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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 271 records · Page 15
All sourcesNASA NTRSDOE OSTI
NASA NTRS2003

Stage Separation Wind Tunnel Tests of a Generic Two-Stage-to-Orbit Launch Vehicle

In support of NASA s Space Launch Initiative Program, stage separation wind tunnel tests of a generic two-stage-to-orbit (TSTO) launch vehicle were conducted to determine the interference aerodynamic forces and moments and to determine the proximity flow environment. The tests were conducted in the NASA Marshall Space Flight Center s Aerodynamic Research Facility using a manual separation fixture for a Mach number range of 2.74 to 4.96 and separation distances up to 80 percent and 35 percent of the body length in the vehicle X and Z coordinates, respectively. For the TSTO bimese, winged-body vehicle configuration, both wing-to-wing and wing-to-fuselage configurations were tested. Individual-body force and moment, schlieren, and surface pressure data were acquired. The results showed that the proximity aerodynamics were dominated by complex bow shock interactions, and that he booster was statically unstable at several separation positions. As compared to the isolated body, the proximity normal force change with pitch angle was found to be nearly the same, and the proximity axial force increased, in general, by 3% for both bodies.

Bordelon, Wayne J., Jr.↗
NASA NTRS1992

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

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

Wong, Douglas T.↗
NASA NTRS1995

SOFIA 2 model telescope wind tunnel test report

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.

Keas, Paul↗
NASA NTRS1973

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

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

Oldenbuttel, R. H.↗
NASA NTRSDate not supplied

X-59 Air Data Probe Calibration Wind Tunnel Test

NASA's aeronautical innovators are working with Lockheed Martin to design and build a supersonic X-plane, the X-59, under the the Low Boom Flight Demonstrator (LBFD) Project, that produces a gentle thump rather than a loud sonic boom. The plane will be used to collect human response data over select U.S. communities. The data from the X-59 test flights will be provided to U.S. and international regulators to potentially allow supersonic flight over land, drastically reducing travel time within the U.S. and around the world. Two wind tunnel tests were conducted in the NASA Glenn Research Center (GRC) 8- by 6-Foot Supersonic Wind Tunnel to calibrate the nose probe for the X-59. The probe was successfully tested and calibrated at 19 Mach numbers between Mach 0.25 and Mach 1.7. Most data were collected using continuous roll angle sweeps at a set Mach number and pitch angle. About 725 high quality data runs were recorded during testing to be used in the calibration of the probe. A backup nose probe was also run through the same test matrix as the primary probe. The probe will be the primary instrument for the flight test of the X-59 for determining angle of attack, angle of sideslip, airspeed, and pressure altitude of the aircraft in flight. Knowing the aircraft's speed and attitude is critical for flight safety and is also critical data for the Sonic Boom mission. Areas of interest for the test were sea level takeoff and landing at Mach 0.2, subsonic cruise at Mach 0.9, and supersonic cruise at Mach 1.4. During Phase I flight testing, the air data nose probe and air data system as a whole will be further calibrated in flight to account for additional installed effects of the airframe on the air data system.

X-59↗
NASA NTRSDate not supplied

X-59 Air Data Probe Calibration Wind Tunnel Test

NASA's aeronautical innovators are working with Lockheed Martin to design and build a supersonic X-plane, the X-59, under the the Low Boom Flight Demonstrator (LBFD) Project, that produces a gentle thump rather than a loud sonic boom. The plane will be used to collect human response data over select U.S. communities. The data from the X-59 test flights will be provided to U.S. and international regulators to potentially allow supersonic flight over land, drastically reducing travel time within the U.S. and around the world. Two wind tunnel tests were conducted in the NASA Glenn Research Center (GRC) 8- by 6-Foot Supersonic Wind Tunnel to calibrate the nose probe for the X-59. The probe was successfully tested and calibrated at 19 Mach numbers between Mach 0.25 and Mach 1.7. Most data were collected using continuous roll angle sweeps at a set Mach number and pitch angle. About 725 high quality data runs were recorded during testing to be used in the calibration of the probe. A backup nose probe was also run through the same test matrix as the primary probe. The probe will be the primary instrument for the flight test of the X-59 for determining angle of attack, angle of sideslip, airspeed, and pressure altitude of the aircraft in flight. Knowing the aircraft's speed and attitude is critical for flight safety and is also critical data for the Sonic Boom mission. Areas of interest for the test were sea level takeoff and landing at Mach 0.2, subsonic cruise at Mach 0.9, and supersonic cruise at Mach 1.4. During Phase I flight testing, the air data nose probe and air data system as a whole will be further calibrated in flight to account for additional installed effects of the airframe on the air data system.

X-59↗
NASA NTRS1977

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

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

Wentz, W. H., Jr.↗
NASA NTRS1945

High-Speed Wind-Tunnel Tests of a 0.3-Scale Model of the P-47D Airplane

Presented in this report are the results of high-speed wind-tunnel tests of a 0.3-scale model of the P-47D airplane. Included are the aerodynamic characteristics of the model, pressure distributions on the wing and horizontal tail, elevator hinge moments, and tuft photographs showing the flow over the wing and horizontal tail. The results indicate that, at level-flight lift coefficients, the P-47D airplane will not exhibit a diving tendency below a Mach number of 0.78 at an altitude of 20,000 feet or below a Mach number of 0.72 at an altitude of 40,000 feet. The Mach number at which the lift, drag, and pitching-moment coefficients diverged rapidly from the low speed values is about C.15 higher than the critical Mach number. This is an unusually large margin.

