 130

Parachute Testing for Mars Science Laboratory

NASA Technical Reports Server (NTRS)

2007-01-01

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

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

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

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

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

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

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

NASA Technical Reports Server (NTRS)

Hoffman, Thomas R.

2001-01-01

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

40 CFR 53.60 - General provisions.

Code of Federal Regulations, 2013 CFR

2013-07-01

... shall not be subject to the requirements of § 53.62 (full wind tunnel test), provided that it meets all requirements of § 53.63 (wind tunnel inlet aspiration test), § 53.65 (loading test), and § 53.66 (volatility... samplers shall not be subject to the requirements of § 53.62 (full wind tunnel test), provided that it...

40 CFR 53.60 - General provisions.

Code of Federal Regulations, 2012 CFR

2012-07-01

... shall not be subject to the requirements of § 53.62 (full wind tunnel test), provided that it meets all requirements of § 53.63 (wind tunnel inlet aspiration test), § 53.65 (loading test), and § 53.66 (volatility... samplers shall not be subject to the requirements of § 53.62 (full wind tunnel test), provided that it...

40 CFR 53.60 - General provisions.

Code of Federal Regulations, 2014 CFR

2014-07-01

... shall not be subject to the requirements of § 53.62 (full wind tunnel test), provided that it meets all requirements of § 53.63 (wind tunnel inlet aspiration test), § 53.65 (loading test), and § 53.66 (volatility... samplers shall not be subject to the requirements of § 53.62 (full wind tunnel test), provided that it...

Computation of wind tunnel wall effects for complex models using a low-order panel method

NASA Technical Reports Server (NTRS)

Ashby, Dale L.; Harris, Scott H.

1994-01-01

A technique for determining wind tunnel wall effects for complex models using the low-order, three dimensional panel method PMARC (Panel Method Ames Research Center) has been developed. Initial validation of the technique was performed using lift-coefficient data in the linear lift range from tests of a large-scale STOVL fighter model in the National Full-Scale Aerodynamics Complex (NFAC) facility. The data from these tests served as an ideal database for validating the technique because the same model was tested in two wind tunnel test sections with widely different dimensions. The lift-coefficient data obtained for the same model configuration in the two test sections were different, indicating a significant influence of the presence of the tunnel walls and mounting hardware on the lift coefficient in at least one of the two test sections. The wind tunnel wall effects were computed using PMARC and then subtracted from the measured data to yield corrected lift-coefficient versus angle-of-attack curves. The corrected lift-coefficient curves from the two wind tunnel test sections matched very well. Detailed pressure distributions computed by PMARC on the wing lower surface helped identify the source of large strut interference effects in one of the wind tunnel test sections. Extension of the technique to analysis of wind tunnel wall effects on the lift coefficient in the nonlinear lift range and on drag coefficient will require the addition of boundary-layer and separated-flow models to PMARC.

Pressure distribution on the roof of a model low-rise building tested in a boundary layer wind tunnel

NASA Astrophysics Data System (ADS)

Goliber, Matthew Robert

With three of the largest metropolitan areas in the United States along the Gulf coast (Houston, Tampa, and New Orleans), residential populations ever increasing due to the subtropical climate, and insured land value along the coast from Texas to the Florida panhandle greater than $500 billion, hurricane related knowledge is as important now as ever before. This thesis focuses on model low-rise building wind tunnel tests done in connection with full-scale low-rise building tests. Mainly, pressure data collection equipment and methods used in the wind tunnel are compared to pressure data collection equipment and methods used in the field. Although the focus of this report is on the testing of models in the wind tunnel, the low-rise building in the field is located in Pensacola, Florida. It has a wall length of 48 feet, a width of 32 feet, a height of 10 feet, and a gable roof with a pitch of 1:3 and 68 pressure ports strategically placed on the surface of the roof. Built by Forest Products Laboratory (FPL) in 2002, the importance of the test structure has been realized as it has been subjected to numerous hurricanes. In fact, the validity of the field data is so important that the following thesis was necessary. The first model tested in the Bill James Wind Tunnel for this research was a rectangular box. It was through the testing of this box that much of the basic wind tunnel and pressure data collection knowledge was gathered. Knowledge gained from Model 1 tests was as basic as how to: mount pressure tubes on a model, use a pressure transducer, operate the wind tunnel, utilize the pitot tube and reference pressure, and measure wind velocity. Model 1 tests also showed the importance of precise construction to produce precise pressure coefficients. Model 2 was tested in the AABL Wind Tunnel at Iowa State University. This second model was a 22 inch cube which contained a total of 11 rows of pressure ports on its front and top faces. The purpose of Model 2 was to validate the tube length, tube diameter, port diameter, and pressure transducer used in the field. Also, Model 2 was used to study the effects of surface roughness on pressure readings. A partial roof and wall of the low-rise building in the field was used as the third model. Similar to the second model, Model 3 was tested in the AABL Wind Tunnel. Initially, the objectives of the third model were to validate the pressure port protection device (PPPD) being used in the field and test the possibility of interpolating between pressure ports. But in the end, Model 3 was best used to validate the inconsistencies of the full-scale PPPD, validate the transducers used in the field, and prove the importance of scaling either all or none of the model. Fourthly, Model 4 was a 1:16 model of the low-rise building itself. Based on the three previous model tests, Model 4 was instrumented with 202 pressure transducers to better understand: (1) the pressure distribution on the roof of the structure, (2) the affects of the fundamental test variables such as tube length, tube diameter, port diameter, transducer type, and surface roughness, (3) the affects of a scaled PPPD, (4) the importance of wind angle of attack, and (5) the possibility of measuring pressure data and load data simultaneously. In the end, the combination of all four model tests proved to be helpful in understanding the pressure data gathered on the roof of the low-rise building in the field. The two main recommendations for the field structure are for reevaluation of the PPPD design and slight redistribution of the pressure ports. The wind tunnel model tests show a need for these two modifications in order to gather more accurate field pressure data. Other than these two adjustments, the model tests show that the remaining data gathering system is currently accurate.

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

NASA Technical Reports Server (NTRS)

Shinoda, Patrick M.

1996-01-01

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

Around Marshall

NASA Image and Video Library

1988-01-01

This photograph shows an overall view of the Marshall Space Flight Center's (MSFC's) 14x14-Inch Trisonic Wind Tunnel. The 14-Inch Wind Tunnel is a trisonic wind tunnel. This means it is capable of running subsonic, below the speed of sound; transonic, at or near the speed of sound (Mach 1, 760 miles per hour at sea level); or supersonic, greater than Mach 1 up to Mach 5. It is an intermittent blowdown tunnel that operates by high pressure air flowing from storage to either vacuum or atmospheric conditions. The MSFC 14x14-Inch Trisonic Wind Tunnel has been an integral part of the development of the United States space program Rocket and launch vehicles from the Jupiter-C in 1958, through the Saturn family up to the current Space Shuttle and beyond have been tested in this Wind Tunnel. MSFC's 14x14-Inch Trisonic Wind Tunnel, as with most other wind tunnels, is named after the size of the test section. The 14-Inch Wind Tunnel, as in the past, will continue to play a large but unseen role in the development of America's space program.

Altitude Wind Tunnel at the NACA’s Aircraft Engine Research Laboratory

NASA Image and Video Library

1945-06-21

Two men on top of the Altitude Wind Tunnel (AWT) at the National Advisory Committee for Aeronautics (NACA) Aircraft Engine Research Laboratory. The tunnel was a massive rectangular structure, which for years provided one of the highest vantage points on the laboratory. The tunnel was 263 feet long on the north and south legs and 121 feet long on the east and west sides. The larger west end of the tunnel, seen here, was 51 feet in diameter. The east side of the tunnel was 31 feet in diameter at the southeast corner and 27 feet in diameter at the northeast. The throat section, which connected the northwest corner to the test section, narrowed sharply from 51 to 20 feet in diameter. The AWT’s altitude simulation required temperature and pressure fluctuations that made the design of the shell more difficult than other tunnels. The simultaneous decrease in both pressure and temperature inside the facility produced uneven stress loads, particularly on the support rings. The steel used in the primary tunnel structure was one inch thick to ensure that the shell did not collapse as the internal air pressure was dropped to simulate high altitudes. It was a massive amount of steel considering the World War II shortages. The shell was covered with several inches of fiberglass insulation to retain the refrigerated air and a thinner outer steel layer to protect the insulation against the weather. A unique system of rollers was used between the shell and its support piers. These rollers allowed for movement as the shell expanded or contracted during the altitude simulations. Certain sections would move as much as five inches during operation.

Investigation of the General Electric I-40 Jet-Propulsion Engine in the Cleveland Altitude Wind Tunnel. 2 - Analysis of Compressor Performance Characteristics

DTIC Science & Technology

1946-11-18

INVESTIGATION OF THE GENERAL ELECTRIC 1-40 JET -PROPULSION ENGINE IN THE CLEVELAND ALTITUDE WIND TUNNEL .; II - ANALYSIS OF COMPRESSOR PERFORMANCE...CHARACTERISTICS By Robert 0. Dietz, Jr. and Robert M. Gelsenheyner Aircraft Engine Research Laboratory 1 Cleveland, Ohio !f -NOT FM ED", P 0 W DESTROY...Command, Army Air Forces INVESTIGATION OF THE GENERAL ELECTRIC 1-40 JET -PROPULSION ENGINE IN THE CLEVELAND ALTITUDE WIND TUNNEL II - ANALYSIS OF

Mars Parachute Testing in World Largest Wind Tunnel

NASA Image and Video Library

2009-04-22

The parachute for NASA next mission to Mars passed flight-qualification testing in March and April 2009 inside the world largest wind tunnel, at NASA Ames Research Center, Moffett Field, Calif. NASA's Mars Science Laboratory mission, to be launched in 2011 and land on Mars in 2012, will use the largest parachute ever built to fly on an extraterrestrial mission. This image shows a duplicate qualification-test parachute inflated in an 80-mile-per-hour (36-meter-per-second) wind inside the test facility. The parachute uses a configuration called disk-gap-band. It has 80 suspension lines, measures more than 50 meters (165 feet) in length, and opens to a diameter of nearly 16 meters (51 feet). Most of the orange and white fabric is nylon, though a small disk of heavier polyester is used near the vent in the apex of the canopy due to higher stresses there. It is designed to survive deployment at Mach 2.2 in the Martian atmosphere, where it will generate up to 65,000 pounds of drag force. The wind tunnel is 24 meters (80 feet) tall and 37 meters (120 feet) wide, big enough to house a Boeing 737. It is part of the National Full-Scale Aerodynamics Complex, operated by the Arnold Engineering Development Center of the U.S. Air Force. http://photojournal.jpl.nasa.gov/catalog/PIA11995

Rolls Royce Avon RA-14 Engine in the Altitude Wind Tunnel

NASA Image and Video Library

1956-03-21

A Rolls Royce Avon RA-14 engine was tested in the Altitude Wind Tunnel at the National Advisory Committee for Aeronautics’ (NACA) Lewis Flight Propulsion Laboratory. The Avon RA-14 engine was a 16-stage axial-flow compressor turbojet capable of producing 9,500 pounds of thrust. The Avon replaced Rolls Royce’s successful Nene engine in 1950 and remained in service until 1974. It was one of several British engines studied in the tunnel during the 1950s. The Altitude Wind Tunnel went through a series of modifications in 1951 to increase its capabilities. An annex was attached to the Exhauster Building to house three new Ingersoll-Rand compressors. The wooden blades on the tunnel’s 31-foot diameter fan were replaced, a pump house and exhaust cooler were constructed underneath the tunnel, and two new cells were added to the cooling tower. The modified wind tunnel continued to analyze jet engines in the 1950s, although the engines, like the RA-14 seen here, were much more powerful than those studied several years before. Lewis researchers studied the RA-14 turbojet engine in the Altitude Wind Tunnel for 11 months in 1956. The engine was mounted on a stand capable of gauging engine thrust, and the tunnel’s air was ducted to the engine through a venturi and bellmouth inlet, seen in this photograph. The initial studies established the engine’s performance characteristics with a fixed-area nozzle and its acceleration characteristics. The researchers also used the tunnel to investigate windmilling of the compressor blades, restarting at high altitudes, and the engine’s performance limits at altitude.

Prospects for Nonlinear Laser Diagnostics in the Jet Noise Laboratory

NASA Technical Reports Server (NTRS)

Herring, Gregory C.; Hart, Roger C.; Fletcher, mark T.; Balla, R. Jeffrey; Henderson, Brenda S.

2007-01-01

Two experiments were conducted to test whether optical methods, which rely on laser beam coherence, would be viable for off-body flow measurement in high-density, compressible-flow wind tunnels. These tests measured the effects of large, unsteady density gradients on laser diagnostics like laser-induced thermal acoustics (LITA). The first test was performed in the Low Speed Aeroacoustics Wind Tunnel (LSAWT) of NASA Langley Research Center's Jet Noise Laboratory (JNL). This flow facility consists of a dual-stream jet engine simulator (with electric heat and propane burners) exhausting into a simulated flight stream, reaching Mach numbers up to 0.32. A laser beam transited the LSAWT flow field and was imaged with a high-speed gated camera to measure beam steering and transverse mode distortion. A second, independent test was performed on a smaller laboratory jet (Mach number < 1.2 and mass flow rate < 0.1 kg/sec). In this test, time-averaged LITA velocimetry and thermometry were performed at the jet exit plane, where the effect of unsteady density gradients is observed on the LITA signal. Both experiments show that LITA (and other diagnostics relying on beam overlap or coherence) faces significant hurdles in the high-density, compressible, and turbulent flow environments similar to those of the JNL.

Ceramic and coating applications in the hostile environment of a high temperature hypersonic wind tunnel. [Langley 8-foot high temperature structures tunnel

NASA Technical Reports Server (NTRS)

Puster, R. L.; Karns, J. R.; Vasquez, P.; Kelliher, W. C.

1981-01-01

A Mach 7, blowdown wind tunnel was used to investigate aerothermal structural phenomena on large to full scale high speed vehicle components. The high energy test medium, which provided a true temperature simulation of hypersonic flow at 24 to 40 km altitude, was generated by the combustion of methane with air at high pressures. Since the wind tunnel, as well as the models, must be protected from thermally induced damage, ceramics and coatings were used extensively. Coatings were used both to protect various wind tunnel components and to improve the quality of the test stream. Planned modifications for the wind tunnel included more extensive use of ceramics in order to minimize the number of active cooling systems and thus minimize the inherent operational unreliability and cost that accompanies such systems. Use of nonintrusive data acquisition techniques, such as infrared radiometry, allowed more widespread use of ceramics for models to be tested in high energy wind tunnels.

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

NASA Astrophysics Data System (ADS)

Huang, Yifeng; Yang, Jixin

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

Space Launch System Ascent Static Aerodynamic Database Development

NASA Technical Reports Server (NTRS)

Pinier, Jeremy T.; Bennett, David W.; Blevins, John A.; Erickson, Gary E.; Favaregh, Noah M.; Houlden, Heather P.; Tomek, William G.

2014-01-01

This paper describes the wind tunnel testing work and data analysis required to characterize the static aerodynamic environment of NASA's Space Launch System (SLS) ascent portion of flight. Scaled models of the SLS have been tested in transonic and supersonic wind tunnels to gather the high fidelity data that is used to build aerodynamic databases. A detailed description of the wind tunnel test that was conducted to produce the latest version of the database is presented, and a representative set of aerodynamic data is shown. The wind tunnel data quality remains very high, however some concerns with wall interference effects through transonic Mach numbers are also discussed. Post-processing and analysis of the wind tunnel dataset are crucial for the development of a formal ascent aerodynamics database.

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.

Air Force Officers Visit Aircraft Engine Research Laboratory

NASA Image and Video Library

1945-08-21

A group of 60 Army Air Forces officers visited the National Advisory Committee for Aeronautics (NACA) Aircraft Engine Research Laboratory on August 27, 1945. The laboratory enacted strict security regulations throughout World War II. During the final months of the war, however, the NACA began opening its doors to groups of writers, servicemen, and aviation industry leaders. These events were the first exposure of the new engine laboratory to the outside world. Grandstands were built alongside the Altitude Wind Tunnel specifically for group photographs. George Lewis, Raymond Sharp, and Addison Rothrock (right to left) addressed this group of officers in the Administration Building auditorium. Lewis was the NACA’s Director of Aeronautical Research, Sharp was the lab’s manager, and Rothrock was the lab’s chief of research. Abe Silverstein, Jesse Hall and others watch from the rear of the room. The group toured several facilities after the talks, including the Altitude Wind Tunnel and a new small supersonic wind tunnel. The visit concluded with a NACA versus Army baseball game and cookout.

Build an Inexpensive Wind Tunnel to Test CO2 Cars

ERIC Educational Resources Information Center

McCormick, Kevin

2012-01-01

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

Calibration of the Flow in the Test Section of the Research Wind Tunnel at DST Group

DTIC Science & Technology

2015-10-01

calibration of the flow in the test section of the Research Wind Tunnel at DST Group. The calibration was performed to establish the flow quality and to...of the Flow in the Test Section of the Research Wind Tunnel at DST Group Executive Summary The Defence Science and Technology Group (DST

Development of 3D Ice Accretion Measurement Method

NASA Technical Reports Server (NTRS)

Lee, Sam; Broeren, Andy P.; Addy, Harold E., Jr.; Sills, Robert; Pifer, Ellen M.

2012-01-01

Icing wind tunnels are designed to simulate in-flight icing environments. The chief product of such facilities is the ice accretion that forms on various test articles. Documentation of the resulting ice accretion key piece of data in icing-wind-tunnel tests. Number of currently used options for documenting ice accretion in icing-wind-tunnel testing.

Python Turboprop Prepared for a Test in the Altitude Wind Tunnel

NASA Image and Video Library

1949-08-21

A 3670-horsepower Armstrong-Siddeley Python turboprop being prepared for tests in the Altitude Wind Tunnel at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory. In 1947 Lewis researcher Walter Olsen led a group of representatives from the military, industry, and the NACA on a fact finding mission to investigate the technological progress of British turbojet manufacturers. Afterwards several British engines, including the Python, were brought to Cleveland for testing in Lewis’s altitude facilities. The Python was a 14-stage axial-flow compressor turboprop with a fixed-area nozzle and contra-rotating propellers. Early turboprops combined the turbojet and piston engine technologies. They could move large quantities of air so required less engine speed and thus less fuel. This was very appealing to the military for some applications. The military asked the NACA to compare the Python’s performance at sea to that at high altitudes. The NACA researchers studied the Python in the Altitude Wind Tunnel from July 1949 through January 1950. It was the first time the tunnel was used to study an engine with the sole purpose of learning about, not improving it. They analyzed the engine’s dynamic response using a frequency response method at altitudes between 10,000 to 30,000 feet. Lewis researchers found that they could predict the dynamic response characteristics at any altitude from the data obtained from any other specific altitude. This portion of the testing was completed during a single test run.

A Cold Temperature Evaluation of the Bonding Adhesives Used for the MUST Inflatable Shelters

DTIC Science & Technology

1975-04-01

been known as the US Army Natick Laboratories (NLABS), and the US Army Natick Development Center ( NDC ). AMIN 1W 5Th Whit# WIN~ zD DI• ~kti u...section throughout the test period. The arctic climatic chamber is a closed ci;•uit wind tunnel in which the temperature can be lowered, controlled...and maintained at any level down to -54WC (-65°F). The shelter section was too large to fit in the test section of the tunnel . The shelter had to be

Inspection of the New 10- by 10-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1956-05-21

Attendees listen during the May 22, 1956 Inspection of the new 10- by 10-Foot Supersonic Wind Tunnel at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory. The facility, known at the time as the Lewis Unitary Plan Tunnel, was in its initial stages of operation. The $33 million 10- by 10 was the most powerful wind tunnel in the nation. Over 150 guests from industry, other NACA laboratories, and the media attended the event. The speakers, from left to right in the front row, addressed the crowd before the tour. Lewis Director Raymond Sharp began the event by welcoming the visitors to the laboratory. NACA Director Hugh Dryden discussed Congress’ Unitary Plan Act and its effect on the creation of the facility. Lewis Associate Director Abe Silverstein discussed the need for research tools and the 10- by 10’s place among the NACA’s other research facilities. Lewis Assistant Director Eugene Wasielewski described the detailed design work that went into the facility. Carl Schueller, Chief of the 10- by 10, described the tunnel’s components and how the facility operated. Robert Godman led the tour afterwards. The 10- by 10 can test engines up to five feet in diameter at supersonic speeds and simulated altitudes of 30 miles. Its main purpose is to investigate problems relating to engine inlet and outlet geometry, engine matching and interference effects, and overall drag. The tunnel’s 250,000-horsepower electric motor drive, the most powerful of its kind in the world, creates air speeds between Mach 2.0 and 3.5.

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

NASA Astrophysics Data System (ADS)

Petronevich, V. V.

2016-10-01

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

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.

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

NASA Technical Reports Server (NTRS)

Ulbrich, N.

2016-01-01

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

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

NASA Technical Reports Server (NTRS)

DeLoach, Richard; Micol, John R.

2011-01-01

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

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

NASA Technical Reports Server (NTRS)

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

2005-01-01

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

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...

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.

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

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

21. Photocopy of drawing (original in possession of Naval Surface Warfare Center Carderock Division, Bethesda, MD) 7 X 10 FOOT SONIC WIND TUNNEL, LABORATORY AND OFFICE BUILDING, 1950 - Naval Surface Warfare Center, Transonic Wind Tunnel Building, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD

A simplified method for calculating temperature time histories in cryogenic wind tunnels

NASA Technical Reports Server (NTRS)

Stallings, R. L., Jr.; Lamb, M.

1976-01-01

Average temperature time history calculations of the test media and tunnel walls for cryogenic wind tunnels have been developed. Results are in general agreement with limited preliminary experimental measurements obtained in a 13.5-inch pilot cryogenic wind tunnel.

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

Quantification of the Uncertainties for the Ares I A106 Ascent Aerodynamic Database

NASA Technical Reports Server (NTRS)

Houlden, Heather P.; Favaregh, Amber L.

2010-01-01

A detailed description of the quantification of uncertainties for the Ares I ascent aero 6-DOF wind tunnel database is presented. The database was constructed from wind tunnel test data and CFD results. The experimental data came from tests conducted in the Boeing Polysonic Wind Tunnel in St. Louis and the Unitary Plan Wind Tunnel at NASA Langley Research Center. The major sources of error for this database were: experimental error (repeatability), database modeling errors, and database interpolation errors.

Measurement of Vibrations from the 8- by 6-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1950-07-21

Reverend Henry Birkenhauer and E.F. Carome measure ground vibrations on West 220th Street caused by the operation of the 8- by 6-Foot Supersonic Wind Tunnel at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory. The 8- by 6 was the laboratory’s first large supersonic wind tunnel. It was also the NACA’s most powerful supersonic tunnel, and the NACA’s first facility capable of running an engine at supersonic speeds. The 8- by 6 was originally an open-throat and non-return tunnel. This meant that the supersonic air flow was blown through the test section and out the other end into the atmosphere. Complaints from the local community led to the installation of a muffler at the tunnel exit and the eventual addition of a return leg. Reverend Brikenhauer, a seismologist, and Carome, an electrical technician were brought in from John Carroll University to take vibration measurements during the 8- by 6 tunnel’s first run with a supersonic engine. They found that the majority of the vibrations came from the air and not the ground. The tunnel’s original muffler offered some relief during the facility checkout runs, but it proved inadequate during the operation of an engine in the test section. Tunnel operation was suspended until a new muffler was designed and installed. The NACA researchers, however, were pleased with the tunnel’s operation. They claimed it was the first time a jet engine was operated in an airflow faster than Mach 2.

Portable Test And Monitoring System For Wind-Tunnel Models

NASA Technical Reports Server (NTRS)

Poupard, Charles A.

1987-01-01

Portable system developed to test and monitor instrumentation used in wind-tunnel models. Self-contained and moves easily to model, either before or after model installed in wind tunnel. System is 44 1/2 in. high, 22 in. wide, and 17 in. deep and weighs 100 lb. Primary benefits realized with portable test and monitoring system associated with saving of time.

Analytical Models for Rotor Test Module, Strut, and Balance Frame Dynamics in the 40 by 80 Ft Wind Tunnel

NASA Technical Reports Server (NTRS)

Johnson, W.

1976-01-01

A mathematical model is developed for the dynamics of a wind tunnel support system consisting of a balance frame, struts, and an aircraft or test module. Data are given for several rotor test modules in the Ames 40 by 80 ft wind tunnel. A model for ground resonance calculations is also described.

Analysis of subsonic wind tunnel with variation shape rectangular and octagonal on test section

NASA Astrophysics Data System (ADS)

Rhakasywi, D.; Ismail; Suwandi, A.; Fadhli, A.

2018-02-01

The need for good design in the aerodynamics field required a wind tunnel design. The wind tunnel design required in this case is capable of generating laminar flow. In this research searched for wind tunnel models with rectangular and octagonal variations with objectives to generate laminar flow in the test section. The research method used numerical approach of CFD (Computational Fluid Dynamics) and manual analysis to analyze internal flow in test section. By CFD simulation results and manual analysis to generate laminar flow in the test section is a design that has an octagonal shape without filled for optimal design.

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

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

NASA Technical Reports Server (NTRS)

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

1973-01-01

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

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.

Strain actuated aeroelastic control

NASA Technical Reports Server (NTRS)

Lazarus, Kenneth B.

1992-01-01

Viewgraphs on strain actuated aeroelastic control are presented. Topics covered include: structural and aerodynamic modeling; control law design methodology; system block diagram; adaptive wing test article; bench-top experiments; bench-top disturbance rejection: open and closed loop response; bench-top disturbance rejection: state cost versus control cost; wind tunnel experiments; wind tunnel gust alleviation: open and closed loop response at 60 mph; wind tunnel gust alleviation: state cost versus control cost at 60 mph; wind tunnel command following: open and closed loop error at 60 mph; wind tunnel flutter suppression: open loop flutter speed; and wind tunnel flutter suppression: closed loop state cost curves.

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.

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.

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.

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.

Design of experiments enhanced statistical process control for wind tunnel check standard testing

NASA Astrophysics Data System (ADS)

Phillips, Ben D.

