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.

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.

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.

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.

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

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...

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.

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.

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.

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.

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

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.

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.

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.

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.

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.

Advancing Test Capabilities at NASA Wind Tunnels

NASA Technical Reports Server (NTRS)

Bell, James

2015-01-01

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

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.

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.

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

NASA Technical Reports Server (NTRS)

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

2014-01-01

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

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.

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

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.

Correlation of the Drag Characteristics of a Typical Pursuit Airplane Obtained from High-Speed Wind-Tunnel and Flight Tests

NASA Technical Reports Server (NTRS)

Nissen, James M; Gadebero, Burnett L; Hamilton, William T

1948-01-01

In order to obtain a correlation of drag data from wind-tunnel and flight tests at high Mach numbers, a typical pursuit airplane, with the propeller removed, was tested in flight at Mach numbers up to 0.755, and the results were compared with wind-tunnel tests of a 1/3-scale model of the airplane. The tests results show that the drag characteristics of the test airplane can be predicted with satisfactory accuracy from tests in the Ames 16-foot high-speed wind tunnel of the Ames Aeronautical Laboratory at both high and low Mach numbers. It is considered that this result is not unique with the airplane.

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.

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.

Aeroservoelastic Wind-Tunnel Tests of a Free-Flying, Joined-Wing SensorCraft Model for Gust Load Alleviation

NASA Technical Reports Server (NTRS)

Scott, Robert C.; Castelluccio, Mark A.; Coulson, David A.; Heeg, Jennifer

2011-01-01

A team comprised of the Air Force Research Laboratory (AFRL), Boeing, and the NASA Langley Research Center conducted three aeroservoelastic wind-tunnel tests in the Transonic Dynamics Tunnel to demonstrate active control technologies relevant to large, exible vehicles. In the first of these three tests, a full-span, aeroelastically scaled, wind-tunnel model of a joined-wing SensorCraft vehicle was mounted to a force balance to acquire a basic aerodynamic data set. In the second and third tests, the same wind-tunnel model was mated to a new, two-degree-of-freedom, beam mount. This mount allowed the full-span model to translate vertically and pitch. Trimmed flight at -10% static margin and gust load alleviation were successfully demonstrated. The rigid body degrees of freedom required that the model be own in the wind tunnel using an active control system. This risky mode of testing necessitated that a model arrestment system be integrated into the new mount. The safe and successful completion of these free-flying tests required the development and integration of custom hardware and software. This paper describes the many systems, software, and procedures that were developed as part of this effort. The balance and free ying wind-tunnel tests will be summarized. The design of the trim and gust load alleviation control laws along with the associated results will also be discussed.

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.

Two-dimensional wind tunnel

NASA Technical Reports Server (NTRS)

1982-01-01

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

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.

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.

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.

Recent research on V/STOL test limits at the University of Washington aeronautical laboratory

NASA Technical Reports Server (NTRS)

Shindo, S.; Rae, W. H., Jr.

1980-01-01

The occurence of flow breakdown during the wind tunnel testing of a powered V/STOL aircraft was studied. Flow breakdown is the low forward speed test limit in a solid wall wind tunnel and is characterized by a vortex which forms on the floor and walls of the wind tunnel thereby failing to simulate free air conditions. The flow is caused by the interaction of the model wake and tunnel boundary layer and affects the model's aerodynamic characteristics in such fashion as to negate their reliability as correctable wind tunnel data. The low speed test limit was examined using a model that possessed a discretely concentrated powered lift system using a pair of lift jets. The system design is discussed and the tests and results which show that flow breakdown occurs at a velocity ratio of approximately 0.20 are reported.

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.

Experimental Results from the Active Aeroelastic Wing Wind Tunnel Test Program

NASA Technical Reports Server (NTRS)

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

2005-01-01

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

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 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.

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...

Hardening Doppler Global Velocimetry Systems for Large Wind Tunnel Applications

NASA Technical Reports Server (NTRS)

Meyers, James F.; Lee, Joseph W.; Fletcher, Mark T.; South, Bruce W.

2004-01-01

The development of Doppler Global Velocimetry from a laboratory curiosity to a wind tunnel instrumentation system is discussed. This development includes system advancements from a single velocity component to simultaneous three components, and from a steady state to instantaneous measurement. Improvements to system control and stability are discussed along with solutions to real world problems encountered in the wind tunnel. This on-going development program follows the cyclic evolution of understanding the physics of the technology, development of solutions, laboratory and wind tunnel testing, and reevaluation of the physics based on the test results.

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.

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.

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.

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.

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.

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.

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.

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.

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.

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 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.

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.

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.

Wind Tunnel Complex at the Aircraft Engine Research Laboratory

NASA Image and Video Library

1945-09-21

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

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.

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.

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.

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.

Michigan/Air Force Research Laboratory (AFRL) Collaborative Center in Aeronautical Sciences (MACCAS)

DTIC Science & Technology

2013-09-01

Interactions - PIV Database for the Second SBLI Workshop”  “Design of a Glass Supersonic Wind Tunnel Experiment for Mixed Compression Inlet Investigations...or small-scale wind tunnel tests. Some of the discipline components have also been compared against well-established numerical solutions (e.g...difficult to test in a wind tunnel environment. The choice of construction, materials, and geometry were such that they allow accurate characterization of

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.

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.

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.

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.

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.

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.

Evaluation of an I-box wind tunnel model for assessment of behavioral responses of blow flies.

PubMed

Moophayak, Kittikhun; Sukontason, Kabkaew L; Kurahashi, Hiromu; Vogtsberger, Roy C; Sukontason, Kom

2013-11-01

The behavioral response of flies to olfactory cues remains the focus of many investigations, and wind tunnels have sometimes been employed for assessment of this variable in the laboratory. In this study, our aim was to design, construct, and operate a new model of I-box wind tunnel with improved efficacy, highlighting the use of a new wind tunnel model to investigate the behavioral response of the medically important blow fly, Chrysomya megacephala (Fabricius). The I-box dual-choice wind tunnel designed for this study consists of seven conjoined compartments that resulted in a linear apparatus with clear glass tunnel of 30 × 30 × 190 cm ended both sides with wooden "fan compartments" which are equipped with adjustable fans as wind source. The clear glass tunnel consisted of two "stimulus compartments" with either presence or absence (control) of bait; two "trap compartments" where flies were attracted and allowed to reside; and one central "release compartment" where flies were introduced. Wind tunnel experiments were carried out in a temperature-controlled room, with a room light as a light source and a room-ventilated fan as odor-remover from tunnel out. Evaluation of testing parameters revealed that the highest attractive index was achieved with the use of 300 g of 1-day tainted pork scrap (pork meat mixed with offal) as bait in wind tunnel settings wind speed of 0.58 m/s, during 1.00-5.00 PM with light intensity of 341.33 lux from vertical light and 135.93 lux from horizontal light for testing a group of 60 flies. In addition, no significant response of well-fed and 24 h staved flies to this bait under these conditions was found. Results of this study supported this new wind tunnel model as a suitable apparatus for investigation of behavioral response of blow flies to bait chemical cues in the laboratory.

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

Force instrumentation for cryogenic wind tunnels using one-piece strain-gage balances

NASA Technical Reports Server (NTRS)

Ferris, A. T.

1980-01-01

The use of cryogenic temperatures in wind tunnels to achieve high Reynolds numbers has imposed a harsh operating environment on the force balance. Laboratory tests were conducted to study the effect cryogenic temperatures have on balance materials, gages, wiring, solder, adhesives, and moisture proofing. Wind tunnel tests were conducted using a one piece three component balance to verify laboratory results. These initial studies indicate that satisfactory force data can be obtained under steady state conditions.

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.

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.

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.

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.

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.

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.

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.

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

NASA Technical Reports Server (NTRS)

Lewis, M. C.

1988-01-01

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

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.

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.

Testing a Parachute for Mars in World Largest Wind Tunnel

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.

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.

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.

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.

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.

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.

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.

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.

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.

Air Flow Modeling in the Wind Tunnel of the FHWA Aerodynamics Laboratory at Turner-Fairbank Highway Research Center

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

Sitek, M. A.; Lottes, S. A.; Bojanowski, C.

Computational fluid dynamics (CFD) modeling is widely used in industry for design and in the research community to support, compliment, and extend the scope of experimental studies. Analysis of transportation infrastructure using high performance cluster computing with CFD and structural mechanics software is done at the Transportation Research and Analysis Computing Center (TRACC) at Argonne National Laboratory. These resources, available at TRACC, were used to perform advanced three-dimensional computational simulations of the wind tunnel laboratory at the Turner-Fairbank Highway Research Center (TFHRC). The goals were to verify the CFD model of the laboratory wind tunnel and then to use versionsmore » of the model to provide the capability to (1) perform larger parametric series of tests than can be easily done in the laboratory with available budget and time, (2) to extend testing to wind speeds that cannot be achieved in the laboratory, and (3) to run types of tests that are very difficult or impossible to run in the laboratory. Modern CFD software has many physics models and domain meshing options. Models, including the choice of turbulence and other physics models and settings, the computational mesh, and the solver settings, need to be validated against measurements to verify that the results are sufficiently accurate for use in engineering applications. The wind tunnel model was built and tested, by comparing to experimental measurements, to provide a valuable tool to perform these types of studies in the future as a complement and extension to TFHRC’s experimental capabilities. Wind tunnel testing at TFHRC is conducted in a subsonic open-jet wind tunnel with a 1.83 m (6 foot) by 1.83 m (6 foot) cross section. A three component dual force-balance system is used to measure forces acting on tested models, and a three degree of freedom suspension system is used for dynamic response tests. Pictures of the room are shown in Figure 1-1 to Figure 1-4. A detailed CAD geometry and CFD model of the wind tunnel laboratory at TFHRC was built and tested. Results were compared against experimental wind velocity measurements at a large number of locations around the room. This testing included an assessment of the air flow uniformity provided by the tunnel to the test zone and assessment of room geometry effects, such as influence of the proximity the room walls, the non-symmetrical position of the tunnel in the room, and the influence of the room setup on the air flow in the room. This information is useful both for simplifying the computational model and in deciding whether or not moving, or removing, some of the furniture or other movable objects in the room will change the flow in the test zone.« less

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...

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.

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

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

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.

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.

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.

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.

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.

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.

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.

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.

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.

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.

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.

Research and test facilities

NASA Technical Reports Server (NTRS)

1993-01-01

A description is given of each of the following Langley research and test facilities: 0.3-Meter Transonic Cryogenic Tunnel, 7-by 10-Foot High Speed Tunnel, 8-Foot Transonic Pressure Tunnel, 13-Inch Magnetic Suspension & Balance System, 14-by 22-Foot Subsonic Tunnel, 16-Foot Transonic Tunnel, 16-by 24-Inch Water Tunnel, 20-Foot Vertical Spin Tunnel, 30-by 60-Foot Wind Tunnel, Advanced Civil Transport Simulator (ACTS), Advanced Technology Research Laboratory, Aerospace Controls Research Laboratory (ACRL), Aerothermal Loads Complex, Aircraft Landing Dynamics Facility (ALDF), Avionics Integration Research Laboratory, Basic Aerodynamics Research Tunnel (BART), Compact Range Test Facility, Differential Maneuvering Simulator (DMS), Enhanced/Synthetic Vision & Spatial Displays Laboratory, Experimental Test Range (ETR) Flight Research Facility, General Aviation Simulator (GAS), High Intensity Radiated Fields Facility, Human Engineering Methods Laboratory, Hypersonic Facilities Complex, Impact Dynamics Research Facility, Jet Noise Laboratory & Anechoic Jet Facility, Light Alloy Laboratory, Low Frequency Antenna Test Facility, Low Turbulence Pressure Tunnel, Mechanics of Metals Laboratory, National Transonic Facility (NTF), NDE Research Laboratory, Polymers & Composites Laboratory, Pyrotechnic Test Facility, Quiet Flow Facility, Robotics Facilities, Scientific Visualization System, Scramjet Test Complex, Space Materials Research Laboratory, Space Simulation & Environmental Test Complex, Structural Dynamics Research Laboratory, Structural Dynamics Test Beds, Structures & Materials Research Laboratory, Supersonic Low Disturbance Pilot Tunnel, Thermal Acoustic Fatigue Apparatus (TAFA), Transonic Dynamics Tunnel (TDT), Transport Systems Research Vehicle, Unitary Plan Wind Tunnel, and the Visual Motion Simulator (VMS).

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.

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.

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.

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.

The Altitude Wind Tunnel (AWT): A unique facility for propulsion system and adverse weather testing

NASA Technical Reports Server (NTRS)

Chamberlin, R.

1985-01-01

A need has arisen for a new wind tunnel facility with unique capabilities for testing propulsion systems and for conducting research in adverse weather conditions. New propulsion system concepts, new aircraft configurations with an unprecedented degree of propulsion system/aircraft integration, and requirements for aircraft operation in adverse weather dictate the need for a new test facility. Required capabilities include simulation of both altitude pressure and temperature, large size, full subsonic speed range, propulsion system operation, and weather simulation (i.e., icing, heavy rain). A cost effective rehabilitation of the NASA Lewis Research Center's Altitude Wind Tunnel (AWT) will provide a facility with all these capabilities.

A numerical study of the controlled flow tunnel for a high lift model

NASA Technical Reports Server (NTRS)

Parikh, P. C.

1984-01-01

A controlled flow tunnel employs active control of flow through the walls of the wind tunnel so that the model is in approximately free air conditions during the test. This improves the wind tunnel test environment, enhancing the validity of the experimentally obtained test data. This concept is applied to a three dimensional jet flapped wing with full span jet flap. It is shown that a special treatment is required for the high energy wake associated with this and other V/STOL models. An iterative numerical scheme is developed to describe the working of an actual controlled flow tunnel and comparisons are shown with other available results. It is shown that control need be exerted over only part of the tunnel walls to closely approximate free air flow conditions. It is concluded that such a tunnel is able to produce a nearly interference free test environment even with a high lift model in the tunnel.

