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At least 775 records · Page 43
All sourcesNASA NTRSDOE OSTI
DOE OSTI2022

UAE6 - Wind Tunnel Tests Data - UAE6 - Sequence M - Raw Data

Sequence M: Transition Fixed (P) Test sequence M used an upwind, rigid turbine with a 0° cone angle. The wind speed ranged from 5 m/s to 15 m/s. Yaw angles ranged from 0° to 90°. The blade tip pitch was 3°. The rotor rotated at 72 RPM. Blade pressure measurements were collected. The five-hole probes were removed and the plugs were installed. Plastic tape 0.03-mm-thick was used to smooth the interface between the plugs and the blade. The teeter dampers were replaced with rigid links, and these two channels were flagged as not applicable by setting the measured values in the data file to –99999.99 Nm. The teeter link load cell was pre-tensioned to 40,000 N. During post-processing, the probe channels were set to read –99999.99. Zigzag tape was installed near the leading edge of the instrumented blade on both the upper and lower surfaces as described in Appendix K. In addition to the standard 30-second campaigns, yaw sweeps were done at 7 m/s and 10 m/s. These 90-second campaigns were collected while the yaw drive rotated the turbine 90° at a rate of 1°/s. The file names for these campaigns use the letter designation, followed by two digits for wind speed, followed by YSU, followed by 00.

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DOE OSTI2022

UAE6 - Wind Tunnel Tests Data - UAE6 - Sequence 7 - Raw Data

Sequence 7: Shroud Operating (P) This test sequence used a downwind, rigid turbine with a 3.4° cone angle. The wind speed ranged from 5 m/s to 25 m/s. Turntable angles from 0° to 30° were achieved, but the yaw error angle was maintained at 0° by adjusting the nacelle yaw. The blade tip pitch was 3°. The rotor rotated at 72 RPM. Blade and probe pressure measurements were collected. The teeter dampers were replaced with rigid links, and these two channels were flagged as not applicable by setting the measured values in the data file to –99999.99 Nm. The teeter link load cell was pre-tensioned to 40,000 N. An airfoil-shaped aluminum shroud was placed on the tower below the 34% span five-hole probe. It extended below the 91% span five-hole probe. The turntable angle was adjusted to position the shroud at various angles relative to the flow, but the turbine yaw error angle was maintained at 0°. Two additional points were collected where the turntable angle was 0°, and the turbine yaw angle was 20° (71000200) or -20° (710M0200).

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DOE OSTI2022

UAE6 - Wind Tunnel Tests Data - UAE6 - Sequence V - Raw Data

Sequence V: Tip Plate (F) This test sequence used an upwind, rigid turbine with a 0° cone angle. The wind speed ranged from 5 m/s to 25 m/s, and the yaw angle was held at 0°. The blade pitch angle was 3°. The rotor rotated at 72 RPM. Blade pressure measurements were collected. The five-hole probes were removed and the plugs were installed. Plastic tape 0.03-mm-thick was used to smooth the interface between the plugs and the blade. The teeter dampers were replaced with rigid links, and these two channels were flagged as not applicable by setting the measured values in the data file to –99999.99 Nm. The teeter link load cell was pre-tensioned to 40,000 N. During post-processing, the probe channels were set to read -99999.99. The standard tip blocks were replaced with a tip plate to simulate the effect of an undeployed tip-mounted aerodynamic brake as shown in Appendix A. Note that the blade radius was not changed during post-processing so the pressure tap locations are at the same radial location, but the reference to 30% represents 30% of 5.029 m, not 4.943 m. Throughout this report, references to the blade span are made for the 5.029- m radius, not the 4.943-m radius.