The current wind tunnel check standard testing program at NASA Langley Research Center is focused on increasing data quality, uncertainty quantification and overall control and improvement of wind tunnel measurement processes. The statistical process control (SPC) methodology employed in the check standard testing program allows for the tracking of variations in measurements over time as well as an overall assessment of facility health. While the SPC approach can and does provide researchers with valuable information, it has certain limitations in the areas of process improvement and uncertainty quantification. It is thought by utilizing design of experiments methodology in conjunction with the current SPC practices that one can efficiently and more robustly characterize uncertainties and develop enhanced process improvement procedures. In this research, methodologies were developed to generate regression models for wind tunnel calibration coefficients, balance force coefficients and wind tunnel flow angularities. The coefficients of these regression models were then tracked in statistical process control charts, giving a higher level of understanding of the processes. The methodology outlined is sufficiently generic such that this research can be applicable to any wind tunnel check standard testing program.

Compiling the space shuttle wind tunnel data base: An exercise in technical and managerial innovators

NASA Technical Reports Server (NTRS)

Kemp, N. D.

1983-01-01

Engineers evaluating Space Shuttle flight data and performance results are using a massive data base of wind tunnel test data. A wind tunnel test data base of the magnitude attained is a major accomplishment. The Apollo program spawned an automated wind tunnel data analysis system called SADSAC developed by the Chrysler Space Division. An improved version of this system renamed DATAMAN was used by Chrysler to document analyzed wind tunnel data and data bank the test data in standardized formats. These analysis documents, associated computer graphics and standard formatted data were disseminated nationwide to the Shuttle technical community. These outputs became the basis for substantiating and certifying the flight worthiness of the Space Shuttle and for improving future designs. As an aid to future programs this paper documents the lessons learned in compiling the massive wind tunnel test data base for developing the Space Shuttle. In particular, innovative managerial and technical concepts evolved in the course of conceiving and developing this successful DATAMAN system and the methods and organization for applying the system are presented.

Airloads Correlation of the UH-60A Rotor Inside the 40- by 80-Foot Wind Tunnel

NASA Technical Reports Server (NTRS)

Chang, I-Chung; Norman, Thomas R.; Romander, Ethan A.

2013-01-01

The presented research validates the capability of a loosely-coupled computational fluid dynamics (CFD) and comprehensive rotorcraft analysis (CRA) code to calculate the flowfield around a rotor and test stand mounted inside a wind tunnel. The CFD/CRA predictions for the full-scale UH-60A Airloads Rotor inside the National Full-Scale Aerodynamics Complex (NFAC) 40- by 80-Foot Wind Tunnel at NASA Ames Research Center are compared with the latest measured airloads and performance data. The studied conditions include a speed sweep at constant lift up to an advance ratio of 0.4 and a thrust sweep at constant speed up to and including stall. For the speed sweep, wind tunnel modeling becomes important at advance ratios greater than 0.37 and test stand modeling becomes increasingly important as the advance ratio increases. For the thrust sweep, both the wind tunnel and test stand modeling become important as the rotor approaches stall. Despite the beneficial effects of modeling the wind tunnel and test stand, the new models do not completely resolve the current airload discrepancies between prediction and experiment.

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.

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.

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

NASA Technical Reports Server (NTRS)

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

1974-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1990-01-01

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

A wall interference assessment/correction system

NASA Technical Reports Server (NTRS)

Lo, Ching F.; Ulbrich, N.; Sickles, W. L.; Qian, Cathy X.

1992-01-01

A Wall Signature method, the Hackett method, has been selected to be adapted for the 12-ft Wind Tunnel wall interference assessment/correction (WIAC) system in the present phase. This method uses limited measurements of the static pressure at the wall, in conjunction with the solid wall boundary condition, to determine the strength and distribution of singularities representing the test article. The singularities are used in turn for estimating wall interferences at the model location. The Wall Signature method will be formulated for application to the unique geometry of the 12-ft Tunnel. The development and implementation of a working prototype will be completed, delivered and documented with a software manual. The WIAC code will be validated by conducting numerically simulated experiments rather than actual wind tunnel experiments. The simulations will be used to generate both free-air and confined wind-tunnel flow fields for each of the test articles over a range of test configurations. Specifically, the pressure signature at the test section wall will be computed for the tunnel case to provide the simulated 'measured' data. These data will serve as the input for the WIAC method-Wall Signature method. The performance of the WIAC method then may be evaluated by comparing the corrected parameters with those for the free-air simulation. Each set of wind tunnel/test article numerical simulations provides data to validate the WIAC method. A numerical wind tunnel test simulation is initiated to validate the WIAC methods developed in the project. In the present reported period, the blockage correction has been developed and implemented for a rectangular tunnel as well as the 12-ft Pressure Tunnel. An improved wall interference assessment and correction method for three-dimensional wind tunnel testing is presented in the appendix.

Ramjet Model and Technicians in the 8- by 6-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1952-02-21

A researcher at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory checks the setup of a RJM-2 ramjet model in the test section of the 8- by 6-Foot Supersonic Wind Tunnel. The 8- by 6 was not only the laboratory’s first large supersonic wind tunnel, but it was also the NACA’s first facility capable of testing an operating engine at supersonic speeds. The 8- by 6-foot tunnel has been used to study engine inlets, fuel injectors, flameholders, exit nozzles, and controls on ramjet and turbojet propulsion systems. The 8-foot wide and 6-foot tall test section consisted of 1-inch thick steel plates with hatches on the floor and ceiling to facilitate the installation of the test article. The two windows seen on the right wall allowed photographic equipment to be set up. The test section was modified in 1956 to accommodate transonic research. NACA engineers drilled 4,700 holes into the test section walls to reduce transonic pressure disturbances and shock waves. NACA Lewis undertook an extensive research program on ramjets in the 1940s using several of its facilities. Ramjets provide a very simple source of propulsion. They are basically a tube which ingests high speed air, ignites it, and then expels the heated air at a significantly higher velocity. Ramjets are extremely efficient and powerful but can only operate at high speeds. Therefore, they require a booster rocket or aircraft drop to accelerate them to high speeds before they can operate.

Design, construction and calibration of a portable boundary layer wind tunnel for field use

USDA-ARS?s Scientific Manuscript database

Wind tunnels have been used for several decades to study wind erosion processes. Portable wind tunnels offer the advantage of testing natural surfaces in the field, but they must be carefully designed to insure that a logarithmic boundary layer is formed and that wind erosion processes may develop ...

Development and Operation of an Automatic Rotor Trim Control System for use During the UH-60 Individual Blade Control Wind Tunnel Test

NASA Technical Reports Server (NTRS)

Theodore, Colin R.

2010-01-01

A full-scale wind tunnel test to evaluate the effects of Individual Blade Control (IBC) on the performance, vibration, noise and loads of a UH-60A rotor was recently completed in the National Full-Scale Aerodynamics Complex (NFAC) 40- by 80-Foot Wind Tunnel [1]. A key component of this wind tunnel test was an automatic rotor trim control system that allowed the rotor trim state to be set more precisely, quickly and repeatably than was possible with the rotor operator setting the trim condition manually. The trim control system was also able to maintain the desired trim condition through changes in IBC actuation both in open- and closed-loop IBC modes, and through long-period transients in wind tunnel flow. This ability of the trim control system to automatically set and maintain a steady rotor trim enabled the effects of different IBC inputs to be compared at common trim conditions and to perform these tests quickly without requiring the rotor operator to re-trim the rotor. The trim control system described in this paper was developed specifically for use during the IBC wind tunnel test

Test techniques: A survey paper on cryogenic tunnels, adaptive wall test sections, and magnetic suspension and balance systems

NASA Technical Reports Server (NTRS)

Kilgore, Robert A.; Dress, David A.; Wolf, Stephen W. D.; Britcher, Colin P.

1989-01-01

The ability to get good experimental data in wind tunnels is often compromised by things seemingly beyond our control. Inadequate Reynolds number, wall interference, and support interference are three of the major problems in wind tunnel testing. Techniques for solving these problems are available. Cryogenic wind tunnels solve the problem of low Reynolds number. Adaptive wall test sections can go a long way toward eliminating wall interference. A magnetic suspension and balance system (MSBS) completely eliminates support interference. Cryogenic tunnels, adaptive wall test sections, and MSBS are surveyed. A brief historical overview is given and the present state of development and application in each area is described.

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.

Large-Scale Wind Turbine Testing in the NASA 24.4m (80) by 36.6m(120) Wind Tunnel

NASA Technical Reports Server (NTRS)

Zell, Peter T.; Imprexia, Cliff (Technical Monitor)

2000-01-01

The 80- by 120-Foot Wind Tunnel at NASA Ames Research Center in California provides a unique capability to test large-scale wind turbines under controlled conditions. This special capability is now available for domestic and foreign entities wishing to test large-scale wind turbines. The presentation will focus on facility capabilities to perform wind turbine tests and typical research objectives for this type of testing.

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.

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.

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

NASA Technical Reports Server (NTRS)

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

2007-01-01

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

Comparison of acoustic data from a 102 mm conic nozzle as measured in the RAE 24-foot wind tunnel and the NASA Ames 40- by 80-foot wind tunnel

NASA Technical Reports Server (NTRS)

Atencio, A., Jr.; Mckie, J.

1982-01-01

A cooperative program between the Royal Aircraft Establishment (RAE), England, and the NASA Ames Research Center was initiated to compare acoustic measurements made in the RAE 24-foot wind tunnel and in the Ames 40- by 80-foot wind tunnel. The acoustic measurements were made in both facilities using the same 102 mm conical nozzle supplied by the RAE. The nozzle was tested by each organization using its respective jet test rig. The mounting hardware and nozzle exit conditions were matched as closely as possible. The data from each wind tunnel were independently analyzed by the respective organization. The results from these tests show good agreement. In both facilities, interference with acoustic measurement is evident at angles in the forward quadrant.

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...

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

NASA Technical Reports Server (NTRS)

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

1977-01-01

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

Experimental Study of Homogeneous Isotropic Slowly-Decaying Turbulence in Giant Grid-Wind Tunnel Set Up

NASA Astrophysics Data System (ADS)

Aliseda, Alberto; Bourgoin, Mickael; Eswirp Collaboration

2014-11-01

We present preliminary results from a recent grid turbulence experiment conducted at the ONERA wind tunnel in Modane, France. The ESWIRP Collaboration was conceived to probe the smallest scales of a canonical turbulent flow with very high Reynolds numbers. To achieve this, the largest scales of the turbulence need to be extremely big so that, even with the large separation of scales, the smallest scales would be well above the spatial and temporal resolution of the instruments. The ONERA wind tunnel in Modane (8 m -diameter test section) was chosen as a limit of the biggest large scales achievable in a laboratory setting. A giant inflatable grid (M = 0.8 m) was conceived to induce slowly-decaying homogeneous isotropic turbulence in a large region of the test section, with minimal structural risk. An international team or researchers collected hot wire anemometry, ultrasound anemometry, resonant cantilever anemometry, fast pitot tube anemometry, cold wire thermometry and high-speed particle tracking data of this canonical turbulent flow. While analysis of this large database, which will become publicly available over the next 2 years, has only started, the Taylor-scale Reynolds number is estimated to be between 400 and 800, with Kolmogorov scales as large as a few mm . The ESWIRP Collaboration is formed by an international team of scientists to investigate experimentally the smallest scales of turbulence. It was funded by the European Union to take advantage of the largest wind tunnel in Europe for fundamental research.

Pose measurement method and experiments for high-speed rolling targets in a wind tunnel.

PubMed

Jia, Zhenyuan; Ma, Xin; Liu, Wei; Lu, Wenbo; Li, Xiao; Chen, Ling; Wang, Zhengqu; Cui, Xiaochun

2014-12-12

High-precision wind tunnel simulation tests play an important role in aircraft design and manufacture. In this study, a high-speed pose vision measurement method is proposed for high-speed and rolling targets in a supersonic wind tunnel. To obtain images with high signal-to-noise ratio and avoid impacts on the aerodynamic shape of the rolling targets, a high-speed image acquisition method based on ultrathin retro-reflection markers is presented. Since markers are small-sized and some of them may be lost when the target is rolling, a novel markers layout with which markers are distributed evenly on the surface is proposed based on a spatial coding method to achieve highly accurate pose information. Additionally, a pose acquisition is carried out according to the mentioned markers layout after removing mismatching points by Case Deletion Diagnostics. Finally, experiments on measuring the pose parameters of high-speed targets in the laboratory and in a supersonic wind tunnel are conducted to verify the feasibility and effectiveness of the proposed method. Experimental results indicate that the position measurement precision is less than 0.16 mm, the pitching and yaw angle precision less than 0.132° and the roll angle precision 0.712°.

Pose Measurement Method and Experiments for High-Speed Rolling Targets in a Wind Tunnel

PubMed Central

Jia, Zhenyuan; Ma, Xin; Liu, Wei; Lu, Wenbo; Li, Xiao; Chen, Ling; Wang, Zhengqu; Cui, Xiaochun

2014-01-01

High-precision wind tunnel simulation tests play an important role in aircraft design and manufacture. In this study, a high-speed pose vision measurement method is proposed for high-speed and rolling targets in a supersonic wind tunnel. To obtain images with high signal-to-noise ratio and avoid impacts on the aerodynamic shape of the rolling targets, a high-speed image acquisition method based on ultrathin retro-reflection markers is presented. Since markers are small-sized and some of them may be lost when the target is rolling, a novel markers layout with which markers are distributed evenly on the surface is proposed based on a spatial coding method to achieve highly accurate pose information. Additionally, a pose acquisition is carried out according to the mentioned markers layout after removing mismatching points by Case Deletion Diagnostics. Finally, experiments on measuring the pose parameters of high-speed targets in the laboratory and in a supersonic wind tunnel are conducted to verify the feasibility and effectiveness of the proposed method. Experimental results indicate that the position measurement precision is less than 0.16 mm, the pitching and yaw angle precision less than 0.132° and the roll angle precision 0.712°. PMID:25615732

Preliminary Tests in the NACA Free-Spinning Wind Tunnel

NASA Technical Reports Server (NTRS)

Zimmerman, C H

1937-01-01

Typical models and the testing technique used in the NACA free-spinning wind tunnel are described in detail. The results of tests on two models afford a comparison between the spinning characteristics of scale models in the tunnel and of the airplanes that they represent.

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.

Scaling between Wind Tunnels-Results Accuracy in Two-Dimensional Testing

NASA Astrophysics Data System (ADS)

Rasuo, Bosko

The establishment of exact two-dimensional flow conditions in wind tunnels is a very difficult problem. This has been evident for wind tunnels of all types and scales. In this paper, the principal factors that influence the accuracy of two-dimensional wind tunnel test results are analyzed. The influences of the Reynolds number, Mach number and wall interference with reference to solid and flow blockage (blockage of wake) as well as the influence of side-wall boundary layer control are analyzed. Interesting results are brought to light regarding the Reynolds number effects of the test model versus the Reynolds number effects of the facility in subsonic and transonic flow.

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.

An Overview of the Semi-Span Super-Sonic Transport (S4T) Wind-Tunnel Model Program

NASA Technical Reports Server (NTRS)

Silva, Walter A.; Perry, Boyd, III; Florance, James R.; Sanetrik, Mark D.; Wieseman, Carol D.; Stevens, William L.; Funk, Christie J.; Christhilf, David M.; Coulson, David A.

2012-01-01

A summary of computational and experimental aeroelastic (AE) and aeroservoelastic (ASE) results for the Semi-Span Super-Sonic Transport (S4T) wind-tunnel model is presented. A broad range of analyses and multiple AE and ASE wind-tunnel tests of the S4T wind-tunnel model have been performed in support of the ASE element in the Supersonics Program, part of the NASA Fundamental Aeronautics Program. This paper is intended to be an overview of multiple papers that comprise a special S4T technical session. Along those lines, a brief description of the design and hardware of the S4T wind-tunnel model will be presented. Computational results presented include linear and nonlinear aeroelastic analyses, and rapid aeroelastic analyses using CFD-based reduced-order models (ROMs). A brief survey of some of the experimental results from two open-loop and two closed-loop wind-tunnel tests performed at the NASA Langley Transonic Dynamics Tunnel (TDT) will be presented as well.

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.

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.

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

NASA Technical Reports Server (NTRS)

Wong, Douglas T.

1992-01-01

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

Numerical Analysis of Mixed-Phase Icing Cloud Simulations in the NASA Propulsion Systems Laboratory

NASA Technical Reports Server (NTRS)

Bartkus, Tadas; Tsao, Jen-Ching; Struk, Peter; Van Zante, Judith

2017-01-01

This presentation describes the development of a numerical model that couples the thermal interaction between ice particles, water droplets, and the flowing gas of an icing wind tunnel for simulation of NASA Glenn Research Centers Propulsion Systems Laboratory (PSL). The ultimate goal of the model is to better understand the complex interactions between the test parameters and have greater confidence in the conditions at the test section of the PSL tunnel. The model attempts to explain the observed changes in test conditions by coupling the conservation of mass and energy equations for both the cloud particles and flowing gas mass. Model predictions were compared to measurements taken during May 2015 testing at PSL, where test conditions varied gas temperature, pressure, velocity and humidity levels, as well as the cloud total water content, particle initial temperature, and particle size distribution.

Numerical Analysis of Mixed-Phase Icing Cloud Simulations in the NASA Propulsion Systems Laboratory

NASA Technical Reports Server (NTRS)

Bartkus, Tadas P.; Tsao, Jen-Ching; Struk, Peter M.; Van Zante, Judith F.

2017-01-01

This paper describes the development of a numerical model that couples the thermal interaction between ice particles, water droplets, and the flowing gas of an icing wind tunnel for simulation of NASA Glenn Research Centers Propulsion Systems Laboratory (PSL). The ultimate goal of the model is to better understand the complex interactions between the test parameters and have greater confidence in the conditions at the test section of the PSL tunnel. The model attempts to explain the observed changes in test conditions by coupling the conservation of mass and energy equations for both the cloud particles and flowing gas mass. Model predictions were compared to measurements taken during May 2015 testing at PSL, where test conditions varied gas temperature, pressure, velocity and humidity levels, as well as the cloud total water content, particle initial temperature, and particle size distribution.

76 FR 47430 - Airworthiness Directives; Airbus Model A300 B4-600, A300 B4-600R, and A300 F4-600R Series...

Federal Register 2010, 2011, 2012, 2013, 2014

2011-08-05

...], reported the failure during a wind tunnel test of a balance weight fastening screw on the RAT turbine cover... balance weight fastening screw on the RAT turbine cover during a wind tunnel test. After investigation, it... during a wind tunnel test of a balance weight fastening screw on the RAT turbine cover. After...

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.

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.

Comparison of propeller cruise noise data taken in the NASA Lewis 8- by 6-foot wind tunnel with other tunnel and flight data

NASA Technical Reports Server (NTRS)

Dittmar, James H.

1989-01-01

The noise of advanced high speed propeller models measured in the NASA 8- by 6-foot wind tunnel has been compared with model propeller noise measured in another tunnel and with full-scale propeller noise measured in flight. Good agreement was obtained for the noise of a model counterrotation propeller tested in the 8- by 6-foot wind tunnel and in the acoustically treated test section of the Boeing Transonic Wind Tunnel. This good agreement indicates the relative validity of taking cruise noise data on a plate in the 8- by 6-foot wind tunnel compared with the free-field method in the Boeing tunnel. Good agreement was also obtained for both single rotation and counter-rotation model noise comparisons with full-scale propeller noise in flight. The good scale model to full-scale comparisons indicate both the validity of the 8- by 6-foot wind tunnel data and the ability to scale to full size. Boundary layer refraction on the plate provides a limitation to the measurement of forward arc noise in the 8- by 6-foot wind tunnel at the higher harmonics of the blade passing tone. The use of a validated boundary layer refraction model to adjust the data could remove this limitation.

Comparison of propeller cruise noise data taken in the NASA Lewis 8- by 6-foot wind tunnel with other tunnel and flight data

NASA Technical Reports Server (NTRS)

Dittmar, James

1989-01-01

The noise of advanced high speed propeller models measured in the NASA 8- by 6-foot wind tunnel has been compared with model propeller noise measured in another tunnel and with full-scale propeller noise measured in flight. Good agreement was obtained for the noise of a model counterrotation propeller tested in the 8- by 6-foot wind tunnel and in the acoustically treated test section of the Boeing Transonic Wind Tunnel. This good agreement indicates the relative validity of taking cruise noise data on a plate in the 8- by 6-foot wind tunnel compared with the free-field method in the Boeing tunnel. Good agreement was also obtained for both single rotation and counter-rotation model noise comparisons with full-scale propeller noise in flight. The good scale model to full-scale comparisons indicate both the validity of the 8- by 6-foot wind tunnel data and the ability to scale to full size. Boundary layer refraction on the plate provides a limitation to the measurement of forward arc noise in the 8- by 6-foot wind tunnel at the higher harmonics of the blade passing tone. The sue of a validated boundary layer refraction model to adjust the data could remove this limitation.

ARES I Aerodynamic Testing at the NASA Langley Unitary Plan Wind Tunnel

NASA Technical Reports Server (NTRS)

Erickson, Gary E.; Wilcox, Floyd J.

2011-01-01

Small-scale force and moment and pressure models based on the outer mold lines of the Ares I design analysis cycle crew launch vehicle were tested in the NASA Langley Research Center Unitary Plan Wind Tunnel from May 2006 to September 2009. The test objectives were to establish supersonic ascent aerodynamic databases and to obtain force and moment, surface pressure, and longitudinal line-load distributions for comparison to computational predictions. Test data were obtained at low through high supersonic Mach numbers for ranges of the Reynolds number, angle of attack, and roll angle. This paper focuses on (1) the sensitivity of the supersonic aerodynamic characteristics to selected protuberances, outer mold line changes, and wind tunnel boundary layer transition techniques, (2) comparisons of experimental data to computational predictions, and (3) data reproducibility. The experimental data obtained in the Unitary Plan Wind Tunnel captured the effects of evolutionary changes to the Ares I crew launch vehicle, exhibited good agreement with predictions, and displayed satisfactory within-test and tunnel-to-tunnel data reproducibility.

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.

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

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.

High-speed aerodynamic design of space vehicle and required hypersonic wind tunnel facilities

NASA Astrophysics Data System (ADS)

Sakakibara, Seizou; Hozumi, Kouichi; Soga, Kunio; Nomura, Shigeaki

Problems associated with the aerodynamic design of space vehicles with emphasis of the role of hypersonic wind tunnel facilities in the development of the vehicle are considered. At first, to identify wind tunnel and computational fluid dynamics (CFD) requirements, operational environments are postulated for hypervelocity vehicles. Typical flight corridors are shown with the associated flow density: real gas effects, low density flow, and non-equilibrium flow. Based on an evaluation of these flight regimes and consideration of the operational requirements, the wind tunnel testing requirements for the aerodynamic design are examined. Then, the aerodynamic design logic and optimization techniques to develop and refine the configurations in a traditional phased approach based on the programmatic design of space vehicle are considered. Current design methodology for the determination of aerodynamic characteristics for designing the space vehicle, i.e., (1) ground test data, (2) numerical flow field solutions and (3) flight test data, are also discussed. Based on these considerations and by identifying capabilities and limits of experimental and computational methods, the role of a large conventional hypersonic wind tunnel and the high enthalpy tunnel and the interrelationship of the wind tunnels and CFD methods in actual aerodynamic design and analysis are discussed.

A study of large scale gust generation in a small scale atmospheric wind tunnel with applications to Micro Aerial Vehicles

NASA Astrophysics Data System (ADS)

Roadman, Jason Markos

Modern technology operating in the atmospheric boundary layer can always benefit from more accurate wind tunnel testing. While scaled atmospheric boundary layer tunnels have been well developed, tunnels replicating portions of the atmospheric boundary layer turbulence at full scale are a comparatively new concept. Testing at full-scale Reynolds numbers with full-scale turbulence in an "atmospheric wind tunnel" is sought. Many programs could utilize such a tool including Micro Aerial Vehicle(MAV) development, the wind energy industry, fuel efficient vehicle design, and the study of bird and insect flight, to name just a few. The small scale of MAVs provide the somewhat unique capability of full scale Reynolds number testing in a wind tunnel. However, that same small scale creates interactions under real world flight conditions, atmospheric gusts for example, that lead to a need for testing under more complex flows than the standard uniform flow found in most wind tunnels. It is for these reasons that MAVs are used as the initial testing application for the atmospheric gust tunnel. An analytical model for both discrete gusts and a continuous spectrum of gusts is examined. Then, methods for generating gusts in agreement with that model are investigated. Previously used methods are reviewed and a gust generation apparatus is designed. Expected turbulence and gust characteristics of this apparatus are compared with atmospheric data. The construction of an active "gust generator" for a new atmospheric tunnel is reviewed and the turbulence it generates is measured utilizing single and cross hot wires. Results from this grid are compared to atmospheric turbulence and it is shown that various gust strengths can be produced corresponding to weather ranging from calm to quite gusty. An initial test is performed in the atmospheric wind tunnel whereby the effects of various turbulence conditions on transition and separation on the upper surface of a MAV wing is investigated using the surface oil flow visualization technique.

Planning Image-Based Measurements in Wind Tunnels by Virtual Imaging

NASA Technical Reports Server (NTRS)

Kushner, Laura Kathryn; Schairer, Edward T.

2011-01-01

Virtual imaging is routinely used at NASA Ames Research Center to plan the placement of cameras and light sources for image-based measurements in production wind tunnel tests. Virtual imaging allows users to quickly and comprehensively model a given test situation, well before the test occurs, in order to verify that all optical testing requirements will be met. It allows optimization of the placement of cameras and light sources and leads to faster set-up times, thereby decreasing tunnel occupancy costs. This paper describes how virtual imaging was used to plan optical measurements for three tests in production wind tunnels at NASA Ames.

A Practical Methodology for Quantifying Random and Systematic Components of Unexplained Variance in a Wind Tunnel

NASA Technical Reports Server (NTRS)

Deloach, Richard; Obara, Clifford J.; Goodman, Wesley L.