SMART Rotor Development and Wind Tunnel Test

DTIC Science & Technology

2009-09-01

amplifier and control system , and data acquisition, processing, and display systems . Boeing�s LRTS (Fig. 2), consists of a sled structure that...Support Test Stand Sled Tail Sting Outrigger Arm Figure 2: System integration test at whirl tower Port Rotor Balance Main Strut Flap Tail...demonstrated. Finally, the reliability of the flap actuation system was successfully proven in more than 60 hours of wind tunnel testing

A Method to Estimate Fabric Particle Penetration Performance

DTIC Science & Technology

2014-09-08

may be needed to improve the correlation between wind tunnel component sleeve tests and bench top swatch test. The ability to predict multi-layered...within the fabric/component gap may be needed to improve the correlation between wind tunnel component sleeve tests and bench top swatch test...impermeable garment . Heat stress becomes a major problem with this approach however, as normal physiological heat loss mechanisms (especially sweat

Propulsion and airframe aerodynamic interactions of supersonic V/STOL configurations. Volume 2: Wind tunnel test force and moment data report

NASA Technical Reports Server (NTRS)

Zilz, D. E.

1985-01-01

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

Experimental Investigation of Project Orion Crew Exploration Vehicle Aeroheating in AEDC Tunnel 9

NASA Technical Reports Server (NTRS)

Hollis, Brian R.; Horvath, Thomas J.; Berger, Karen T.; Lillard, Randolph P.; Kirk, Benjamin S.; Coblish, Joseph J.; Norris, Joseph D.

2008-01-01

An investigation of the aeroheating environment of the Project Orion Crew Entry Vehicle has been performed in the Arnold Engineering Development Center Tunnel 9. The goals of this test were to measure turbulent heating augmentation levels on the heat shield and to obtain high-fidelity heating data for assessment of computational fluid dynamics methods. 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 for the computational method. Data from both the wind tunnel test and the computational study are presented herein.

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

NASA Technical Reports Server (NTRS)

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

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

NASA Technical Reports Server (NTRS)

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

Wind tunnel tests on a one-foot diameter SR-7L propfan model

NASA Technical Reports Server (NTRS)

Aljabri, Abdullah S.

1987-01-01

Wind tunnel tests have been conducted on a one-foot diameter model of the SR-7L propfan in the Langley 16-Foot and 4 x 7 Meter Wind Tunnels as part of the Propfan Test Assessment (PTA) Program. The model propfan was sized to be used on a 1/9-scale model of the PTA testbed aircraft. The model propeller was tested in isolation and wing-mounted on the aircraft configuration at various Mach numbers and blade pitch angles. Agreement between data obtained from these tests and data from Hamilton Standard validate that the 1/9-scale propeller accurately simulates the aerodynamics of the SR-7L propfan. Predictions from an analytical computer program are presented and show good agreement with the experimental data.

Fluctuating disturbances in a Mach 5 wind tunnel

NASA Technical Reports Server (NTRS)

Anders, J. B.; Stainback, P. C.; Beckwith, I. E.; Keefe, L. R.

1976-01-01

An experimental investigation has been conducted to determine the source and nature of disturbances in the settling chamber and test section of a Mach 5 wind tunnel. Various changes in the air supply piping to the wind tunnel are shown to influence the disturbance levels in the settling chamber. These levels were reduced by the use of an acoustic muffler section in the settling chamber. Three nozzles were tested with the same settling chamber and hot-wire measurements indicated that the test section disturbances were entirely acoustic. Significant reductions in the test section noise levels were obtained with an electroplated nozzle utilizing boundary-layer removal upstream of the throat. The source of test section noise is shown to be different for laminar and turbulent nozzle-wall boundary layers.

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

NASA Technical Reports Server (NTRS)

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

1980-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1986-01-01

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

Correlations of Platooning Track Test and Wind Tunnel Data

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

Lammert, Michael P.; Kelly, Kenneth J.; Yanowitz, Janet

In this report, the National Renewable Energy Laboratory analyzed results from multiple, independent truck platooning projects to compare and contrast track test results with wind tunnel test results conducted by Lawrence Livermore National Laboratory (LLNL). Some highlights from the report include compiled data, and results from four independent SAE J1321 full-size track test campaigns that were compared to LLNL wind tunnel testing results. All platooning scenarios tested demonstrated significant fuel savings with good correlation relative to following distances, but there are still unanswered questions and clear opportunities for system optimization. NOx emissions showed improvements from NREL tests in 2014 tomore » Auburn tests in 2015 with respect to J1321 platooning track testing of Peloton system. NREL evaluated data from Volpe's Naturalistic Study of Truck Following Behavior, which showed minimal impact of naturalistic background platooning. We found significant correlation between multiple track studies, wind tunnel tests, and computational fluid dynamics, but also showed that there is more to learn regarding close formation and longer-distance effects. We also identified potential areas for further research and development, including development of advanced aerodynamic designs optimized for platooning, measurement of platoon system performance in traffic conditions, impact of vehicle lateral offsets on platooning performance, and characterization of the national potential for platooning based on fleet operational characteristics.« less

Low Drag Airfoil Design Utilizing Passive Laminar Flow and Coupled Diffusion Control Techniques.

DTIC Science & Technology

1980-09-01

number and real flow environment Influence on transition are pre- sented. Examination of wind tunnel scaling Influences and real flow environ- ment...effects on the ATC/lamlnar airfoil performance are discussed and summarized for typical low turbulence tunnels ,(e.g., the NASA-Ames 12 foot pressure tunnel ...35 5.0 WIND TUNNEL VALIDATION ASSESSMENT --------------------- 42 5.1 TEST FACILITY EVALUATION ------------------------- 42 5.2

Drag and Cooling Tests in the 24 ft Wind Tunnel on a Centaurus-Buckingham Wing Nacelle Installation. Part 3. Tests with High Speed Cowl Entry (Tempest Type)

DTIC Science & Technology

1946-07-01

ESTABLISHMENT Farnboruugh. Hants. DRAG AND COOLING TESTS IN THE 24 ft. WIND TUNNEL ON A ^ CENTAURUS -BUCKINGHAM WING NACELLE INSTALLATION PART...ILLIETUATICNS Ccntaurus-f uckinghain nacelle inst -illation with high entry velocity cowl ^soolii» ; fan not fitted. Installation of Centaurus .inline

Research in Natural Laminar Flow and Laminar-Flow Control, part 2

NASA Technical Reports Server (NTRS)

Hefner, Jerry N. (Compiler); Sabo, Frances E. (Compiler)

1987-01-01

Part 2 of the Symposium proceedings includes papers addressing various topics in basic wind tunnel research/techniques and computational transitional research. Specific topics include: advanced measurement techniques; laminar flow control; Tollmien-Schlichting wave characteristics; boundary layer transition; flow visualization; wind tunnel tests; flight tests; boundary layer equations; swept wings; and skin friction.

Vanguard 2C VTOL Airplane Tested in the Ames 40x80 Foot Wind Tunnel.

NASA Image and Video Library

1960-02-01

Vanguard 2C vertical take-off and landing (VTOL) airplane, wind tunnel test. Front view from below, model 14 1/2 feet high disk off. Nasa Ames engineer Ralph Maki in photo. Variable height struts and ground plane, low pressure ratio, fan in wing. 02/01/1960.

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

USDA-ARS?s Scientific Manuscript database

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

Comparison of nozzle and afterbody surface pressures from wind tunnel and flight test of the YF-17 aircraft

NASA Technical Reports Server (NTRS)

Lucas, E. J.; Fanning, A. E.; Steers, L. I.

1978-01-01

Results are reported from the initial phase of an effort to provide an adequate technical capability to accurately predict the full scale, flight vehicle, nozzle-afterbody performance of future aircraft based on partial scale, wind tunnel testing. The primary emphasis of this initial effort is to assess the current capability and identify the cause of limitations on this capability. A direct comparison of surface pressure data is made between the results from an 0.1-scale model wind tunnel investigation and a full-scale flight test program to evaluate the current subscale testing techniques. These data were acquired at Mach numbers 0.6, 0.8, 0.9, 1.2, and 1.5 on four nozzle configurations at various vehicle pitch attitudes. Support system interference increments were also documented during the wind tunnel investigation. In general, the results presented indicate a good agreement in trend and level of the surface pressures when corrective increments are applied for known effects and surface differences between the two articles under investigation.

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

NASA Technical Reports Server (NTRS)

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

1975-01-01

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

Analytical and Experimental Evaluation of Digital Control Systems for the Semi-Span Super-Sonic Transport (S4T) Wind Tunnel Model

NASA Technical Reports Server (NTRS)

Wieseman, Carol D.; Christhilf, David; Perry, Boyd, III

2012-01-01

An important objective of the Semi-Span Super-Sonic Transport (S4T) wind tunnel model program was the demonstration of Flutter Suppression (FS), Gust Load Alleviation (GLA), and Ride Quality Enhancement (RQE). It was critical to evaluate the stability and robustness of these control laws analytically before testing them and experimentally while testing them to ensure safety of the model and the wind tunnel. MATLAB based software was applied to evaluate the performance of closed-loop systems in terms of stability and robustness. Existing software tools were extended to use analytical representations of the S4T and the control laws to analyze and evaluate the control laws prior to testing. Lessons were learned about the complex windtunnel model and experimental testing. The open-loop flutter boundary was determined from the closed-loop systems. A MATLAB/Simulink Simulation developed under the program is available for future work to improve the CPE process. This paper is one of a series of that comprise a special session, which summarizes the S4T wind-tunnel program.

System Dynamic Analysis of a Wind Tunnel Model with Applications to Improve Aerodynamic Data Quality

NASA Technical Reports Server (NTRS)

Buehrle, Ralph David

1997-01-01

The research investigates the effect of wind tunnel model system dynamics on measured aerodynamic data. During wind tunnel tests designed to obtain lift and drag data, the required aerodynamic measurements are the steady-state balance forces and moments, pressures, and model attitude. However, the wind tunnel model system can be subjected to unsteady aerodynamic and inertial loads which result in oscillatory translations and angular rotations. The steady-state force balance and inertial model attitude measurements are obtained by filtering and averaging data taken during conditions of high model vibrations. The main goals of this research are to characterize the effects of model system dynamics on the measured steady-state aerodynamic data and develop a correction technique to compensate for dynamically induced errors. Equations of motion are formulated for the dynamic response of the model system subjected to arbitrary aerodynamic and inertial inputs. The resulting modal model is examined to study the effects of the model system dynamic response on the aerodynamic data. In particular, the equations of motion are used to describe the effect of dynamics on the inertial model attitude, or angle of attack, measurement system that is used routinely at the NASA Langley Research Center and other wind tunnel facilities throughout the world. This activity was prompted by the inertial model attitude sensor response observed during high levels of model vibration while testing in the National Transonic Facility at the NASA Langley Research Center. The inertial attitude sensor cannot distinguish between the gravitational acceleration and centrifugal accelerations associated with wind tunnel model system vibration, which results in a model attitude measurement bias error. Bias errors over an order of magnitude greater than the required device accuracy were found in the inertial model attitude measurements during dynamic testing of two model systems. Based on a theoretical modal approach, a method using measured vibration amplitudes and measured or calculated modal characteristics of the model system is developed to correct for dynamic bias errors in the model attitude measurements. The correction method is verified through dynamic response tests on two model systems and actual wind tunnel test data.

Simulation of Flight-Type Engine Fan Noise in the NASA-Lewis 9X15 Anechoic Wind Tunnel

NASA Technical Reports Server (NTRS)

Heidmann, M. F.; Dietrich, D. A.

1976-01-01

Flight type noise as contrasted to the usual ground static test noise exhibits substantial reductions in the time unsteadiness of tone noise, and in the mean level of tones calculated to be nonpropagating or cut-off. A model fan designed with cuttoff of the fundamental tone was acoustically tested in the anechoic wind tunnel under both static and tunnel flow conditions. The properties that characterize flight type noise were progressively simulated with increasing tunnel flow. The distinctly lobed directivity pattern of propagating rotor/stator interaction modes was also observed. Excess noise attributed to the ingestion of the flow disturbances that prevail near most static test facilities is substantially reduced with tunnel flow.

Full-Span Tiltrotor Aeroacoustic Model (TRAM) Overview and 40- by 80-Foot Wind Tunnel Test. [conducted in the 40- by 80-Foot Wind Tunnel at Ames Research Center

NASA Technical Reports Server (NTRS)

McCluer, Megan S.; Johnson, Jeffrey L.; Rutkowski, Michael (Technical Monitor)

2001-01-01

Most helicopter data trends cannot be extrapolated to tiltrotors because blade geometry and aerodynamic behavior, as well as rotor and fuselage interactions, are significantly different for tiltrotors. A tiltrotor model has been developed to investigate the aeromechanics of tiltrotors, to develop a comprehensive database for validating tiltrotor analyses, and to provide a research platform for supporting future tiltrotor designs. The Full-Span Tiltrotor Aeroacoustic Model (FS TRAM) is a dual-rotor, powered aircraft model with extensive instrumentation for measurement of structural and aerodynamic loads. This paper will present the Full-Span TRAM test capabilities and the first set of data obtained during a 40- by 80-Foot Wind Tunnel test conducted in late 2000 at NASA Ames Research Center. The Full-Span TRAM is a quarter-scale representation of the V-22 Osprey aircraft, and a heavily instrumented NASA and U.S. Army wind tunnel test stand. Rotor structural loads are monitored and recorded for safety-of-flight and for information on blade loads and dynamics. Left and right rotor balance and fuselage balance loads are monitored for safety-of-flight and for measurement of vehicle and rotor aerodynamic performance. Static pressure taps on the left wing are used to determine rotor/wing interactional effects and rotor blade dynamic pressures measure blade airloads. All of these measurement capabilities make the FS TRAM test stand a unique and valuable asset for validation of computational codes and to aid in future tiltrotor designs. The Full-Span TRAM was tested in the NASA Ames Research Center 40- by 80-Foot Wind Tunnel from October through December 2000. Rotor and vehicle performance measurements were acquired in addition to wing pressures, rotor acoustics, and Laser Light Sheet (LLS) flow visualization data. Hover, forward flight, and airframe (rotors off) aerodynamic runs were performed. Helicopter-mode data were acquired during angle of attack and thrust sweeps for a variety of tunnel speeds. Wake geometry images were acquired using LLS photographs and suggest dual tip vortex formation at low thrust conditions. The full paper will include comparisons to isolated-rotor TRAM data acquired at the Duits-Nederlandse Windtunnel (DNW) in 1998. The FS TRAM has been established as a valuable national asset for tiltrotor research. Data reduction and analysis of the 40- by 80-Foot Wind Tunnel test results are underway. Follow-on testing of the FS TRAM is currently being planned for the NASA Ames 80- by 120-Foot Wind Tunnel in late 2001.