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DOE OSTI2022

UAE6 - Wind Tunnel Tests Data - UAE6 - Sequence W - Raw Data

Sequence W: Extended Blade (F) This test sequence used an upwind, rigid turbine with a 0° cone angle. The wind speed ranged from 5 m/s to 21 m/s, and the yaw angle was held at 0°. The blade pitch angle was 3°. The rotor rotated at 72 RPM. Blade pressure measurements were collected. The five-hole probes were removed and the plugs were installed. Plastic tape 0.03-mm thick was used to smooth the interface between the plugs and the blade. The teeter dampers were replaced with rigid links, and these two channels were flagged as not applicable by setting the measured values in the data file to –99999.99 Nm. The teeter link load cell was pretensioned to 40,000 N. During postprocessing, the probe channels were set to read –99999.99. The standard tip blocks were replaced with blade extensions that created a 5.532-m radius rotor as shown in Appendix A. The extension used the S809 airfoil throughout, and the linear taper of the blade continued along the extension. Note that the blade radius was not changed during post-processing so the pressure tap locations are at the same radial location, but the reference to 30% represents 30% of 5.029 m, not 5.532 m. Throughout this report, references to the blade span are made for the 5.029-m radius, not the 5.532-m radius.

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NASA NTRS2015

Reducing the Effect of Transducer Mount Induced Noise (XMIN) on Aeroacoustic Wind Tunnel Testing Data with a New Transducer Mount Design

Characterization of flight vehicle unsteady aerodynamics is often studied via large scale wind tunnel testing. Boundary layer noise is measured by miniature pressure transducers installed in a model. Noise levels (2-5 dB ref. 20 μPa) can be induced when transducer is mounted out of flush with model outer surface. This effect must be minimized to accurately determine aerodynamically induced acoustic environments.

Herron, Andrew J.↗
DOE OSTI2022

UAE6 - Wind Tunnel Tests Data - UAE6 - Sequence P - Raw Data

Sequence P: Wake Flow Visualization, Upwind (P) This test sequence used an upwind, rigid turbine with a 0° cone angle. The wind speed ranged from 5 m/s to 15 m/s. Yaw angles of 0° to –60° were achieved. The blade tip pitch was 3°. The rotor rotated at 72 RPM. Blade and probe pressure measurements were collected. The teeter dampers were replaced with rigid links, and these two channels were flagged as not applicable by setting the measured values in the data file to –99999.99 Nm. The teeter link load cell was pre-tensioned to 40,000 N. The aluminum blade tip designed to contain a smoke generator was installed, and counterweights were installed in the non-instrumented blade tip to compensate. The turbine was positioned at the appropriate yaw angle, and the smoke generator was ignited remotely. The campaign duration was 3 minutes for all tests except P1000000, which was 2 minutes. The file name convention was the standard format except for P10000A0, which indicated a 3° pitch angle. File P1000000 used a 12° pitch angle. After these two campaigns were collected, it was determined that all subsequent data should be collected with a 3° pitch angle. Pressure data were not acquired during this sequence, so all associated data values were flagged as not applicable by setting the measured values in the data file to 0.000 Pa. Corresponding pressure data are available from Sequence H for the 3° pitch angle test points. Flow visualization data obtained from wall- and ceiling-mounted video cameras were recorded to videotape. The camera locations and calibration procedures are described in Appendix J.

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NASA NTRS1999

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

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

Allen, Jerry↗
DOE OSTI2022

UAE6 - Wind Tunnel Tests Data - UAE6 - Sequence 3 - Raw Data

Sequence 3: Tower Wake Measure (P) Sequence 3 used a downwind, rigid turbine with a 3.4° cone angle. The wind speeds for this sequence corresponded to subcritical [7 m/s], transitional [15 m/s], and supercritical [20 m/s] Reynolds number regimes for the circular cross-section tower. The yaw angle and blade tip pitch angles were selected to traverse the probe tip across the tower wake, while keeping the probe axis approximately parallel to the tunnel centerline. The rotor was parked with the instrumented blade at 180° azimuth. Blade and probe pressure measurements were collected.

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DOE OSTI2022

UAE6 - Wind Tunnel Tests Data - UAE6 - Sequence G - Raw Data

Sequence G: Upwind Teetered (F) Test sequence G used an upwind, teetered turbine with a 0° cone angle. The wind speeds ranged from 5 m/s to 25 m/s. Yaw angles of ±180° were achieved at low wind speeds and angles of ±10° were achieved at the high wind speeds. The blade tip pitch was 3°. The rotor rotated at 72 RPM. Blade and probe pressure measurements were collected.