2012-01-01

This paper documents a check standard wind tunnel test conducted in the Langley 0.3-Meter Transonic Cryogenic Tunnel (0.3M TCT) that was designed and analyzed using the Modern Design of Experiments (MDOE). The test designed to partition the unexplained variance of typical wind tunnel data samples into two constituent components, one attributable to ordinary random error, and one attributable to systematic error induced by covariate effects. Covariate effects in wind tunnel testing are discussed, with examples. The impact of systematic (non-random) unexplained variance on the statistical independence of sequential measurements is reviewed. The corresponding correlation among experimental errors is discussed, as is the impact of such correlation on experimental results generally. The specific experiment documented herein was organized as a formal test for the presence of unexplained variance in representative samples of wind tunnel data, in order to quantify the frequency with which such systematic error was detected, and its magnitude relative to ordinary random error. Levels of systematic and random error reported here are representative of those quantified in other facilities, as cited in the references.

Wind Tunnel Model Design for Sonic Boom Studies of Nozzle Jet with Shock Interactions

NASA Technical Reports Server (NTRS)

Cliff, Susan E.; Denison, Marie; Sozer, Emre; Moini-Yekta, Shayan

2016-01-01

NASA and Industry are performing vehicle studies of configurations with low sonic boom pressure signatures. The computational analyses of modern configuration designs have matured to the point where there is confidence in the prediction of the pressure signature from the front of the vehicle, but uncertainty in the aft signatures with often greater boundary layer effects and nozzle jet pressures. Wind tunnel testing at significantly lower Reynolds numbers than in flight and without inlet and nozzle jet pressures make it difficult to accurately assess the computational solutions of flight vehicles. A wind tunnel test in the NASA Ames 9- by 7-Foot Supersonic Wind Tunnel from Mach 1.6 to 2.0 will be used to assess the effects of shocks from components passing through nozzle jet plumes on the sonic boom pressure signature and provide datasets for comparison with CFD codes. A large number of high-fidelity numerical simulations of wind tunnel test models with a variety of shock generators that simulate horizontal tails and aft decks have been studied to provide suitable models for sonic boom pressure measurements using a minimally intrusive pressure rail in the wind tunnel. The computational results are presented and the evolution of candidate wind tunnel models is summarized and discussed in this paper.

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.

Wind tunnel experiments to study chaparral crown fires

Treesearch

Jeanette Cobian-Iñiguez; AmirHessam Aminfar; Joey Chong; Gloria Burke; Albertina Zuniga; David R. Weise; Marko Princevac

2017-01-01

The present protocol presents a laboratory technique designed to study chaparral crown fire ignition and spread. Experiments were conducted in a low velocity fire wind tunnel where two distinct layers of fuel were constructed to represent surface and crown fuels in chaparral. Chamise, a common chaparral shrub, comprised the live crown layer. The dead fuel surface layer...

27. VIEW OF EXHAUST AND DEFLECTOR FOR SUBSONIC AERODYNAMICS RESEARCH ...

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

27. VIEW OF EXHAUST AND DEFLECTOR FOR SUBSONIC AERODYNAMICS RESEARCH LABORATORY, BUILDING 25C, WHICH REPLACED THE 10-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

26. VIEW OF EXHAUST AND DEFLECTOR FOR SUBSONIC AERODYNAMICS RESEARCH ...

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

26. VIEW OF EXHAUST AND DEFLECTOR FOR SUBSONIC AERODYNAMICS RESEARCH LABORATORY, BUILDING 25C, WHICH REPLACED THE 10-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

28. VIEW OF EXHAUST AND DEFLECTOR FOR SUBSONIC AERODYNAMICS RESEARCH ...

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

28. VIEW OF EXHAUST AND DEFLECTOR FOR SUBSONIC AERODYNAMICS RESEARCH LABORATORY, BUILDING 25C, WHICH REPLACED THE 10-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

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.

76 FR 25259 - Airworthiness Directives; Airbus Model A300 B4-600, A300 B4-600R, and A300 F4-600R Series...

Federal Register 2010, 2011, 2012, 2013, 2014

2011-05-04

...], reported the failure during a wind tunnel test of a balance weight fastening screw on the RAT turbine cover... balance weight fastening screw on the RAT turbine cover during a wind tunnel test. After investigation, it... failure during a wind tunnel test of a balance weight fastening screw on the RAT turbine cover. After...

The Modern Design of Experiments: A Technical and Marketing Framework

NASA Technical Reports Server (NTRS)

DeLoach, R.

2000-01-01

A new wind tunnel testing process under development at NASA Langley Research Center, called Modern Design of Experiments (MDOE), differs from conventional wind tunnel testing techniques on a number of levels. Chief among these is that MDOE focuses on the generation of adequate prediction models rather than high-volume data collection. Some cultural issues attached to this and other distinctions between MDOE and conventional wind tunnel testing are addressed in this paper.

Altitude Wind Tunnel Drive Fan being Assembled

NASA Image and Video Library

1943-07-21

National Advisory Committee for Aeronautics (NACA) engineers assembled the Altitude Wind Tunnel’s (AWT) large wooden drive fan inside the hangar at the Aircraft Engine Research Laboratory. When it was built at the in the early 1940s the AWT was among the most complex test facilities ever designed. It was the first wind tunnel capable of operating full-scale engines under realistic flight conditions. This simulation included the reduction of air temperature, a decrease in air pressure, and the creation of an airstream velocity of up to 500 miles per hour. The AWT was constructed in 1942 and 1943. This photograph shows NACA engineers Lou Hermann and Jack Aust assembling the tunnel’s drive fan inside the hangar. The 12-bladed, 31-foot-diameter spruce wood fan would soon be installed inside the wind tunnel to create the high-speed airflow. This massive propeller was designed and constructed by the engine lab's design team at Langley Field. John Breisch, a Langley technician with several years of wind tunnel installation experience, arrived in Cleveland at the time of this photograph to supervise the fan assembly inside the hangar. He would return several weeks later to oversee the actual installation in the tunnel. The fan was driven at 410 revolutions per minute by an 18,000-horsepower General Electric induction motor that was located in the rear corner of the Exhauster Building. An extension shaft connected the motor to the fan. A bronze screen protected the fan against damage from failed engine parts sailing through the tunnel. Despite this screen the blades did become worn or cracked over time and had to be replaced. An entire new fan was installed in 1951.

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.

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.

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

NASA Technical Reports Server (NTRS)

Schwing, Alan

2012-01-01

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

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.

40 CFR 53.60 - General provisions.

Code of Federal Regulations, 2011 CFR

2011-07-01

... samplers are subject to the additional tests and performance requirements specified in § 53.62 (full wind... shall not be subject to the requirements of § 53.62 (full wind tunnel test), provided that it meets all requirements of § 53.63 (wind tunnel inlet aspiration test), § 53.65 (loading test), and § 53.66 (volatility...

40 CFR 53.60 - General provisions.

Code of Federal Regulations, 2010 CFR

2010-07-01

... samplers are subject to the additional tests and performance requirements specified in § 53.62 (full wind... shall not be subject to the requirements of § 53.62 (full wind tunnel test), provided that it meets all requirements of § 53.63 (wind tunnel inlet aspiration test), § 53.65 (loading test), and § 53.66 (volatility...

Experiences with a high-blockage model tested in the NASA Ames 12-foot pressure wind tunnel

NASA Technical Reports Server (NTRS)

Coder, D. W.

1984-01-01

Representation of the flow around full-scale ships was sought in the subsonic wind tunnels in order to a Hain Reynolds numbers as high as possible. As part of the quest to attain the largest possible Reynolds number, large models with high blockage are used which result in significant wall interference effects. Some experiences with such a high blockage model tested in the NASA Ames 12-foot pressure wind tunnel are summarized. The main results of the experiment relating to wind tunnel wall interference effects are also presented.

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.

Heat-flux gage measurements on a flat plate at a Mach number of 4.6 in the VSD high speed wind tunnel, a feasibility test (LA28). [wind tunnel tests of measuring instruments for boundary layer flow

NASA Technical Reports Server (NTRS)

1975-01-01

The feasibility of employing thin-film heat-flux gages was studied as a method of defining boundary layer characteristics at supersonic speeds in a high speed blowdown wind tunnel. Flow visualization techniques (using oil) were employed. Tabulated data (computer printouts), a test facility description, and photographs of test equipment are given.

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.

Altitude Test Cell in the Four Burner Area

NASA Image and Video Library

1947-10-21

One of the two altitude simulating-test chambers in Engine Research Building at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory. The two chambers were collectively referred to as the Four Burner Area. NACA Lewis’ Altitude Wind Tunnel was the nation’s first major facility used for testing full-scale engines in conditions that realistically simulated actual flight. The wind tunnel was such a success in the mid-1940s that there was a backlog of engines waiting to be tested. The Four Burner chambers were quickly built in 1946 and 1947 to ease the Altitude Wind Tunnel’s congested schedule. The Four Burner Area was located in the southwest wing of the massive Engine Research Building, across the road from the Altitude Wind Tunnel. The two chambers were 10 feet in diameter and 60 feet long. The refrigeration equipment produced the temperatures and the exhauster equipment created the low pressures present at altitudes up to 60,000 feet. In 1947 the Rolls Royce Nene was the first engine tested in the new facility. The mechanic in this photograph is installing a General Electric J-35 engine. Over the next ten years, a variety of studies were conducted using the General Electric J-47 and Wright Aeronautical J-65 turbojets. The two test cells were occasionally used for rocket engines between 1957 and 1959, but other facilities were better suited to the rocket engine testing. The Four Burner Area was shutdown in 1959. After years of inactivity, the facility was removed from the Engine Research Building in late 1973 in order to create the High Temperature and Pressure Combustor Test Facility.

Wind-tunnel investigation of the thrust augmentor performance of a large-scale swept wing model. [in the Ames 40 by 80 foot wind tunnel

NASA Technical Reports Server (NTRS)

Koenig, D. G.; Falarski, M. D.

1979-01-01

Tests were made in the Ames 40- by 80-foot wind tunnel to determine the forward speed effects on wing-mounted thrust augmentors. The large-scale model was powered by the compressor output of J-85 driven viper compressors. The flap settings used were 15 deg and 30 deg with 0 deg, 15 deg, and 30 deg aileron settings. The maximum duct pressure, and wind tunnel dynamic pressure were 66 cmHg (26 in Hg) and 1190 N/sq m (25 lb/sq ft), respectively. All tests were made at zero sideslip. Test results are presented without analysis.

PTERA - Modular Aircraft Flight Test

NASA Image and Video Library

2016-01-13

Aerospace testing can be costly and time consuming but a new modular, subscale remotely piloted aircraft offers NASA researchers more affordable options for developing a wide range of cutting edge aviation and space technologies. The Prototype-Technology Evaluation and Research Aircraft (PTERA), developed by Area-I, Inc., of Kennesaw, Georgia, is an extremely versatile and high quality, yet inexpensive, flying laboratory bridging the gap between wind tunnels and crewed flight testing.

Analysis of the Flight Motions of a Small Deployable Glider Configuration

NASA Technical Reports Server (NTRS)

Coe, Paul L., Jr.

1975-01-01

An investigation was conducted at the request of the U.S. Air Force Avionics Laboratory to analyze the flight characteristics of a small uncontrolled glider with folding wings. The study consisted of wind-tunnel tests of an actual glider and a theoretical analysis of the performance, stability, and trimmability of the configuration.

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

NASA Technical Reports Server (NTRS)

Wolf, S. W. D.

1978-01-01

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

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.

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

NASA Technical Reports Server (NTRS)

Lamar, John E. (Editor)

2001-01-01

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

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.

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.

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.

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.

Dynamic wind-tunnel testing of active controls by the NASA Langley Research Center

NASA Technical Reports Server (NTRS)

Abel, I.; Doggett, R. V.; Newsom, J. R.; Sandford, M.

1984-01-01

Dynamic wind-tunnel testing of active controls by the NASA Langley Research Center is presented. Seven experimental studies that were accomplished to date are described. Six of the studies focus on active flutter suppression. The other focuses on active load alleviation. In addition to presenting basic results for these experimental studies, topics including model design and construction, control law synthesis, active control system implementation, and wind-tunnel test techniques are discussed.

Aeroelastic Analysis of a Flexible Wing Wind Tunnel Model with Variable Camber Continuous Trailing Edge Flap Design

NASA Technical Reports Server (NTRS)

Nguyen, Nhan; Ting, Eric; Lebofsky, Sonia

2015-01-01

This paper presents data analysis of a flexible wing wind tunnel model with a variable camber continuous trailing edge flap (VCCTEF) design for drag minimization tested at the University of Washington Aeronautical Laboratory (UWAL). The wind tunnel test was designed to explore the relative merit of the VCCTEF concept for improved cruise efficiency through the use of low-cost aeroelastic model test techniques. The flexible wing model is a 10%-scale model of a typical transport wing and is constructed of woven fabric composites and foam core. The wing structural stiffness in bending is tailored to be half of the stiffness of a Boeing 757-era transport wing while the torsional stiffness is about the same. This stiffness reduction results in a wing tip deflection of about 10% of the wing semi-span. The VCCTEF is a multi-segment flap design having three chordwise camber segments and five spanwise flap sections for a total of 15 individual flap elements. The three chordwise camber segments can be positioned appropriately to create a desired trailing edge camber. Elastomeric material is used to cover the gaps in between the spanwise flap sections, thereby creating a continuous trailing edge. Wind tunnel data analysis conducted previously shows that the VCCTEF can achieve a drag reduction of up to 6.31% and an improvement in the lift-to-drag ratio (L=D) of up to 4.85%. A method for estimating the bending and torsional stiffnesses of the flexible wingUWAL wind tunnel model from static load test data is presented. The resulting estimation indicates that the stiffness of the flexible wing is significantly stiffer in torsion than in bending by as much as 9 to 1. The lift prediction for the flexible wing is computed by a coupled aerodynamic-structural model. The coupled model is developed by coupling a conceptual aerodynamic tool Vorlax with a finite-element model of the flexible wing via an automated geometry deformation tool. Based on the comparison of the lift curve slope, the lift prediction for the rigid wing is in good agreement with the estimated lift coefficients derived from the wind tunnel test data. Due to the movement of the VCCTEF during the wind tunnel test, uncertainty in the lift prediction due to the indicated variations of the VCCTEF deflection is studied. The results show a significant spread in the lift prediction which contradicts the consistency in the aerodynamic measurements, thus suggesting that the indicated variations as measured by the VICON system may not be reliable. The lift prediction of the flexible wing agrees very well with the measured lift curve for the baseline configuration. The computed bending deflection and wash-out twist of the flexible wing also match reasonably well with the aeroelastic deflection measurements. The results demonstrate the validity of the aerodynamic-structural tool for use to analyze aerodynamic performance of flexible wings.

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.

Improved Correction System for Vibration Sensitive Inertial Angle of Attack Measurement Devices

NASA Technical Reports Server (NTRS)

Crawford, Bradley L.; Finley, Tom D.

2000-01-01

Inertial angle of attack (AoA) devices currently in use at NASA Langley Research Center (LaRC) are subject to inaccuracies due to centrifugal accelerations caused by model dynamics, also known as sting whip. Recent literature suggests that these errors can be as high as 0.25 deg. With the current AoA accuracy target at LaRC being 0.01 deg., there is a dire need for improvement. With other errors in the inertial system (temperature, rectification, resolution, etc.) having been reduced to acceptable levels, a system is currently being developed at LaRC to measure and correct for the sting-whip-induced errors. By using miniaturized piezoelectric accelerometers and magnetohydrodynamic rate sensors, not only can the total centrifugal acceleration be measured, but yaw and pitch dynamics in the tunnel can also be characterized. These corrections can be used to determine a tunnel's past performance and can also indicate where efforts need to be concentrated to reduce these dynamics. Included in this paper are data on individual sensors, laboratory testing techniques, package evaluation, and wind tunnel test results on a High Speed Research (HSR) model in the Langley 16-Foot Transonic Wind Tunnel.

Jason Roadman | NREL

Science.gov Websites

using trained raptors, metrology, storm chasing, and wind tunnel testing of atmospheric turbulence. He Wind tunnel testing of atmospheric turbulence Education B.S. in Aerospace Engineering from Virginia

Investigation on aerodynamic characteristics of baseline-II E-2 blended wing-body aircraft with canard via computational simulation

NASA Astrophysics Data System (ADS)

Nasir, Rizal E. M.; Ali, Zurriati; Kuntjoro, Wahyu; Wisnoe, Wirachman

2012-06-01

Previous wind tunnel test has proven the improved aerodynamic charasteristics of Baseline-II E-2 Blended Wing-Body (BWB) aircraft studied in Universiti Teknologi Mara. The E-2 is a version of Baseline-II BWB with modified outer wing and larger canard, solely-designed to gain favourable longitudinal static stability during flight. This paper highlights some results from current investigation on the said aircraft via computational fluid dynamics simulation as a mean to validate the wind tunnel test results. The simulation is conducted based on standard one-equation turbulence, Spalart-Allmaras model with polyhedral mesh. The ambience of the flight simulation is made based on similar ambience of wind tunnel test. The simulation shows lift, drag and moment results to be near the values found in wind tunnel test but only within angles of attack where the lift change is linear. Beyond the linear region, clear differences between computational simulation and wind tunnel test results are observed. It is recommended that different type of mathematical model be used to simulate flight conditions beyond linear lift region.

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

NASA Technical Reports Server (NTRS)

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

1994-01-01

The development of a large-scale anechoic test facility where large models of engine/airframe/high-lift systems can be tested for both improved noise reduction and minimum performance degradation is described. The facility development is part of the effort to investigate economically viable methods of reducing second generation high speed civil transport noise during takeoff and climb-out that is now under way in the United States. This new capability will be achieved through acoustic modifications of NASA's second largest subsonic wind tunnel: the 40-by 80-Foot Wind Tunnel at the NASA Ames Research Center. Three major items are addressed in the design of this large anechoic and quiet wind tunnel: a new deep (42 inch (107 cm)) test section liner, expansion of the wind tunnel drive operating envelope at low rpm to reduce background noise, and other promising methods of improving signal-to-noise levels of inflow microphones. Current testing plans supporting the U.S. high speed civil transport program are also outlined.

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.

Hypersonic nozzle design

NASA Technical Reports Server (NTRS)

Griffith, Wayland C.

1989-01-01

Possible experimental facilities appropriate to a university environment that could make meaningful contributions to the solution of problems in hypersonic aerodynamics are investigated. Needs for the National Aerospace Plane and interplanetary flights with atmospheric aerobraking are used to scope the problem. Relevant events of the past two decades in universities and at the national laboratories are examined for their implications regarding both problems and prospects. Most striking is the emergence of computational fluid dynamics, which is viewed here as an equal partner with laboratory experimentation and flight test in relating theory with reality. Also significant are major advances in instrumentation and data processing methods, especially optical techniques. The direction of the study was guided by the concept of a companion program, i.e., the university effort should complement a major area of endeavor at NASA-Langley. Through this, both faculty and student participants gain a natural and effective working relationship. Existing and proposed major hypersonic aerodynamic facilities in industry and at the national laboratories are examined by type; hypersonic wind tunnels, arc-heated tunnels, shock tubes and tunnels, and ballistic ranges. Of these, the free piston tunnel and shock tube/tunnel are most appropriate for a university.

Analysis of Wind Tunnel Polar Replicates Using the Modern Design of Experiments

NASA Technical Reports Server (NTRS)

Deloach, Richard; Micol, John R.

2010-01-01

The role of variance in a Modern Design of Experiments analysis of wind tunnel data is reviewed, with distinctions made between explained and unexplained variance. The partitioning of unexplained variance into systematic and random components is illustrated, with examples of the elusive systematic component provided for various types of real-world tests. The importance of detecting and defending against systematic unexplained variance in wind tunnel testing is discussed, and the random and systematic components of unexplained variance are examined for a representative wind tunnel data set acquired in a test in which a missile is used as a test article. The adverse impact of correlated (non-independent) experimental errors is described, and recommendations are offered for replication strategies that facilitate the quantification of random and systematic unexplained variance.

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.

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.

Investing American Recovery and Reinvestment Act Funds to Advance Capability, Reliability, and Performance in NASA Wind Tunnels

NASA Technical Reports Server (NTRS)

Sydnor, Goerge H.

2010-01-01

The National Aeronautics and Space Administration's (NASA) Aeronautics Test Program (ATP) is implementing five significant ground-based test facility projects across the nation with funding provided by the American Recovery and Reinvestment Act (ARRA). The projects were selected as the best candidates within the constraints of the ARRA and the strategic plan of ATP. They are a combination of much-needed large scale maintenance, reliability, and system upgrades plus creating new test beds for upcoming research programs. The projects are: 1.) Re-activation of a large compressor to provide a second source for compressed air and vacuum to the Unitary Plan Wind Tunnel at the Ames Research Center (ARC) 2.) Addition of high-altitude ice crystal generation at the Glenn Research Center Propulsion Systems Laboratory Test Cell 3, 3.) New refrigeration system and tunnel heat exchanger for the Icing Research Tunnel at the Glenn Research Center, 4.) Technical viability improvements for the National Transonic Facility at the Langley Research Center, and 5.) Modifications to conduct Environmentally Responsible Aviation and Rotorcraft research at the 14 x 22 Subsonic Tunnel at Langley Research Center. The selection rationale, problem statement, and technical solution summary for each project is given here. The benefits and challenges of the ARRA funded projects are discussed. Indirectly, this opportunity provides the advantages of developing experience in NASA's workforce in large projects and maintaining corporate knowledge in that very unique capability. It is envisioned that improved facilities will attract a larger user base and capabilities that are needed for current and future research efforts will offer revenue growth and future operations stability. Several of the chosen projects will maximize wind tunnel reliability and maintainability by using newer, proven technologies in place of older and obsolete equipment and processes. The projects will meet NASA's goal of integrating more efficient, environmentally safer, and less energy consuming hardware and processes into existing tunnel systems. These include Environmental Protection Agency-approved refrigerants, energy efficient motors, and faster, flexible tunnel data systems.

Forces on Elliptic Cylinders in Uniform Air Stream

NASA Technical Reports Server (NTRS)

Zahm, A F; Smith, R H; Louden, F A

1929-01-01

This report presents the results of wind tunnel tests on four elliptic cylinders with various fineness ratios, conducted in the Navy Aerodynamic Laboratory, Washington. The object of the tests was to investigate the characteristics of sections suitable for streamline wire which normally has an elliptic section with a fineness ratio of 4.0; also to learn whether a reduction in fineness ratio would result in improvement; also to determine the pressure distribution on the model of fineness ratio of 4. Four elliptic cylinders with fineness ratios of 2.5, 3.0, 3.5, and 4.0 were made and then tested in the 8 by 8 wind tunnel; first, for cross-wind force, drag, and yawing moment at 30 miles an hour and various angles of yaw; next for drag 0 degree pitch and 0 degree yaw and various wind speeds; then for end effect on the smallest and largest models; and lastly for pressure distribution over the surface of the largest model at 0 degree pitch and 0 degree yaw and various wind speeds. In all tests, the length of the model was transverse to the current. The results are given for standard air density, p = .002378 slug per cubic foot. This account is a slight revised form of report no. 315. A summary of conclusions is given at the end of the text. (author)

Acoustic Survey of a 3/8-Scale Automotive Wind Tunnel

NASA Technical Reports Server (NTRS)

Booth, Earl R., Jr.; Romberg, Gary; Hansen, Larry; Lutz, Ron

1996-01-01

An acoustic survey that consists of insertion loss and flow noise measurements was conducted at key locations around the circuit of a 3/8-scale automotive acoustic wind tunnel. Descriptions of the test, the instrumentation, and the wind tunnel facility are included in the current report, along with data obtained in the test in the form of 1/3-octave-band insertion loss and narrowband flow noise spectral data.

Flight and wind-tunnel calibrations of a flush airdata sensor at high angles of attack and sideslip and at supersonic Mach numbers

NASA Technical Reports Server (NTRS)

Moes, Timothy R.; Whitmore, Stephen A.; Jordan, Frank L., Jr.

1993-01-01

A nonintrusive airdata-sensing system was calibrated in flight and wind-tunnel experiments to an angle of attack of 70 deg and to angles of sideslip of +/- 15 deg. Flight-calibration data have also been obtained to Mach 1.2. The sensor, known as the flush airdata sensor, was installed on the nosecap of an F-18 aircraft for flight tests and on a full-scale F-18 forebody for wind-tunnel tests. Flight tests occurred at the NASA Dryden Flight Research Facility, Edwards, California, using the F-18 High Alpha Research Vehicle. Wind-tunnel tests were conducted in the 30- by 60-ft wind tunnel at the NASA LaRC, Hampton, Virginia. The sensor consisted of 23 flush-mounted pressure ports arranged in concentric circles and located within 1.75 in. of the tip of the nosecap. An overdetermined mathematical model was used to relate the pressure measurements to the local airdata quantities. The mathematical model was based on potential flow over a sphere and was empirically adjusted based on flight and wind-tunnel data. For quasi-steady maneuvering, the mathematical model worked well throughout the subsonic, transonic, and low supersonic flight regimes. The model also worked well throughout the angle-of-attack and sideslip regions studied.

Flight and wind-tunnel calibrations of a flush airdata sensor at high angles of attack and sideslip and at supersonic Mach numbers

NASA Technical Reports Server (NTRS)

Moes, Timothy R.; Whitmore, Stephen A.; Jordan, Frank L., Jr.

1993-01-01

A nonintrusive airdata-sensing system was calibrated in flight and wind-tunnel experiments to an angle of attack of 70 deg and to angles of sideslip of +/- 15 deg. Flight-calibration data have also been obtained to Mach 1.2. The sensor, known as the flush airdata sensor, was installed on the nosecap of an F-18 aircraft for flight tests and on a full-scale F-18 forebody for wind-tunnel tests. Flight tests occurred at the NASA Dryden Flight Research Facility, Edwards, California, using the F-18 High Alpha Research Vehicle. Wind-tunnel tests were conducted in the 30- by 60-ft wind tunnel at the NASA LaRC, Hampton, Virginia. The sensor consisted of 23 flush-mounted pressure ports arranged in concentric circles and located within 1.75 in. of the tip of the nosecap. An overdetermined mathematical model was used to relate the pressure measurements to the local airdata quantities. The mathematical model was based on potential flow over a sphere and was empirically adjusted based on flight and wind-tunnel data. For quasi-steady maneuvering, the mathematical model worked well throughout the subsonic, transonic, and low supersonic flight regimes. The model also worked well throughout the angles-of-attack and -sideslip regions studied.