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)

Cryogenic Model Materials

NASA Technical Reports Server (NTRS)

Kimmel, W. M.; Kuhn, N. S.; Berry, R. F.; Newman, J. A.

2001-01-01

An overview and status of current activities seeking alternatives to 200 grade 18Ni Steel CVM alloy for cryogenic wind tunnel models is presented. Specific improvements in material selection have been researched including availability, strength, fracture toughness and potential for use in transonic wind tunnel testing. Potential benefits from utilizing damage tolerant life-prediction methods, recently developed fatigue crack growth codes and upgraded NDE methods are also investigated. Two candidate alloys are identified and accepted for cryogenic/transonic wind tunnel models and hardware.

Space capsule mounted in the Full Scale Wind Tunnel

NASA Image and Video Library

1959-01-22

The Mercury space capsule undergoing tests in Full Scale Wind Tunnel, January 1959. Photograph published in Winds of Change, 75th Anniversary NASA publication, page 75, by James Schultz. Also Photograph published in Engineer in Charge: A History of the Langley Aeronautical Laboratory, 1917-1958, page 389, by James R. Hansen.

Aerodynamic characteristics of the modified 40- by 80-foot wind tunnel as measured in a 1/50th-scale model

NASA Technical Reports Server (NTRS)

Smith, Brian E.; Naumowicz, Tim

1987-01-01

The aerodynamic characteristics of the 40- by 80-Foot Wind Tunnel at Ames Research Center were measured by using a 1/50th-scale facility. The model was configured to closely simulate the features of the full-scale facility when it became operational in 1986. The items measured include the aerodynamic effects due to changes in the total-pressure-loss characteristics of the intake and exhaust openings of the air-exchange system, total-pressure distributions in the flow field at locations around the wind tunnel circuit, the locations of the maximum total-pressure contours, and the aerodynamic changes caused by the installation of the acoustic barrier in the southwest corner of the wind tunnel. The model tests reveal the changes in the aerodynamic performance of the 1986 version of the 40- by 80-Foot Wind Tunnel compared with the performance of the 1982 configuration.

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

NASA Astrophysics Data System (ADS)

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

2016-09-01

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

Development of the NASA-Ames low disturbance supersonic wind tunnel for transition research up to Mach 2.5

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 aerodynamic features of this new quiet tunnel will be a low-disturbance settling chamber, laminar boundary layers on the nozzle walls and steady supersonic diffuser flow. Furthermore, this new wind tunnel will operate continuously at uniquely low compression ratios (less than unity). This feature allows an existing non-specialist compressor to be used as a major part of the drive system. In this paper, we highlight activities associated with drive system development, the establishment of natural laminar flow on the test section walls, and instrumentation development for transition detection. Experimental results from an 1/8th-scale model of the supersonic wind tunnel are presented and discussed in association with theoretical predictions. Plans are progressing to build the full-scale wind tunnel by the end of 1993.

Introduction to cryogenic wind tunnels

NASA Technical Reports Server (NTRS)

Goodyer, M. J.

1985-01-01

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

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.

Assessment of Scaled Rotors for Wind Tunnel Experiments.

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

Maniaci, David Charles; Kelley, Christopher Lee; Chiu, Phillip

2015-07-01

Rotor design and analysis work has been performed to support the conceptualization of a wind tunnel test focused on studying wake dynamics. This wind tunnel test would serve as part of a larger model validation campaign that is part of the Department of Energy Wind and Water Power Program’s Atmosphere to electrons (A2e) initiative. The first phase of this effort was directed towards designing a functionally scaled rotor based on the same design process and target full-scale turbine used for new rotors for the DOE/SNL SWiFT site. The second phase focused on assessing the capabilities of an already available rotor,more » the G1, designed and built by researchers at the Technical University of München.« less

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.

Performance evaluation of Space Shuttle SRB parachutes from air drop and scaled model wind tunnel tests. [Solid Rocket Booster recovery system

NASA Technical Reports Server (NTRS)

Moog, R. D.; Bacchus, D. L.; Utreja, L. R.

1979-01-01

The aerodynamic performance characteristics have been determined for the Space Shuttle Solid Rocket Booster drogue, main, and pilot parachutes. The performance evaluation on the 20-degree conical ribbon parachutes is based primarily on air drop tests of full scale prototype parachutes. In addition, parametric wind tunnel tests were performed and used in parachute configuration development and preliminary performance assessments. The wind tunnel test data are compared to the drop test results and both sets of data are used to determine the predicted performance of the Solid Rocket Booster flight parachutes. Data from other drop tests of large ribbon parachutes are also compared with the Solid Rocket Booster parachute performance characteristics. Parameters assessed include full open terminal drag coefficients, reefed drag area, opening characteristics, clustering effects, and forebody interference.

Application of interactive computer graphics in wind-tunnel dynamic model testing

NASA Technical Reports Server (NTRS)

Doggett, R. V., Jr.; Hammond, C. E.

1975-01-01

The computer-controlled data-acquisition system recently installed for use with a transonic dynamics tunnel was described. This includes a discussion of the hardware/software features of the system. A subcritical response damping technique, called the combined randomdec/moving-block method, for use in windtunnel-model flutter testing, that has been implemented on the data-acquisition system, is described in some detail. Some results using the method are presented and the importance of using interactive graphics in applying the technique in near real time during wind-tunnel test operations is discussed.

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

NASA Technical Reports Server (NTRS)

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

1942-01-01

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

Wind-tunnel investigation of a flush airdata system at Mach numbers from 0.7 to 1.4

NASA Technical Reports Server (NTRS)

Larson, Terry J.; Moes, Timothy R.; Siemers, Paul M., III

1990-01-01

Flush pressure orifices installed on the nose section of a 1/7-scale model of the F-14 airplane were evaluated for use as a flush airdata system (FADS). Wing-tunnel tests were conducted in the 11- by 11-ft Unitary Wind Tunnel at NASA Ames Research Center. A full-scale FADS of the same configuration was previously tested using an F-14 aircraft at the Dryden Flight Research Facility of NASA Ames Research Center (Ames-Dryden). These tests, which were published, are part of a NASA program to assess accuracies of FADS for use on aircraft. The test program also provides data to validate algorithms for the shuttle entry airdata system developed at the NASA Langley Research Center. The wind-tunnel test Mach numbers were 0.73, 0.90, 1.05, 1.20, and 1.39. Angles of attack were varied in 2 deg increments from -4 deg to 20 deg. Sideslip angles were varied in 4 deg increments from -8 deg to 8 deg. Airdata parameters were evaluated for determination of free-stream values of stagnation pressure, static pressure, angle of attack, angle of sideslip, and Mach number. These parameters are, in most cases, the same as the parameters investigated in the flight test program. The basic FADS wind-tunnel data are presented in tabular form. A discussion of the more accurate parameters is included.

Control Surface Interaction Effects of the Active Aeroelastic Wing Wind Tunnel Model

NASA Technical Reports Server (NTRS)

Heeg, Jennifer

2006-01-01

This paper presents results from testing the Active Aeroelastic Wing wind tunnel model in NASA Langley s Transonic Dynamics Tunnel. The wind tunnel test provided an opportunity to study aeroelastic system behavior under combined control surface deflections, testing for control surface interaction effects. Control surface interactions were observed in both static control surface actuation testing and dynamic control surface oscillation testing. The primary method of evaluating interactions was examination of the goodness of the linear superposition assumptions. Responses produced by independently actuating single control surfaces were combined and compared with those produced by simultaneously actuating and oscillating multiple control surfaces. Adjustments to the data were required to isolate the control surface influences. Using dynamic data, the task increases, as both the amplitude and phase have to be considered in the data corrections. The goodness of static linear superposition was examined and analysis of variance was used to evaluate significant factors influencing that goodness. The dynamic data showed interaction effects in both the aerodynamic measurements and the structural measurements.

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.

A Novel Method to Predict Circulation Control Noise

DTIC Science & Technology

2016-03-17

Semi-empirical aeracoustic prediction code for wind turbines . In NREL/ TP-500-34478, National Wind Technology Center. MOSHER, M. 1983 Acoustics of...velocimetry, unsteady pressure and phased-acoustic- array data are acquired simultaneously in an aeroacoustic wind -tunnel facility. The velocity field...her open-jet wind tunnels or flight testing which makes noise prediction for underwater vehicles especially difficult . 1 In this document , a

Computational Results for the KTH-NASA Wind-Tunnel Model Used for Acquisition of Transonic Nonlinear Aeroelastic Data

NASA Technical Reports Server (NTRS)

Silva, Walter A.; Chwalowski, Pawel; Wieseman, Carol D.; Eller, David; Ringertz, Ulf

2017-01-01

A status report is provided on the collaboration between the Royal Institute of Technology (KTH) in Sweden and the NASA Langley Research Center regarding the aeroelastic analyses of a full-span fighter configuration wind-tunnel model. This wind-tunnel model was tested in the Transonic Dynamics Tunnel (TDT) in the summer of 2016. Large amounts of data were acquired including steady/unsteady pressures, accelerations, strains, and measured dynamic deformations. The aeroelastic analyses presented include linear aeroelastic analyses, CFD steady analyses, and analyses using CFD-based reduced-order models (ROMs).

Validation of the Lockheed Martin Morphing Concept with Wind Tunnel Testing

NASA Technical Reports Server (NTRS)

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

2007-01-01

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

Static and Wind-on Performance of Polymer-Based Pressure-Sensitive Paints Using Platinum and Ruthenium as the Luminophore

PubMed Central

Lo, Kin Hing; Kontis, Konstantinos

2016-01-01

An experimental study has been conducted to investigate the static and wind-on performance of two in-house-developed polymer-based pressure-sensitive paints. Platinum tetrakis (pentafluorophenyl) porphyrin and tris-bathophenanthroline ruthenium II are used as the luminophores of these two polymer-based pressure-sensitive paints. The pressure and temperature sensitivity and the photo-degradation rate of these two pressure-sensitive paints have been investigated. In the wind tunnel test, it was observed that the normalised intensity ratio of both polymer-based pressure-sensitive paints being studied decreases with increasing the number of wind tunnel runs. The exact reason that leads to the occurrence of this phenomenon is unclear, but it is deduced that the luminophore is either removed or deactivated by the incoming flow during a wind tunnel test. PMID:27128913

Blockage Testing in the NASA Glenn 225 Square Centimeter Supersonic Wind Tunnel

NASA Technical Reports Server (NTRS)

Sevier, Abigail; Davis, David O.; Schoenenberger, Mark

2017-01-01

The starting characteristics for three different model geometries were tested in the Glenn Research Center 225 Square Centimeter Supersonic Wind Tunnel. The test models were tested at Mach 2, 2.5 and 3 in a square test section and at Mach 2.5 again in an asymmetric test section. The results gathered in this study will help size the test models and inform other design features for the eventual implementation of a magnetic suspension system.

40 CFR Table F-1 to Subpart F of... - Performance Specifications for PM2.5 Class II Equivalent Samplers

Code of Federal Regulations, 2011 CFR

2011-07-01

... specified in the post-loading evaluation test (§ 53.62, § 53.63, or § 53.64). § 53.66 Volatility Test... Equivalent Samplers Performance test Specifications Acceptance criteria § 53.62 Full Wind Tunnel Evaluation...: 95% ≤ Rc ≤ 105%. § 53.63 Wind Tunnel Inlet Aspiration Test Liquid VOAG produced aerosol at 2 km/hr...

40 CFR Table F-1 to Subpart F of... - Performance Specifications for PM2.5 Class II Equivalent Samplers

Code of Federal Regulations, 2010 CFR

2010-07-01

... specified in the post-loading evaluation test (§ 53.62, § 53.63, or § 53.64). § 53.66 Volatility Test... Equivalent Samplers Performance test Specifications Acceptance criteria § 53.62 Full Wind Tunnel Evaluation...: 95% ≤ Rc ≤ 105%. § 53.63 Wind Tunnel Inlet Aspiration Test Liquid VOAG produced aerosol at 2 km/hr...

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

NASA Technical Reports Server (NTRS)

Mukhopadhyay, Vivek

1999-01-01

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

Key technique study and application of infrared thermography in hypersonic wind tunnel

NASA Astrophysics Data System (ADS)

LI, Ming; Yang, Yan-guang; Li, Zhi-hui; Zhu, Zhi-wei; Zhou, Jia-sui

2014-11-01

The solutions to some key techniques using infrared thermographic technique in hypersonic wind tunnel, such as temperature measurement under great measurement angle, the corresponding relation between model spatial coordinates and the ones in infrared map, the measurement uncertainty analysis of the test data etc., are studied. The typical results in the hypersonic wind tunnel test are presented, including the comparison of the transfer rates on a thin skin flat plate model with a wedge measured with infrared thermography and thermocouple, the experimental study heating effect on the flat plate model impinged by plume flow and the aerodynamic heating on the lift model.