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NASA NTRS1984

Low altitude wind shear statistics derived from measured and FAA proposed standard wind profiles

Wind shear statistics were calculated for a simulated data set using wind profiles proposed as a standard and compared to statistics derived from measured wind profile data. Wind shear values were grouped in altitude bands of 100 ft between 100 and 1400 ft, and in wind shear increments of 0.025 kt/ft between + or - 0.600 kt/ft for the simulated data set and between + or - 0.200 kt/ft for the measured set. No values existed outside the + or - 0.200 kt/ft boundaries for the measured data. Frequency distributions, means, and standard deviations were derived for each altitude band for both data sets, and compared. Also, frequency distributions were derived for the total sample for both data sets and compared. Frequency of occurrence of a given wind shear was about the same for both data sets for wind shears, but less than + or 0.10 kt/ft, but the simulated data set had larger values outside these boundaries. Neglecting the vertical wind component did not significantly affect the statistics for these data sets. The frequency of occurrence of wind shears for the flight measured data was essentially the same for each altitude band and the total sample, but the simulated data distributions were different for each altitude band. The larger wind shears for the flight measured data were found to have short durations.

Dunham, R. E., Jr.↗
NASA NTRS2016

Boeing Smart Rotor Full-scale Wind Tunnel Test Data Report

A full-scale helicopter smart material actuated rotor technology (SMART) rotor test was conducted in the USAF National Full-Scale Aerodynamics Complex 40- by 80-Foot Wind Tunnel at NASA Ames. The SMART rotor system is a five-bladed MD 902 bearingless rotor with active trailing-edge flaps. The flaps are actuated using piezoelectric actuators. Rotor performance, structural loads, and acoustic data were obtained over a wide range of rotor shaft angles of attack, thrust, and airspeeds. The primary test objective was to acquire unique validation data for the high-performance computing analyses developed under the Defense Advanced Research Project Agency (DARPA) Helicopter Quieting Program (HQP). Other research objectives included quantifying the ability of the on-blade flaps to achieve vibration reduction, rotor smoothing, and performance improvements. This data set of rotor performance and structural loads can be used for analytical and experimental comparison studies with other full-scale rotor systems and for analytical validation of computer simulation models. The purpose of this final data report is to document a comprehensive, highquality data set that includes only data points where the flap was actively controlled and each of the five flaps behaved in a similar manner.

Full-Scale Wind Tunnel↗
DOE OSTI2022

UAE6 - Wind Tunnel Tests Data - UAE6 - Sequence 4 - Raw Data

Sequence 4: Static Pressure Calibration (P) This test sequence used a downwind, teetered turbine with a 3.4° cone angle. The wind speed ranged from 5 m/s to 25 m/s. Yaw angles of 0° to 30° were achieved. The blade tip pitch was 3°. The rotor rotated at 72 RPM. Blade and probe pressure measurements were collected. A static pitot probe was installed on a boom 2.3 m upwind of the nacelle at hub height and connected to the reference side of the Mensor digital differential pressure transducer through the pneumatic rotary coupling. This provided the reference pressure for all blade-mounted pressure transducers, as opposed to the NASA atmospheric pressure reference that was used for all other test sequences. This simulates the pressure system connections used in field measurements where the static pressure was unknown.

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DOE OSTI2022

UAE6 - Wind Tunnel Tests Data - UAE6 - Sequence 6 - Raw Data

Sequence 6: Shroud Wake Measure (P) This test sequence used a downwind, rigid turbine with a 3.4° cone angle. The wind speeds for this sequence corresponded to subcritical [7 m/s], transitional [15 m/s], and supercritical [20 m/s] Reynolds number regimes for the circular cross-section tower. The yaw angle and blade tip pitch angles were selected to traverse the probe tip across the tower wake while keeping the probe axis approximately parallel to the tunnel centerline. The rotor was parked with the instrumented blade at 180° azimuth. Blade and probe pressure measurements were collected. An airfoil-shaped aluminum shroud was placed on the tower below the 34% span five-hole probe. It extended below the 91% span five-hole probe. The turntable angle was adjusted to position the shroud at various angles relative to the flow.

