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Results for “sweep testing”

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 NTRS1988

Measured and predicted pressure distributions on the AFTI/F-111 mission adaptive wing

Flight tests have been conducted using an F-111 aircraft modified with a mission adaptive wing (MAW). The MAW has variable-camber leading and trailing edge surfaces that can change the wing camber in flight, while preserving smooth upper surface contours. This paper contains wing surface pressure measurements obtained during flight tests at Dryden Flight Research Facility of NASA Ames Research Center. Upper and lower surface steady pressure distributions were measured along four streamwise rows of static pressure orifices on the right wing for a leading-edge sweep angle of 26 deg. The airplane, wing, instrumentation, and test conditions are discussed. Steady pressure results are presented for selected wing camber deflections flown at subsonic Mach numbers up to 0.90 and an angle-of-attack range of 5 to 12 deg. The Reynolds number was 26 million, based on the mean aerodynamic chord. The MAW flight data are compared to MAW wind tunnel data, transonic aircraft technology (TACT) flight data, and predicted pressure distributions. The results provide a unique database for a smooth, variable-camber, advanced supercritical wing.

Webb, Lannie D.↗
NASA NTRS1976

Pressure distribution on a 1- by 3-meter semispan wing at sweep angles from 0 deg to 40 deg in subsonic flow

A 1- by 3-meter semispan wing of taper ratio 1.0 with NACA 0012 airfoil section contours was tested in the Langley V/STOL tunnel to measure the pressure distribution at five sweep angles, 0 deg, 10 deg, 20 deg, 30 deg, and 40 deg, through an angle-of-attack range from -6 deg to 20 deg. The pressure data are presented as plots of pressure coefficients at each static-pressure tap location on the wing. Flow visualization wing-tuft photographs are also presented for a wing of 40 deg sweep. A comparison between theory and experiment using two inviscid theories and a viscous theory shows good agreement for pressure distributions, normal forces, and pitching moments for the wing at 0 deg sweep.

Yip, L. P.↗
DOE OSTI2024

Roles of temperature, materials, and domain inversion in high-performance, low-bias-drift thin film lithium niobate blue light modulators

We demonstrate a thin film lithium niobate electro-optic modulator operating at 456 nm with an RF voltage-length product of 0.38 V-cm and a bandwidth of 6.9 GHz. We test the dielectric relaxation of the modulator with sweeps of temperature and optical input power, and compare equivalent modulators with electrode materials of Cr-Au, Ti-Au and Al in terms of bias stability and current-voltage characteristics. We demonstrate bias stability over at least 8 hours with Al devices, and show relationships between drift, I-V characteristics and ferroelectric domain switching.

Celik, Oguz Tolga (ORCID:0000000264542306)↗
NASA NTRS1974

Exhaust-nozzle characterisitcs for a twin-jet variable-wing-sweep fighter airplane model at Mach numbers to 2.2

A wind-tunnel investigation has been conducted to determine the exhaust-nozzle aerodynamic and propulsive characteristics for a twin-jet variable-wing-sweep fighter airplane model. The powered model was tested in the Langley 16-foot transonic tunnel and in the Langley 4-foot supersonic pressure tunnel at Mach numbers to 2.2 and at angles of attack from about minus 2 to 6 deg. Compressed air was used to simulate the nozzle exhaust flow at values of jet total-pressure ratio from approximately 1 (jet off) to about 21. Effects of configuration variables such as speed-brake deflection, store installation, and boundary-layer thickness on the the nozzle characteristics were also investigated.

Reubush, D. E.↗
NASA NTRS1978

Subsonic longitudinal and lateral-directional static aerodynamic characteristics of a general research fighter model employing a strake-wing concept

A general research fighter model was tested in the Langley 7 by 10 foot high speed tunnel at a Mach number of 0.3. Strakes with exposed semi-spans of 10 percent, 20 percent, and 30 percent of the wing reference semi-span were tested in combination with wings having leading edge sweep angles of 30, 44, and 60 degrees. The angle of attack range was from -4 degrees to approximately 48 degrees at sideslip angles of 0, -5, and 5 degrees. The data are presented without analysis in order to expedite publication.

Fox, C. H., Jr.↗
NASA NTRS1981

Self-induced wing rock of slender delta wings

As part of a research program aimed at exploring basic mechanisms that cause wing rock in combat aircraft, an investigation was conducted to study the aerodynamic factors which cause the low-speed wing rock exhibited by slender delta wings. A flat-plate delta wing with 80 deg leading-edge sweep was subjected to conventional static-force tests and dynamic wind-tunnel experiments which included forced-oscillation, rotary, and free-to-roll tests. In addition, visualization of the flow phenomena involved was obtained by observing tuft patterns and using a helium-bubble technique. This paper summarizes the results of this study. Fundamental information is presented on the aerodynamic mechanisms that cause the wing rock and the problem of mathematically modeling the aerodynamics and motions is discussed.