Wind Engineering

NASA Technical Reports Server (NTRS)

1983-01-01

Dr. Jack Cermak, Director of Fluid Dynamics and Diffusion Laboratory, developed the first wind tunnel to simulate the changing temperatures, directions and velocities of natural winds. In this work, Cermak benefited from NASA technology related to what is known as the atmospheric boundary layer (ABL).

An Overview of Preliminary Computational and Experimental Results for the Semi-Span Super-Sonic Transport (S4T) Wind-Tunnel Model

NASA Technical Reports Server (NTRS)

Silva, Walter A.; Perry, Boyd, III; Florance, James R.; Sanetrik, Mark D.; Wieseman, Carol D.; Stevens, William L.; Funk, Christie J.; Hur, Jiyoung; Christhilf, David M.; Coulson, David A.

2011-01-01

A summary of computational and experimental aeroelastic and aeroservoelastic (ASE) results for the Semi-Span Super-Sonic Transport (S4T) wind-tunnel model is presented. A broad range of analyses and multiple ASE wind-tunnel tests of the S4T have been performed in support of the ASE element in the Supersonics Program, part of NASA's Fundamental Aeronautics Program. The computational results to be presented include linear aeroelastic and ASE analyses, nonlinear aeroelastic analyses using an aeroelastic CFD code, and rapid aeroelastic analyses using CFD-based reduced-order models (ROMs). Experimental results from two closed-loop wind-tunnel tests performed at NASA Langley's Transonic Dynamics Tunnel (TDT) will be presented as well.

General Electric TG-100A Turboprop in the Altitude Wind Tunnel

NASA Image and Video Library

1946-12-21

A General Electric TG-100A seen from the rear in the test section of the Altitude Wind Tunnel at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory in Cleveland, Ohio. The Altitude Wind Tunnel was used to study almost every model of US turbojet that emerged in the 1940s, as well as some ramjets and turboprops. In the early 1940s the military was interested in an engine that would use less fuel than the early jets but would keep up with them performance-wise. Turboprops seemed like a plausible solution. They could move a large volume of air and thus required less engine speed and less fuel. Researchers at General Electric’s plant in Schenectady, New York worked on the turboprop for several years in the 1930s. They received an army contract in 1941 to design a turboprop engine using an axial-flow compressor. The result was the 14-stage TG-100, the nation's first turboprop aircraft engine. Development of the engine was slow, however, and the military asked NACA Lewis to analyze the engine’s performance. The TG-100A was tested in the Altitude Wind Tunnel and it was determined that the compressors, combustion chamber, and turbine were impervious to changes in altitude. The researchers also established the optimal engine speed and propeller angle at simulated altitudes up to 35,000 feet. Despite these findings, development of the TG-100 was cancelled in May 1947. Twenty-eight of the engines were produced, but they were never incorporated into production aircraft.

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

NASA Technical Reports Server (NTRS)

Deloach, Richard

2012-01-01

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

A Wind-Tunnel Investigation of Tilt-Rotor Gust Alleviation Systems

NASA Technical Reports Server (NTRS)

Ham, N. D.; Whitaker, H. P.

1978-01-01

The alleviation of the effects of gusts on tilt rotor aircraft by means of active control systems was investigated. The gust generator, the derivation of the equations of motion of the rotor wing combination, the correlation of these equations with the results of wind tunnel model tests, the use of the equations to design various gust alleviating active control systems, and the testing and evaluation of these control systems by means of wind tunnel model tests were developed.

Ground testing and simulation. II - Aerodynamic testing and simulation: Saving lives, time, and money

NASA Technical Reports Server (NTRS)

Dayman, B., Jr.; Fiore, A. W.

1974-01-01

The present work discusses in general terms the various kinds of ground facilities, in particular, wind tunnels, which support aerodynamic testing. Since not all flight parameters can be simulated simultaneously, an important problem consists in matching parameters. It is pointed out that there is a lack of wind tunnels for a complete Reynolds-number simulation. Using a computer to simulate flow fields can result in considerable reduction of wind-tunnel hours required to develop a given flight vehicle.

Analysis of the performance of the drive system and diffuser of the Langley unitary plan wind tunnel

NASA Technical Reports Server (NTRS)

Hasel, L. E.; Stallings, R. L.

1981-01-01

A broad program was initiated at the Langley Research Center in 1973 to reduce the energy consumption of the laboratory. As a part of this program, the performance characteristics of the Unitary Plan Wind Tunnel were reexamined to determine if potential methods for incresing the operating efficiencies of the tunnel could be formulated. The results of that study are summarized. The performance characteristics of the drive system components and the variable-geometry diffuser system of the tunnel are documented and analyzed. Several potential methods for reducing the energy requirements of the facility are discussed.

Analysis and correlation of the test data from an advanced technology rotor system

NASA Technical Reports Server (NTRS)

Jepson, D.; Moffitt, R.; Hilzinger, K.; Bissell, J.

1983-01-01

Comparisons were made of the performance and blade vibratory loads characteristics for an advanced rotor system as predicted by analysis and as measured in a 1/5 scale model wind tunnel test, a full scale model wind tunnel test and flight test. The accuracy with which the various tools available at the various stages in the design/development process (analysis, model test etc.) could predict final characteristics as measured on the aircraft was determined. The accuracy of the analyses in predicting the effects of systematic tip planform variations investigated in the full scale wind tunnel test was evaluated.

A study of sedimentation and aggregation of volcanic particles based on experiments carried out with a vertical wind tunnel

NASA Astrophysics Data System (ADS)

Bagheri, G.; Bonadonna, C.; Manzella, I.; Pontelandolfo, P.; Haas, P.

2012-12-01

A complete understanding and parameterization of both particle sedimentation and particle aggregation require systematic and detailed laboratory investigations performed in controlled conditions. For this purpose, a dedicated 4-meter-high vertical wind tunnel has been designed and constructed at the University of Geneva in collaboration with the Groupe de compétence en mécanique des fluides et procédés énergétiques (CMEFE). Final design is a result of Computational Fluid Dynamics simulations combined with laboratory tests. With its diverging test section, the tunnel is designed to suspend particles of different shapes and sizes in order to study the aero-dynamical behavior of volcanic particles and their collision and aggregation. In current set-up, velocities between 5.0 to 27 ms-1 can be obtained, which correspond to typical volcanic particles with diameters between 10 to 40 mm. A combination of Particle Tracking Velocimetry (PTV) and statistical methods is used to derive particle terminal velocity. The method is validated using smooth spherical particles with known drag coefficient. More than 120 particles of different shapes (i.e. spherical, regular and volcanic) and compositions are 3D-scanned and almost 1 million images of their suspension in the test section of wind tunnel are recorded by a high speed camera and analyzed by a PTV code specially developed for the wind tunnel. Measured values of terminal velocity for tested particles are between 3.6 and 24.9 ms-1 which corresponds to Reynolds numbers between 8×103 and 1×105. In addition to the vertical wind tunnel, an apparatus with height varying between 0.5 and 3.5 m has been built to measure terminal velocity of micrometric particles in Reynolds number between 4 and 100. In these experiments, particles are released individually in the air at top of the apparatus and their terminal velocities are measured at the bottom of apparatus by a combination of high-speed camera imaging and PTV post-analyzing. Effects of shape, porosity and orientation of the particles on their terminal velocity are studied. Various shape factors are measured based on different methods, such as 3D-scanning, 2D-image processing, SEM image analysis, caliper measurements, pycnometer and buoyancy tests. Our preliminary experiments on non-smooth spherical particles and irregular particles reveal some interesting aspects. First, the effect of surface roughness and porosity is more important for spherical particles than for regular non-spherical and irregular particles. Second, results underline how, the aero-dynamical behavior of individual irregular particles is better characterized by a range of values of drag coefficients instead of a single value. Finally, since all the shape factors are calculated precisely for each individual particle, the resulted database can provide important information to benchmark and improve existing terminal-velocity models. Modifications of the wind tunnel, i.e. very low air speed (0.03-5.0 ms-1) for suspension of micrometric particles, and of the PTV code, i.e. multiple particle tracking and collision counting, have also been performed in combination to the installation of a particle charging device, a controlled humidifier and a high-power chiller (to reach values down to -20 °C) in order to investigate both wet and dry aggregation of volcanic particles.

Countermeasures for Reducing Unsteady Aerodynamic Force Acting on High-Speed Train in Tunnel by Use of Modifications of Train Shapes

NASA Astrophysics Data System (ADS)

Suzuki, Masahiro; Nakade, Koji; Ido, Atsushi

As the maximum speed of high-speed trains increases, flow-induced vibration of trains in tunnels has become a subject of discussion in Japan. In this paper, we report the result of a study on use of modifications of train shapes as a countermeasure for reducing an unsteady aerodynamic force by on-track tests and a wind tunnel test. First, we conduct a statistical analysis of on-track test data to identify exterior parts of a train which cause the unsteady aerodynamic force. Next, we carry out a wind tunnel test to measure the unsteady aerodynamic force acting on a train in a tunnel and examined train shapes with a particular emphasis on the exterior parts identified by the statistical analysis. The wind tunnel test shows that fins under the car body are effective in reducing the unsteady aerodynamic force. Finally, we test the fins by an on-track test and confirmed its effectiveness.

Development of a process control computer device for the adaptation of flexible wind tunnel walls

NASA Technical Reports Server (NTRS)

Barg, J.

1982-01-01

In wind tunnel tests, the problems arise of determining the wall pressure distribution, calculating the wall contour, and controlling adjustment of the walls. This report shows how these problems have been solved for the high speed wind tunnel of the Technical University of Berlin.

Validation of a Compact Isokinetic Total Water Content Probe for Wind Tunnel Characterization at NASA Glenn Icing Research Tunnel and at NRC Ice Crystal Tunnel

NASA Technical Reports Server (NTRS)

Davison, Craig R.; Landreville, Charles; Ratvasky, Thomas P.

2017-01-01

A new compact isokinetic probe to measure total water content in a wind tunnel environment has been developed. The probe has been previously tested under altitude conditions. This paper presents a comprehensive validation of the probe under a range of liquid water conditions at sea level in the NASA Glenn Icing Research Tunnel and with ice crystals at sea level at the NRC wind tunnel. The compact isokinetic probe is compared to tunnel calibrations and other probes.

A Study of the Effects of Large Scale Gust Generation in a Small Scale Atmospheric Wind Tunnel: Application to Micro Aerial Vehicles

NASA Astrophysics Data System (ADS)

Roadman, Jason; Mohseni, Kamran

2009-11-01

Modern technology operating in the atmospheric boundary layer could benefit from more accurate wind tunnel testing. While scaled atmospheric boundary layer tunnels have been well developed, tunnels replicating portions of the turbulence of the atmospheric boundary layer at full scale are a comparatively new concept. Testing at full-scale Reynolds numbers with full-scale turbulence in an ``atmospheric wind tunnel'' is sought. Many programs could utilize such a tool including that of Micro Aerial Vehicles (MAVs) and other unmanned aircraft, the wind energy industry, fuel efficient vehicles, and the study of bird and insect fight. The construction of an active ``gust generator'' for a new atmospheric tunnel is reviewed and the turbulence it generates is measured utilizing single and cross hot wires. Results from this grid are compared to atmospheric turbulence and it is shown that various gust strengths can be produced corresponding to days ranging from calm to quite gusty. An initial test is performed in the atmospheric wind tunnel whereby the effects of various turbulence conditions on transition and separation on the upper surface of a MAV wing is investigated using oil flow visualization.

Reflection-Type Oil-Film Skin-Friction Meter

NASA Technical Reports Server (NTRS)

Bandyopadhyay, Promode R.; Weinstein, Leonard M.

1993-01-01

Oil-film skin-friction meter for both flight and wind-tunnel applications uses internal reflection and is self-contained, compact unit. Contained in palm-sized housing, in which source of light, mirrors, and sensor mounted rigidly in alignment. Entire unit mounted rigidly under skin of aircraft or wind tunnel, eliminating any relative vibration between optical elements and skin of aircraft or wind tunnel. Meter primarily applicable to flight and wind-tunnel tests, also used in chemical-processing plants.

Design features of a low-disturbance supersonic wind tunnel for transition research at low supersonic Mach numbers

NASA Technical Reports Server (NTRS)

Wolf, Stephen W. D.; Laub, James A.; King, Lyndell S.; Reda, Daniel C.

1992-01-01

A unique, low-disturbance supersonic wind tunnel is being developed at NASA-Ames to support supersonic laminar flow control research at cruise Mach numbers of the High Speed Civil Transport (HSCT). The distinctive design features of this new quiet tunnel are a low-disturbance settling chamber, laminar boundary layers along the nozzle/test section walls, and steady supersonic diffuser flow. This paper discusses these important aspects of our quiet tunnel design and the studies necessary to support this design. Experimental results from an 1/8th-scale pilot supersonic wind tunnel are presented and discussed in association with theoretical predictions. Natural laminar flow on the test section walls is demonstrated and both settling chamber and supersonic diffuser performance is examined. The full-scale wind tunnel should be commissioned by the end of 1993.

Hyper-X Engine Design and Ground Test Program

NASA Technical Reports Server (NTRS)

Voland, R. T.; Rock, K. E.; Huebner, L. D.; Witte, D. W.; Fischer, K. E.; McClinton, C. R.

1998-01-01

The Hyper-X Program, NASA's focused hypersonic technology program jointly run by NASA Langley and Dryden, is designed to move hypersonic, air-breathing vehicle technology from the laboratory environment to the flight environment, the last stage preceding prototype development. The Hyper-X research vehicle will provide the first ever opportunity to obtain data on an airframe integrated supersonic combustion ramjet propulsion system in flight, providing the first flight validation of wind tunnel, numerical and analytical methods used for design of these vehicles. A substantial portion of the integrated vehicle/engine flowpath development, engine systems verification and validation and flight test risk reduction efforts are experimentally based, including vehicle aeropropulsive force and moment database generation for flight control law development, and integrated vehicle/engine performance validation. The Mach 7 engine flowpath development tests have been completed, and effort is now shifting to engine controls, systems and performance verification and validation tests, as well as, additional flight test risk reduction tests. The engine wind tunnel tests required for these efforts range from tests of partial width engines in both small and large scramjet test facilities, to tests of the full flight engine on a vehicle simulator and tests of a complete flight vehicle in the Langley 8-Ft. High Temperature Tunnel. These tests will begin in the summer of 1998 and continue through 1999. The first flight test is planned for early 2000.

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.

A Numerical Comparison of Symmetric and Asymmetric Supersonic Wind Tunnels

NASA Astrophysics Data System (ADS)

Clark, Kylen D.

Supersonic wind tunnels are a vital aspect to the aerospace industry. Both the design and testing processes of different aerospace components often include and depend upon utilization of supersonic test facilities. Engine inlets, wing shapes, and body aerodynamics, to name a few, are aspects of aircraft that are frequently subjected to supersonic conditions in use, and thus often require supersonic wind tunnel testing. There is a need for reliable and repeatable supersonic test facilities in order to help create these vital components. The option of building and using asymmetric supersonic converging-diverging nozzles may be appealing due in part to lower construction costs. There is a need, however, to investigate the differences, if any, in the flow characteristics and performance of asymmetric type supersonic wind tunnels in comparison to symmetric due to the fact that asymmetric configurations of CD nozzle are not as common. A computational fluid dynamics (CFD) study has been conducted on an existing University of Michigan (UM) asymmetric supersonic wind tunnel geometry in order to study the effects of asymmetry on supersonic wind tunnel performance. Simulations were made on both the existing asymmetrical tunnel geometry and two axisymmetric reflections (of differing aspect ratio) of that original tunnel geometry. The Reynolds Averaged Navier Stokes equations are solved via NASAs OVERFLOW code to model flow through these configurations. In this way, information has been gleaned on the effects of asymmetry on supersonic wind tunnel performance. Shock boundary layer interactions are paid particular attention since the test section integrity is greatly dependent upon these interactions. Boundary layer and overall flow characteristics are studied. The RANS study presented in this document shows that the UM asymmetric wind tunnel/nozzle configuration is not as well suited to producing uniform test section flow as that of a symmetric configuration, specifically one that has been scaled to have equal aspect ratio. Comparisons of numerous parameters, such as flow angles, pressure (both static and stagnation), entropy, boundary layers and displacement thickness, vorticity, etc. paint a picture that shows the symmetric equal aspect ratio configuration to be decidedly better at producing desirable test section flow. It has been shown that virtually all parameters of interest are both more consistent and have lower deviation from ideal conditions for the symmetric equal area configuration.

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

NASA Technical Reports Server (NTRS)

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

1988-01-01

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

Schlieren Movies of the 8-Inch Diameter Rigid Parachute Model of the Cook Research Laboratory Taken During the Fourth Phase of Testing in the Langley Unitary Plan Wind Tunnel

NASA Technical Reports Server (NTRS)

1958-01-01

Canopy Model IV was tested in four different configuration series. Shroud lines were used in the first three series of tests; none were used in the fourth series. Other variables were Mach number (1.77, 2.17, 2.76), dynamic pressure (290, 250, 155 lb per sq ft), camera speed, and attitude.

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.

Design, manufacturing and characterization of aero-elastically scaled wind turbine blades for testing active and passive load alleviation techniques within a ABL wind tunnel

NASA Astrophysics Data System (ADS)

Campagnolo, Filippo; Bottasso, Carlo L.; Bettini, Paolo

2014-06-01

In the research described in this paper, a scaled wind turbine model featuring individual pitch control (IPC) capabilities, and equipped with aero-elastically scaled blades featuring passive load reduction capabilities (bend-twist coupling, BTC), was constructed to investigate, by means of wind tunnel testing, the load alleviation potential of BTC and its synergy with active load reduction techniques. The paper mainly focus on the design of the aero-elastic blades and their dynamic and static structural characterization. The experimental results highlight that manufactured blades show desired bend-twist coupling behavior and are a first milestone toward their testing in the wind tunnel.

Ares I Aerodynamic Testing at the Boeing Polysonic Wind Tunnel

NASA Technical Reports Server (NTRS)

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

2011-01-01

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

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.

Novel, Post-Stall, Thrust-Vectored F-15 RPVs: Laboratory and Flight Tests

DTIC Science & Technology

1990-04-24

Flight Tests Program Manager : Douglas Bowers 1ST-Year Report Principal Investigator: Benjamin 6al-Or April 24, 1990 DTIC.LECTE AUG201990 i/ E...constructed. The geometry, dimensions and preliminary wind-tunnel test data for such a design are provided In Appendix A. If funded, such a 3rd...Preliminary Calibration Flight Test Data Obtained from the Onboard Computer ........ 33 Talless, PST-RaNPAS, Roll-Yaw-Pitch, Thrust-Vectored, PST F-15 (Cf. ADp

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.

Field and Wind Tunnel Testing on Natural Ventilation Cooling Effects on Three Navy Buildings.

DTIC Science & Technology

1984-12-01

ORGANIZATION NAME AND ADDRESS 10 PROGRAM ELEMENT. PROJECT, TASK NAVAL CIVIL ENGINEERING LABORATORY AE OEUI UBR Port Hueneme, CA 93043S02-13A 7Y/ 11I...base of pressure difference coefficients for a variety of buildings with spe- cific architechtural features. Values from this data base can be used to

YF-12 Experiments Symposium, Volume 1

NASA Technical Reports Server (NTRS)

1978-01-01

Papers presented by personnel from the Dryden Flight Research Center, the Lewis Research Center, and the Ames Research Center are presented. Topics cover propulsion system performance, inlet time varying distortion, structures, aircraft controls, propulsion controls, and aerodynamics. The reports were based on analytical studies, laboratory experiments, wind tunnel tests, and extensive flight research with two YF-12 airplanes.

Numerical investigation of air flow in a supersonic wind tunnel

NASA Astrophysics Data System (ADS)

Drozdov, S. M.; Rtishcheva, A. S.

2017-11-01

In the framework of TsAGI’s supersonic wind tunnel modernization program aimed at improving flow quality and extending the range of test regimes it was required to design and numerically validate a new test section and a set of shaped nozzles: two flat nozzles with flow Mach number at nozzle exit M=4 and M=5 and two axisymmetric nozzles with M=5 and M=6. Geometric configuration of the nozzles, the test section (an Eiffel chamber) and the diffuser was chosen according to the results of preliminary calculations of two-dimensional air flow in the wind tunnel circuit. The most important part of the work are three-dimensional flow simulation results obtained using ANSYS Fluent software. The following flow properties were investigated: Mach number, total and static pressure, total and static temperature and turbulent viscosity ratio distribution, heat flux density at wind tunnel walls (for high-temperature flow regimes). It is demonstrated that flow perturbations emerging from the junction of the nozzle with the test section and spreading down the test section behind the boundaries of characteristic rhomb’s reverse wedge are nearly impossible to eliminate. Therefore, in order to perform tests under most uniform flow conditions, the model’s center of rotation and optical window axis should be placed as close to the center of the characteristic rhomb as possible. The obtained results became part of scientific and technical basis of supersonic wind tunnel design process and were applied to a generalized class of similar wind tunnels.

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.

Study of optical techniques for the Ames unitary wind tunnels. Part 3: Angle of attack

NASA Technical Reports Server (NTRS)

Lee, George

1992-01-01

A review of optical sensors that are capable of accurate angle of attack measurements in wind tunnels was conducted. These include sensors being used or being developed at NASA Ames and Langley Research Centers, Boeing Airplane Company, McDonald Aircraft Company, Arnold Engineering Development Center, National Aerospace Laboratory of the Netherlands, National Research Council of Canada, and the Royal Aircraft Establishment of England. Some commercial sensors that may be applicable to accurate angle measurements were also reviewed. It was found that the optical sensor systems were based on interferometers, polarized light detector, linear or area photodiode cameras, position sensing photodetectors, and laser scanners. Several of the optical sensors can meet the requirements of the Ames Unitary Plan Wind Tunnel. Two of these, the Boeing interferometer and the Complere lateral effect photodiode sensors are being developed for the Ames Unitary Plan Wind Tunnel.

Relation between the Fluctuating Wall Pressure and the Turbulent Structure of a Boundary Layer on a Cylinder in Axial Flow

DTIC Science & Technology

1993-08-12

Shop for their expert assistance during thze design ard development ur the wind tunnel and experimental apparatus; Drs. Alan L. Kistler, Seth Lichter...vertical wind tunnel was designed and built for this research. I With the test section in a vertical orientation, gravity effects leading to cylinder sag...were eliminated. The overall design and layout of the wind tunnel, as well as specific design features incorporated into the wind tunnel to satisfy

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

NASA Technical Reports Server (NTRS)

Bell, James H.; Burner, Alpheus W.

2004-01-01

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

Computed and Experimental Flutter/LCO Onset for the Boeing Truss-Braced Wing Wind-Tunnel Model

NASA Technical Reports Server (NTRS)

Bartels, Robert E.; Scott, Robert C.; Funk, Christie J.; Allen, Timothy J.; Sexton, Bradley W.

2014-01-01

This paper presents high fidelity Navier-Stokes simulations of the Boeing Subsonic Ultra Green Aircraft Research truss-braced wing wind-tunnel model and compares the results to linear MSC. Nastran flutter analysis and preliminary data from a recent wind-tunnel test of that model at the NASA Langley Research Center Transonic Dynamics Tunnel. The simulated conditions under consideration are zero angle of attack, so that structural nonlinearity can be neglected. It is found that, for Mach number greater than 0.78, the linear flutter analysis predicts flutter onset dynamic pressure below the wind-tunnel test and that predicted by the Navier-Stokes analysis. Furthermore, the wind-tunnel test revealed that the majority of the high structural dynamics cases were wing limit cycle oscillation (LCO) rather than flutter. Most Navier-Stokes simulated cases were also LCO rather than hard flutter. There is dip in the wind-tunnel test flutter/LCO onset in the Mach 0.76-0.80 range. Conditions tested above that Mach number exhibited no aeroelastic instability at the dynamic pressures reached in the tunnel. The linear flutter analyses do not show a flutter/LCO dip. The Navier-Stokes simulations also do not reveal a dip; however, the flutter/LCO onset is at a significantly higher dynamic pressure at Mach 0.90 than at lower Mach numbers. The Navier-Stokes simulations indicate a mild LCO onset at Mach 0.82, then a more rapidly growing instability at Mach 0.86 and 0.90. Finally, the modeling issues and their solution related to the use of a beam and pod finite element model to generate the Navier-Stokes structure mode shapes are discussed.

Digital control of wind tunnel magnetic suspension and balance systems

NASA Technical Reports Server (NTRS)

Britcher, Colin P.; Goodyer, Michael J.; Eskins, Jonathan; Parker, David; Halford, Robert J.

1987-01-01

Digital controllers are being developed for wind tunnel magnetic suspension and balance systems, which in turn permit wind tunnel testing of aircraft models free from support interference. Hardware and software features of two existing digital control systems are reviewed. Some aspects of model position sensing and system calibration are also discussed.

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.

Turbine endwall two-cylinder program. [wind tunnel and water tunnel investigation of three dimensional separation of fluid flow

NASA Technical Reports Server (NTRS)

Langston, L. S.

1980-01-01

Progress is reported in an effort to study the three dimensional separation of fluid flow around two isolated cylinders mounted on an endwall. The design and performance of a hydrogen bubble generator for water tunnel tests to determine bulk flow properties and to measure main stream velocity and boundary layer thickness are described. Although the water tunnel tests are behind schedule because of inlet distortion problems, tests are far enough along to indicate cylinder spacing, wall effects and low Reynolds number behavior, all of which impacted wind tunnel model design. The construction, assembly, and operation of the wind tunnel and the check out of its characteristics are described. An off-body potential flow program was adapted to calculate normal streams streamwise pressure gradients at the saddle point locations.

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

NASA Technical Reports Server (NTRS)

Brooks, J. D.

1976-01-01

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

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.