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DOE OSTI2022

UAE6 - Wind Tunnel Tests Data - UAE6 - Sequence N - Raw Data

Sequence N: Sin AOA, Rotating (P) This sequence was designed to quantify the blade 3-D unsteady aerodynamic response in the presence of rotational influences by varying blade pitch angle. Test sequence N used an upwind, rigid turbine with a 0° cone angle. The wind speed was 15 m/s, and the yaw angle was 0° throughout the sequence. The rotor rotated at 72 RPM. Blade and probe pressure measurements were collected. The teeter dampers were replaced with rigid links, and these two channels were flagged as not applicable by setting the measured values in the data file to –99999.99 Nm. The teeter link load cell was pre-tensioned to 40,000 N. Sinusoidal angle-of-attack histories were designed to emulate 2-D S809 dynamic data obtained at the Ohio State University (OSU) wind tunnel (Reuss Ramsey et al. 1995) and the 1P yawed flow angle-of-attack variations the turbine could be expected to encounter in routine operation. The reduced frequency (K), mean angle of attack (αm), and oscillation amplitude (αω) were computed for each of the five primary span locations. The file lengths varied according to the time required to complete 40 oscillations at the specified frequency. The file name convention used the sequence designation N, followed by two digits for the span location (30, 47, 63, 80, or 95), followed by two digits representing the condition specified in the test matrix, followed by the repetition digit. Due to time constraints, most of the test conditions corresponding to the 63% span location were eliminated. Test points with a static pitch angle of 3° were collected throughout the sequence to ascertain instrumentation fidelity.

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NASA NTRS2004

Superposed Epoch Analysis of Ring Current Geoeffectiveness Related to Solar Wind and Plasma Sheet Drivers

The goal of that proposal was to examine the relationship between solar wind drivers and ring current dynamics through data analysis and numerical simulations. The data analysis study was a statistical examination (via superposed epoch analyses) of a solar cycle's worth of storm data. Solar wind data, geophysical indices, and geosynchronous plasma data were collected for every time period with Dst< -50 nT from 1989 through 2002, and the storm list now exceeds 400 entries. This work was first conducted by a summer undergraduate student, Mr. John Vann (University of Kansas), with funding from the NSF Research Experience for Undergraduates program. It was then continued by a University of Michigan graduate student, Mr. Jichun Zhang. Mr. Zhang is now in his fourth year at U-M and is progressing very well toward a PhD in space science. His dissertation will be based on his data analysis and modeling efforts using this geomagnetic storm database. The results of the data analysis study have been the focus of several conference presentations, and the first manuscript has just been published. Two additional papers are presently being prepared, one on average (superposed) solar wind features for various storm subsets (e.g., intense storms at solar maximum), and another on geosynchronous plasma features for these same storm subsets. The latter result was highlighted by the TR&T program director in his presentation at the COSPAR meeting this summer.

Liemohm, M. W.↗
NASA NTRSDate not supplied

Analysis and Comparison of Surface Roughness Effects on Pressure Data from SLS Wind Tunnel Test

Unsteady flows are becoming a more characterized field of study as advancements in technology is allowing for more high-speed, time-resolved data acquisition. It is important to understand these flows as the shock-wave boundary-layer interactions and separation they are involved with can generate extreme loads caused by high pressure and temperature on the body of a flight vehicle. It is vital to characterize these loads, as they can impede upon the structural integrity of a high-speed system. A common instrument utilized today to characterize pressure distributions caused by unsteady flows are unsteady pressure transducers. Set flush with the surface of a flight vehicle or wind tunnel model, these transducers can provide a pressure distribution, capturing the frequency of unsteady loads that could impinge upon the test article. A couple negative aspects of this instrumentation is that they are expensive, have long lead times, and a pressure distribution can only be determined in one discrete location where the transducer is installed. An optical diagnostic known as unsteady pressure-sensitive paint (uPSP) can help to satisfy these deficiencies. This is a spray paint that can be applied to a surface of a model to provide a global pressure distribution across the whole painted region.2 A complete picture of the fluid dynamic pressures occurring on the surface of a test article can be analyzed, especially in locations where it may have been impossible to install a transducer. Pressure-sensitive paint is also less expensive than traditional pressure transducers, which are, on average, about $2,000 each including the cost of installation.3 Unsteady PSP will not replace unsteady pressure transducers, but the transducers can be used to validate the PSP response to the flow and measure frequencies higher than the frequency response of the paint. One issue that can arise using these tools simultaneously, is that the surface roughness of the uPSP can affect the pressure readings of the transducers. Applying the uPSP to the surface of the model is going to affect the flow in some way as the topography of the model is now different than if it was not painted. One study with steady PSP conducted by Amer, Obara, and Liu at NASA Langley concluded that the PSP surface roughness was minimally intrusive at different test conditions with various lift, drag, etc., however, it was noted that data from some pressure transducers had decreased, likely due to the uneven application of the PSP in the localized region of the transducers.9 Sugioka et. al. tested various formulations of uPSP to study surface roughness effects on a Common Research Model (CRM) in transonic flow. The influence of uPSP on the surface of the model, depending on how it was formulated, was noted to have the possibility of relocating the shock wave present in the flow. This study was conducted to examine and compare transducer data acquired in December 2017 of a 4% scale model of the Space Launch System (SLS) performed at NASA Ames Research Center (ARC) Unitary Plan Wind Tunnel (UPWT). One goal of the test was to demonstrate to the customer the potential of the uPSP system developed at NASA ARC. The effect of uPSP on transducer signals also became an interest in this experiment and is the reason for this data analysis. The results of this study were discussed with the SLS team to help them make decisions regarding the next test on an SLS model with the PSP and Unitary Plan Wind Tunnel teams.