Evaluation of Simultaneous Multisine Excitation of the Joined Wing SensorCraft Aeroelastic Wind Tunnel Model

NASA Technical Reports Server (NTRS)

Heeg, Jennifer; Morelli, Eugene A.

2011-01-01

Multiple mutually orthogonal signals comprise excitation data sets for aeroservoelastic system identification. A multisine signal is a sum of harmonic sinusoid components. A set of these signals is made orthogonal by distribution of the frequency content such that each signal contains unique frequencies. This research extends the range of application of an excitation method developed for stability and control flight testing to aeroservoelastic modeling from wind tunnel testing. Wind tunnel data for the Joined Wing SensorCraft model validates this method, demonstrating that these signals applied simultaneously reproduce the frequency response estimates achieved from one-at-a-time excitation.

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

NASA Technical Reports Server (NTRS)

Mukhopadhyay, Vivek

1994-01-01

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

Prediction of Unsteady Aerodynamic Coefficients at High Angles of Attack

NASA Technical Reports Server (NTRS)

Pamadi, Bandu N.; Murphy, Patrick C.; Klein, Vladislav; Brandon, Jay M.

2001-01-01

The nonlinear indicial response method is used to model the unsteady aerodynamic coefficients in the low speed longitudinal oscillatory wind tunnel test data of the 0.1 scale model of the F-16XL aircraft. Exponential functions are used to approximate the deficiency function in the indicial response. Using one set of oscillatory wind tunnel data and parameter identification method, the unknown parameters in the exponential functions are estimated. The genetic algorithm is used as a least square minimizing algorithm. The assumed model structures and parameter estimates are validated by comparing the predictions with other sets of available oscillatory wind tunnel test data.

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

NASA Technical Reports Server (NTRS)

Mukhopadhyay, Vivek

1992-01-01

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

Low and high speed propellers for general aviation: Performance potential and recent wind tunnel test results

NASA Technical Reports Server (NTRS)

Jeracki, R. J.; Mitchell, G. A.

1981-01-01

The performance of lower speed, 5 foot diameter model general aviation propellers, was tested in the Lewis wind tunnel. Performance was evaluated for various levels of airfoil technology and activity factor. The difference was associated with inadequate modeling of blade and spinner losses for propellers round shank blade designs. Suggested concepts for improvement are: (1) advanced blade shapes (airfoils and sweep); (2) tip devices (proplets); (3) integrated propeller/nacelles; and (4) composites. Several advanced aerodynamic concepts were evaluated in the Lewis wind tunnel. Results show that high propeller performance can be obtained to at least Mach 0.8.

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

NASA Technical Reports Server (NTRS)

Mukhopadhyay, Vivek

1992-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1984-01-01

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

Results of a Pressure Loads Investigation on a 0.030-scale Model (47-OTS) of the Integrated Space Shuttle Vehicle Configuration 5 in the NASA Ames Research Center 11 by 11 Foot Leg of the Unitary Plan Wind Tunnel (IA81A), Volume 1

NASA Technical Reports Server (NTRS)

Chee, E.

1975-01-01

Results of wind tunnel tests on a 0.030-scale model of the integrated space shuttle vehicle configuration 5 are presented. Testing was conducted in the NASA Ames Research Center 11 x 11 foot leg of the Unitary Plan Wind Tunnel to investigate pressure distributions for airloads analyses at Mach numbers from 0.9 through 1.4. Angles of attack and sideslip were varied from -6 to +6 degrees.

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.

Models of Lift and Drag Coefficients of Stalled and Unstalled Airfoils in Wind Turbines and Wind Tunnels

NASA Technical Reports Server (NTRS)

Spera, David A.

2008-01-01

Equations are developed with which to calculate lift and drag coefficients along the spans of torsionally-stiff rotating airfoils of the type used in wind turbine rotors and wind tunnel fans, at angles of attack in both the unstalled and stalled aerodynamic regimes. Explicit adjustments are made for the effects of aspect ratio (length to chord width) and airfoil thickness ratio. Calculated lift and drag parameters are compared to measured parameters for 55 airfoil data sets including 585 test points. Mean deviation was found to be -0.4 percent and standard deviation was 4.8 percent. When the proposed equations were applied to the calculation of power from a stall-controlled wind turbine tested in a NASA wind tunnel, mean deviation from 54 data points was -1.3 percent and standard deviation was 4.0 percent. Pressure-rise calculations for a large wind tunnel fan deviated by 2.7 percent (mean) and 4.4 percent (standard). The assumption that a single set of lift and drag coefficient equations can represent the stalled aerodynamic behavior of a wide variety of airfoils was found to be satisfactory.

Comparison of model and flight test data for an augmented jet flap STOL research aircraft

NASA Technical Reports Server (NTRS)

Cook, W. L.; Whittley, D. C.

1975-01-01

Aerodynamic design data for the Augmented Jet Flap STOL Research Aircraft or commonly known as the Augmentor-Wing Jet-STOL Research Aircraft was based on results of tests carried out on a large scale research model in the NASA Ames 40- by 80-Foot Wind Tunnel. Since the model differs in some respects from the aircraft, precise correlation between tunnel and flight test is not expected, however the major areas of confidence derived from the wind tunnel tests are delineated, and for the most part, tunnel results compare favorably with flight experience. In some areas the model tests were known to be nonrepresentative so that a degree of uncertainty remained: these areas of greater uncertainty are identified, and discussed in the light of subsequent flight tests.

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

NASA Technical Reports Server (NTRS)

Allmaras, S. R.

1986-01-01

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

Wind tunnel testing of low-drag airfoils

NASA Technical Reports Server (NTRS)

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

1986-01-01

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

Wind tunnel performance results of an aeroelastically scaled 2/9 model of the PTA flight test prop-fan

NASA Technical Reports Server (NTRS)

Stefko, George L.; Rose, Gayle E.; Podboy, Gary G.

1987-01-01

High speed wind tunnel aerodynamic performance tests of the SR-7A advanced prop-fan have been completed in support of the Prop-Fan Test Assessment (PTA) flight test program. The test showed that the SR-7A model performed aerodynamically very well. At the cruise design condition, the SR-7A prop fan had a high measured net efficiency of 79.3 percent.

Engineering changes to the 0.1m cryogenic wind tunnel at Southampton University

NASA Technical Reports Server (NTRS)

Goodyer, M. J.

1984-01-01

The more important changes to the 0.1 m cryogenic wind tunnel since its completion in 1977 are outlined. These include detailed improvements in the fan drive to allow higher speeds, and the provision of a test section leg suitable for use with a magnetic suspension and balance system. The instrumentation, data logging, data reduction and tunnel controls were also improved and modernized. A tunnel performance summary is given.

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.

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

NASA Astrophysics Data System (ADS)

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

2016-09-01

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

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

NASA Technical Reports Server (NTRS)

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

2006-01-01

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

A New Method of Testing in Wind Tunnels

NASA Technical Reports Server (NTRS)

Margoulis, W

1921-01-01

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

Wind Tunnel Database Development using Modern Experiment Design and Multivariate Orthogonal Functions

NASA Technical Reports Server (NTRS)

Morelli, Eugene A.; DeLoach, Richard

2003-01-01

A wind tunnel experiment for characterizing the aerodynamic and propulsion forces and moments acting on a research model airplane is described. The model airplane called the Free-flying Airplane for Sub-scale Experimental Research (FASER), is a modified off-the-shelf radio-controlled model airplane, with 7 ft wingspan, a tractor propeller driven by an electric motor, and aerobatic capability. FASER was tested in the NASA Langley 12-foot Low-Speed Wind Tunnel, using a combination of traditional sweeps and modern experiment design. Power level was included as an independent variable in the wind tunnel test, to allow characterization of power effects on aerodynamic forces and moments. A modeling technique that employs multivariate orthogonal functions was used to develop accurate analytic models for the aerodynamic and propulsion force and moment coefficient dependencies from the wind tunnel data. Efficient methods for generating orthogonal modeling functions, expanding the orthogonal modeling functions in terms of ordinary polynomial functions, and analytical orthogonal blocking were developed and discussed. The resulting models comprise a set of smooth, differentiable functions for the non-dimensional aerodynamic force and moment coefficients in terms of ordinary polynomials in the independent variables, suitable for nonlinear aircraft simulation.

Development of an Intelligent Videogrammetric Wind Tunnel Measurement System

NASA Technical Reports Server (NTRS)

Graves, Sharon S.; Burner, Alpheus W.

2004-01-01

A videogrammetric technique developed at NASA Langley Research Center has been used at five NASA facilities at the Langley and Ames Research Centers for deformation measurements on a number of sting mounted and semispan models. These include high-speed research and transport models tested over a wide range of aerodynamic conditions including subsonic, transonic, and supersonic regimes. The technique, based on digital photogrammetry, has been used to measure model attitude, deformation, and sting bending. In addition, the technique has been used to study model injection rate effects and to calibrate and validate methods for predicting static aeroelastic deformations of wind tunnel models. An effort is currently underway to develop an intelligent videogrammetric measurement system that will be both useful and usable in large production wind tunnels while providing accurate data in a robust and timely manner. Designed to encode a higher degree of knowledge through computer vision, the system features advanced pattern recognition techniques to improve automated location and identification of targets placed on the wind tunnel model to be used for aerodynamic measurements such as attitude and deformation. This paper will describe the development and strategy of the new intelligent system that was used in a recent test at a large transonic wind tunnel.

Review Report on Design Study and Economic Assessment of Multi-Unit Offshore Wind Energy Conversion Systems Applications,

DTIC Science & Technology

1977-03-21

meter turbine . Available from NTIS; $6.50. 113 pages. 7. SAND-76-0130 Wind Tunnel Performance Data for the Darrieus Wind Tur- bine with NACA-0012...2-meter-diameter Darrieus wind turbine have been tested in a low speed wind tunnel. The airfoil section for all configurations was NACA 0012. The... Darrieus Vertical-Axis Wind Turbine Program at Sandia Laboratories, Kadlec, E.G., published by Sandia Laboratories 1976. Contract No. AT(29-1)-789. From

The elaboration of a new family of helicopter blade profiles

NASA Technical Reports Server (NTRS)

Thibert, J. J.

1981-01-01

An airfoil family of helicopter rotor blades was designed. Three airfoils with thickness to chord ratios of 12, 9, and 7% were designed. Their improved performance in two dimensional rotor mockup wind tunnel tests led to testing of the tapered blades on four bladed rotors in a wind tunnel and flight tests on the Dauphin series of helicopters, confirming the expected gains.

Investigation of a Technique for Measuring Dynamic Ground Effect in a Subsonic Wind Tunnel

NASA Technical Reports Server (NTRS)

Graves, Sharon S.

1999-01-01

To better understand the ground effect encountered by slender wing supersonic transport aircraft, a test was conducted at NASA Langley Research Center's 14 x 22 foot Subsonic Wind Tunnel in October, 1997. Emphasis was placed on improving the accuracy of the ground effect data by using a "dynamic" technique in which the model's vertical motion was varied automatically during wind-on testing. This report describes and evaluates different aspects of the dynamic method utilized for obtaining ground effect data in this test. The method for acquiring and processing time data from a dynamic ground effect wind tunnel test is outlined with details of the overall data acquisition system and software used for the data analysis. The removal of inertial loads due to sting motion and the support dynamics in the balance force and moment data measurements of the aerodynamic forces on the model is described. An evaluation of the results identifies problem areas providing recommendations for future experiments. Test results are validated by comparing test data for an elliptical wing planform with an Elliptical wing planform section with a NACA 0012 airfoil to results found in current literature. Major aerodynamic forces acting on the model in terms of lift curves for determining ground effect are presented. Comparisons of flight and wind tunnel data for the TU-144 are presented.

A new position measurement system using a motion-capture camera for wind tunnel tests.

PubMed

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

2013-09-13

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

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

PubMed Central

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

2013-01-01

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

Acquisition of Turbulence Data Using the DST Group Constant-Temperature Hot-Wire Anemometer System

DTIC Science & Technology

2015-10-01

fluctuations in the low-speed wind tunnel at DST Group. The use of both single- wire and crossed- wire (2 wire ) probes is described. Areas covered include a...fluid-flow studies, including testing of models of aircraft, ships and submarines in wind and water tunnels. Hot- wire anemometers and associated hot...spectra of velocity fluctuations in the low-speed wind tunnel at DST Group. The use of both single- wire and crossed- wire (2 wires ) probes is

Quantification of the Uncertainties for the Space Launch System Liftoff/Transition and Ascent Databases

NASA Technical Reports Server (NTRS)

Favaregh, Amber L.; Houlden, Heather P.; Pinier, Jeremy T.

2016-01-01

A detailed description of the uncertainty quantification process for the Space Launch System Block 1 vehicle configuration liftoff/transition and ascent 6-Degree-of-Freedom (DOF) aerodynamic databases is presented. These databases were constructed from wind tunnel test data acquired in the NASA Langley Research Center 14- by 22-Foot Subsonic Wind Tunnel and the Boeing Polysonic Wind Tunnel in St. Louis, MO, respectively. The major sources of error for these databases were experimental error and database modeling errors.

Numerical calculation of transonic flow about slender bodies of revolution

NASA Technical Reports Server (NTRS)

Bailey, F. R.

1971-01-01

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

A comparison of Wortmann airfoil computer-generated lift and drag polars with flight and wind tunnel results

NASA Technical Reports Server (NTRS)

Bowers, A. H.; Sim, A. G.