Abe Silverstein Leads Tour of the 10- by 10-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1955-11-21

Abe Silverstein, Associate Director of the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory, provides a personal tour of the new 10- by 10-Foot Supersonic Wind Tunnel for US Senator George Bender (hat in hand) and General Lemuel Shepherd. Shepherd was Commandant of the Marine Corps and had served in World War I, World War II, and the Korean War. The general was accompanied by Admiral Herbert Leary, in dark uniform. Bender was a Republican Senator from Ohio. Behind Bender is President of the Cleveland Chamber of Commerce Curtis Smith. NACA Lewis managers Eugene Manganiello and Wilson Hunter assist with the tour. Abe Silverstein oversaw all research at the laboratory. Upon taking his post in 1952 he reorganized the research staff and began shifting the focus away from airbreathing aircraft engines to new fields such as high energy fuels, electric propulsion, and nuclear power and propulsion. He was an early advocate of the NACA’s involvement in the space program and crucial to the founding of National Aeronautics and Space Administration in 1958. Silverstein began his career helping design and conduct research in the Full Scale Tunnel in 1929 at the Langley Memorial Aeronautical Laboratory. Silverstein advocated a series of increasingly large supersonic wind tunnels after the war, culminating in the 10- by 10.

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.

A procedure for predicting internal and external noise fields of blowdown wind tunnels

NASA Technical Reports Server (NTRS)

Hosier, R. N.; Mayes, W. H.

1972-01-01

The noise generated during the operation of large blowdown wind tunnels is considered. Noise calculation procedures are given to predict the test-section overall and spectrum level noise caused by both the tunnel burner and turbulent boundary layer. External tunnel noise levels due to the tunnel burner and circular jet exhaust flow are also calculated along with their respective cut-off frequency and spectrum peaks. The predicted values are compared with measured data, and the ability of the prediction procedure to estimate blowdown-wind-tunnel noise levels is shown.

An assessment of wind tunnel test data on flexible thermal protection materials and results of new fatigue tests of threads

NASA Technical Reports Server (NTRS)

Coe, Charles F.

1985-01-01

Advanced Flexible Reusable Surface Insulation (AFRSI) was developed as a replacement for the low-temperature (white) tiles on the Space Shuttle. The first use of the AFRSI for an Orbiter flight was on the OMS POD of Orbiter (OV-099) for STS-6. Post flight examination after STS-6 showed that damage had occurred to the AFRSI during flight. The failure anomaly between previous wind-tunnel tests and STS-6 prompted a series of additional wind tunnel tests to gain an insight as to the cause of the failure. An assessment of all the past STS-6 wind tunnel tests pointed out the sensitivity of the test results to scaling of dynamic loads due to the difference of boundary layer thickness, and the material properties as a result of exposure to heating. The thread component of the AFRSI was exposed to fatigue testing using an apparatus that applied pulsating aerodynamic loads on the threads similar to the loads caused by an oscillating shock. Comparison of the mean values of the number-of-cycles to failure showed that the history of the thread was the major factor in its performance. The thread and the wind tunnel data suggests a mechanism of failure for the AFRSI.

Building 865 Hypersonic Wind Tunnel Power System Analysis

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

Schneider, Larry X.

2015-02-01

This report documents the characterization and analysis of a high current power supply for the building 865 Hypersonic Wind Tunnel at Sandia National Laboratories. The system described in this report became operational in 2013, replacing the original 1968 system which employed an induction voltage regulator. This analysis and testing was completed to help the parent organization understand why an updated and redesigned power system was not delivering adequate power to resistive heater elements in the HWT. This analysis led to an improved understanding of the design and operation of the revised 2013 power supply system and identifies several reasons themore » revised system failed to achieve the performance of the original power supply installation. Design modifications to improve the performance of this system are discussed.« less

Evaluation of electrolytic tilt sensors for wind tunnel model angle-of-attack (AOA) measurements

NASA Technical Reports Server (NTRS)

Wong, Douglas T.

1991-01-01

The results of a laboratory evaluation of three types of electrolytic tilt sensors as potential candidates for model attitude or angle of attack (AOA) measurements in wind tunnel tests are presented. Their performance was also compared with that from typical servo accelerometers used for AOA measurements. Model RG-37 electrolytic tilt sensors were found to have the highest overall accuracy among the three types. Compared with the servo accelerometer, their accuracies are about one order of magnitude worse and each of them cost about two-thirds less. Therefore, the sensors are unsuitable for AOA measurements although they are less expensive. However, the potential for other applications exists where the errors resulting from roll interaction, vibration, and response time are less, and sensor temperature can be controlled.

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.

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

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

NASA Technical Reports Server (NTRS)

Braddock, W. F.

1979-01-01

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

F/A-18 1/9th scale model tail buffet measurements

NASA Technical Reports Server (NTRS)

Martin, C. A.; Glaister, M. K.; Maclaren, L. D.; Meyn, L. A.; Ross, J.

1991-01-01

Wind tunnel tests were carried out on a 1/9th scale model of the F/A-18 at high angles of attack to investigate the characteristics of tail buffet due to bursting of the wing leading edge extension (LEX) vortices. The tests were carried out at the Aeronautical Research Laboratory low-speed wind tunnel facility and form part of a collaborative activity with NASA Ames Research Center, organized by The Technical Cooperative Program (TTCP). Information from the program will be used in the planning of similar collaborative tests, to be carried out at NASA Ames, on a full-scale aircraft. The program covered the measurement of unsteady pressures and fin vibration for cases with and without the wing LEX fences fitted. Fourier transform methods were used to analyze the unsteady data, and information on the spatial and temporal content of the vortex burst pressure field was obtained. Flow visualization of the vortex behavior was carried out using smoke and a laser light sheet technique.

Aerodynamic results of wind tunnel tests on a 0.010-scale model (32-QTS) space shuttle integrated vehicle in the AEDC VKF-40-inch supersonic wind tunnel (IA61)

NASA Technical Reports Server (NTRS)

Daileda, J. J.

1976-01-01

Plotted and tabulated aerodynamic coefficient data from a wind tunnel test of the integrated space shuttle vehicle are presented. The primary test objective was to determine proximity force and moment data for the orbiter/external tank and solid rocket booster (SRB) with and without separation rockets firing for both single and dual booster runs. Data were obtained at three points (t = 0, 1.25, and 2.0 seconds) on the nominal SRB separation trajectory.

Wind Tunnel Model and Test to Evaluate the Effectiveness of a Passive Gust Alleviation Device for a Flying Wing Aircraft

DTIC Science & Technology

2016-10-04

model of 1.24 m with the PGAD and control surface 3 1.2. Design and manufacture of the gust generator (frame, blades , actuation and control system...Chapter 3, a gust generator with two rotating blades was designed and manufactured to induce a transverse turbulence for wind tunnel test. A CFD...velocity at 8C (eight times of blade chord length) achieved 1.3%. In Chapter 4, the wind tunnel test of the scaled wing model is presented, including the

Intelligent hypertext systems for aerospace engineering applications

NASA Technical Reports Server (NTRS)

Lo, Ching F.

1989-01-01

This paper is a progress report on the utilization of AI technology for assisting users locating and understanding technical information in manuals used for planning and conducting wind tunnel test. The specific goal is to create an Intelligent Hypertext System (IHS) for wind tunnel testing which combines the computerized manual in the form of hypertext and an advisory system that stores experts' knowledge and experiences. A prototype IHS for conducting transonic wind tunnel testing has been constructed with limited knowledge base. The prototype is being evaluated by potential users.

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

NASA Technical Reports Server (NTRS)

Kmak, Frank J.

2000-01-01

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

Control law parameterization for an aeroelastic wind-tunnel model equipped with an active roll control system and comparison with experiment

NASA Technical Reports Server (NTRS)

Perry, Boyd, III; Dunn, H. J.; Sandford, Maynard C.

1988-01-01

Nominal roll control laws were designed, implemented, and tested on an aeroelastically-scaled free-to-roll wind-tunnel model of an advanced fighter configuration. The tests were performed in the NASA Langley Transonic Dynamics Tunnel. A parametric study of the nominal roll control system was conducted. This parametric study determined possible control system gain variations which yielded identical closed-loop stability (roll mode pole location) and identical roll response but different maximum control-surface deflections. Comparison of analytical predictions with wind-tunnel results was generally very good.

Real-Gas Flow Properties for NASA Langley Research Center Aerothermodynamic Facilities Complex Wind Tunnels

NASA Technical Reports Server (NTRS)

Hollis, Brian R.

1996-01-01

A computational algorithm has been developed which can be employed to determine the flow properties of an arbitrary real (virial) gas in a wind tunnel. A multiple-coefficient virial gas equation of state and the assumption of isentropic flow are used to model the gas and to compute flow properties throughout the wind tunnel. This algorithm has been used to calculate flow properties for the wind tunnels of the Aerothermodynamics Facilities Complex at the NASA Langley Research Center, in which air, CF4. He, and N2 are employed as test gases. The algorithm is detailed in this paper and sample results are presented for each of the Aerothermodynamic Facilities Complex wind tunnels.

Experimental and analytical research on the aerodynamics of wind driven turbines. Final report

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

Rohrbach, C.; Wainauski, H.; Worobel, R.

1977-12-01

This aerodynamic research program was aimed at providing a reliable, comprehensive data base on a series of wind turbine models covering a broad range of the prime aerodynamic and geometric variables. Such data obtained under controlled laboratory conditions on turbines designed by the same method, of the same size, and tested in the same wind tunnel had not been available in the literature. Moreover, this research program was further aimed at providing a basis for evaluating the adequacy of existing wind turbine aerodynamic design and performance methodology, for assessing the potential of recent advanced theories and for providing a basismore » for further method development and refinement.« 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.

Payload bay atmospheric vent airflow testing at the Vibration and Acoustic Test Facility

NASA Technical Reports Server (NTRS)

Johnston, James D., Jr.

1988-01-01

Several concerns related to venting the Space Shuttle Orbiter payload bay during launch led to laboratory experiments with a flight-type vent box installed in the wall of a subsonic wind tunnel. This report describes the test setups and procedures used to acquire data for characterization of airflow through the vent box and acoustic tones radiated from the vent-box cavity. A flexible boundary-layer spoiler which reduced the vent-tone amplitude is described.

A flight investigation of the spinning of the NY-1 airplane with varied mass distribution and other modifications, and an analysis based on wind-tunnel tests

NASA Technical Reports Server (NTRS)

Scudder, Nathan F

1934-01-01

This report presents the results of an investigation of the spinning characteristics of NY-1 naval training biplane. The results of flight tests and an analysis based on wind-tunnel test data are given and compared. The primary purpose of the investigation was the determination in flight of the effect of changes in mass distribution along the longitudinal axis, without change of mass quantity or centroid. Other effects were also investigated, such as those due to wing loading, center-of-gravity position, dihedral of wings, control setting, and the removal of a large portion of the fabric from the fin and rudder. The wind tunnel test results used in the numerical analysis were obtained in the 7 by 10 foot wind tunnel through an angle-of-attack.

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

NASA Technical Reports Server (NTRS)

Potsdam, Mark; Roberts, Leonard

1990-01-01

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

Investigation of water droplet trajectories within the NASA icing research tunnel

NASA Technical Reports Server (NTRS)

Reehorst, Andrew; Ibrahim, Mounir

1995-01-01

Water droplet trajectories within the NASA Lewis Research Center's Icing Research Tunnel (IRT) were studied through computer analysis. Of interest was the influence of the wind tunnel contraction and wind tunnel model blockage on the water droplet trajectories. The computer analysis was carried out with a program package consisting of a three-dimensional potential panel code and a three-dimensional droplet trajectory code. The wind tunnel contraction was found to influence the droplet size distribution and liquid water content distribution across the test section from that at the inlet. The wind tunnel walls were found to have negligible influence upon the impingement of water droplets upon a wing model.

Flowfield measurements in a model scramjet combustion using laser-induced iodine fluorescence

NASA Technical Reports Server (NTRS)

Mcdaniel, J. C., Jr.

1984-01-01

Preliminary designs were completed for an iodine mixing chamber and the optical setup to be used with a modified wind tunnel in obtaining accurate, spatially resolved measurements of variables in the flowfield of a model nonreacting scramjet combustor. Schematics of the iodine-seeded wind tunnel and a sketch of the charcoal filter for removing the iodine are included along with a cutaway section of the laboratory.

Research done at DERAT (October 1982 through September 1983); summary of principal results obtained

NASA Technical Reports Server (NTRS)

1984-01-01

The progress in the following areas is described: measurement equipment; F2 FAUGA wind tunnel tests; unsteady boundary layers; body and axisymmetrical boundary layers; wing fuselage interactions; turbulence; subsonic-transonic flow; cryogenic wind tunnel tests; and profile testing.

Determination of Airloads on Certain Airport Towers Due to Addition of Radomes.

DOT National Transportation Integrated Search

1981-08-01

This report presents the results of wind tunnel tests conducted on three different radome configurations mounted atop three different airport tower designs. The wind tunnel tests were conducted on 1/12th scale models. Field tests were conducted on on...

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

NASA Technical Reports Server (NTRS)

Wolf, S. W. D.

1977-01-01

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

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.

SOFIA 2 model telescope wind tunnel test report

NASA Technical Reports Server (NTRS)

Keas, Paul

1995-01-01

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

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.

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.

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.

Low-cost wind tunnel for aerosol inhalation studies.

PubMed

Chung, I P; Dunn-Rankin, D; Phalen, R F; Oldham, M J

1992-04-01

A low-cost wind tunnel for aerosol studies has been designed, constructed, and evaluated for aerosol uniformity with 2- and 0.46-micron particles. A commercial nebulizer was used to produce the suspended test particles, and a custom-made, four-hole injector was used to introduce the aerosol into the wind tunnel. A commercially available optical particle counter measured the particle concentration. Performance tests of the velocity profile and particle concentration distribution at two flow rates showed that the system performs well for small particles.

Low-Pressure Capability of NASA Glenn's 10- by 10-Foot Supersonic Wind Tunnel Expanded

NASA Technical Reports Server (NTRS)

Roeder, James W.

2004-01-01

Extremely low dynamic pressure Q conditions are desired for space-related research including the testing of parachute designs and other decelerator concepts for future vehicles landing on Mars. Therefore, the low-pressure operating capability of the Abe Silverstein 10- by 10-foot Supersonic Wind Tunnel (10 10 SWT) at NASA Glenn Research Center was recently increased. Successful checkout tests performed in the fall of 2002 showed significantly reduced minimum operating pressures in the wind tunnel.

Passive Turbulence Generating Grid Arrangements in a Turbine Cascade Wind Tunnel

DTIC Science & Technology

2014-04-02

root mean square of free stream velocity flow viscosity Turbine Cascade Wind Tunnels ( CWT ) are similar to conventional wind tunnels except the test...section o f interest is in a corner. Figure I shows the United States Air Force Academy (USAF A) closed-loop CWT . Turbine cascade facilities are used...evaluating only the middle third span of the blade, the ceiling and floor effects in the tunne l can be mitigated. A CWT test section inlet must have

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

NASA Technical Reports Server (NTRS)

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

1988-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1988-01-01

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

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.

Testing and Performance Verification of a High Bypass Ratio Turbofan Rotor in an Internal Flow Component Test Facility

NASA Technical Reports Server (NTRS)

VanZante, Dale E.; Podboy, Gary G.; Miller, Christopher J.; Thorp, Scott A.

2009-01-01

A 1/5 scale model rotor representative of a current technology, high bypass ratio, turbofan engine was installed and tested in the W8 single-stage, high-speed, compressor test facility at NASA Glenn Research Center (GRC). The same fan rotor was tested previously in the GRC 9x15 Low Speed Wind Tunnel as a fan module consisting of the rotor and outlet guide vanes mounted in a flight-like nacelle. The W8 test verified that the aerodynamic performance and detailed flow field of the rotor as installed in W8 were representative of the wind tunnel fan module installation. Modifications to W8 were necessary to ensure that this internal flow facility would have a flow field at the test package that is representative of flow conditions in the wind tunnel installation. Inlet flow conditioning was designed and installed in W8 to lower the fan face turbulence intensity to less than 1.0 percent in order to better match the wind tunnel operating environment. Also, inlet bleed was added to thin the casing boundary layer to be more representative of a flight nacelle boundary layer. On the 100 percent speed operating line the fan pressure rise and mass flow rate agreed with the wind tunnel data to within 1 percent. Detailed hot film surveys of the inlet flow, inlet boundary layer and fan exit flow were compared to results from the wind tunnel. The effect of inlet casing boundary layer thickness on fan performance was quantified. Challenges and lessons learned from testing this high flow, low static pressure rise fan in an internal flow facility are discussed.

Wind tunnel productivity status and improvement activities at NASA Langley Research Center

NASA Technical Reports Server (NTRS)

Putnam, Lawrence E.

1996-01-01

Over the last three years, a major effort has been underway to re-engineering the way wind tunnel testing is accomplished at the NASA Langley Research Center. This effort began with the reorganization of the LaRC and the consolidation of the management of the wind tunnels in the Aerodynamics Division under one operations branch. This paper provides an overview of the re-engineering activities and gives the status of the improvements in the wind tunnel productivity and customer satisfaction that have resulted from the new ways of working.

Design and testing of an oblique all-wing supersonic transport

NASA Technical Reports Server (NTRS)

Lee, Christopher A.

1994-01-01

This report describes the preliminary design of an Oblique All-Wing (OAW) supersonic transport and a corresponding wind-tunnel model that was tested in the NASA Ames 9- by 7-Foot supersonic wind tunnel. The main goal was the determination of the cruise performance (lift/drag ratio) of a realistically configured OAW. To achieve an acceptable level of realism, it was necessary to consider many issues of design practicality such as the need for a viable propulsion system, adequate control surfaces, landing gear, provisions for 450 passengers, and fuel to fly 5,000 nautical miles. The aircraft had to be stable, structurally sound, and needed to fit into airports across the world. Support was directed primarily towards integration of the propulsion system, however, there were notable contributions to many aspects of the configuration design, wind tunnel model, and wind tunnel test.

JPL noise control program

NASA Technical Reports Server (NTRS)

Klascius, A. F.

1975-01-01

Exposures of personnel to noise pollution at the Jet Propulsion Laboratories, Pasadena, California, were investigated. As a result of the study several protective measures were taken: (1) employees exposed to noise hazards were required to wear ear-protection devices, (2) mufflers and air diversion devices were installed around the wind tunnels; and (3) all personnel that are required to wear ear protection are given annual audimeter tests.

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

NASA Technical Reports Server (NTRS)

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

2004-01-01

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

Development and Operation of an Automatic Rotor Trim Control System for the UH-60 Individual Blade Control Wind Tunnel Test

NASA Technical Reports Server (NTRS)

Theodore, Colin R.; Tischler, Mark B.

2010-01-01

An automatic rotor trim control system was developed and successfully used during a wind tunnel test of a full-scale UH-60 rotor system with Individual Blade Control (IBC) actuators. The trim control system allowed rotor trim to be set more quickly, precisely and repeatably than in previous wind tunnel tests. This control system also allowed the rotor trim state to be maintained during transients and drift in wind tunnel flow, and through changes in IBC actuation. The ability to maintain a consistent rotor trim state was key to quickly and accurately evaluating the effect of IBC on rotor performance, vibration, noise and loads. This paper presents details of the design and implementation of the trim control system including the rotor system hardware, trim control requirements, and trim control hardware and software implementation. Results are presented showing the effect of IBC on rotor trim and dynamic response, a validation of the rotor dynamic simulation used to calculate the initial control gains and tuning of the control system, and the overall performance of the trim control system during the wind tunnel test.

What scaling means in wind engineering: Complementary role of the reduced scale approach in a BLWT and the full scale testing in a large climatic wind tunnel

NASA Astrophysics Data System (ADS)

Flamand, Olivier

2017-12-01

Wind engineering problems are commonly studied by wind tunnel experiments at a reduced scale. This introduces several limitations and calls for a careful planning of the tests and the interpretation of the experimental results. The talk first revisits the similitude laws and discusses how they are actually applied in wind engineering. It will also remind readers why different scaling laws govern in different wind engineering problems. Secondly, the paper focuses on the ways to simplify a detailed structure (bridge, building, platform) when fabricating the downscaled models for the tests. This will be illustrated by several examples from recent engineering projects. Finally, under the most severe weather conditions, manmade structures and equipment should remain operational. What “recreating the climate” means and aims to achieve will be illustrated through common practice in climatic wind tunnel modelling.

Acoustical characteristics of the NASA Langley full scale wind tunnel test section

NASA Technical Reports Server (NTRS)

Abrahamson, A. L.; Kasper, P. K.; Pappa, R. S.

1975-01-01

The full-scale wind tunnel at NASA-Langley Research Center was designed for low-speed aerodynamic testing of aircraft. Sound absorbing treatment has been added to the ceiling and walls of the tunnel test section to create a more anechoic condition for taking acoustical measurements during aerodynamic tests. The results of an experimental investigation of the present acoustical characteristics of the tunnel test section are presented. The experimental program included measurements of ambient nosie levels existing during various tunnel operating conditions, investigation of the sound field produced by an omnidirectional source, and determination of sound field decay rates for impulsive noise excitation. A comparison of the current results with previous measurements shows that the added sound treatment has improved the acoustical condition of the tunnel test section. An analysis of the data indicate that sound reflections from the tunnel ground-board platform could create difficulties in the interpretation of actual test results.

On the skin friction drag reduction in large wind turbines using sharp V-grooved riblets. Application to a 2.5 MW Clipper wind turbine section

NASA Astrophysics Data System (ADS)

Arndt, Roger; Chamorro, Leonardo; Sotiropoulos, Fotis

2010-11-01

Skin friction drag reduction through the use of riblets has been a topic of intensive research during the last decades. Main efforts have been placed on both numerical (mainly DNS) and experimental approaches. In spite of the valuable efforts, the fundamental mechanisms that induce drag reduction are not well established. In this study, wind tunnel experiments were performed to quantify the drag reduction in a wind turbine airfoil using different V-groove riblet structures. A full-scale 2.5MW Clipper wind turbine airfoil section (of 1 meter chord length, typical of the 88% blade span), was placed in the freestream flow of the wind tunnel at the Saint Anthony Falls Laboratory, University of Minnesota. Four different sizes of V-groove riblets were tested at different angles of attack at full scale Reynolds number of Re=2.67x106 (based on the airfoil chord length). Force sensors were used to measure Lift and Drag. A combination of single and cross-wire anemometers were also used to study the turbulent scale-to-scale interaction in the near wall region to better understand the physical mechanisms of drag reduction and flow characteristics in that region. The measurements will be used to develop and test the performance of near-wall boundary conditions in the context of RANS and hybrid RANS/LES models.

16-Inch Diameter Ramjet Prepared for Flight Test

NASA Image and Video Library

1947-07-21

A NACA researcher prepares a 16-inch diameter and 16-foot long ramjet for a launch over Wallops Island in July 1947. The Lewis Flight Propulsion Laboratory conducted a wide variety of studies on ramjets in the 1940s and 1960s to determine the basic operational data necessary to design missiles. Although wind tunnel and test stand investigations were important first steps in determining these factors, actual flight tests were required. Lewis possessed several aircraft for the ramjet studies, including North American F-82 Mustangs, a Northrup P-61 Black Widow, and a Boeing B-29 Superfortress, which was used for this particular ramjet. This was Lewis’ first flight at over the experimental testing ground at Wallops Island. The NACA’s Langley laboratory established the station on the Virginia coast in 1945 to conduct early missile tests. This ramjet-powered missile was affixed underneath the B-29’s left wing and flown up to 29,000 feet. The ramjet was ignited as the aircraft reached Mach 0.5 and released. The flight went well, but a problem with the data recording prevented a successful mission. Nonetheless additional flights in November 1947 provided researchers with data on the engine’s combustion efficiency at different levels of fuel-air ratios, thrust coefficients, temperatures, and drag. Transonic flight data such as the rapid acceleration through varying flight conditions could not be easily captured in wind tunnels.

Aeroelastic Analyses of the SemiSpan SuperSonic Transport (S4T) Wind Tunnel Model at Mach 0.95

NASA Technical Reports Server (NTRS)

Hur, Jiyoung

2014-01-01

Detailed aeroelastic analyses of the SemiSpan SuperSonic Transport (S4T) wind tunnel model at Mach 0.95 with a 1.75deg fixed angle of attack are presented. First, a numerical procedure using the Computational Fluids Laboratory 3-Dimensional (CFL3D) Version 6.4 flow solver is investigated. The mesh update method for structured multi-block grids was successfully applied to the Navier-Stokes simulations. Second, the steady aerodynamic analyses with a rigid structure of the S4T wind tunnel model are reviewed in transonic flow. Third, the static analyses were performed for both the Euler and Navier-Stokes equations. Both the Euler and Navier-Stokes equations predicted a significant increase of lift forces, compared to the results from the rigid structure of the S4T wind-tunnel model, over various dynamic pressures. Finally, dynamic aeroelastic analyses were performed to investigate the flutter condition of the S4T wind tunnel model at the transonic Mach number. The condition of flutter was observed at a dynamic pressure of approximately 75.0-psf for the Navier-Stokes simulations. However, it was observed that the flutter condition occurred a dynamic pressure of approximately 47.27-psf for the Euler simulations. Also, the computational efficiency of the aeroelastic analyses for the S4T wind tunnel model has been assessed.

Operating safety of a hot-shot wind tunnel with combined test gas heating in stabilization mode

NASA Astrophysics Data System (ADS)

Shumskii, V. V.; Yaroslavtsev, M. I.

2017-07-01

In the present paper, we analyze emergency situations typical of short-duration wind tunnels with electric-arc or combined test-gas heating in the presence of stabilization and diaphragm-rupturing systems, which occur in the case of no discharge initiation in the settling chamber, with the capacitor battery having remained charged during the start of wind-tunnel systems. For avoiding such emergency situations, some additional changes based on using feedback elements are introduced into the wind-tunnel design: the piston of the fast-response valve is made hollow for increasing the volume of the shutoff cavity and for making the release of pressure from this cavity unnecessary; the high-pressure channel, which connects the piston and the piston rod with the settling-chamber cavity, is filled with a liquid and is closed from the side of the settling chamber with a piston; the device for controlled diaphragm breakdown is provided with an external electric circuit intended to control the diaphragm-rupturing process. Those modifications allow subsequent functioning of the wind-tunnel systems only in the presence of heat-supply-induced pressure growth in the settling chamber of the wind tunnel.