pressure-sensitive paint↗
NASA NTRS2015

Reducing the Effect of Transducer Mount Induced Noise on Aeroacoustic Wind Tunnel Testing Data with a New Transducer Mount Design

Flight vehicle aeroacoustic environments induced during transonic and supersonic flight are usually predicted by subscale wind tunnel testing utilizing high frequency miniature pressure transducers. In order to minimize noise induced by the measurement itself, transducer flush mounting with the model surface is very important. The National Aeronautics and Space Administration (NASA) has accomplished flushness in recent testing campaigns via use of a transducer holder that can be machined and sanded. A single hole in the holder allows the flow medium to interact with the transducer diaphragm. Noise is induced by the resulting cavity however, and is a challenge to remove in post-processing. A new holder design has been developed that minimizes the effects of this transducer mount induced noise (XMIN) by reducing the resonance amplitude or increasing its resonance frequency beyond the range of interest. This paper describes a test conducted at the NASA/George C. Marshall Space Flight Center Trisonic Wind Tunnel intended to verify the effectiveness of this design. The results from this test show that this new transducer holder design does significantly reduce the influence of XMIN on measured fluctuating pressure levels without degrading a transducer's ability to accurately measure the noise external to the model.

Herron, Andrew J.↗
DOE OSTI2022

UAE6 - Wind Tunnel Tests Data - UAE6 - Sequence 8 - Raw Data

Sequences 8 and 9: Downwind Sonics (F,P) and Downwind Sonics Parked (P) This test sequence used an upwind, rigid turbine with a 0° cone angle. The wind speed ranged from 5 m/s to 25 m/s. Yaw angles of 0° to 60° were achieved. The blade tip pitch was 3°. The rotor rotated at 72 RPM during Sequence 8, but it was parked during Sequence 9. Blade pressure measurements were collected. The five-hole probes were removed and the plugs were installed. Plastic tape 0.03-mm-thick was used to smooth the interface between the plugs and the blade. The teeter dampers were replaced with rigid links, and these two channels were flagged as not applicable by setting the measured values in the data file to –99999.99 Nm. The teeter link load cell was pre-tensioned to 40,000 N. During post-processing, the probe channels were set to read -99999.99. Sonic anemometers were mounted on a strut downwind of the turbine. The strut was mounted to the T-frame, which was rotated to align the anemometers aft of the 9% and 49% radius locations at hub height. Because of this configuration, the tunnel balance data are considered invalid. Sequence 9 was designed to compare the downwind sonic anemometer readings with the upwind sonic anemometers without interference from the turbine. The rotor was parked with the instrumented blade at 0° azimuth. All pressure measurements obtained in Sequence 9 are invalid because sufficient time for temperature stabilization did not occur, thus all associated data values were flagged as not applicable by setting the measured values in the data file to 0.0000 Pa. This test is further described in Appendix G.

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