1984-01-01

Computations of drag polars for a low-speed Wortmann sailplane airfoil are compared with both wind tunnel and flight test results. Excellent correlation was shown to exist between computations and flight results except when separated flow regimes were encountered. Smoothness of the input coordinates to the PROFILE computer program was found to be essential to obtain accurate comparisons of drag polars or transition location to either the flight or wind tunnel flight results.

A Hydrogen Peroxide Hot-Jet Simulator for Wind-Tunnel Tests of Turbojet-Exit Models

NASA Technical Reports Server (NTRS)

Runckel, Jack F.; Swihart, John M.

1959-01-01

A turbojet-engine-exhaust simulator which utilizes a hydrogen peroxide gas generator has been developed for powered-model testing in wind tunnels with air exchange. Catalytic decomposition of concentrated hydrogen peroxide provides a convenient and easily controlled method of providing a hot jet with characteristics that correspond closely to the jet of a gas turbine engine. The problems associated with simulation of jet exhausts in a transonic wind tunnel which led to the selection of a liquid monopropellant are discussed. The operation of the jet simulator consisting of a thrust balance, gas generator, exit nozzle, and auxiliary control system is described. Static-test data obtained with convergent nozzles are presented and shown to be in good agreement with ideal calculated values.

Wing configuration on Wind Tunnel Testing of an Unmanned Aircraft Vehicle

NASA Astrophysics Data System (ADS)

Daryanto, Yanto; Purwono, Joko; Subagyo

2018-04-01

Control surface of an Unmanned Aircraft Vehicle (UAV) consists of flap, aileron, spoiler, rudder, and elevator. Every control surface has its own special functionality. Some particular configurations in the flight mission often depend on the wing configuration. Configuration wing within flap deflection for takeoff setting deflection of flap 20° but during landing deflection of flap set on the value 40°. The aim of this research is to get the ultimate CLmax for take-off flap deflection setting. It is shown from Wind Tunnel Testing result that the 20° flap deflection gives optimum CLmax with moderate drag coefficient. The results of Wind Tunnel Testing representing by graphic plots show good performance as well as the stability of UAV.

Full-scale wind-tunnel test of the aeroelastic stability of a bearingless main rotor

NASA Technical Reports Server (NTRS)

Warmbrodt, W.; Mccloud, J., III; Sheffler, M.; Staley, J.

1981-01-01

The rotor studied in the wind tunnel had previously been flight tested on a BO-105 helicopter. The investigation was conducted to determine the rotor's aeroelastic stability characteristics in hover and at airspeeds up to 143 knots. These characteristics are compared with those obtained from whirl-tower and flight tests and predictions from a digital computer simulation. It was found that the rotor was stable for all conditions tested. At constant tip speed, shaft angle, and airspeed, stability increases with blade collective pitch setting. No significant change in system damping occurred that was attributable to frequency coalescence between the rotor inplane regressing mode and the support modes. Stability levels determined in the wind tunnel were of the same magnitude and yielded the same trends as data obtained from whirl-tower and flight tests.

Practical Applications of a Building Method to Construct Aerodynamic Database of Guided Missile Using Wind Tunnel Test Data

NASA Astrophysics Data System (ADS)

Kim, Duk-hyun; Lee, Hyoung-Jin

2018-04-01

A study of efficient aerodynamic database modeling method was conducted. A creation of database using periodicity and symmetry characteristic of missile aerodynamic coefficient was investigated to minimize the number of wind tunnel test cases. In addition, studies of how to generate the aerodynamic database when the periodicity changes due to installation of protuberance and how to conduct a zero calibration were carried out. Depending on missile configurations, the required number of test cases changes and there exist tests that can be omitted. A database of aerodynamic on deflection angle of control surface can be constituted using phase shift. A validity of modeling method was demonstrated by confirming that the result which the aerodynamic coefficient calculated by using the modeling method was in agreement with wind tunnel test results.

Scaling Methods for Simulating Aircraft In-Flight Icing Encounters

NASA Technical Reports Server (NTRS)

Anderson, David N.; Ruff, Gary A.

1997-01-01

This paper discusses scaling methods which permit the use of subscale models in icing wind tunnels to simulate natural flight in icing. Natural icing conditions exist when air temperatures are below freezing but cloud water droplets are super-cooled liquid. Aircraft flying through such clouds are susceptible to the accretion of ice on the leading edges of unprotected components such as wings, tailplane and engine inlets. To establish the aerodynamic penalties of such ice accretion and to determine what parts need to be protected from ice accretion (by heating, for example), extensive flight and wind-tunnel testing is necessary for new aircraft and components. Testing in icing tunnels is less expensive than flight testing, is safer, and permits better control of the test conditions. However, because of limitations on both model size and operating conditions in wind tunnels, it is often necessary to perform tests with either size or test conditions scaled. This paper describes the theoretical background to the development of icing scaling methods, discusses four methods, and presents results of tests to validate them.

Aeronautical Facilities Catalogue. Volume 1: Wind Tunnels

NASA Technical Reports Server (NTRS)

Penaranda, F. E. (Compiler); Freda, M. S. (Compiler)

1985-01-01

Domestic and foreign wind tunnel facilities are enumerated and their technical parameters are described. Data pertinent to managers and engineers are presented. Facilities judged comparable in testing capability are noted and grouped together. Several comprehensive cross-indexes and charts are included.

Development of a 5-Component Balance for Water Tunnel Applications

NASA Technical Reports Server (NTRS)

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

1999-01-01

The principal objective of this research/development effort was to develop a multi-component strain gage balance to measure both static and dynamic forces and moments on models tested in flow visualization water tunnels. A balance was designed that allows measuring normal and side forces, and pitching, yawing and rolling moments (no axial force). The balance mounts internally in the model and is used in a manner typical of wind tunnel balances. The key differences between a water tunnel balance and a wind tunnel balance are the requirement for very high sensitivity since the loads are very low (typical normal force is 90 grams or 0.2 lbs), the need for water proofing the gage elements, and the small size required to fit into typical water tunnel models. The five-component balance was calibrated and demonstrated linearity in the responses of the primary components to applied loads, very low interactions between the sections and no hysteresis. Static experiments were conducted in the Eidetics water tunnel with delta wings and F/A-18 models. The data were compared to forces and moments from wind tunnel tests of the same or similar configurations. The comparison showed very good agreement, providing confidence that loads can be measured accurately in the water tunnel with a relatively simple multi-component internal balance. The success of the static experiments encouraged the use of the balance for dynamic experiments. Among the advantages of conducting dynamic tests in a water tunnel are less demanding motion and data acquisition rates than in a wind tunnel test (because of the low-speed flow) and the capability of performing flow visualization and force/moment (F/M) measurements simultaneously with relative simplicity. This capability of simultaneous flow visualization and for F/M measurements proved extremely useful to explain the results obtained during these dynamic tests. In general, the development of this balance should encourage the use of water tunnels for a wider range of quantitative and qualitative experiments, especially during the preliminary phase of aircraft design.

Space Launch System Liftoff and Transition Aerodynamic Characterization in the NASA Langley 14- by 22-Foot Subsonic Wind Tunnel

NASA Technical Reports Server (NTRS)

Pinier, Jeremy T.; Erickson, Gary E.; Paulson, John W.; Tomek, William G.; Bennett, David W.; Blevins, John A.

2015-01-01

A 1.75% scale force and moment model of the Space Launch System was tested in the NASA Langley Research Center 14- by 22-Foot Subsonic Wind Tunnel to quantify the aerodynamic forces that will be experienced by the launch vehicle during its liftoff and transition to ascent flight. The test consisted of two parts: the first was dedicated to measuring forces and moments for the entire range of angles of attack (0deg to 90deg) and roll angles (0 deg. to 360 deg.). The second was designed to measure the aerodynamic effects of the liftoff tower on the launch vehicle for ground winds from all azimuthal directions (0 deg. to 360 deg.), and vehicle liftoff height ratios from 0 to 0.94. This wind tunnel model also included a set of 154 surface static pressure ports. Details on the experimental setup, and results from both parts of testing are presented, along with a description of how the wind tunnel data was analyzed and post-processed in order to develop an aerodynamic database. Finally, lessons learned from experiencing significant dynamics in the mid-range angles of attack due to steady asymmetric vortex shedding are presented.

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

NASA Technical Reports Server (NTRS)

Hayden, R. E.

1984-01-01

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

Wind tunnel test of Teledyne Geotech model 1564B cup anemometer

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

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

1991-04-04

The Department of Energy (DOE) Environment, Safety and Health Compliance Assessment (Tiger Team) of the Savannah River Site (SRS) questioned the method by which wind speed sensors (cup anemometers) are calibrated by the Environmental Technology Section (ETS). The Tiger Team member was concerned that calibration data was generated by running the wind tunnel to only 26 miles per hour (mph) when speeds exceeding 50 mph are readily obtainable. A wind tunnel experiment was conducted and confirmed the validity of the practice. Wind speeds common to SRS (6 mph) were predicted more accurately by 0--25 mph regression equations than 0--50 mphmore » regression equations. Higher wind speeds were slightly overpredicted by the 0--25 mph regression equations when compared to 0--50 mph regression equations. However, the greater benefit of more accurate lower wind speed predictions accuracy outweight the benefit of slightly better high (extreme) wind speed predictions. Therefore, it is concluded that 0--25 mph regression equations should continue to be utilized by ETS at SRS. During the Department of Energy Tiger Team audit, concerns were raised about the calibration of SRS cup anemometers. Wind speed is measured by ETS with Teledyne Geotech model 1564B cup anemometers, which are calibrated in the ETS wind tunnel. Linear regression lines are fitted to data points of tunnel speed versus anemometer output voltages up to 25 mph. The regression coefficients are then implemented into the data acquisition computer software when an instrument is installed in the field. The concern raised was that since the wind tunnel at SRS is able to generate a maximum wind speed higher than 25 mph, errors may be introduced in not using the full range of the wind tunnel.« less

Wind tunnel test of Teledyne Geotech model 1564B cup anemometer

NASA Astrophysics Data System (ADS)

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

1991-04-01

The Department of Energy (DOE) Environment, Safety, and Health Compliance Assessment (Tiger Team) of the Savannah River Site (SRS) questioned the method by which wind speed sensors (cup anemometers) are calibrated by the Environmental Technology Section (ETS). The Tiger Team member was concerned that calibration data was generated by running the wind tunnel to only 26 miles per hour (mph) when speeds exceeding 50 mph are readily obtainable. A wind tunnel experiment was conducted and confirmed the validity of the practice. Wind speeds common to SRS (6 mph) were predicted more accurately by 0-25 mph regression equations than 0-50 mph regression equations. Higher wind speeds were slightly overpredicted by the 0-25 mph regression equations when compared to 0-50 mph regression equations. However, the greater benefit of more accurate lower wind speed predictions accuracy outweigh the benefit of slightly better high (extreme) wind speed predictions. Therefore, it is concluded that 0-25 mph regression equations should continue to be utilized by ETS at SRS. During the Department of Energy Tiger Team audit, concerns were raised about the calibration of SRS cup anemometers. Wind speed is measured by ETS with Teledyne Geotech model 1564B cup anemometers, which are calibrated in the ETS wind tunnel. Linear regression lines are fitted to data points of tunnel speed versus anemometer output voltages up to 25 mph. The regression coefficients are then implemented into the data acquisition computer software when an instrument is installed in the field. The concern raised was that since the wind tunnel at SRS is able to generate a maximum wind speed higher than 25 mph, errors may be introduced in not using the full range of the wind tunnel.

Prediction of Flutter Boundary Using Flutter Margin for The Discrete-Time System

NASA Astrophysics Data System (ADS)

Dwi Saputra, Angga; Wibawa Purabaya, R.

2018-04-01

Flutter testing in a wind tunnel is generally conducted at subcritical speeds to avoid damages. Hence, The flutter speed has to be predicted from the behavior some of its stability criteria estimated against the dynamic pressure or flight speed. Therefore, it is quite important for a reliable flutter prediction method to estimates flutter boundary. This paper summarizes the flutter testing of a wing cantilever model in a wind tunnel. The model has two degree of freedom; they are bending and torsion modes. The flutter test was conducted in a subsonic wind tunnel. The dynamic data responses was measured by two accelerometers that were mounted on leading edge and center of wing tip. The measurement was repeated while the wind speed increased. The dynamic responses were used to determine the parameter flutter margin for the discrete-time system. The flutter boundary of the model was estimated using extrapolation of the parameter flutter margin against the dynamic pressure. The parameter flutter margin for the discrete-time system has a better performance for flutter prediction than the modal parameters. A model with two degree freedom and experiencing classical flutter, the parameter flutter margin for the discrete-time system gives a satisfying result in prediction of flutter boundary on subsonic wind tunnel test.

Visual display and alarm system for wind tunnel static and dynamic loads

NASA Technical Reports Server (NTRS)

Hanly, Richard D.; Fogarty, James T.

1987-01-01

A wind tunnel balance monitor and alarm system developed at NASA Ames Research Center will produce several beneficial results. The costs of wind tunnel delays because of inadvertent balance damage and the costs of balance repair or replacement can be greatly reduced or eliminated with better real-time information on the balance static and dynamic loading. The wind tunnel itself will have enhanced utility with the elimination of overly cautious limits on test conditions. The microprocessor-based system features automatic scaling and 16 multicolored LED bargraphs to indicate both static and dynamic components of the signals from eight individual channels. Five individually programmable alarm levels are available with relay closures for internal or external visual and audible warning devices and other functions such as automatic activation of external recording devices, model positioning mechanisms, or tunnel shutdown.