The rationale and design features for the 40 by 80/80 by 120 foot wind tunnel

NASA Technical Reports Server (NTRS)

Mort, K. W.; Kelly, M. W.; Hickey, D. H.

1976-01-01

A substantial increase in the test capability of full scale wind tunnels is considered. In order to determine the most cost effective means for providing this desired increase in test capability, a series of design studies were conducted of various new facilities as well as of major modifications to the existing 40- by 80-foot wind tunnel. The most effective trade between test capability and facility cost was provided by repowering the existing 40 by 80 foot wind tunnel to increase the maximum speed from 200 knots to 300 knots and by the addition of a new 80- by 120-foot test section having a 110 knot maximum speed. The design of the facility is described with special emphasis on the unique features, such as the drive system which absorbs nearly four times the power without an increase in noise, and the large flow diversion devices required to interface the two test sections to a single drive.

Selected results of the F-15 propulsion interactions program

NASA Technical Reports Server (NTRS)

Webb, L. D.; Nugent, J.

1982-01-01

A better understanding of propulsion system/airframe flow interactions could aid in the reduction of aircraft drag. For this purpose, NASA and the United States Air Force have conducted a series of wind-tunnel and flight tests on the F-15 airplane. This paper presents a correlation of flight test data from tests conducted at the NASA Dryden Flight Research Facility of the Ames Research Center, with data obtained from wind-tunnel tests. Flights were made at stabilized Mach numbers around 0.6, 0.9, 1.2, and 1.5 with accelerations up to near Mach number 2. Wind-tunnel tests used a 7.5 percent-scale F-15 inlet/airframe model. Flight and wind-tunnel pressure coefficients showed good agreement in most cases. Correlation of interaction effects caused by changes in cowl angle, angle-of-attack, and Mach number are presented. For the afterbody region, the pressure coefficients on the nozzle surfaces were influenced by boattail angles and Mach number. Boundary-layer thickness decreased as angle of attack increased above 4 deg.

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.

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.

Calculation of wall effects of flow on a perforated wall with a code of surface singularities

NASA Astrophysics Data System (ADS)

Piat, J. F.

1994-07-01

Simplifying assumptions are inherent in the analytic method previously used for the determination of wall interferences on a model in a wind tunnel. To eliminate these assumptions, a new code based on the vortex lattice method was developed. It is suitable for processing any shape of test sections with limited areas of porous wall, the characteristic of which can be nonlinear. Calculation of wall effects in S3MA wind tunnel, whose test section is rectangular 0.78 m x 0.56 m, and fitted with two or four perforated walls, have been performed. Wall porosity factors have been adjusted to obtain the best fit between measured and computed pressure distributions on the test section walls. The code was checked by measuring nearly equal drag coefficients for a model tested in S3MA wind tunnel (after wall corrections) and in S2MA wind tunnel whose test section is seven times larger (negligible wall corrections).

Wind-tunnel and Flight Investigations of the Use of Leading-Edge Area Suction for the Purpose of Increasing the Maximum Lift Coefficient of a 35 Degree Swept-Wing Airplane

NASA Technical Reports Server (NTRS)

Holzhauser, Curt A; Bray, Richard S

1956-01-01

An investigation was undertaken to determine the increase in maximum lift coefficient that could be obtained by applying area suction near the leading edge of a wing. This investigation was performed first with a 35 degree swept-wing model in the wind tunnel, and then with an operational 35 degree swept-wing airplane which was modified in accord with the wind-tunnel results. The wind-tunnel and flight tests indicated that the maximum lift coefficient was increased more than 50 percent by the use of area suction. Good agreement was obtained in the comparison of the wind-tunnel results with those measured in flight.

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.

The Unitary Plan Wind Tunnel(UPWT) Test 1891 Space Launch System

NASA Image and Video Library

2014-10-15

Stage Separation Test of the Space Launch System(SLS) in the Langley Unitary Plan Wind Tunnel (UPWT). The model used High Pressure air blown through the solid rocket boosters. (SRB) to simulate the booster separation motors (BSM) firing.

The Unitary Plan Wind Tunnel(UPWT) Test 1891 Space Launch System

NASA Image and Video Library

2014-10-14

Stage Separation Test of the Space Launch System(SLS) in the Langley Unitary Plan Wind Tunnel (UPWT). The model used High Pressure air blown through the solid rocket boosters. (SRB) to simulate the booster separation motors (BSM) firing.

Transition heating rates determined on a 0.006 scale space shuttle orbiter model (no. 50-0) in the NASA/LaRC Mach 8 variable density wind tunnel test (OH14)

NASA Technical Reports Server (NTRS)

Cummings, J.

1976-01-01

Data obtained from wind tunnel tests of an .006-scale space shuttle orbiter model in the 18 in. Variable Density Wind Tunnel are presented. The tests, denoted as OH14, were performed to determine transition heating rates using thin skin thermocouples located at various locations on the space shuttle orbiter. The model was tested at M = 8.0 for a range of Reynolds numbers per foot varying from 1.0 to 10.0 million with angles-of-attack from 20 to 35 degrees incremented by 5 degrees.

A wind tunnel study of gaseous tracer dispersion in the convective boundary layer capped by a temperature inversion

NASA Astrophysics Data System (ADS)

Fedorovich, E.; Thäter, J.

Results are presented from wind tunnel simulations of gaseous pollutant dispersion in the atmospheric convective boundary layer (CBL) capped by a temperature inversion. The experiments were performed in the thermally stratified wind tunnel of the University of Karlsruhe, Germany. In the tunnel, the case of horizontally evolving, sheared CBL is reproduced. This distinguishes the employed experimental setup from the preceding laboratory and numerical CBL dispersion studies. The diffusive and mixing properties of turbulence in the studied CBL case have been found to be essentially dependent on the stage of the CBL evolution. Effects of the point source elevation on the horizontal variability of the concentration field, and on the ground level concentration as function of distance from the source have been investigated. The applicability of bottom-up/top-down diffusion concept in the simulated CBL case has been evaluated. The influence of surface wind shear and capping inversion strength on the pollutant dispersion and turbulent exchange across the CBL top has been demonstrated. The imposed positive shear across the inversion has been identified as inhibitor of the CBL growth. Comparisons of concentration patterns from the wind tunnel with water tank data are presented.

Recent Cycle Time Reduction at Langley Research Center

NASA Technical Reports Server (NTRS)

Kegelman, Jerome T.

2000-01-01

The NASA Langley Research Center (LaRC) has been engaged in an effort to reduce wind tunnel test cycle time in support of Agency goals and to satisfy the wind tunnel testing needs of the commercial and military aerospace communities. LaRC has established the Wind Tunnel Enterprise (WTE), with goals of reducing wind tunnel test cycle time by an order of magnitude by 2002, and by two orders of magnitude by 2010. The WTE also plans to meet customer expectations for schedule integrity, as well as data accuracy and quality assurance. The WTE has made progress towards these goals over the last year with a focused effort on technological developments balanced by attention to process improvements. This paper presents a summary of several of the WTE activities over the last year that are related to test cycle time reductions at the Center. Reducing wind tunnel test cycle time, defined here as the time between the freezing of loft lines and delivery of test data, requires that the relationship between high productivity and data quality assurance be considered. The efforts have focused on all of the drivers for test cycle time reduction, including process centered improvements, facility upgrades, technological improvements to enhance facility readiness and productivity, as well as advanced measurement techniques. The application of internet tools and computer modeling of facilities to allow a virtual presence of the customer team is also presented.

Field-testing a portable wind tunnel for fine dust emissions

USDA-ARS?s Scientific Manuscript database

A protable wind tunnel has been developed to allow erodibility and dust emissions testing of soil surfaces with the premise that dust concentration and properties are highly correlated with surface soil properties, as modified by crop management system. In this study we report on the field-testing ...

Aerodynamic forces on freight trains : volume 1. wind tunnel tests of containers and trailers on flatcars

DOT National Transportation Integrated Search

1976-12-01

The aerodynamic forces on trailers and containers on flatcars have been measured in wind tunnel tests. The forces were measured on the central car of a five-car train consisting of a locomotive, three flatcars with various loadings and a boxcar. Test...

Investigation of correlation between full-scale and fifth-scale wind tunnel tests of a Bell helicopter Textron Model 222

NASA Technical Reports Server (NTRS)

Squires, P. K.

1982-01-01

Reasons for lack of correlation between data from a fifth-scale wind tunnel test of the Bell Helicopter Textron Model 222 and a full-scale test of the model 222 prototype in the NASA Ames 40-by 80-foot tunnel were investigated. This investigation centered around a carefully designed fifth-scale wind tunnel test of an accurately contoured model of the Model 222 prototype mounted on a replica of the full-scale mounting system. The improvement in correlation for drag characteristics in pitch and yaw with the fifth-scale model mounted on the replica system is shown. Interference between the model and mounting system was identified as a significant effect and was concluded to be a primary cause of the lack of correlation in the earlier tests.

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.

Heat transfer phase change paint tests of 0.0175-scale models (nos. 21-0 and 46-0) of the Rockwell International space shuttle orbiter in the AEDC tunnel B hypersonic wind tunnel (test OH25A)

NASA Technical Reports Server (NTRS)

Dye, W. H.

1975-01-01

Tests were conducted in a hypersonic wind tunnel using various truncated space shuttle orbiter configurations in an attempt to establish the optimum model size for other tests examining body shock-wing leading edge interference effects. The tests were conducted at Mach number 8 using the phase change paint technique. A test description, tabulated data, and tracings of isotherms made from photographs taken during the test are presented.

Determination of HART I Blade Structural Properties by Laboratory Testing

NASA Technical Reports Server (NTRS)

Jung, Sung N.; Lau, Benton H.

2012-01-01

The structural properties of higher harmonic Aeroacoustic Rotor Test (HART I) blades were measured using the original set of blades tested in the German-dutch wind tunnel (DNW) in 1994. the measurements include bending and torsion stiffness, geometric offsets, and mass and inertia properties of the blade. the measured properties were compared to the estimated values obtained initially from the blade manufacturer. The previously estimated blade properties showed consistently higher stiffness, up to 30 percent for the flap bending in the blade inboard root section.

Evaluation of flow quality in two large NASA wind tunnels at transonic speeds

NASA Technical Reports Server (NTRS)

Harvey, W. D.; Stainback, P. C.; Owen, F. K.

1980-01-01

Wind tunnel testing of low drag airfoils and basic transition studies at transonic speeds are designed to provide high quality aerodynamic data at high Reynolds numbers. This requires that the flow quality in facilities used for such research be excellent. To obtain a better understanding of the characteristics of facility disturbances and identification of their sources for possible facility modification, detailed flow quality measurements were made in two prospective NASA wind tunnels. Experimental results are presented of an extensive and systematic flow quality study of the settling chamber, test section, and diffuser in the Langley 8 foot transonic pressure tunnel and the Ames 12 foot pressure wind tunnel. Results indicate that the free stream velocity and pressure fluctuation levels in both facilities are low at subsonic speeds and are so high as to make it difficult to conduct meaningful boundary layer control and transition studies at transonic speeds.

Model Deformation Measurements at NASA Langley Research Center

NASA Technical Reports Server (NTRS)

Burner, A. W.

1998-01-01

Only recently have large amounts of model deformation data been acquired in NASA wind tunnels. This acquisition of model deformation data was made possible by the development of an automated video photogrammetric system to measure the changes in wing twist and bending under aerodynamic load. The measurement technique is based upon a single view photogrammetric determination of two dimensional coordinates of wing targets with a fixed third dimensional coordinate, namely the spanwise location. A major consideration in the development of the measurement system was that use of the technique must not appreciably reduce wind tunnel productivity. The measurement technique has been used successfully for a number of tests at four large production wind tunnels at NASA and a dedicated system is nearing completion for a fifth facility. These facilities are the National Transonic Facility, the Transonic Dynamics Tunnel, and the Unitary Plan Wind Tunnel at NASA Langley, and the 12-FT Pressure Tunnel at NASA Ames. A dedicated system for the Langley 16-Foot Transonic Tunnel is scheduled to be used for the first time for a test in September. The advantages, limitations, and strategy of the technique as currently used in NASA wind tunnels are presented. Model deformation data are presented which illustrate the value of these measurements. Plans for further enhancements to the technique are presented.

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 complexity of developing a model that can be used for successfully implementing a standardized management planning tool. The objective of this study was to implement an Integrated Wind Tunnel Planning System to improve the operations within the aeronautics testing and research group, in particular Wind Tunnel Enterprise. The study included following steps: Conducted literature search and expert discussions (NASA and Old Dominion University faculty), Performed environmental scan of NASA Langley wind tunnel operations as foundation for problem definition. Established operation requirements and evaluation methodologies. Examined windtunnel operations to map out the common characteristics, critical components, and system structure. Reviewed and evaluated various project scheduling and management systems for implementation, Evaluated and implemented "Theory of Constraints (TOC)" project scheduling methodology at NASA Langley wind tunnel operations together with NASA staff.

Mercury Capsule Model in the 1- by 1-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1959-10-21

National Aeronautics and Space Administration (NASA) researchers install a small-scale model of the capsule for Project Mercury in the 1- by 1-Foot Supersonic Wind Tunnel at the Lewis Research Center. NASA Lewis conducted a variety of tests for Project Mercury, including retrorocket calibration, escape tower engine performance, and separation of the capsule from simulated Atlas and Redstone boosters. The test of this capsule and escape tower model in the 1- by 1-foot tunnel were run in January and February 1960. The 1-by 1-Foot Supersonic Wind Tunnel had a 15-inch long test section, seen here, that was one foot wide and one foot high. The sides were made of glass to allow cameras to capture the supersonic air flow over the models. The tunnel could generate air flows from Mach 1.3 to 3.0. At the time, it was one of nine small supersonic wind tunnels at Lewis. These tunnels used the exhauster and compressor equipment of the larger facilities. The 1- by 1 tunnel, which began operating in the early 1950s, was built inside a test cell in the expansive Engine Research Building. During the 1950s the 1- by 1 was used to study a variety of inlets, nozzles, and cones for missiles and scramjets. The Mercury capsule tests were among the last at the facility for many years. The tunnel was mothballed in 1960. The 1- by 1 was briefly restored in 1972, then brought back online for good in 1979. The facility has maintained a brisk operating schedule ever since.

Sediment-transport experiments in zero-gravity

NASA Technical Reports Server (NTRS)

Iversen, James D.; Greeley, Ronald

1987-01-01

One of the important parameters in the analysis of sediment entrainment and transport is gravitational attraction. The availability of a laboratory in earth orbit would afford an opportunity to conduct experiments in zero and variable gravity environments. Elimination of gravitational attraction as a factor in such experiments would enable other critical parameters (such as particle cohesion and aerodynamic forces) to be evaluated much more accurately. A Carousel Wind Tunnel (CWT) is proposed for use in conducting experiments concerning sediment particle entrainment and transport in a space station. In order to test the concept of this wind tunnel design a one third scale model CWT was constructed and calibrated. Experiments were conducted in the prototype to determine the feasibility of studying various aeolian processes and the results were compared with various numerical analysis. Several types of experiments appear to be feasible utilizing the proposed apparatus.

Sediment-transport experiments in zero-gravity

NASA Technical Reports Server (NTRS)

Iversen, J. D.; Greeley, R.

1986-01-01

One of the important parameters in the analysis of sediment entrainment and transport is gravitational attraction. The availability of a laboratory in Earth orbit would afford an opportunity to conduct experiments in zero and variable gravity environments. Elimination of gravitational attraction as a factor in such experiments would enable other critical parameters (such as particle cohesion and aerodynamic forces) to be evaluated much more accurately. A Carousel Wind Tunnel (CWT) is proposed for use in conducting experiments concerning sediment particle entrainment and transport in a space station. In order to test the concept of this wind tunnel design a one third scale model CWT was constructed and calibrated. Experiments were conducted in the prototype to determine the feasibility of studying various aeolian processes and the results were compared with various numerical analysis. Several types of experiments appear to be feasible utilizing the proposed apparatus.

NACA Computer Operates an IBM Telereader

NASA Image and Video Library

1952-02-21

A staff member from the Computing Section at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory operates an International Business Machines (IBM) telereader at the 8- by 6-Foot Supersonic Wind Tunnel. The telereader was used to measure recorded data from motion picture film or oscillographs. The machine could perform 50 measurements per minute. The component to her right is a telerecordex that was used convert the telereader measurements into decimal form and record the data on computer punch cards. During test runs in the 8- by 6-foot tunnel, or the other large test facilities, pressure sensors on the test article were connected to mercury-filled manometer tubes located below the test section. The mercury would rise or fall in relation to the pressure fluctuations in the test section. Initially, female staff members, known as “computers,” transcribed all the measurements by hand. The process became automated with the introduction of the telereader and other data reduction equipment in the early 1950s. The Computer Section staff members were still needed to operate the machines. The Computing Section was introduced during World War II to relieve short-handed research engineers of some of the tedious work. The computers made the initial computations and plotted the data graphically. The researcher then analyzed the data and either summarized the findings in a report or made modifications or ran the test again. The computers and analysts were located in the Altitude Wind Tunnel Shop and Office Building office wing during the 1940s. They were transferred to the new facility when the 8- by 6-Foot tunnel began operations in 1948.

Wind Tunnel Model Design for the Study of Plume Effects on Sonic Boom for Isolated Exhaust Nozzles

NASA Technical Reports Server (NTRS)

Castner, Raynold S.

2010-01-01

A low cost test capability was developed at the NASA Glenn Research Center 1- by 1-Foot Supersonic Wind Tunnel (SWT), with a goal to reduce the disturbance caused by supersonic aircraft flight over populated areas. This work focused on the shock wave structure caused by the exhaust nozzle plume. Analysis and design was performed on a new rig to test exhaust nozzle plume effects on sonic boom signature. Test capability included a baseline nozzle test article and a wind tunnel model consisting of a strut, a nosecone and an upper plenum. Analysis was performed on the external and internal aerodynamic configuration, including the shock reflections from the wind tunnel walls caused by the presence of the model nosecone. This wind tunnel model was designed to operate from Mach 1.4 to Mach 3.0 with nozzle pressure ratios from 6 to 12 and altitudes from 30,000 ft (4.36 psia) to 50,000 ft (1.68 psia). The model design was based on a 1 in. outer diameter, was 9 in. in overall length, and was mounted in the wind tunnel on a 3/8 in. wide support strut. For test conditions at 50,000 ft the strut was built to supply 90 psia of pressure, and to achieve 20 psia at the nozzle inlet with a maximum nozzle pressure of 52 psia. Instrumentation was developed to measure nozzle pressure ratio, and an external static pressure probe was designed to survey near field static pressure profiles at one nozzle diameter above the rig centerline. Model layout placed test nozzles between two transparent sidewalls in the 1 1 SWT for Schlieren photography and comparison to CFD analysis.

Wind Tunnel Model Design for the Study of Plume Effects on Sonic Boom for Isolated Exhaust Nozzles

NASA Technical Reports Server (NTRS)

Castner, Raymond S.

2009-01-01

A low cost test capability was developed at the NASA Glenn Research Center 1- by 1-Foot Supersonic Wind Tunnel (SWT), with a goal to reduce the disturbance caused by supersonic aircraft flight over populated areas. This work focused on the shock wave structure caused by the exhaust nozzle plume. Analysis and design was performed on a new rig to test exhaust nozzle plume effects on sonic boom signature. Test capability included a baseline nozzle test article and a wind tunnel model consisting of a strut, a nose cone and an upper plenum. Analysis was performed on the external and internal aerodynamic configuration, including the shock reflections from the wind tunnel walls caused by the presence of the model nosecone. This wind tunnel model was designed to operate from Mach 1.4 to Mach 3.0 with nozzle pressure ratios from 6 to 12 and altitudes from 30,000 ft (4.36 psia) to 50,000 ft (1.68 psia). The model design was based on a 1 in. outer diameter, was 9 in. in overall length, and was mounted in the wind tunnel on a 3/8 in. wide support strut. For test conditions at 50,000 ft the strut was built to supply 90 psia of pressure, and to achieve 20 psia at the nozzle inlet with a maximum nozzle pressure of 52 psia. Instrumentation was developed to measure nozzle pressure ratio, and an external static pressure probe was designed to survey near field static pressure profiles at one nozzle diameter above the rig centerline. Model layout placed test nozzles between two transparent sidewalls in the 1x1 SWT for Schlieren photography and comparison to CFD analysis.

Videogrammetric Model Deformation Measurement Technique

NASA Technical Reports Server (NTRS)

Burner, A. W.; Liu, Tian-Shu

2001-01-01

The theory, methods, and applications of the videogrammetric model deformation (VMD) measurement technique used at NASA for wind tunnel testing are presented. The VMD technique, based on non-topographic photogrammetry, can determine static and dynamic aeroelastic deformation and attitude of a wind-tunnel model. Hardware of the system includes a video-rate CCD camera, a computer with an image acquisition frame grabber board, illumination lights, and retroreflective or painted targets on a wind tunnel model. Custom software includes routines for image acquisition, target-tracking/identification, target centroid calculation, camera calibration, and deformation calculations. Applications of the VMD technique at five large NASA wind tunnels are discussed.

Advances in Projection Moire Interferometry Development for Large Wind Tunnel Applications

NASA Technical Reports Server (NTRS)

Fleming, Gary A.; Soto, Hector L.; South, Bruce W.; Bartram, Scott M.

1999-01-01

An instrument development program aimed at using Projection Moire Interferometry (PMI) for acquiring model deformation measurements in large wind tunnels was begun at NASA Langley Research Center in 1996. Various improvements to the initial prototype PMI systems have been made throughout this development effort. This paper documents several of the most significant improvements to the optical hardware and image processing software, and addresses system implementation issues for large wind tunnel applications. The improvements have increased both measurement accuracy and instrument efficiency, promoting the routine use of PMI for model deformation measurements in production wind tunnel tests.

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

NASA Technical Reports Server (NTRS)

Binion, T. W., Jr.

1975-01-01

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

A scale model wind tunnel study of dispersion in the Cleveland area. Laboratory simulation of lake breeze effects on diffusion from ground level emissions

NASA Technical Reports Server (NTRS)

Hoydysh, W. G.

1974-01-01

A wind tunnel simulation of the diffusion patterns in a sea breeze was attempted. The results indicate that the low level onshore flow was well simulated for neutral, stable, unstable, and elevated inversion conditions. Velocity, turbulence, shear stress, and temperature data were taken, and the spread of emissions from ground level sources was investigated. Comparison is made with theoretical predictions by E. Inoue and with the open, homogeneous plane field results of Pasquill. Agreement with the predictions by Inoue is good, and the comparison with Pasquill's results shows that the wind tunnel flows are shifted two categories towards more stable. The discrepancy may be explained as a matter of averaging time.

A laboratory measurement of drop impact on a water surface in the presence of wind

NASA Astrophysics Data System (ADS)

Liu, Xinan; Liu, Ren

2014-11-01

The impact of single water drops on a water surface was studied experimentally in a wind tunnel. Water drops were generated from a needle oriented vertically from the top of the wind tunnel test section. After leaving the needle, the drops move downward due to gravity and downstream due to the effect of the wind, eventually impinging obliquely on the surface of a pool of water on the bottom of the test section. The vertical velocities of drops were about 2.0 m/s and the wind speeds varied from 0 to 6.4 m/s. The drop impacts were recorded simultaneously from the side and above with two high-speed movie cameras with frame rates of 1,000 Hz. Our measurements show that both wind speed and initial drop size dramatically affect the drop impacts and subsequent generation of crowns, secondary drops, stalks and ring waves. In the presence of wind, an asymmetric crown forms after the drop hits the water surface and secondary drops are generated from the fragmentation of the leeward side of the crown rim. This is followed by a stalk formation and ring waves at the location of the water drop impact. It is found that the stalks tilt to leeward and the ring waves in the windward direction are stronger than that in those in the leeward. This work is supported by National Science Foundation, Division of Ocean Sciences.

Measurement of attachment-line location in a wind-tunnel and in supersonic flight

NASA Technical Reports Server (NTRS)

Agarwal, Naval K.; Miley, Stan J.; Fisher, Michael C.; Anderson, Bianca T.; Geenen, Robert J.

1992-01-01

For the supersonic laminar flow control research program, tests are being conducted to measure the attachment-line flow characteristics and its location on a highly swept aircraft wing. Subsonic wind tunnel experiments were conducted on 2D models to develop sensors and techniques for the flight application. Representative attachment-line data are discussed and results from the wind tunnel investigation are presented.

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.

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.

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

NASA Technical Reports Server (NTRS)

Rivers, Melissa; Quest, Juergen; Rudnik, Ralf

2015-01-01

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

General Electric TG-180 Turbojet in the Altitude Wind Tunnel

NASA Image and Video Library

1947-09-21

A General Electric TG-180 turbojet installed in the Altitude Wind Tunnel at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory. In 1943 the military asked General Electric to develop an axial-flow jet engine which became the TG-180. The military understood that the TG-180 would not be ready during World War II but recognized the axial-flow compressor’s long-term potential. Although the engine was bench tested in April 1944, it was not flight tested until February 1946. The TG-180 was brought to the Altitude Wind Tunnel in 1945 for a series of investigations. The studies, which continued intermittently into 1948, analyzed an array of performance issues. NACA modifications steadily improved the TG-180’s performance, including the first successful use of an afterburner. The Lewis researchers studied a 29-inch diameter afterburner over a range of altitude conditions using several different types of flameholders and fuel systems. Lewis researchers concluded that a three-stage flameholder with its largest stage upstream was the best burner configuration. Although the TG-180 (also known as the J35) was not the breakthrough engine that the military had hoped for, it did power the Douglas D-558-I Skystreak to a world speed record on August 20, 1947. The engines were also used on the Republic F-84 Thunderjet and the Northrup F-89 Scorpion.