William T Hamilton↗
NASA NTRS2017

Computational Support of 9x7 Wind Tunnel Test of Sonic Boom Models with Plumes

NASA and its industry partners are performing studies of supersonic aircraft concepts with low sonic boom pressure signatures. The interaction of the nozzle jet flow with the aircrafts' aft components is typically where the greatest uncertainly in the pressure signature is observed with high-fidelity numerical simulations. An extensive wind tunnel test was conducted in February 2016 in the NASA Ames 9- by 7- Foot Supersonic Wind Tunnel to help address the nozzle jet effects on sonic boom. Five test models with a variety of shock generators of differing waveforms and strengths were tested with a convergent-divergent nozzle for a wide range of nozzle pressure ratios. The LAVA unstructured flow solver was used to generate first CFD comparisons with the new experimental database using best practice meshing and analysis techniques for sonic boom vehicle design for all five different configurations. LAVA was also used to redesign the internal flow path of the nozzle and to better understand the flow field in the test section, both of which significantly improved the quality of the test data.

9x7 Wind Tunnel↗
NASA NTRS2000

Contributions of Transonic Dynamics Tunnel Testing to Airplane Flutter Clearance

The Transonic Dynamics Tunnel (TDT) became in operational in 1960, and since that time has achieved the status of the world's premier wind tunnel for testing large in aeroelastically scaled models at transonic speeds. The facility has many features that contribute to its uniqueness for aeroelastic testing. This paper will briefly describe these capabilities and features, and their relevance to aeroelastic testing. Contributions to specific airplane configurations and highlights from the flutter tests performed in the TDT aimed at investigating the aeroelastic characteristics of these configurations are presented.

Rivera, Jose A.↗
NASA NTRS2013

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

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.

Schauerhamer, Daniel G.↗
NASA NTRSDate not supplied

Wind Tunnel Testing of AFC over a Deflected Aileron on the High-Lift Common Research Model

Active flow control (AFC) using discrete fluidic actuators distributed along the span, just upstream of the deflected aileron of the 10% scale high-lift version of the Common Research Model (CRM-HL), was evaluated during a wind tunnel test in the NASA Langley 14- by 22-Foot Subsonic Tunnel. For this set of experiments, a new outboard section was fabricated incorporating a deflectable aileron. Aileron deflection angles of 0, 7.5, 16, and 25 were investigated. This experimental investigation is in response to a recent study by Boeing potential to use AFC with a drooped aileron during takeoff to improve lift-to-drag ratio (L/D) by as much as 5%, depending on AFC mass flow rate. AFC is used at deflection angles above the nominal deflection of 7.5 to control the resulting flow separation that occurs and potentially improve L/D. Improvements in aircraft low-speed L/D can affect vehicle range and/or payload. Tuft flow visualization data, steady and unsteady surface pressure data, and force and moment data are used to compare the flowfields with and without AFC. Parameters varied include actuator momentum coefficient and aileron deflection angle. The surface pressure and tuft flow visualization results indicate that without AFC, the flow over a majority of the aileron upper surface is separated for aileron deflection angles larger than 7.5. When AFC is applied, the flow is reattached to the aileron. Force and moment results show that a local increase in lift leads to an increase in L/D of at least 3.5% using mass flow rates thought to be available from air sources onboard a commercial transport.

AFC↗
NASA NTRSDate not supplied

Wind Tunnel Testing of AFC over a Deflected Aileron on the High-Lift Common Research Model

Active flow control (AFC) using discrete fluidic actuators distributed along the span, just upstream of the deflected aileron of the 10% scale high-lift version of the Common Research Model (CRM-HL), was evaluated during a wind tunnel test in the NASA Langley 14- by 22-Foot Subsonic Tunnel. For this set of experiments, a new outboard section was fabricated incorporating a deflectable aileron. Aileron deflection angles of 0, 7.5, 16, and 25 were investigated. This experimental investigation is in response to a recent study by Boeing potential to use AFC with a drooped aileron during takeoff to improve lift-to-drag ratio (L/D) by as much as 5%, depending on AFC mass flow rate. AFC is used at deflection angles above the nominal deflection of 7.5 to control the resulting flow separation that occurs and potentially improve L/D. Improvements in aircraft low-speed L/D can affect vehicle range and/or payload. Tuft flow visualization data, steady and unsteady surface pressure data, and force and moment data are used to compare the flowfields with and without AFC. Parameters varied include actuator momentum coefficient and aileron deflection angle. The surface pressure and tuft flow visualization results indicate that without AFC, the flow over a majority of the aileron upper surface is separated for aileron deflection angles larger than 7.5. When AFC is applied, the flow is reattached to the aileron. Force and moment results show that a local increase in lift leads to an increase in L/D of at least 3.5% using mass flow rates thought to be available from air sources onboard a commercial transport.

AFC↗