Visual display and alarm system for wind tunnel static and dynamic loads

NASA Technical Reports Server (NTRS)

Hanly, Richard D.; Fogarty, James T.

1987-01-01

A wind tunnel balance monitor and alarm system developed at NASA Ames Research Center will produce several beneficial results. The costs of wind tunnel delays because of inadvertent balance damage and the costs of balance repair or replacement can be greatly reduced or eliminated with better real-time information on the balance static and dynamic loading. The wind tunnel itself will have enhanced utility with the elimination of overly cautious limits on test conditions. The microprocessor-based system features automatic scaling and 16 multicolored LED bargraphs to indicate both static and dynamic components of the signals from eight individual channels. Five individually programmable alarm levels are available with relay closures for internal or external visual and audible warning devices and other functions such as automatic activation of external recording devices, model positioning mechanism, or tunnel shutdown.

New methodologies for calculation of flight parameters on reduced scale wings models in wind tunnel =

NASA Astrophysics Data System (ADS)

Ben Mosbah, Abdallah

In order to improve the qualities of wind tunnel tests, and the tools used to perform aerodynamic tests on aircraft wings in the wind tunnel, new methodologies were developed and tested on rigid and flexible wings models. A flexible wing concept is consists in replacing a portion (lower and/or upper) of the skin with another flexible portion whose shape can be changed using an actuation system installed inside of the wing. The main purpose of this concept is to improve the aerodynamic performance of the aircraft, and especially to reduce the fuel consumption of the airplane. Numerical and experimental analyses were conducted to develop and test the methodologies proposed in this thesis. To control the flow inside the test sections of the Price-Paidoussis wind tunnel of LARCASE, numerical and experimental analyses were performed. Computational fluid dynamics calculations have been made in order to obtain a database used to develop a new hybrid methodology for wind tunnel calibration. This approach allows controlling the flow in the test section of the Price-Paidoussis wind tunnel. For the fast determination of aerodynamic parameters, new hybrid methodologies were proposed. These methodologies were used to control flight parameters by the calculation of the drag, lift and pitching moment coefficients and by the calculation of the pressure distribution around an airfoil. These aerodynamic coefficients were calculated from the known airflow conditions such as angles of attack, the mach and the Reynolds numbers. In order to modify the shape of the wing skin, electric actuators were installed inside the wing to get the desired shape. These deformations provide optimal profiles according to different flight conditions in order to reduce the fuel consumption. A controller based on neural networks was implemented to obtain desired displacement actuators. A metaheuristic algorithm was used in hybridization with neural networks, and support vector machine approaches and their combination was optimized, and very good results were obtained in a reduced computing time. The validation of the obtained results has been made using numerical data obtained by the XFoil code, and also by the Fluent code. The results obtained using the methodologies presented in this thesis have been validated with experimental data obtained using the subsonic Price-Paidoussis blow down wind tunnel.

Investigation of Rotor Performance and Loads of a UH-60A Individual Blade Control System

NASA Technical Reports Server (NTRS)

Yeo, Hyeonsoo; Romander, Ethan A.; Norman, Thomas R.

2010-01-01

A full-scale wind tunnel test was recently conducted (March 2009) in the National Full-Scale Aerodynamics Complex (NFAC) 40- by 80-FootWind Tunnel to evaluate the potential of an individual blade control (IBC) system to improve rotor performance and reduce vibrations, loads, and noise for a UH-60A rotor system [1]. This test was the culmination of a long-termcollaborative effort between NASA, U.S. Army, Sikorsky Aircraft Corporation, and ZF Luftfahrttechnik GmbH (ZFL) to demonstrate the benefits of IBC for a UH-60Arotor. Figure 1 shows the UH-60Arotor and IBC system mounted on the NFAC Large Rotor Test Apparatus (LRTA). The IBC concept used in the current study utilizes actuators placed in the rotating frame, one per blade. In particular, the pitch link of the rotor blade was replacedwith an actuator, so that the blade root pitch can be changed independently. This concept, designed for a full-scale UH-60A rotor, was previously tested in the NFAC 80- by 120-FootWind Tunnel in September 2001 at speeds up to 85 knots [2]. For the current test, the same UH-60A rotor and IBC system were tested in the 40- by 80-FootWind Tunnel at speeds up to 170 knots. Figure 2 shows the servo-hydraulic IBC actuator installed between the swashplate and the blade pitch horn. Although previous wind tunnel experiments [3, 4] and analytical studies on IBC [5, 6] have shown the promise to improve the rotor s performance, in-depth correlation studies have not been performed. Thus, the current test provides a unique resource that can be used to assess the accuracy and reliability of prediction methods and refine theoretical models, with the ultimate goal of providing the technology for timely and cost-effective design and development of new rotors. In this paper, rotor performance and loads calculations are carried out using the analyses CAMRAD II and coupled OVERFLOW-2/CAMRAD II and the results are compared with these UH-60A/IBC wind tunnel test data.

A measurement of forward-flight effects on the noise from a JT15D-1 turbofan engine in the NASA-Ames 40- by 80-Foot Wind Tunnel

NASA Technical Reports Server (NTRS)

Ahtye, W. F.

1980-01-01

A Pratt and Whitney JT15D-1 turbofan engine was tested in two facilities at Ames Research Center: the outdoor Static Test Facility and the 40- by 80-Foot Wind Tunnel. The primary purposes of the test were to determine the effects of forward velocity on the turbofan spectra in the forward quadrant for the cruise inlet and to compare these wind-tunnel spectra with outdoor spectra to determine the possibility of simulating forward-velocity effects from purely outdoor measurements. The wind-tunnel data show a reduction in the blade-passage frequency tones of the order of 10 dB with increasing forward velocity at subsonic fan-tip speeds. No forward-velocity variation was observed at supersonic tip speeds. Comparison of in-duct spectra for the cruise inlet at forward velocity, with spectra from outdoor tests with a distortion-control inlet shows excellent agreement for the in-duct data when allowance is made for different in-duct volumes. This is also reflected in good agreement for the far-field spectra at small forward angles. The comparisons of wind-tunnel and outdoor data also indicate that at least for the JT15D-1, it may be possible to approximate the shape of the far-field spectra at large directivity angles from an outdoor measurement with the cruise inlet, providing an effective inflow control device is used.

Wind Tunnel Studies in Aerodynamic Phenomena at High Speed

NASA Technical Reports Server (NTRS)

Caldwell, F W; Fales, E N

1921-01-01

A great amount of research and experimental work has been done and fair success obtained in an effort to place airplane and propeller design upon an empirical basis. However, one can not fail to be impressed by the apparent lack of data available toward establishing flow phenomena upon a rational basis, such that they may be interpreted in terms of the laws of physics. With this end in view it was the object of the authors to design a wind tunnel differing from the usual type especially in regard to large power and speed of flow. This report describes the wind tunnel at Mccook Field and gives the results of experiments conducted in testing the efficiency of the wind tunnel.

KC-135 Winglet Program Review

NASA Technical Reports Server (NTRS)

1982-01-01

The results of a joint NASA/USAF program to develop flight test winglets on a KC-135 aircraft are reviewed. The winglet development from concept through wind tunnel and flight tests is discussed. Predicted, wind tunnel, and flight test results are compared for the performance, loads and flutter characteristics of the winglets. The flight test winglets had a variable winglet cant and incidence angle capability which enabled a limited evaluation of the effects of these geometry changes.

Laser velocimeter data acquisition system for the Langley 14- by 22-foot subsonic tunnel. Software reference guide version 3.3

NASA Technical Reports Server (NTRS)

Jumper, Judith K.

1994-01-01

The Laser Velocimeter Data Acquisition System (LVDAS) in the Langley 14- by 22-Foot Tunnel is controlled by a comprehensive software package. The software package was designed to control the data acquisition process during wind tunnel tests which employ a laser velocimeter measurement system. This report provides detailed explanations on how to configure and operate the LVDAS system to acquire laser velocimeter and static wind tunnel data.

26. Photocopy of photograph (original in the Langley Research Center ...

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

26. Photocopy of photograph (original in the Langley Research Center Archives, Hampton, VA LaRC) (L64792) ALBACORE SUBMARINE DRAG TESTS IN THE FULL-SCALE WIND TUNNEL. - NASA Langley Research Center, Full-Scale Wind Tunnel, 224 Hunting Avenue, Hampton, Hampton, VA

Program of Research in Flight Dynamics in the JIAFS, George Washington University at NASA Langley Research Center

NASA Technical Reports Server (NTRS)

Klein, Vladislav

2002-01-01

The program objectives are fully defined in the original proposal entitled 'Program of Research in Flight Dynamics in GW at NASA Langley Research Center,' which was originated March 20, 1975, and in the renewals of the research program from December 1, 2000 to November 30, 2001. The program in its present form includes three major topics: 1) the improvement of existing methods and development of new methods for wind tunnel and flight test data analysis, 2) the application of these methods to wind tunnel and flight test data obtained from advanced airplanes, 3) the correlation of flight results with wind tunnel measurements, and theoretical predictions. The Principal Investigator of the program is Dr. Vladislav Klein. Three Graduate Research Scholar Assistants (K. G. Mas, M. M. Eissa and N. M. Szyba) also participated in the program. Specific developments in the program during the period Dec. 1, 2001 through Nov. 30, 2002 included: 1) Data analysis of highly swept delta wing aircraft from wind and water tunnel data, and 2) Aerodynamic characteristics of the radio control aircraft from flight test.

Wind tunnel test of the 0.019 (2A configuration) jet plume space shuttle integrated vehicle in the ARC 9- by 7-foot unitary wind tunnel (IA12B)

NASA Technical Reports Server (NTRS)

Hardin, R. B.; Burrows, R. R.

1974-01-01

The wind tunnel test of the 0.019 jet plume space shuttle integrated vehicle in the Ames 9 ft by 7 ft unitary wind tunnel was conducted at Mach numbers of 1.55 and 2.0 over a Reynolds number range from 3.5 million to 4.1 million/ft. Data were obtained at angles of attack from minus 8 deg to plus 8 deg at 0 deg sideslip and at angles of sideslip from minus 9 deg to plus 8 deg at 0 deg angle of attack. The basic configuration tested was the 2A vehicle with the orbiter at 0 deg angle of incidence with respect to the external tank. The other deviations to the 2A configuration were the solid rocket motor shrouds, which were designed to vehicle '3' lines, and the tank nose, which consisted of the retro-package being removed and replaced by a 16.5 inch full scale radius nose.

Aerodynamic Analyses and Database Development for Lift-Off/Transition and First Stage Ascent of the Ares I A106 Vehicle

NASA Technical Reports Server (NTRS)

Pamadi, Bandu N.; Pei, Jing; Covell, Peter F.; Favaregh, Noah M.; Gumbert, Clyde R.; Hanke, Jeremy L.

2011-01-01

NASA Langley Research Center, in partnership with NASA Marshall Space Flight Center and NASA Ames Research Center, was involved in the aerodynamic analyses, testing, and database development for the Ares I A106 crew launch vehicle in support of the Ares Design and Analysis Cycle. This paper discusses the development of lift-off/transition and ascent databases. The lift-off/transition database was developed using data from tests on a 1.75% scale model of the A106 configuration in the NASA Langley 14x22 Subsonic Wind Tunnel. The power-off ascent database was developed using test data on a 1% A106 scale model from two different facilities, the Boeing Polysonic Wind Tunnel and the NASA Langley Unitary Plan Wind Tunnel. The ascent database was adjusted for differences in wind tunnel and flight Reynolds numbers using USM3D CFD code. The aerodynamic jet interaction effects due to first stage roll control system were modeled using USM3D and OVERFLOW CFD codes.

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

NASA Technical Reports Server (NTRS)

Koning, Witold J. F.

2015-01-01

Experimental techniques to measure rotorcraft aerodynamic performance are widely used. However, most of them are either unable to capture interference effects from bodies, or require an extremely large computational budget. The objective of the present research is to develop an XV-15 Tilt Rotor Research Aircraft rotor model for investigation of wind tunnel wall interference using a novel Computational Fluid Dynamics (CFD) solver for rotorcraft, RotCFD. In RotCFD, a mid-fidelity URANS solver is used with an incompressible flow model and a realizable k-e turbulence model. The rotor is, however, not modeled using a computationally expensive, unsteady viscous body-fitted grid, but is instead modeled using a blade element model with a momentum source approach. Various flight modes of the XV-15 isolated rotor, including hover, tilt and airplane mode, have been simulated and correlated to existing experimental and theoretical data. The rotor model is subsequently used for wind tunnel wall interference simulations in the National Full-Scale Aerodynamics Complex (NFAC) at NASA Ames Research Center in California. The results from the validation of the isolated rotor performance showed good correlation with experimental and theoretical data. The results were on par with known theoretical analyses. In RotCFD the setup, grid generation and running of cases is faster than many CFD codes, which makes it a useful engineering tool. Performance predictions need not be as accurate as high-fidelity CFD codes, as long as wall effects can be properly simulated. For both test sections of the NFAC wall interference was examined by simulating the XV-15 rotor in the test section of the wind tunnel and with an identical grid but extended boundaries in free field. Both cases were also examined with an isolated rotor or with the rotor mounted on the modeled geometry of the Tiltrotor Test Rig (TTR). A 'quasi linear trim' was used to trim the thrust for the rotor to compare the power as a unique variable. Power differences between free field and wind tunnel cases were found from -7 % to 0 % in the 80- by 120-Foot Wind Tunnel test section and -1.6 % to 4.8 % in the 40- by 80-Foot Wind Tunnel, depending on the TTR orientation, tunnel velocity and blade setting. The TTR will be used in 2016 to test the Bell 609 rotor in a similar fashion to the research in this report.

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

NASA Technical Reports Server (NTRS)

Nutt, K. W.