Nguyen, L. T.↗
NASA NTRS1985

A faster 'transition' to laminar flow

A discussion is given of the ongoing research related to laminar flow airfoils, nacelles, and wings where the laminar flow is maintained by a favorable pressure gradient, surface suction or a combination of the two. Design methologies for natural laminar flow airfoil sections and wings for both low and high speed applications are outlined. Tests of a 7-foot chord, 23-deg sweep laminar flow-control airfoil at high subsonic Mach numbers are described, along with the associated stability theory used to design the suction system. The state-of-the-art of stability theory is simply stated and a typical calculation illustrated. In addition, recent computer simulations of transition using the time dependent Navier-Stokes equations are briefly described. Advances in wind tunnel capabilities and instrumentation will be reviewed, followed by the presentation of a few results from both wind tunnels and flight. Finally, some suggestions for future work will complete the paper.

Bobbitt, P. J.↗
NASA NTRS1987

Using frequency-domain methods to identify XV-15 aeroelastic modes

The XV-15 Tilt-Rotor wing has six major aeroelastic modes that are close in frequency. To precisely excite individual modes during flight test, dual flaperon exciters with automatic frequency-sweep controls were installed. The resulting structural data were analyzed in the frequency domain (Fourier transformed) with cross spectral and transfer function methods. Modal frequencies and damping were determined by performing curve fits to transfer function magnitude and phase data and to cross spectral magnitude data. Results are given for the XV-15 with its original metal rotor blades. Frequency and damping values are also compared with earlier predictions.

Acree, C. W., Jr.↗
NASA NTRS1988

Using frequency-domain methods to identify XV-15 aeroelastic modes

The XV-15 Tilt-Rotor wing has six major aeroelastic modes that are close in frequency. To precisely excite individual modes during flight test, dual flaperon exciters with automatic frequency-sweep controls were installed. The resulting structural data were analyzed in the frequency domain (Fourier transformed) with cross spectral and transfer function methods. Modal frequencies and damping were determined by performing curve fits to transfer function magnitude and phase data and to cross spectral magnitude data. Results are given for the XV-15 with its original metal rotor blades. Frequency and damping values are also compared with earlier predictions.

Acree, C. W., Jr.↗
NASA NTRS1989

Identification of XV-15 aeroelastic modes using frequency-domain methods

The XV-15 Tilt-Rotor wing has six major aeroelastic modes that are close in frequency. To precisely excite individual modes during flight test, dual flaperon exciters with automatic frequency-sweep controls were installed. The resulting structural data were analyzed in the frequency domain (Fourier transformed) with cross spectral and transfer function methods. Modal frequencies and damping were determined by performing curve fits to transfer function magnitude and phase data and to cross spectral magnitude data. Results are given for the XV-15 with its original metal rotor blades. Frequency and damping values are also compared with earlier predictions.

Acree, C. W., Jr.↗
NASA NTRS1989

Identification of XV-15 aeroelastic modes using frequency-domain methods

The XV-15 Tilt-Rotor wing has six major aeroelastic modes that are close in frequency. To precisely excite individual modes during flight test, dual flaperon exciters with automatic frequency-sweep controls were installed. The resulting structural data were analyzed in the frequency domain (Fourier transformed) with cross spectral and transfer function methods. Modal frequencies and damping were determined by performing curve fits to transfer function magnitude and phase data and to cross spectral magnitude data. Results are given for the XV-15 with its original metal rotor blades. Frequency and damping values are also compared with earlier predictions.

Acree, Cecil W., Jr.↗
NASA NTRS1993

Behavior of spinning laminated composite plates with initial twist-experimental vibrations, strain, and deflection results

The experimental behavior of spinning, pre-twisted laminated composite plates was investigated. The purpose of these experiments was to establish an experimental database consisting of strain, deflections, and natural frequencies as a function of rotational velocity. Six different plate sets were tested, that included three different stacking sequences (two symmetric, one asymmetric), two different initial twist levels (0 deg, 30 deg), and two different initial twist axis locations (midchord, quarter-chord). The plates were spin tested at four different combinations of pitch and sweep. It was observed that the location of the pretwist axis and the level of pretwist greatly affects the strain and deflections of the spinning plate, while only the pretwist level affects the measured natural frequencies.

Lapid, A. J.↗
NASA NTRS1988

Identification of 15-15 Aeroelastic Modes using Frequency-Domain Methods

The XV-15 Tilt-Rotor wing has six major aeroelastic modes that are close in frequency. To precisely excite individual modes during flight test, dual flaperon exciters with automatic frequency-sweep controls were installed. The resulting structural data were analyzed in the frequency domain (Fourier-transformed). Modal frequencies and damping were determined by performing curve fits to frequency-response magnitude and phase data. Results are given for the XV-15 with its original metal rotor blades. Frequency and damping values are also compared with new predictions by two different programs, CAMRAD and ASAP.