Wind Tunnel Model Design for Sonic Boom Studies of Nozzle Jet Flows with Shock Interactions

NASA Technical Reports Server (NTRS)

Cliff, Susan E.; Denison, Marie; Moini-Yekta, Shayan; Morr, Donald E.; Durston, Donald A.

2016-01-01

NASA and the U.S. aerospace industry are performing studies of supersonic aircraft concepts with low sonic boom pressure signatures. The computational analyses of modern aircraft designs have matured to the point where there is confidence in the prediction of the pressure signature from the front of the vehicle, but uncertainty remains in the aft signatures due to boundary layer and nozzle exhaust jet effects. Wind tunnel testing without inlet and nozzle exhaust jet effects at lower Reynolds numbers than in-flight make it difficult to accurately assess the computational solutions of flight vehicles. A wind tunnel test in the NASA Ames 9- by 7-Foot Supersonic Wind Tunnel is planned for February 2016 to address the nozzle jet effects on sonic boom. The experiment will provide pressure signatures of test articles that replicate waveforms from aircraft wings, tails, and aft fuselage (deck) components after passing through cold nozzle jet plumes. The data will provide a variety of nozzle plume and shock interactions for comparison with computational results. A large number of high-fidelity numerical simulations of a variety of shock generators were evaluated to define a reduced collection of suitable test models. The computational results of the candidate wind tunnel test models as they evolved are summarized, and pre-test computations of the final designs are provided.

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.

Wind tunnel-sidewall-boundary-layer effects in transonic airfoil testing-some correctable, but some not

NASA Technical Reports Server (NTRS)

Lynch, F. T.; Johnson, C. B.

1988-01-01

The need to correct transonic airfoil wind tunnel test data for the influence of the tunnel sidewall boundary layers, in addition to the wall accepted corrections for the analytical investigation was carried out in order to evaluate sidewall boundary layer effects on transonic airfoil characteristics, and to validate proposed correction and the limit to their applications. This investigation involved testing of modern airfoil configurations in two different transonic airfoil test facilities, the 15 x 60 inch two-dimensional insert of the National Aeronautical Establishment (NAE) 5 foot tunnel in Ottawa, Canada, and the two-dimensional test section of the NASA Langley 0.3 m Transonic Cryogenic Tunnel (TCT). Results presented included effects of variations in sidewall-boundary layer bleed in both facilities, different sidewall boundary layer correction procedures, tunnel-to tunnel comparisons of correcte results, and flow conditions with and without separation.

Large scale wind tunnel investigation of a folding tilt rotor

NASA Technical Reports Server (NTRS)

1972-01-01

A twenty-five foot diameter folding tilt rotor was tested in a large scale wind tunnel to determine its aerodynamic characteristics in unfolded, partially folded, and fully folded configurations. During the tests, the rotor completed over forty start/stop sequences. After completing the sequences in a stepwise manner, smooth start/stop transitions were made in approximately two seconds. Wind tunnel speeds up through seventy-five knots were used, at which point the rotor mast angle was increased to four degrees, corresponding to a maneuver condition of one and one-half g.

Radionuclides in Soils Along a Mountain-Basin Transect in the Koratepa Mountains of Uzbekistan

USDA-ARS?s Scientific Manuscript database

Wind tunnels have been used for several decades to study wind erosion processes. Portable wind tunnels offer the advantage of testing natural surfaces in the field, but they must be carefully designed to insure that a logarithmic boundary layer is formed and that wind erosion processes may develop ...

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

NASA Image and Video Library

1978-04-21

Technicians examine a scale model of the space shuttle used to obtain pressure data during tests in the 10- by 10-Foot Supersonic Wind Tunnel at the National Aeronautics and Space Administration (NASA) Lewis Research Center. Lewis researchers used the 10- by 10 tunnel extensively in the 1970s to study shuttle configurations in order to forecast conditions during an actual flight. These tests included analysis of the solid rocket boosters’ aerodynamics, orbiter forebody angle -of -attack and air speed, base heating for entire shuttle, and engine-out loads. The test seen in this photograph used a 3.5- percent scale aluminum alloy model of the entire launch configuration. The program was designed to obtain aerodynamic pressure data. The tests were part of a larger program to study possible trouble areas for the shuttle’s new Advanced Flexible Reusable Surface Insulation. The researchers obtained aeroacoustic data and pressure distributions from five locations on the model. Over 100 high-temperature pressure transducers were attached to the model. Other portions of the test program were conducted at Lewis’ 8- by 6-Foot Supersonic Wind Tunnel and the 11- by 11-Foot Transonic Wind Tunnel at Ames Research Center.

Dot Projection Photogrammetric Technique for Shape Measurements of Aerospace Test Articles

NASA Technical Reports Server (NTRS)

Jones, Thomas W.; Pappa, Richard S.

2002-01-01

Results from initial laboratory investigations with the dot projection photogrammetric technique are presented for three wind-tunnel test articles with a range of surface scattering and reflection properties. These test articles are a semispan model and a micro air vehicle with a latex wing that are both diffusely reflecting, and a highly polished specularly reflecting model used for high Reynolds number testing. Results using both white light and laser illumination are presented. Some of the advantages and limitations of the dot projection technique are discussed. Although a desirable final outcome of this research effort is the characterization of dynamic behavior, only static laboratory results are presented in this preliminary effort.

40 CFR 53.40 - General provisions.

Code of Federal Regulations, 2013 CFR

2013-07-01

... 50 percent cutpoint of a test sampler shall be determined in a wind tunnel using 10 particle sizes... particle sampling effectiveness of a test sampler shall be determined in a wind tunnel using 25 µm... Protection of Environment ENVIRONMENTAL PROTECTION AGENCY (CONTINUED) AIR PROGRAMS (CONTINUED) AMBIENT AIR...

40 CFR 53.40 - General provisions.

Code of Federal Regulations, 2014 CFR

2014-07-01

... 50 percent cutpoint of a test sampler shall be determined in a wind tunnel using 10 particle sizes... particle sampling effectiveness of a test sampler shall be determined in a wind tunnel using 25 µm... Protection of Environment ENVIRONMENTAL PROTECTION AGENCY (CONTINUED) AIR PROGRAMS (CONTINUED) AMBIENT AIR...

40 CFR 53.40 - General provisions.

Code of Federal Regulations, 2012 CFR

2012-07-01

... 50 percent cutpoint of a test sampler shall be determined in a wind tunnel using 10 particle sizes... particle sampling effectiveness of a test sampler shall be determined in a wind tunnel using 25 µm... Protection of Environment ENVIRONMENTAL PROTECTION AGENCY (CONTINUED) AIR PROGRAMS (CONTINUED) AMBIENT AIR...

LNG vapor barrier and obstacle evaluation: Wind-tunnel simulation of 1987 Falcon Spill Series. Final report, July 1987-February 1991

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

Shin, S.H.; Meroney, R.N.; Neff, D.E.

1991-03-01

Measurements of the behavior of simulated liquefied natural gas clouds dispersing over small-scale model placed in environmental wind tunnels permits evaluations of the fluid physics of dense cloud movement and dispersion in a controlled environment. A large data base on the interaction of simulated LNG plumes with the Falcon test configuration of vapor barrier fences and vortex generators was obtained. The purpose of the reported test program is to provide post-field-spill wind tunnel experiments to augment the LNG Vapor Fence Field Program data obtained during the Falcon Test Series in 1987. The goal of the program is to determine themore » probable response of a dense LNG Vapor cloud to vortex inducer obstacles and fences, examine the sensitivity of results to various scaling arguments which might augment limit, or extend the value of the field and wind-tunnel tests, and identify important details of the spill behavior which were not predicted during the pretest planning phase.« less

Air Force Research Laboratory Resident Associateship Program Continuation

DTIC Science & Technology

2014-12-04

2011-7/17/2012 United States Received Veremyev, Alexander Fedorovich Pasiliao, Eduardo Lewis 8/1/2012-7/31/2013 Russia Sensors Directorate Aga...mass and damping on their modal characteristics. 5 Aerodynamic loads were estimated from the wind -tunnel test data, where the angle of attack of the... Wireless Networks; Throughput Optimization for Cognitive Radio Network with Slowly Varying Channels. 2 Capacity Optimization of MIMO Links with

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

Code of Federal Regulations, 2010 CFR

2010-07-01

... the sampler inlet opening centered in the sampling zone. To meet the maximum blockage limit of § 53.62(c)(1) or for convenience, part of the test sampler may be positioned external to the wind tunnel... = reference method sampler volumetric flow rate; and t = sampling time. (iii) Remove the reference method...

Comparison of aerodynamic coefficients obtained from theoretical calculations wind tunnel tests and flight tests data reduction for the alpha jet aircraft

NASA Technical Reports Server (NTRS)

Guiot, R.; Wunnenberg, H.

1980-01-01

The methods by which aerodynamic coefficients are determined and discussed. These include: calculations, wind tunnel experiments and experiments in flight for various prototypes of the Alpha Jet. A comparison of obtained results shows good correlation between expectations and in-flight test results.

Correlation of Fin Buffet Pressures on an F/A-18 with Scaled Wind-Tunnel Measurements

NASA Technical Reports Server (NTRS)

Moses, Robert W.; Shah, Gautam H.

1999-01-01

Buffeting is an aeroelastic phenomenon occurring at high angles of attack that plagues high performance aircraft, especially those with twin vertical tails. Previous wind-tunnel and flight tests were conducted to characterize the buffet loads on the vertical tails by measuring surface pressures, bending moments, and accelerations. Following these tests, buffeting responses were computed using the measured buffet pressures and compared to the measured buffeting responses. The calculated results did not match the measured data because the assumed spatial correlation of the buffet pressures was not correct. A better understanding of the partial (spatial) correlation of the differential buffet pressures on the tail was necessary to improve the buffeting predictions. Several wind-tunnel investigations were conducted for this purpose. When compared, the results of these tests show that the partial correlation scales with flight conditions. One of the remaining questions is whether the wind-tunnel data is consistent with flight data. Presented herein, cross-spectra and coherence functions calculated from pressures that were measured on the High Alpha Research Vehicle indicate that the partial correlation of the buffet pressures in flight agrees with the partial correlation observed in the wind tunnel.

On-road and wind-tunnel measurement of motorcycle helmet noise.

PubMed

Kennedy, J; Carley, M; Walker, I; Holt, N

2013-09-01

The noise source mechanisms involved in motorcycling include various aerodynamic sources and engine noise. The problem of noise source identification requires extensive data acquisition of a type and level that have not previously been applied. Data acquisition on track and on road are problematic due to rider safety constraints and the portability of appropriate instrumentation. One way to address this problem is the use of data from wind tunnel tests. The validity of these measurements for noise source identification must first be demonstrated. In order to achieve this extensive wind tunnel tests have been conducted and compared with the results from on-track measurements. Sound pressure levels as a function of speed were compared between on track and wind tunnel tests and were found to be comparable. Spectral conditioning techniques were applied to separate engine and wind tunnel noise from aerodynamic noise and showed that the aerodynamic components were equivalent in both cases. The spectral conditioning of on-track data showed that the contribution of engine noise to the overall noise is a function of speed and is more significant than had previously been thought. These procedures form a basis for accurate experimental measurements of motorcycle noise.

Development and testing of a unique carousel wind tunnel to experimentally determine the effect of gravity and the interparticle force on the physics of wind-blown particles

NASA Technical Reports Server (NTRS)

Leach, R. N.; Greeley, Ronald; White, Bruce R.; Iversen, James D.

1987-01-01

In the study of planetary aeolian processes the effect of gravity is not readily modeled. Gravity appears in the equations of particle motion along with the interparticle forces but the two are not separable. A wind tunnel that perimits multiphase flow experiments with wind blown particles at variable gravity was built and experiments were conducted at reduced gravity. The equations of particle motion initiation (saltation threshold) with variable gravity were experimentally verified and the interparticle force was separated. A uniquely design Carousel Wind Tunnel (CWT) allows for the long flow distance in a small sized tunnel since the test section if a continuous loop and develops the required turbulent boundary layer. A prototype model of the tunnel where only the inner drum rotates was built and tested in the KC-135 Weightless Wonder 4 zero-g aircraft. Future work includes further experiments with walnut shell in the KC-135 which sharply graded particles of widely varying median sizes including very small particles to see how interparticle force varies with particle size, and also experiments with other aeolian material.

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

NASA Technical Reports Server (NTRS)

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

1996-01-01

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

Turboprop Model in the 8- by 6-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1976-08-21

National Aeronautics and Space Administration (NASA) engineer Robert Jeracki prepares a Hamilton Standard SR-1 turboprop model in the test section of the 8- by 6-Foot Supersonic Wind Tunnel at the Lewis Research Center. Lewis researchers were analyzing a series of eight-bladed propellers in their wind tunnels to determine their operating characteristics at speeds up to Mach 0.8. The program, which became the Advanced Turboprop, was part of a NASA-wide Aircraft Energy Efficiency Program which was designed to reduce aircraft fuel costs by 50 percent. The ATP concept was different from the turboprops in use in the 1950s. The modern versions had at least eight blades and were swept back for better performance. After Lewis researchers developed the advanced turboprop theory and established its potential performance capabilities, they commenced an almost decade-long partnership with Hamilton Standard to develop, verify, and improve the concept. A series of 24-inch scale models of the SR-1 with different blade shapes and angles were tested in Lewis’ wind tunnels. A formal program was established in 1978 to examine associated noise levels, aerodynamics, and the drive system. The testing of the large-scale propfan was done on test rigs, in large wind tunnels, and, eventually, on aircraft.

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.

Cyber-Physical Systems to Understand the Dynamics of Nonlinear Aeroelastic Systems for Flexible MAVs and Energy Harvesting Applications

DTIC Science & Technology

2015-09-28

release. Rotary encoder Brushless servo motor Wind tunnel bottom wall Stainless steel shaft Shaft coupling Wind tunnel top wall Titanium flat plate...illustrating the flat plate mounted to a virtual spring-damper system in the wind tunnel test section. 2 DISTRIBUTION A: Distribution approved for...non-dimensional ratios. For example the non-dimensional stiffness, k∗ = 2k/(ρU2∞c 2h), can be kept constant even if the wind speed, U∞, chord, c, and

Application of Stereo PIV on a Supersonic Parachute Model

NASA Technical Reports Server (NTRS)

Wernet, Mark P.; Locke, Randy J.; Wroblewski, Adam; Sengupta, Anita

2009-01-01

The Mars Science Laboratory (MSL) is the next step in NASA's Mars Exploration Program, currently scheduled for 2011. The spacecraft's descent into the Martian atmosphere will be slowed from Mach 2 to subsonic speeds via a large parachute system with final landing under propulsive control. A Disk-Band-Gap (DBG) parachute will be used on MSL similar to the designs that have been used on previous missions, however; the DBG parachute used by MSL will be larger (21.5 m) than in any of the previous missions due to the weight of the payload and landing site requirements. The MSL parachute will also deploy at higher Mach number (M 2) than previous parachutes, which can lead to instabilities in canopy performance. Both the increased size of the DBG above previous demonstrated configurations and deployment at higher Mach numbers add uncertainty to the deployment, structural integrity and performance of the parachute. In order to verify the performance of the DBG on MSL, experimental testing, including acquisition of Stereo Particle Imaging Velocimetry (PIV) measurements were required for validating CFD predictions of the parachute performance. A rigid model of the DBG parachute was tested in the 10x10 foot wind tunnel at GRC. Prior to the MSL tests, a PIV system had never been used in the 10x10 wind tunnel. In this paper we discuss some of the technical challenges overcome in implementing a Stereo PIV system with a 750x400 mm field-of-view in the 10x10 wind tunnel facility and results from the MSL hardshell canopy tests.

Features of the Upgraded Imaging for Hypersonic Experimental Aeroheating Testing (IHEAT) Software

NASA Technical Reports Server (NTRS)

Mason, Michelle L.; Rufer, Shann J.

2016-01-01

The Imaging for Hypersonic Experimental Aeroheating Testing (IHEAT) software is used at the NASA Langley Research Center to analyze global aeroheating data on wind tunnel models tested in the Langley Aerothermodynamics Laboratory. One-dimensional, semi-infinite heating data derived from IHEAT are used in the design of thermal protection systems for hypersonic vehicles that are exposed to severe aeroheating loads, such as reentry vehicles during descent and landing procedures. This software program originally was written in the PV-WAVE(Registered Trademark) programming language to analyze phosphor thermography data from the two-color, relative-intensity system developed at Langley. To increase the efficiency, functionality, and reliability of IHEAT, the program was migrated to MATLAB(Registered Trademark) syntax and compiled as a stand-alone executable file labeled version 4.0. New features of IHEAT 4.0 include the options to perform diagnostic checks of the accuracy of the acquired data during a wind tunnel test, to extract data along a specified multi-segment line following a feature such as a leading edge or a streamline, and to batch process all of the temporal frame data from a wind tunnel run. Results from IHEAT 4.0 were compared on a pixel level to the output images from the legacy software to validate the program. The absolute differences between the heat transfer data output from the two programs were on the order of 10(exp -5) to 10(exp -7). IHEAT 4.0 replaces the PV-WAVE(Registered Trademark) version as the production software for aeroheating experiments conducted in the hypersonic facilities at NASA Langley.

Half Wing N219 Aircraft Model Clean Configuration for Flutter Test On Low Speed Wind Tunnel

NASA Astrophysics Data System (ADS)

Syamsuar, Sayuti; Sampurno, Budi; Mayang Mahasti, Katia; Bayu Sakti Pratama, Muchamad; Widi Sasongko, Triyono; Kartika, Nina; Suksmono, Adityo; Aji Saputro, Mohamad Ivan; Bahtera Eskayudha, Dimas

2018-04-01

Flutter is a rapid self-feeding motion which is caused by the interaction of aerodynamic, structural and inertial forces. Flutter can cause major damage on aircraft structure which can lead to fatal accident in aviation. Several methods have been evolved to avoid the flutter phenomena occur during the flight envelope of aircraft design. On this study, method was developed by Indonesian Aerospace which consist of Finite Element Method (FEM) analysis, Ground Vibration Test (GVT), and Wind Tunnel Flutter Test (WTT). Based on the study, FEM have similar results toward to Wind Tunnel Flutter Test conjunction the clean configuration of N219 aircraft half wing model.

Infrared Images of Boundary Layer Transition on the D8 Transport Configuration in the LaRC 14- by 22-Foot Subsonic Tunnel

NASA Technical Reports Server (NTRS)

Mason, Michelle L.; Gatlin, Gregory M.

2015-01-01

Grit, trip tape, or trip dots are routinely applied on the leading-edge regions of the fuselage, wings, tails or nacelles of wind tunnel models to trip the flow from laminar to turbulent. The thickness of the model's boundary layer is calculated for nominal conditions in the wind tunnel test to determine the effective size of the trip dots, but the flow over the model may not transition as intended for runs with different flow conditions. Temperature gradients measured with an infrared camera can be used to detect laminar to turbulent boundary layer transition on a wind tunnel model. This non-intrusive technique was used in the NASA Langley 14- by 22-Foot Subsonic Tunnel to visualize the behavior of the flow over a D8 transport configuration model. As the flow through the wind tunnel either increased to or decreased from the run conditions, a sufficient temperature difference existed between the air and the model to visualize the transition location (due to different heat transfer rates through the laminar and the turbulent boundary layers) for several runs in this test. Transition phenomena were visible without active temperature control in the atmospheric wind tunnel, whether the air was cooler than the model or vice-versa. However, when the temperature of the model relative to the air was purposely changed, the ability to detect transition in the infrared images was enhanced. Flow characteristics such as a wing root horseshoe vortex or the presence of fore-body vortical flows also were observed in the infrared images. The images of flow features obtained for this study demonstrate the usefulness of current infrared technology in subsonic wind tunnel tests.

Development of an intelligent hypertext system for wind tunnel testing

NASA Technical Reports Server (NTRS)

Lo, Ching F.; Shi, George Z.; Steinle, Frank W.; Wu, Y. C. L. Susan; Hoyt, W. Andes

1991-01-01

This paper summarizes the results of a system utilizing artificial intelligence technology to improve the productivity of project engineers who conduct wind tunnel tests. The objective was to create an intelligent hypertext system which integrates a hypertext manual and expert system that stores experts' knowledge and experience. The preliminary (Phase I) effort implemented a prototype IHS module encompassing a portion of the manuals and knowledge used for wind tunnel testing. The effort successfully demonstrated the feasibility of the intelligent hypertext system concept. A module for the internal strain gage balance, implemented on both IBM-PC and Macintosh computers, is presented. A description of the Phase II effort is included.

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

NASA Technical Reports Server (NTRS)

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

2013-01-01

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

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.

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.

MHK Hydrofoils Design, Wind Tunnel Optimization and CFD Analysis Report for the Aquantis 2.5MW Ocean Current Generation Device

DOE Data Explorer

Shiu, Henry; Swales, Henry; Van Damn, Case

2015-06-03

Dataset contains MHK Hydrofoils Design and Optimization and CFD Analysis Report for the Aquantis 2.5 MW Ocean Current Generation Device, as well as MHK Hydrofoils Wind Tunnel Test Plan and Checkout Test Report.

NASA's Planetary Aeolian Laboratory: Status and Update

NASA Astrophysics Data System (ADS)

Williams, D. A.; Smith, J. K.

2017-05-01

This presentation provides a status update on the operational capabilities and funding plans by NASA for the Planetary Aeolian Laboratory located at NASA Ames Research Center, including details for those proposing future wind tunnel experiments.

Static and wind tunnel near-field/far-field jet noise measurements from model scale single-flow base line and suppressor nozzles. Summary report. [conducted in the Boeing large anechoic test chamber and the NASA-Ames 40by 80-foot wind tunnel

NASA Technical Reports Server (NTRS)

Jaeck, C. L.

1977-01-01

A test program was conducted in the Boeing large anechoic test chamber and the NASA-Ames 40- by 80-foot wind tunnel to study the near- and far-field jet noise characteristics of six baseline and suppressor nozzles. Static and wind-on noise source locations were determined. A technique for extrapolating near field jet noise measurements into the far field was established. It was determined if flight effects measured in the near field are the same as those in the far field. The flight effects on the jet noise levels of the baseline and suppressor nozzles were determined. Test models included a 15.24-cm round convergent nozzle, an annular nozzle with and without ejector, a 20-lobe nozzle with and without ejector, and a 57-tube nozzle with lined ejector. The static free-field test in the anechoic chamber covered nozzle pressure ratios from 1.44 to 2.25 and jet velocities from 412 to 594 m/s at a total temperature of 844 K. The wind tunnel flight effects test repeated these nozzle test conditions with ambient velocities of 0 to 92 m/s.

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

NASA Technical Reports Server (NTRS)

Oliver, Anthony Brandon

2010-01-01

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

Propulsion and airframe aerodynamic interactions of supersonic V/STOL configurations. Volume 1: Wind tunnel test pressure data report

NASA Technical Reports Server (NTRS)

Zilz, D. E.; Devereaux, P. A.

1985-01-01

A wind tunnel model of a supersonic V/STOL fighter configuration has been tested to measure the aerodynamic interaction effects which can result from geometrically close-coupled propulsion system/airframe components. The approach was to configure the model to represent two different test techniques. One was a conventional test technique composed of two test modes. In the Flow-Through mode, absolute configuration aerodynamics are measured, including inlet/airframe interactions. In the Jet-Effects mode, incremental nozzle/airframe interactions are measured. The other test technique is a propulsion simulator approach, where a sub-scale, externally powered engine is mounted in the model. This allows proper measurement of inlet/airframe and nozzle/airframe interactions simultaneously. This is Volume 1 of 2: Wind Tunnel Test Pressure Data Report.

Effect of thermal stability/complex terrain on wind turbine model(s): a wind tunnel study to address complex atmospheric conditions

NASA Astrophysics Data System (ADS)

Guala, M.; Hu, S. J.; Chamorro, L. P.

2011-12-01

Turbulent boundary layer measurements in both wind tunnel and in the near-neutral atmospheric surface layer revealed in the last decade the significant contribution of the large scales of motions to both turbulent kinetic energy and Reynolds stresses, for a wide range of Reynolds number. These scales are known to grow throughout the logarithmic layer and to extend several boundary layer heights in the streamwise direction. Potentially, they are a source of strong unsteadiness in the power output of wind turbines and in the aerodynamic loads of wind turbine blades. However, the large scales in realistic atmospheric conditions deserves further study, with well controlled boundary conditions. In the atmospheric wind tunnel of the St. Anthony Falls Laboratory, with a 16 m long test section and independently controlled incoming flow and floor temperatures, turbulent boundary layers in a range of stability conditions, from the stratified to the convective case, can be reproduced and monitored. Measurements of fluctuating temperature, streamwise and wall normal velocity components are simultaneously obtained by an ad hoc calibrated and customized triple-wire sensor. A wind turbine model with constant loading DC motor, constant tip speed ratio, and a rotor diameter of 0.128m is used to mimic a large full scale turbine in the atmospheric boundary layer. Measurements of the fluctuating voltage generated by the DC motor are compared with measurements of the blade's angular velocity by laser scanning, and eventually related to velocity measurements from the triple-wire sensor. This study preliminary explores the effect of weak stability and complex terrain (through a set of spanwise aligned topographic perturbations) on the large scales of the flow and on the fluctuations in the wind turbine(s) power output.

Ice Accretions and Icing Effects for Modern Airfoils

NASA Technical Reports Server (NTRS)

Addy, Harold E., Jr.

2000-01-01

Icing tests were conducted to document ice shapes formed on three different two-dimensional airfoils and to study the effects of the accreted ice on aerodynamic performance. The models tested were representative of airfoil designs in current use for each of the commercial transport, business jet, and general aviation categories of aircraft. The models were subjected to a range of icing conditions in an icing wind tunnel. The conditions were selected primarily from the Federal Aviation Administration's Federal Aviation Regulations 25 Appendix C atmospheric icing conditions. A few large droplet icing conditions were included. To verify the aerodynamic performance measurements, molds were made of selected ice shapes formed in the icing tunnel. Castings of the ice were made from the molds and placed on a model in a dry, low-turbulence wind tunnel where precision aerodynamic performance measurements were made. Documentation of all the ice shapes and the aerodynamic performance measurements made during the icing tunnel tests is included in this report. Results from the dry, low-turbulence wind tunnel tests are also presented.