1984-01-01

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

Space shuttle solid rocket booster sting interference wind tunnel test analysis

NASA Technical Reports Server (NTRS)

Conine, B.; Boyle, W.

1981-01-01

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

Wind tunnel investigation of aerodynamic characteristics of scale models of three rectangular shaped cargo containers

NASA Technical Reports Server (NTRS)

Laub, G. H.; Kodani, H. M.

1972-01-01

Wind tunnel tests were conducted on scale models of three rectangular shaped cargo containers to determine the aerodynamic characteristics of these typical externally-suspended helicopter cargo configurations. Tests were made over a large range of pitch and yaw attitudes at a nominal Reynolds number per unit length of 1.8 x one million. The aerodynamic data obtained from the tests are presented.

Wind tunnel test results for the direction controlled antitank DCAT missile at Mach numbers from 0.64 to 2.50

NASA Technical Reports Server (NTRS)

Martin, T. A.; Spring, D. J.

1973-01-01

Wind tunnel test results are presented to show aerodynamic characteristics over the Mach number range of 0.64 to 2.50 of the DCAT missile. Data are presented showing the interference created by the rear mounted reaction control system. Two candidate fins were installed on the model during tests: a flat folding fin and a curved wrap around fin.

Analysis and design of quiet hypersonic wind tunnels

NASA Astrophysics Data System (ADS)

Naiman, Hadassah

The purpose of the present work is to integrate CFD into the design of quiet hypersonic wind tunnels and the analysis of their performance. Two specific problems are considered. The first problem is the automated design of the supersonic portion of a quiet hypersonic wind tunnel. Modern optimization software is combined with full Navier-Stokes simulations and PSE stability analysis to design a Mach 6 nozzle with maximum quiet test length. A response surface is constructed from a user-specified set of contour shapes and a genetic algorithm is used to find the "optimal contour", which is defined as the shortest nozzle with the maximum quiet test length. This is achieved by delaying transition along the nozzle wall. It is found that transition is triggered by Goertler waves, which can be suppressed by including a section of convex curvature along the contour. The optimal design has an unconventional shape described as compound curvature, which makes the contour appear slightly wavy. The second problem is the evaluation of a proposed modification of the test section in the Boeing/AFOSR Mach 6 Quiet Tunnel. The new design incorporates a section of increased diameter with the intention of enabling the tunnel to start in the presence of larger blunt models. Cone models with fixed base diameter (and hence fixed blockage ratio) are selected for this study. Cone half-angles from 15° to 75° are examined to ascertain the effect of ii the strength of the test model shock wave on the tunnel startup. The unsteady, laminar, compressible Navier-Stokes equations are solved. The resulting flowfields are analyzed to see what affect the shocks and shear layers have on the quiet test section flow. This study indicates that cone angles ≤20° allow the tunnel to start. Keywords. automated optimization, response surface, parabolized stability equations, compound curvature, laminar, wind tunnel, unstart, test section.

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.

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

NASA Technical Reports Server (NTRS)

Deloach, Richard

2013-01-01

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

Toward Real Time Neural Net Flight Controllers

NASA Technical Reports Server (NTRS)

Jorgensen, C. C.; Mah, R. W.; Ross, J.; Lu, Henry, Jr. (Technical Monitor)

1994-01-01

NASA Ames Research Center has an ongoing program in neural network control technology targeted toward real time flight demonstrations using a modified F-15 which permits direct inner loop control of actuators, rapid switching between alternative control designs, and substitutable processors. An important part of this program is the ACTIVE flight project which is examining the feasibility of using neural networks in the design, control, and system identification of new aircraft prototypes. This paper discusses two research applications initiated with this objective in mind: utilization of neural networks for wind tunnel aircraft model identification and rapid learning algorithms for on line reconfiguration and control. The first application involves the identification of aerodynamic flight characteristics from analysis of wind tunnel test data. This identification is important in the early stages of aircraft design because complete specification of control architecture's may not be possible even though concept models at varying scales are available for aerodynamic wind tunnel testing. Testing of this type is often a long and expensive process involving measurement of aircraft lift, drag, and moment of inertia at varying angles of attack and control surface configurations. This information in turn can be used in the design of the flight control systems by applying the derived lookup tables to generate piece wise linearized controllers. Thus, reduced costs in tunnel test times and the rapid transfer of wind tunnel insights into prototype controllers becomes an important factor in more efficient generation and testing of new flight systems. NASA Ames Research Center is successfully applying modular neural networks as one way of anticipating small scale aircraft model performances prior to testing, thus reducing the number of in tunnel test hours and potentially, the number of intermediate scaled models required for estimation of surface flow effects.

Opening Loads Analyses for Various Disk-Gap-Band Parachutes

NASA Technical Reports Server (NTRS)

Cruz, J. R.; Kandis, M.; Witkowski, A.

2003-01-01

Detailed opening loads data is presented for 18 tests of Disk-Gap-Band (DGB) parachutes of varying geometry with nominal diameters ranging from 43.2 to 50.1 ft. All of the test parachutes were deployed from a mortar. Six of these tests were conducted via drop testing with drop test vehicles weighing approximately 3,000 or 8,000 lb. Twelve tests were conducted in the National Full-Scale Aerodynamics Complex 80- by 120-foot wind tunnel at the NASA Ames Research Center. The purpose of these tests was to structurally qualify the parachute for the Mars Exploration Rover mission. A key requirement of all tests was that peak parachute load had to be reached at full inflation to more closely simulate the load profile encountered during operation at Mars. Peak loads measured during the tests were in the range from 12,889 to 30,027 lb. Of the two test methods, the wind tunnel tests yielded more accurate and repeatable data. Application of an apparent mass model to the opening loads data yielded insights into the nature of these loads. Although the apparent mass model could reconstruct specific tests with reasonable accuracy, the use of this model for predictive analyses was not accurate enough to set test conditions for either the drop or wind tunnel tests. A simpler empirical model was found to be suitable for predicting opening loads for the wind tunnel tests to a satisfactory level of accuracy. However, this simple empirical model is not applicable to the drop tests.

Blockage corrections for three-dimensional-flow closed-throat wind tunnels, with consideration of the effect of compressibility

NASA Technical Reports Server (NTRS)

Herriot, John G

1947-01-01

Theoretical blockage corrections are presented for a body of revolution and for a three-dimensional unswept wing in a circular or rectangular wind tunnel. The theory takes account of the effects of the wake and of the compressibility of the fluid, and is based on the assumption that the dimensions of the model are small in comparison with those of the tunnel throat. Formulas are given for correcting a number of the quantities, such as dynamic pressure and Mach number, measured in wind-tunnel tests. The report presents a summary and unification of the existing literature on the subject.

Design and calibration of the carousel wind tunnel

NASA Technical Reports Server (NTRS)

Leach, R. N.; Greeley, R.; Iversen, J.; White, B.; Marshall, J. R.

1986-01-01

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 interparticle forces but the two terms are not separable. A wind tunnel that would permit variable gravity would allow separation of the forces and aid greatly in understanding planetary aeolian processes. The design Carousel Wind Tunnel (CWT) allows for a long flow distance in a small sized tunnel since the test section is a continuo us circuit and allows for a variable pseudo gravity. A prototype design was built and calibrated to gain some understanding of the characteristics of the design and the results presented.

Design and calibration of the carousel wind tunnel

NASA Technical Reports Server (NTRS)

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

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 interparticle forces but the two terms are not separable. A wind tunnel that would permit variable gravity would allow separation of the forces and aid greatly in understanding planetary aeolian processes. The design of the Carousel Wind Tunnel (CWT) allows for a long flow distance in a small sized tunnel since the test section is a continuous circuit and allows for a variable pseudo-gravity. A prototype design was built and calibrated to gain some understanding of the characteristics of the design and the results presented.

Metric half-span model support system

NASA Technical Reports Server (NTRS)

Jackson, C. M., Jr.; Dollyhigh, S. M.; Shaw, D. S. (Inventor)

1982-01-01

A model support system used to support a model in a wind tunnel test section is described. The model comprises a metric, or measured, half-span supported by a nonmetric, or nonmeasured half-span which is connected to a sting support. Moments and forces acting on the metric half-span are measured without interference from the support system during a wind tunnel test.

Wind-Tunnel Research Comparing Lateral Control Devices Particularly at High Angles of Attack XIII : Auxiliary Airfoils Used as External Ailerons

NASA Technical Reports Server (NTRS)

Weick, Fred E; Noyes, Richard W

1936-01-01

This is the thirteenth report on a series of systematic tests comparing lateral control devices with particular reference to their effectiveness at high angles of attack. The present wind tunnel tests were made to determine the most feasible locations for lateral control surfaces mounted externally to a rectangular Clark y wing.

Computational Design and Analysis of a Transonic Natural Laminar Flow Wing for a Wind Tunnel Model

NASA Technical Reports Server (NTRS)

Lynde, Michelle N.; Campbell, Richard L.

2017-01-01

A natural laminar flow (NLF) wind tunnel model has been designed and analyzed for a wind tunnel test in the National Transonic Facility (NTF) at the NASA Langley Research Center. The NLF design method is built into the CDISC design module and uses a Navier-Stokes flow solver, a boundary layer profile solver, and stability analysis and transition prediction software. The NLF design method alters the pressure distribution to support laminar flow on the upper surface of wings with high sweep and flight Reynolds numbers. The method addresses transition due to attachment line contamination/transition, Gortler vortices, and crossflow and Tollmien-Schlichting modal instabilities. The design method is applied to the wing of the Common Research Model (CRM) at transonic flight conditions. Computational analysis predicts significant extents of laminar flow on the wing upper surface, which results in drag savings. A 5.2 percent scale semispan model of the CRM NLF wing will be built and tested in the NTF. This test will aim to validate the NLF design method, as well as characterize the laminar flow testing capabilities in the wind tunnel facility.

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.

Wind erosion in semiarid landscapes: Predictive models and remote sensing methods for the influence of vegetation

NASA Technical Reports Server (NTRS)

Musick, H. Brad

1993-01-01

The objectives of this research are: to develop and test predictive relations for the quantitative influence of vegetation canopy structure on wind erosion of semiarid rangeland soils, and to develop remote sensing methods for measuring the canopy structural parameters that determine sheltering against wind erosion. The influence of canopy structure on wind erosion will be investigated by means of wind-tunnel and field experiments using structural variables identified by the wind-tunnel and field experiments using model roughness elements to simulate plant canopies. The canopy structural variables identified by the wind-tunnel and field experiments as important in determining vegetative sheltering against wind erosion will then be measured at a number of naturally vegetated field sites and compared with estimates of these variables derived from analysis of remotely sensed data.

Mercury Capsule Retrorocket Test in the Altitude Wind Tunnel

NASA Image and Video Library

1960-09-21

A mechanic at the National Aeronautics and Space Administration (NASA) Lewis Research Center prepares the inverted base of a Mercury capsule for a test of its posigrade retrorockets inside the Altitude Wind Tunnel. In October 1959 NASA’s Space Task Group allocated several Project Mercury assignments to Lewis. The Altitude Wind Tunnel was modified to test the Atlas separation system, study the escape tower rocket plume, train astronauts to bring a spinning capsule under control, and calibrate the capsule’s retrorockets. The turning vanes, makeup air pipes, and cooling coils were removed from the wide western end of the tunnel to create a 51-foot diameter test chamber. The Mercury capsule had a six-rocket retro-package affixed to the bottom of the capsule. Three of these were posigrade rockets used to separate the capsule from the booster and three were retrograde rockets used to slow the capsule for reentry into the earth’s atmosphere. Performance of the retrorockets was vital since there was no backup system. Qualification tests of the retrorockets began in April 1960 on a retrograde thrust stand inside the southwest corner of the Altitude Wind Tunnel. These studies showed that a previous issue concerning the delayed ignition of the propellant had been resolved. Follow-up test runs verified reliability of the igniter’s attachment to the propellant. In addition, the capsule’s retrorockets were calibrated so they would not alter the capsule’s attitude when fired.

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.

Aeroservoelastic wind-tunnel investigations using the Active Flexible Wing Model: Status and recent accomplishments

NASA Technical Reports Server (NTRS)

Noll, Thomas E.; Perry, Boyd, III; Tiffany, Sherwood H.; Cole, Stanley R.; Buttrill, Carey S.; Adams, William M., Jr.; Houck, Jacob A.; Srinathkumar, S.; Mukhopadhyay, Vivek; Pototzky, Anthony S.

1989-01-01

The status of the joint NASA/Rockwell Active Flexible Wing Wind-Tunnel Test Program is described. The objectives are to develop and validate the analysis, design, and test methodologies required to apply multifunction active control technology for improving aircraft performance and stability. Major tasks include designing digital multi-input/multi-output flutter-suppression and rolling-maneuver-load alleviation concepts for a flexible full-span wind-tunnel model, obtaining an experimental data base for the basic model and each control concept and providing comparisons between experimental and analytical results to validate the methodologies. The opportunity is provided to improve real-time simulation techniques and to gain practical experience with digital control law implementation procedures.

Wind-tunnel free-flight investigation of a supersonic persistence fighter

NASA Technical Reports Server (NTRS)

Hahne, David E.; Wendel, Thomas R.; Boland, Joseph R.

1993-01-01

Wind-tunnel free-flight tests have been conducted in the Langley 30- by 60-Foot Wind Tunnel to examine the high-angle-of-attack stability and control characteristics and control law design of a supersonic persistence fighter (SSPF) at 1 g flight conditions. In addition to conventional control surfaces, the SSPF incorporated deflectable wingtips (tiperons) and pitch and yaw thrust vectoring. A direct eigenstructure assignment technique was used to design control laws to provide good flying characteristics well into the poststall angle-of-attack region. Free-flight tests indicated that it was possible to blend effectively conventional and unconventional control surfaces to achieve good flying characteristics well into the poststall angle-of-attack region.