Acree, C. W., Jr.↗
NASA NTRS1992

Aerodynamic Spring and Damping of Free-Pitching Tips on a Semispan Wing

A test was conducted in the NASA Ames 7- by 10-Foot Wind Tunnel to derive aerodynamic spring and damping estimates for free-pitching tips on a semispan wing. The test model was a rectangular planform semispan wing with wing tips that had a single rigid-body pitch degree of freedom with respect to the inboard wing. A number of different tip planform geometries were tested, incorporating a range of quarter-chord sweep angles and taper ratios. The wing-tip dynamic response characteristics were measured at several wing angles of attack and tunnel dynamic pressures. The tip oscillations were initiated by releasing the tips from prescribed angles of attack. A new method to isolate Coulomb damping from aerodynamic damping from these tip-motion time histories is developed and applied. Correlations were performed between the experimentally derived wing-tip aerodynamic spring and damping values and predictions from a semiempirical analysis based on steady-state tip aerodynamic loads.

Young, Larry A.↗
NASA NTRS2001

Vortex Wake Geometry of a Model Tilt Rotor in Forward Flight

A full-span 0.25-scale V-22 tiltrotor was tested in the NASA Ames 40-by 80-Foot Wind Tunnel in November 2000. The main objective of the test was to acquire a comprehensive database to validate tiltrotor analyses. Figure 1 shows the model installed in the Ames 40- by 80-Foot Wind Tunnel. Rotor and vehicle performance measurements were taken in addition to wing pressures, acoustics, and flow visualization. A dual acoustic traverse system was installed to measure blade-vortex interaction (BVI) noise levels and directivity. Test conditions included hover and forward flight in helicopter mode. Angle-of-attack and thrust sweeps for three tunnel speeds were acquired before model problems caused the premature conclusion of the test. The test will resume in the Ames 80- by 120-Foot Wind Tunnel in late 2001. This paper will focus on the wake geometry measurements that were acquired during the test. The wake geometry measurements were a small subset of a larger matrix of planned measurements designed to study the development and structure of the dual vortex system generated during BVI conditions. The present paper will provide wake geometry data for four test conditions. In addition, the data will be compared with previously acquired wake measurements from an isolated tiltrotor

Wadcock, Alan J.↗
NASA NTRS2011

Correlating CFD Simulation with Wind Tunnel Test for the Full-Scale UH-60A Airloads Rotor

Data from the recent UH-60A Airloads Test in the National Full-Scale Aerodynamics Complex 40- by 80- Foot Wind Tunnel at NASA Ames Research Center are presented and compared to predictions computed by a loosely coupled Computational Fluid Dynamics (CFD)/Comprehensive analysis. Primary calculations model the rotor in free-air, but initial calculations are presented including a model of the tunnel test section. The conditions studied include a speed sweep at constant lift up to an advance ratio of 0.4 and a thrust sweep at constant speed into deep stall. Predictions show reasonable agreement with measurement for integrated performance indicators such as power and propulsive but occasionally deviate significantly. Detailed analysis of sectional airloads reveals good correlation in overall trends for normal force and pitching moment but pitching moment mean often differs. Chord force is frequently plagued by mean shifts and an overprediction of drag on the advancing side. Locations of significant aerodynamic phenomena are predicted accurately although the magnitude of individual events is often missed.

Romandr, Ethan↗
NASA NTRS2018

Use of the Ames Check Standard Model for the Validation of Wall Interference Corrections

The new check standard model of the NASA Ames 11-ft Transonic Wind Tunnel was chosen for a future validation of the facility's wall interference correction system. The chosen validation approach takes advantage of the fact that test conditions experienced by a large model in the slotted part of the tunnel's test section will change significantly if a subset of the slots is temporarily sealed. Therefore, the model's aerodynamic coefficients have to be recorded, corrected, and compared for two different test section configurations in order to perform the validation. Test section configurations with highly accurate Mach number and dynamic pressure calibrations were selected for the validation. First, the model is tested with all test section slots in open configuration while keeping the model's center of rotation on the tunnel centerline. In the next step, slots on the test section floor are sealed and the model is moved to a new center of rotation that is 33 inches below the tunnel centerline. Then, the original angle of attack sweeps are repeated. Afterwards, wall interference corrections are applied to both test data sets and response surface models of the resulting aerodynamic coefficients in interference-free flow are generated. Finally, the response surface models are used to predict the aerodynamic coefficients for a family of angles of attack while keeping dynamic pressure, Mach number, and Reynolds number constant. The validation is considered successful if the corrected aerodynamic coefficients obtained from the related response surface model pair show good agreement. Residual differences between the corrected coefficient sets will be analyzed as well because they are an indicator of the overall accuracy of the facility's wall interference correction process.

wind tunnel wall interferenc↗