The 6-foot-4-inch Wind Tunnel at the Washington Navy Yard

NASA Technical Reports Server (NTRS)

Desmond, G L; Mccrary, J A

1935-01-01

The 6-foot-4-inch wind tunnel and its auxiliary equipment has proven itself capable of continuous and reliable output of data. The real value of the tunnel will increase as experience is gained in checking the observed tunnel performance against full-scale performance. Such has been the case of the 8- by 8-foot tunnel, and for that reason the comparison in the calibration tests have been 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.

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

NASA Technical Reports Server (NTRS)

Flemming, R.; Ruddell, A.

1974-01-01

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

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.

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

NASA Technical Reports Server (NTRS)

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

2015-01-01

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

Suppression of background noise in a transonic wind-tunnel test section

NASA Technical Reports Server (NTRS)

Schutzenhofer, L. A.; Howard, P. W.

1975-01-01

Some exploratory tests were recently performed in the transonic test section of the NASA Marshall Space Flight Center 14-in. wind tunnel to suppress the background noise. In these tests, the perforated walls of the test section were covered with fine wire screens. The screens eliminated the edge tones generated by the holes in the perforated walls and significantly reduced the tunnel background noise. The tunnel noise levels were reduced to such a degree by this simple modification at Mach numbers 0.75, 0.9, 1.1, 1.2, and 1.46 that the fluctuating pressure levels of a turbulent boundary layer could be measured on a 5-deg half-angle cone.

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.

FUN3D Airload Predictions for the Full-Scale UH-60A Airloads Rotor in a Wind Tunnel

NASA Technical Reports Server (NTRS)

Lee-Rausch, Elizabeth M.; Biedron, Robert T.

2013-01-01

An unsteady Reynolds-Averaged Navier-Stokes solver for unstructured grids, FUN3D, is used to compute the rotor performance and airloads of the UH-60A Airloads Rotor in the National Full-Scale Aerodynamic Complex (NFAC) 40- by 80-foot Wind Tunnel. The flow solver is loosely coupled to a rotorcraft comprehensive code, CAMRAD-II, to account for trim and aeroelastic deflections. Computations are made for the 1-g level flight speed-sweep test conditions with the airloads rotor installed on the NFAC Large Rotor Test Apparatus (LRTA) and in the 40- by 80-ft wind tunnel to determine the influence of the test stand and wind-tunnel walls on the rotor performance and airloads. Detailed comparisons are made between the results of the CFD/CSD simulations and the wind tunnel measurements. The computed trends in solidity-weighted propulsive force and power coefficient match the experimental trends over the range of advance ratios and are comparable to previously published results. Rotor performance and sectional airloads show little sensitivity to the modeling of the wind-tunnel walls, which indicates that the rotor shaft-angle correction adequately compensates for the wall influence up to an advance ratio of 0.37. Sensitivity of the rotor performance and sectional airloads to the modeling of the rotor with the LRTA body/hub increases with advance ratio. The inclusion of the LRTA in the simulation slightly improves the comparison of rotor propulsive force between the computation and wind tunnel data but does not resolve the difference in the rotor power predictions at mu = 0.37. Despite a more precise knowledge of the rotor trim loads and flight condition, the level of comparison between the computed and measured sectional airloads/pressures at an advance ratio of 0.37 is comparable to the results previously published for the high-speed flight test condition.

Fan Blade Shake Test Results for the 40- by 80-/80- by 120-Foot Wind Tunnel

NASA Technical Reports Server (NTRS)

Warmbrodt, W.; Graham, T.

1983-01-01

This report documents the shake tests performed on the first set of hydulignum fan blades for the 40- by 80-/80- by 120-Foot Wind Tunnel. The purpose of the shake test program is described. The test equipment and test procedures are reviewed. Results from each shake test are presented and the overall findings of the shake test program are discussed.

Orbiter entry leeside heat-transfer data analysis

NASA Technical Reports Server (NTRS)

Throckmorton, D. A.; Zoby, E. V.

1983-01-01

Heat-transfer data measured along the Space Shuttle Orbiter's leeward centerline and over the wing leeside surface during the STS-2 and STS-3 mission entries are presented. The flight data are compared with available wind-tunnel results. Flight heating levels are, in general, lower than those which are inferred from the wind-tunnel results. This result is apparently due to the flight leeside flowfield remaining laminar over a larger Reynolds number range than that of corresponding ground test results. The flight/wind-tunnel data comparisons confirm the adequacy of, and conservatism embodied in, the direct application of wind-tunnel data at flight conditions for the design of Orbiter leeside thermal protection.

2. EXTERIOR VIEW OF BUILDING 25B (TEST CHAMBER BUILDING) AND ...

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

2. EXTERIOR VIEW OF BUILDING 25B (TEST CHAMBER BUILDING) AND 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

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

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

NASA Technical Reports Server (NTRS)

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

1978-01-01

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

Mercury Retrorocket Tests

NASA Image and Video Library

1960-04-01

The retrorockets, which would be fired to return the Mercury Capsule to earth's atmosphere, were tested inside the Altitude Wind Tunnel in April 1960. A retrograde thrust stand was erected in the C corner in the wide end of the Altitude Wind Tunnel. This film contains footage from above and the side of the retrorockets firing.

The requirements for a new full scale subsonic wind tunnel

NASA Technical Reports Server (NTRS)

Kelly, M. W.; Mckinney, M. O.; Luidens, R. W.

1972-01-01

Justification and requirements are presented for a large subsonic wind tunnel capable of testing full scale aircraft, rotor systems, and advanced V/STOL propulsion systems. The design considerations and constraints for such a facility are reviewed, and the trades between facility test capability and costs are discussed.

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

NASA Technical Reports Server (NTRS)

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

1979-01-01

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

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

NASA Technical Reports Server (NTRS)

Kemp, W. B., Jr.

1984-01-01

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

ARC-1957-A-23438

NASA Image and Video Library

1957-12-30

H. Julian 'Harvey' Allen in front of the NASA Ames 8_x_7 foot Supersonic Wind Tunnel test section. A blunt body model mounted in the test section is ready for testing . The 8_X_7_foot is part of the Unitary Plan WInd Tunnel Complex Note: printed in 60 year at NASA Ames Research Center by Glenn Bugos NASA SP-2000-4314

Highlights of experience with a flexible walled test section in the NASA Langley 0.3-meter transonic cryogenic tunnel

NASA Technical Reports Server (NTRS)

Wolf, Stephen W. D.; Ray, Edward J.

1988-01-01

The unique combination of adaptive wall technology with a contonuous flow cryogenic wind tunnel is described. This powerful combination allows wind tunnel users to carry out 2-D tests at flight Reynolds numbers with wall interference essentially eliminated. Validation testing was conducted to support this claim using well tested symmetrical and cambered airfoils at transonic speeds and high Reynolds numbers. The test section hardware has four solid walls, with the floor and ceiling flexible. The method of adapting/shaping the floor and ceiling to eliminate top and bottom wall interference at its source is outlined. Data comparisons for different size models tested and others in several sophisticated 2-D wind tunnels are made. In addition, the effects of Reynolds number, testing at high lift with associated large flexible wall movements, the uniqueness of the adapted wall shapes, and the effects of sidewall boundary layer control are examined. The 0.3-m TCT is now the most advanced 2-D research facility anywhere.

Potential benefits of magnetic suspension and balance systems

NASA Technical Reports Server (NTRS)

Lawing, Pierce L.; Dress, David A.; Kilgore, Robert A.

1987-01-01

The potential of Magnetic Suspension and Balance Systems (MSBS) to improve conventional wind tunnel testing techniques is discussed. Topics include: elimination of model geometry distortion and support interference to improve the measurement accuracy of aerodynamic coefficients; removal of testing restrictions due to supports; improved dynamic stability data; and stores separation testing. Substantial increases in wind tunnel productivity are anticipated due to the coalescence of these improvements. Specific improvements in testing methods for missiles, helicopters, fighter aircraft, twin fuselage transports and bombers, state separation, water tunnels, and automobiles are also forecast. In a more speculative vein, new wind tunnel test techniques are envisioned as a result of applying MSBS, including free-flight computer trajectories in the test section, pilot-in-the-loop and designer-in-the-loop testing, shipboard missile launch simulation, and optimization of hybrid hypersonic configurations. Also addressed are potential applications of MSBS to such diverse technologies as medical research and practice, industrial robotics, space weaponry, and ore processing in space.

Residual interference and wind tunnel wall adaption

NASA Technical Reports Server (NTRS)

Mokry, Miroslav

1989-01-01

Measured flow variables near the test section boundaries, used to guide adjustments of the walls in adaptive wind tunnels, can also be used to quantify the residual interference. Because of a finite number of wall control devices (jacks, plenum compartments), the finite test section length, and the approximation character of adaptation algorithms, the unconfined flow conditions are not expected to be precisely attained even in the fully adapted stage. The procedures for the evaluation of residual wall interference are essentially the same as those used for assessing the correction in conventional, non-adaptive wind tunnels. Depending upon the number of flow variables utilized, one can speak of one- or two-variable methods; in two dimensions also of Schwarz- or Cauchy-type methods. The one-variable methods use the measured static pressure and normal velocity at the test section boundary, but do not require any model representation. This is clearly of an advantage for adaptive wall test section, which are often relatively small with respect to the test model, and for the variety of complex flows commonly encountered in wind tunnel testing. For test sections with flexible walls the normal component of velocity is given by the shape of the wall, adjusted for the displacement effect of its boundary layer. For ventilated test section walls it has to be measured by the Calspan pipes, laser Doppler velocimetry, or other appropriate techniques. The interface discontinuity method, also described, is a genuine residual interference assessment technique. It is specific to adaptive wall wind tunnels, where the computation results for the fictitious flow in the exterior of the test section are provided.

Wall interaction effects for a full-scale helicopter rotor in the NASA Ames 80- by 120-foot wind tunnel

NASA Technical Reports Server (NTRS)

Shinoda, Patrick M.

1994-01-01

A full-scale helicopter rotor test was conducted in the NASA Ames 80- by 120-Foot Wind Tunnel with a four-bladed S-76 rotor system. This wind tunnel test generated a unique and extensive data base covering a wide range of rotor shaft angles-of-attack and rotor thrust conditions from 0 to 100 knots. Three configurations were tested: (1) empty tunnel; (2) test stand body (fuselage) and support system; and (3) fuselage and support system with rotor installed. Empty tunnel wall pressure data are evaluated as a function of tunnel speed to understand the baseline characteristics. Aerodynamic interaction effects between the fuselage and the walls of the tunnel are investigated by comparing wall, ceiling, and floor pressures for various tunnel velocities and fuselage angles-of-attack. Aerodynamic interaction effects between the rotor and the walls of the tunnel are also investigated by comparing wall, ceiling, and floor pressures for various rotor shaft angles, rotor thrust conditions, and tunnel velocities. Empty tunnel wall pressure data show good repeatability and are not affected by tunnel speed. In addition, the tunnel wall pressure profiles are not affected by the presence of the fuselage apart from a pressure shift. Results do not indicate that the tunnel wall pressure profiles are affected by the presence of the rotor. Significant changes in the wall, ceiling, and floor pressure profiles occur with changing tunnel speeds for constant rotor thrust and shaft angle conditions. Significant changes were also observed when varying rotor thrust or rotor shaft angle-of-attack. Other results indicate that dynamic rotor loads and blade motion are influenced by the presence of the tunnel walls at very low tunnel velocity and, together with the wall pressure data, provide a good indication of flow breakdown.

Project FIRE Flight Investigation Reentry Environment- Winds of Change

NASA Image and Video Library

1962-11-21

As part of the project FIRE study, technicians ready materials to be subjected to high temperatures that will simulate the effects of re-entry heating. Tests of various space capsule materials for Project FIRE were conducted. Photographed in the 9 X 6 Foot Thermal Structures Tunnel. Photograph published in Winds of Change, 75th Anniversary NASA publication, by James Schultz (page 78). Photograph also published in Engineer in Charge: A History of the Langley Aeronautical Laboratory, 1917-1958 by James R. Hansen (page 476). Also Published in the book " A Century at Langley" by Joseph Chambers. Pg. 92

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

NASA Technical Reports Server (NTRS)

Kilgore, R. A.

1974-01-01

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

Description and evaluation of an interference assessment for a slotted-wall wind tunnel

NASA Technical Reports Server (NTRS)

Kemp, William B., Jr.

1991-01-01

A wind-tunnel interference assessment method applicable to test sections with discrete finite-length wall slots is described. The method is based on high order panel method technology and uses mixed boundary conditions to satisfy both the tunnel geometry and wall pressure distributions measured in the slotted-wall region. Both the test model and its sting support system are represented by distributed singularities. The method yields interference corrections to the model test data as well as surveys through the interference field at arbitrary locations. These results include the equivalent of tunnel Mach calibration, longitudinal pressure gradient, tunnel flow angularity, wall interference, and an inviscid form of sting interference. Alternative results which omit the direct contribution of the sting are also produced. The method was applied to the National Transonic Facility at NASA Langley Research Center for both tunnel calibration tests and tests of two models of subsonic transport configurations.

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.

Airfoil flutter model suspension system

NASA Technical Reports Server (NTRS)

Reed, Wilmer H. (Inventor)

1987-01-01

A wind tunnel suspension system for testing flutter models under various loads and at various angles of attack is described. The invention comprises a mounting bracket assembly affixing the suspension system to the wind tunnel, a drag-link assembly and a compound spring arrangement comprises a plunge spring working in opposition to a compressive spring so as to provide a high stiffness to trim out steady state loads and simultaneously a low stiffness to dynamic loads. By this arrangement an airfoil may be tested for oscillatory response in both plunge and pitch modes while being held under high lifting loads in a wind tunnel.

A theory for predicting boundary impedance and resonance frequencies of slotted-wall wind tunnels, including plenum effects

NASA Technical Reports Server (NTRS)

Barger, R. L.

1981-01-01

Wave-induced resonance associated with the geometry of wind-tunnel test sections can occur. A theory that uses acoustic impedance concepts to predict resonance modes in a two dimensional, slotted wall wind tunnel with a plenum chamber is described. The equation derived is consistent with known results for limiting conditions. The computed resonance modes compare well with appropriate experimental data. When the theory is applied to perforated wall test sections, it predicts the experimentally observed closely spaced modes that occur when the wavelength is not long compared with he plenum depth.

Development of an apparatus to measure thermophysical properties of wind tunnel heat transfer models

NASA Technical Reports Server (NTRS)

Romanowski, R. F.; Steinberg, I. H.

1974-01-01

The apparatus and technique for measuring the thermophysical properties of models used with the phase-change paint method for obtaining wind tunnel heat transfer data are described. The method allows rapid measurement of the combined properties in a transient manner similar to an actual wind tunnel test. An effective value of the thermophysical properties can be determined which accounts for changes in thermal properties with temperature or with depth into the model surface. The apparatus was successfully tested at various heating rates between 19,000 and 124,000 watts per square meter.

General Dynamics YF-16 Model in the 8- by 6-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1974-01-21

A model of the General Dynamics YF-16 Fighting Falcon in the test section of the 8- by 6-Foot Supersonic Wind Tunnel at the National Aeronautics and Space Administration (NASA) Lewis Research Center. The YF-16 was General Dynamics response to the military’s 1972 request for proposals to design a new 20,000-pound fighter jet with exceptional acceleration, turn rate, and range. The aircraft included innovative design elements to help pilots survive turns up to 9Gs, a new frameless bubble canopy, and a Pratt and Whitney 24,000-pound thrust F-100 engine. The YF-16 made its initial flight in February 1974, just six weeks before this photograph, at Edwards Air Force Base. Less than a year later, the Air Force ordered 650 of the aircraft, designated as F-16 Fighting Falcons. The March and April 1974 tests in the 8- by 6-foot tunnel analyzed the aircraft’s fixed-shroud ejector nozzle. The fixed-nozzle area limited drag, but also limited the nozzle’s internal performance. NASA researchers identified and assessed aerodynamic and aerodynamic-propulsion interaction uncertainties associated the prototype concept. YF-16 models were also tested extensively in the 11- by 11-Foot Transonic Wind Tunnel and 9- by 7-Foot Supersonic Wind Tunnel at Ames Research Center and the 12-Foot Pressure Wind Tunnel at Langley Research Center.

Results of test MA22 in the NASA/LaRC 31-inch CFHT on an 0.010-scale model (32-0) of the space shuttle configuration 3 to determine RCS jet flow field interaction, volume 1. [wind tunnel tests for interactions of aerodynamic heating on jet flow

NASA Technical Reports Server (NTRS)

Kanipe, D. B.

1976-01-01

A wind tunnel test was conducted in the Langley Research Center 31-inch Continuous Flow Hypersonic Wind Tunnel from May 6, 1975 through June 3, 1975. The primary objectives of this test were the following: (1) to study the ability of the wind tunnel to repeat, on a run-to-run basis, data taken for identical configurations to determine if errors in repeatability could have a significant effect on jet interaction data, (2) to determine the effect of aerodynamic heating of the scale model on jet interaction, (3) to investigate the effects of elevon and body flap deflections on jet interaction, (4) to determine if the effects from jets fired separately along different axes can be added to equal the effects of the jets fired simultaneously (super position effects), (5) to study multiple jet effects, and (6) to investigate area ratio effects, i.e., the effect on jet interaction measurements of using wind tunnel nozzles with different area ratios in the same location. The model used in the test was a .010-scale model of the Space Shuttle Orbiter Configuration 3. The test was conducted at Mach 10.3 and a dynamic pressure of 150 psf. RCS chamber pressure was varied to simulate free flight dynamic pressures of 5, 7.5, 10, and 20 psf.

NASA Researcher Examines an Aircraft Model with a Four-Fan Thrust Reverser

NASA Image and Video Library

1972-03-21

National Aeronautics and Space Administration (NASA) researcher John Carpenter inspects an aircraft model with a four-fan thrust reverser which would be studied in the 9- by 15-Foot Low Speed Wind Tunnel at the Lewis Research Center. Thrust reversers were introduced in the 1950s as a means for slowing high-speed jet aircraft during landing. Engineers sought to apply the technology to Vertical and Short Takeoff and Landing (VSTOL) aircraft in the 1970s. The new designs would have to take into account shorter landing areas, noise levels, and decreased thrust levels. A balance was needed between the thrust reverser’s efficiency, its noise generation, and the engine’s power setting. This model underwent a series of four tests in the 9- by 15-foot tunnel during April and May 1974. The model, with a high-wing configuration and no tail, was equipped with four thrust-reverser engines. The investigations included static internal aerodynamic tests on a single fan/reverser, wind tunnel isolated fan/reverser thrust tests, installation effects on a four-fan airplane model in a wind tunnel, and single reverser acoustic tests. The 9-by 15 was built inside the return leg of the 8- by 6-Foot Supersonic Wind Tunnel in 1968. The facility generates airspeeds from 0 to 175 miles per hour to evaluate the aerodynamic performance and acoustic characteristics of nozzles, inlets, and propellers, and investigate hot gas re-ingestion of advanced VSTOL concepts. John Carpenter was a technician in the Wind Tunnels Service Section of the Test Installations Division.

Traveling Crossow Instability for HIFiRE-5 in a Quiet Hypersonic Wind Tunnel (Postprint)

DTIC Science & Technology

2013-06-01

scale model of the 2:1 elliptic cone HIFiRE-5 flight vehicle was used to investigate the traveling crossflow instability at Mach 6 in Purdue...Force Research Laboratory, Air Vehicles Directorate 2130 8th St., WPAFB, OH 45433-7542, USA Abstract A scale model of the 2:1 elliptic cone HIFiRE-5...flight vehicle was used to investigate the traveling crossflow instability at Mach 6 in Purdue University’s Mach-6 quiet wind tunnel. Traveling crossflow

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

NASA Technical Reports Server (NTRS)

Kemp, W. B., Jr.

1980-01-01

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

LACIS-T - A moist air wind tunnel for investigating the interactions between cloud microphysics and turbulence

NASA Astrophysics Data System (ADS)

Niedermeier, Dennis; Voigtländer, Jens; Siebert, Holger; Desai, Neel; Shaw, Raymond; Chang, Kelken; Krueger, Steven; Schumacher, Jörg; Stratmann, Frank

2017-11-01

Turbulence - cloud droplet interaction processes have been investigated primarily through numerical simulation and field measurements over the last ten years. However, only in the laboratory we can be confident in our knowledge of initial and boundary conditions, and are able to measure for extended times under statistically stationary and repeatable conditions. Therefore, the newly built turbulent wind tunnel LACIS-T (Turbulent Leipzig Aerosol Cloud Interaction Simulator) is an ideal facility for pursuing mechanistic understanding of these processes. Within the tunnel we are able to adjust precisely controlled turbulent temperature and humidity fields so as to achieve supersaturation levels allowing for detailed investigations of the interactions between cloud microphysical processes (e.g., cloud droplet activation) and the turbulent flow, under well-defined and reproducible laboratory conditions. We will present the fundamental operating principle, first results from ongoing characterization efforts, numerical simulations as well as first droplet activation experiments.

NASA Lewis 8- by 6-foot supersonic wind tunnel user manual

NASA Technical Reports Server (NTRS)

Soeder, Ronald H.

1993-01-01

The 8- by 6-Foot Supersonic Wind Tunnel (SWT) at Lewis Research Center is available for use by qualified researchers. This manual contains tunnel performance maps which show the range of total temperature, total pressure, static pressure, dynamic pressure, altitude, Reynolds number, and mass flow as a function of test section Mach number. These maps are applicable for both the aerodynamic and propulsion cycle. The 8- by 6-Foot Supersonic Wind Tunnel is an atmospheric facility with a test section Mach number range from 0.36 to 2.0. General support systems (air systems, hydraulic system, hydrogen system, infrared system, laser system, laser sheet system, and schlieren system are also described as are instrumentation and data processing and acquisition systems. Pretest meeting formats are outlined. Tunnel user responsibility and personal safety requirements are also stated.

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.

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

NASA Technical Reports Server (NTRS)

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

2016-01-01

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

On-line analysis capabilities developed to support the AFW wind-tunnel tests

NASA Technical Reports Server (NTRS)

Wieseman, Carol D.; Hoadley, Sherwood T.; Mcgraw, Sandra M.

1992-01-01

A variety of on-line analysis tools were developed to support two active flexible wing (AFW) wind-tunnel tests. These tools were developed to verify control law execution, to satisfy analysis requirements of the control law designers, to provide measures of system stability in a real-time environment, and to provide project managers with a quantitative measure of controller performance. Descriptions and purposes of the developed capabilities are presented along with examples. Procedures for saving and transferring data for near real-time analysis, and descriptions of the corresponding data interface programs are also presented. The on-line analysis tools worked well before, during, and after the wind tunnel test and proved to be a vital and important part of the entire test effort.

Spectral Analysis and Experimental Modeling of Ice Accretion Roughness

NASA Technical Reports Server (NTRS)

Orr, D. J.; Breuer, K. S.; Torres, B. E.; Hansman, R. J., Jr.

1996-01-01

A self-consistent scheme for relating wind tunnel ice accretion roughness to the resulting enhancement of heat transfer is described. First, a spectral technique of quantitative analysis of early ice roughness images is reviewed. The image processing scheme uses a spectral estimation technique (SET) which extracts physically descriptive parameters by comparing scan lines from the experimentally-obtained accretion images to a prescribed test function. Analysis using this technique for both streamwise and spanwise directions of data from the NASA Lewis Icing Research Tunnel (IRT) are presented. An experimental technique is then presented for constructing physical roughness models suitable for wind tunnel testing that match the SET parameters extracted from the IRT images. The icing castings and modeled roughness are tested for enhancement of boundary layer heat transfer using infrared techniques in a "dry" wind tunnel.

Application of Wind Tunnel Free-Flight Technique for Wake Vortex Encounters

NASA Technical Reports Server (NTRS)

Brandon, Jay M.; Jordan, Frank L., Jr.; Stuever, Robert A.; Buttrill, Catherine W.

1997-01-01

A wind tunnel investigation was conducted in the Langley 30- by 60-Foot Tunnel to assess the free-flight test technique as a tool in research on wake vortex encounters. A typical 17.5-percent scale business-class jet airplane model was flown behind a stationary wing mounted in the forward portion of the wind tunnel test section. The span ratio (model span-generating wingspan) was 0.75. The wing angle of attack could be adjusted to produce a vortex of desired strength. The test airplane model was successfully flown in the vortex and through the vortex for a range of vortex strengths. Data obtained included the model airplane body axis accelerations, angular rates, attitudes, and control positions as a function of vortex strength and relative position. Pilot comments and video records were also recorded during the vortex encounters.

Wind Tunnel Test Technique and Instrumentation Development at LaRC

NASA Technical Reports Server (NTRS)

Putnam, Lawrence E.

1999-01-01

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

Some theoretical considerations of longitudinal stability in power-on flight with special reference to wind-tunnel testing, November 1942

NASA Technical Reports Server (NTRS)

Donlan, C. J.

1976-01-01

Some problems relating to longitudinal stability in power-on flight are considered. A derivation is included which shows that, under certain conditions, the rate of change of the pitching moment coefficient with lift coefficient as obtained in wind tunnel tests simulating constant power operation is directly proportional to one of the indices of stability commonly associated with flight analysis, (the slope of the curve relating the elevator angle for trim and lift coefficient). The necessity of analyzing power-on wind tunnel data for trim conditions is emphasized, and a method is provided for converting data obtained from constant thrust tests to simulated constant throttle flight conditions.

U.S. aerospace industry opinion of the effect of computer-aided prediction-design technology on future wind-tunnel test requirements for aircraft development programs

NASA Technical Reports Server (NTRS)

Treon, S. L.

1979-01-01

A survey of the U.S. aerospace industry in late 1977 suggests that there will be an increasing use of computer-aided prediction-design technology (CPD Tech) in the aircraft development process but that, overall, only a modest reduction in wind-tunnel test requirements from the current level is expected in the period through 1995. Opinions were received from key spokesmen in 23 of the 26 solicited major companies or corporate divisions involved in the design and manufacture of nonrotary wing aircraft. Development programs for nine types of aircraft related to test phases and wind-tunnel size and speed range were considered.