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.

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

NASA Astrophysics Data System (ADS)

Gorbushin, A. R.

2017-11-01

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

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)

A full-scale wind tunnel investigation of a helicopter bearingless main rotor. [Ames 40 by 80 Wind Tunnel

NASA Technical Reports Server (NTRS)

Warmbrodt, W.; Mccloud, J. L., II

1981-01-01

A helicopter bearingless main rotor was tested. Areas of investigation included aeroelastic stability, aerodynamic performance, and rotor loads as a function of collective pitch setting, RPM, airspeed and shaft angle. The rotor/support system was tested with the wind tunnel balance dampers installed and, subsequently, removed. Modifications to the rotor hub were tested. These included a reduction in the rotor control system stiffness and increased flexbeam structural damping. The primary objective of the test was to determine aeroelastic stability of the fundamental flexbeam/blade chordwise bending mode. The rotor was stable for all conditions. Damping of the rotor chordwise bending mode increases with increased collective pitch angle at constant operating conditions. No significant decrease in rotor damping occured due to frequency coalescence between the blade chordwise fundamental bending mode and the support system.

Dynamic response of NASA Rotor Test Apparatus and Sikorsky S-76 hub mounted in the 80- by 120-Foot Wind Tunnel

NASA Technical Reports Server (NTRS)

Peterson, Randall L.; Hoque, Muhammed S.

1994-01-01

A shake test was conducted in the 80- by 120-Foot Wind Tunnel at NASA Ames Research Center, using the NASA Ames Rotor Test Apparatus (RTA) and the Sikorsky S-76 rotor hub. The primary objective of this shake test was to determine the modal properties of the RTA, the S-76 rotor hub, and the model support system installed in the wind tunnel. Random excitation was applied at the rotor hub, and vibration responses were measured using accelerometers mounted at various critical locations on the model and the model support system. Transfer functions were computed using the load cell data and the accelerometer responses. The transfer function data were used to compute the system modal parameters with the aid of modal analysis software.

Laminar Flow Supersonic Wind Tunnel primary air injector

NASA Technical Reports Server (NTRS)

Smith, Brooke Edward

1993-01-01

This paper describes the requirements, design, and prototype testing of the flex-section and hinge seals for the Laminar Flow Supersonic Wind Tunnel Primary Injector. The supersonic atmospheric primary injector operates between Mach 1.8 and Mach 2.2 with mass-flow rates of 62 to 128 lbm/s providing the necessary pressure reduction to operate the tunnel in the desired Reynolds number (Re) range.

A method for the modelling of porous and solid wind tunnel walls in computational fluid dynamics codes

NASA Technical Reports Server (NTRS)

Beutner, Thomas John

1993-01-01

Porous wall wind tunnels have been used for several decades and have proven effective in reducing wall interference effects in both low speed and transonic testing. They allow for testing through Mach 1, reduce blockage effects and reduce shock wave reflections in the test section. Their usefulness in developing computational fluid dynamics (CFD) codes has been limited, however, by the difficulties associated with modelling the effect of a porous wall in CFD codes. Previous approaches to modelling porous wall effects have depended either upon a simplified linear boundary condition, which has proven inadequate, or upon detailed measurements of the normal velocity near the wall, which require extensive wind tunnel time. The current work was initiated in an effort to find a simple, accurate method of modelling a porous wall boundary condition in CFD codes. The development of such a method would allow data from porous wall wind tunnels to be used more readily in validating CFD codes. This would be beneficial when transonic validations are desired, or when large models are used to achieve high Reynolds numbers in testing. A computational and experimental study was undertaken to investigate a new method of modelling solid and porous wall boundary conditions in CFD codes. The method utilized experimental measurements at the walls to develop a flow field solution based on the method of singularities. This flow field solution was then imposed as a pressure boundary condition in a CFD simulation of the internal flow field. The effectiveness of this method in describing the effect of porosity changes on the wall was investigated. Also, the effectiveness of this method when only sparse experimental measurements were available has been investigated. The current work demonstrated this approach for low speed flows and compared the results with experimental data obtained from a heavily instrumented variable porosity test section. The approach developed was simple, computationally inexpensive, and did not require extensive or intrusive measurements of the boundary conditions during the wind tunnel test. It may be applied to both solid and porous wall wind tunnel tests.

An investigation of turbulence structure in a low-Reynolds-number incompressible turbulent boundary

NASA Technical Reports Server (NTRS)

White, B. R.; Strataridakis, C. J.

1987-01-01

An existing high turbulence intensity level (5%) atmospheric boundary-layer wind tunnel has been successfully converted to a relatively low level turbulence (0.3%) wind tunnel through extensive modification, testing, and calibration. A splitter plate was designed, built, and installed into the wind-tunnel facility to create thick, mature, two-dimensional turbulent boundary layer flow at zero pressure gradient. Single and cross hot-wire measurements show turbulent boundary layer characteristics of good quality with unusually large physical size, i.e., viscous sublayer of the order of 1 mm high. These confirm the potential ability of the tunnel to be utilized for future high-quality near-wall turbulent boundary layer measurements. It compares very favorably with many low turbulence research tunnels.

Control of Wind Tunnel Operations Using Neural Net Interpretation of Flow Visualization Records

NASA Technical Reports Server (NTRS)

Buggele, Alvin E.; Decker, Arthur J.

1994-01-01

Neural net control of operations in a small subsonic/transonic/supersonic wind tunnel at Lewis Research Center is discussed. The tunnel and the layout for neural net control or control by other parallel processing techniques are described. The tunnel is an affordable, multiuser platform for testing instrumentation and components, as well as parallel processing and control strategies. Neural nets have already been tested on archival schlieren and holographic visualizations from this tunnel as well as recent supersonic and transonic shadowgraph. This paper discusses the performance of neural nets for interpreting shadowgraph images in connection with a recent exercise for tuning the tunnel in a subsonic/transonic cascade mode of operation. That mode was operated for performing wake surveys in connection with NASA's Advanced Subsonic Technology (AST) noise reduction program. The shadowgraph was presented to the neural nets as 60 by 60 pixel arrays. The outputs were tunnel parameters such as valve settings or tunnel state identifiers for selected tunnel operating points, conditions, or states. The neural nets were very sensitive, perhaps too sensitive, to shadowgraph pattern detail. However, the nets exhibited good immunity to variations in brightness, to noise, and to changes in contrast. The nets are fast enough so that ten or more can be combined per control operation to interpret flow visualization data, point sensor data, and model calculations. The pattern sensitivity of the nets will be utilized and tested to control wind tunnel operations at Mach 2.0 based on shock wave patterns.

Wind Tunnel Investigation of Ground Wind Loads for Ares Launch Vehicle

NASA Technical Reports Server (NTRS)

Keller, Donald F.; Ivanco, Thomas G.

2010-01-01

A three year program was conducted at the NASA Langley Research Center (LaRC) Aeroelasticity Branch (AB) and Transonic Dynamics Tunnel (TDT) with the primary objective to acquire scaled steady and dynamic ground-wind loads (GWL) wind-tunnel data for rollout, on-pad stay, and on-pad launch configurations for the Ares I-X Flight Test Vehicle (FTV). The experimental effort was conducted to obtain an understanding of the coupling of aerodynamic and structural characteristics that can result in large sustained wind-induced oscillations (WIO) on such a tall and slender launch vehicle and to generate a unique database for development and evaluation of analytical methods for predicting steady and dynamic GWL, especially those caused by vortex shedding, and resulting in significant WIO. This paper summarizes the wind-tunnel test program that employed two dynamically-aeroelastically scaled GWL models based on the Ares I-X Flight Test Vehicle. The first model tested, the GWL Checkout Model (CM), was a relatively simple model with a secondary objective of restoration and development of processes and methods for design, fabrication, testing, and data analysis of a representative ground wind loads model. In addition, parametric variations in surface roughness, Reynolds number, and protuberances (on/off) were investigated to determine effects on GWL characteristics. The second windtunnel model, the Ares I-X GWL Model, was significantly more complex and representative of the Ares I-X FTV and included the addition of simplified rigid geometrically-scaled models of the Kennedy Space Center (KSC) Mobile Launch Platform (MLP) and Launch Complex 39B primary structures. Steady and dynamic base bending moment as well as model response and steady and unsteady pressure data was acquired during the testing of both models. During wind-tunnel testing of each model, flow conditions (speed and azimuth) where significant WIO occurred, were identified and thoroughly investigated. Scaled data from the Ares I-X GWL model test was used in the determination of worst-case loads for the analysis of Ares I-X FTV design wind conditions. Finally, this paper includes a brief discussion of the limited full-scale GWL data acquired during the rollout and on-pad stay of the Ares I-X FTV that was launched from KSC on October 28, 2009.

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

NASA Technical Reports Server (NTRS)

Wolf, Stephen W. D.

1988-01-01

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

Wind-Tunnel and Flight Test Results for the Measurements of Flow Variables at Supersonic Speeds Using Improved Wedge and Conical Probes

NASA Technical Reports Server (NTRS)

Bobbitt, Percy J.; Maglieri, Domenic J.; Banks, Daniel W.; Frederick, Michael A.; Fuchs, Aaron W.

2012-01-01

The results of supersonic wind-tunnel tests on three probes at nominal Mach numbers of 1.6, 1.8 and 2.0 and flight tests on two of these probes up to a Mach number of 1.9 are described. One probe is an 8 deg. half-angle wedge with two total-pressure measurements and one static. The second, a conical probe, is a cylinder that has a 15 deg., semi-angle cone tip with one total-pressure orifice at the apex and four static-pressure orifices on the surface of the cone, 90 deg. apart, and about two-thirds of the distance from the cone apex to the base of the cone. The third is a 2 deg. semi-angle cone that has two static ports located 180 deg. apart about 1.5 inches behind the apex of the cone. The latter probe was included since it has been the "probe of choice" for wind-tunnel flow-field pressure measurements (or one similar to it) for the past half-century. The wedge and 15 deg. conical probes used in these tests were designed for flight diagnostic measurements for flight Mach numbers down to 1.35 and 1.15 respectively, and have improved capabilities over earlier probes of similar shape. The 15. conical probe also has a temperature sensor that is located inside the cylindrical part of the probe that is exposed to free-stream flow through an annulus at the apex of the cone. It enables the determination of free-stream temperature, density, speed of sound, and velocity, in addition to free-stream pressure, Mach number, angle of attack and angle of sideslip. With the time-varying velocity, acceleration can be calculated. Wind-tunnel tests of the two probes were made in NASA Langley Research Center fs Unitary Plan Wind Tunnel (UPWT) at Mach numbers of 1.6, 1.8, and 2.0. Flight tests were carried out at the NASA Dryden Flight Research Center (DFRC) on its F-15B aircraft up to Mach numbers of 1.9. The probes were attached to a fixture, referred to as the Centerline Instrumented Pylon (CLIP), under the fuselage of the aircraft. Problems controlling the velocity of the flow through the conical probe required for accurate temperature measurements are noted, as well as some calibration problems of the miniature pressure sensors that required a re-calculation of the flow variables. Data are presented for angle of attack, pressure and Mach number obtained in the wind tunnel and in flight. In the wind tunnel some transient data were obtained by translating the probes through the shock flow field created by a bump on the wind-tunnel wall.

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

NASA Technical Reports Server (NTRS)

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

1988-01-01

Archived wind tunnel data are available for flyback booster or other alternative recoverable configurations as well as reusable orbiters studied during initial development (Phase B) of the Space Shuttle. Considerable wind tunnel data was acquired by the competing contractors and the NASA Centers for an extensive variety of configurations with an array of wing and body planforms. All contractor and NASA wind tunnel data acquired in the Phase B development have been compiled into a data base and are available for application to current winged flyback or recoverable booster aerodynamic studies. The Space Shuttle Phase B Wind Tunnel Database is structured by vehicle component and configuration type. Basic components include booster, orbiter and launch vehicle. Booster configuration types include straight and delta wings, canard, cylindrical, retroglide and twin body. Orbital configuration types include straight and delta wings, lifting body, drop tanks and double delta wings. This is Volume 3 (Part 2) of the report -- Launch Configuration -- which includes booster and orbiter components in various stacked and tandem combinations.

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

NASA Technical Reports Server (NTRS)

Becks, Edward A.

1998-01-01

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

Design and Characterization of the UTIAS Anechoic Wind Tunnel

NASA Astrophysics Data System (ADS)

Chow, Derrick H. F.

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

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

NASA Technical Reports Server (NTRS)

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

1993-01-01

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

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

NASA Technical Reports Server (NTRS)

Kingsland, R. B.

1976-01-01

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

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.

Wind-tunnel tests on a 3-dimensional fixed-geometry scramjet inlet at M = 2.30 to 4.60

NASA Technical Reports Server (NTRS)

Mueller, J. N.; Trexler, C. A.; Souders, S. W.

1977-01-01

Wind-tunnel tests were conducted on a baseline scramjet inlet model having fixed geometry and swept leading edges at M = 2.30, 2.96, 3.95, and 4.60 in the Langley unitary plan wind tunnel. The unit Reynolds number of the tests was held constant at 6.56 million per meter (2 million per foot). The objectives of the tests were to establish inlet performance and starting characteristics in the lower Mach number range of operation (less than M = 5). Surface pressures obtained on the inlet components are presented, along with the results of the internal flow surveys made at the throat and capture stations of the inlet. Contour plots of the inlet-flow-field parameters such as Mach numbers, pressure recovery, flow capture, local static and total pressure ratios at the survey stations are shown for the test Mach numbers.

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

NASA Technical Reports Server (NTRS)

Stroub, Robert H.; Young, Larry A.

1985-01-01

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

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

NASA Technical Reports Server (NTRS)

Oliver, W. R.

1980-01-01

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