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At least 739 records · Page 41

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NASA NTRS2000

Dynamic Stability Testing of the Genesis Sample Return Capsule

This paper documents a series of free flight tests of a scale model of the Genesis Sample Return Capsule. These tests were conducted in the Aeroballistic Research Facility (ARF), located at Eglin AFB, FL, during April 1999 and were sponsored by NASA Langley Research Center. Because these blunt atmospheric entry shapes tend to experience small angle of attack dynamic instabilities (frequently leading to limit cycle motions), the primary purpose of the present tests was to determine the dynamic stability characteristics of the Genesis configuration. The tests were conducted over a Mach number range of 1.0 to 4.5. The results for this configuration indicate that the models were dynamically unstable at low angles of attack for all Mach numbers tested. At Mach numbers below 2.5, the models were also unstable at the higher angles of attack (above 15 deg), and motion amplitudes of up to 40 deg were experienced. Above Mach 2.5, the models were dynamically stable at the higher angles of attack.

Cheatwood, F. McNeil↗

NASA NTRS1972

Low speed aerodynamic characteristics of a large scale model with a thin, highly swept, 2.67 aspect ratio wing having a cranked leading edge

The low speed aerodynamic characteristics of a large-scale model intended to represent advanced fixed-wing fighters have been investigated in the Ames 40 by 80 foot wind tunnel. The model possessed positive static longitudinal stability to nearly 28 deg angle of attack, the maximum tested, both with and without leading-edge flaps deflected. Lateral control with differentially deflected ailerons and a right wing spoiler simultaneously deployed as combined controls was only slightly greater than that with the differentially deflected ailerons deployed as a separate control without the spoiler. Measured lift and drag were in close agreement with that predicted by theory to about 14 deg angle of attack, including the prediction of lift due to 30 deg of trailing edge flap deflection. Estimated takeoff performance of an aircraft 4/3 the scale of the test model showed takeoff distances of less than 2000 ft.

Giulianetti, D. J.↗

NASA NTRS1986

Subsonic flow investigations on a cranked wing designed for high maneuverability

The characteristic pitching moment nonlinearity of cranked wings limits their usable lift coefficient well below C sub L max. The potential of several aerodynamic devices, viz., fences, pylon vortex generators (PVG), mid-span strakes and cavity flaps, in delaying the pitch up onset on a 70/50 deg cranked wing was explored in low speed tunnel tests. Upper surface pressure measurements and low visualizations were conducted on a semi-span wing model to observe the vortex flow development with increasing angle of attack, and then to assess the effectiveness of the devices in controlling the collapse of vortex lift over the wing panel outboard of the crank. Force tests on a full span wing and body model were also conducted to assess the fence and PVG in improving the usable C sub L.

Rao, D. M.↗

NASA NTRS1957

Static Longitudinal and Lateral Stability Characteristics at Low Speed of 45 Degree Sweptback-midwing Models Having Wings with an Aspect Ratio of 2, 4, or 6

Results are presented of tests conducted in the Langley stability tunnel to determine the effects of various components and combinations of components on the static longitudinal and lateral stability characteristics at low speed of a series of 45 degree sweptback-midwing-airplane configurations having wings with an aspect ratio of 2, 4, or 6. The tests were made at a dynamic pressure of 24.9 pounds per square foot which corresponds to a Mach number of 0.13 and Reynolds numbers of 1.00 x 10(exp 6), 0.71 x 10 (exp 6), and 0.58 x 10 (exp 6) based on the respective wing mean aerodynamic chords. The angle-of-attack range covered was from -4 degrees to 32 degrees and the sideslip angles used for the lateral-derivative tests were 5 degrees and -5 degrees. An increasing loss in tail contribution to directional stability both with increasing wing aspect ratio and increasing angle of attack was found to be a principal cause of all the complete models becoming directionally unstable in the high angle-of-attack range.

STABILITY, LONGITUDINAL - STATIC↗

NASA NTRSDate not supplied

Wall-Modeled Large-Eddy Simulations for High-Lift Configurations using FUN3D

Wall-modeled large-eddy simulation (WMLES) capability has recently been implemented into FUN3D, an unstructured, node-centered, finite-volume solver developed at the NASA Langley Research Center. In this paper, WMLES is assessed for two configurations that are representative for high-lift applications. The first configuration is a nominal two-dimensional multielement airfoil that has been extensively studied in the literature. WMLES solutions are computed for four angles of attack and compared with previously reported solutions. Good agreement of integrated forces, surface pressures, and boundary-layer velocity profiles is shown with available experimental data especially at lower angles of attack. WMLES solutions are also computed for the NASA High-Lift Common Research Model over a large range of angles of attack. Forces, pitching moments, and pressure distributions are favorably compared with the experimental data up to the maximum lift, including the angle of attack where the maximum lift is obtained experimentally. Eddy visualization techniques of q-criterion and density-gradient magnitude illustrate the resolved content.

High-Lift flows↗

NASA NTRS1981

Unpowered Aerodynamic Characteristics of a 15-Percent Scale Model of a Twin-Engine Commuter Aircraft

An experimental investigation was conducted in the Ames 12-Foot Pressure Wind Tunnel to determine the unpowered aerodynamic characteristics of a 15-percent-scale model of a twin-engine commuter aircraft. Model longitudinal aerodynamic characteristics were examined at discrete flap deflections for various angle-of-attack and wind-tunnel-velocity ranges with the empennage on and off. Data are presented for the basic model configuration consisting of the fuselage, wing, basic wing leading edge, double slotted flaps, midengine nacelles, and empennage. Other configurations tested include a particle-span drooped leading edge (dropped outboard of the engine nacelles), a full-span drooped leading edge, low- and high-mounted engine nacelles, and a single-slotted flap. An evaluation was made of the model mounting system by comparing data obtained with the model mounted conventionally on the wind-tunnel model-support struts and the model inverted.

Morgan, D. G.↗

NASA NTRS1985

Angle-of-attack estimation for analysis of CAT encounters

Recent studies of clear-air turbulence (CAT) encounters involving wide-body airliners have been based upon flight-path wind estimates made by analyzing digital flight-data-recorder (DFDR) records and radar records. Such estimates require a time history of the aircraft angle of attack, a record that is not usually included in the DFDR measurement set. This paper describes a method for reconstructing angle of attack that utilizes available flight record and aircraft-specific information associated with an aerodynamic model of the lift coefficient. Results from two wide-body incidents in which vane measurements of angle of attack were recorded show good agreement between measured and calculated time histories. This research has been performed in cooperation with the National Transportation Safety Board to provide a better understanding of the CAT phenomenon.

Bach, R. E., Jr.↗

NASA NTRS2024

Computational Analysis of the X-57 Maxwell Airplane, the Landing Configuration with High-Lift Blowing and Aileron Deflections (Preliminary Fuselage)

The X-57 Maxwell is an all-electric airplane with a distributed electric propulsion system used for a high-lift system at takeoff and landing conditions. The Kestrel and USM3D flow solvers were used at NASA Langley to investigate the performance of the X-57 Maxwell in the development of an aerodynamic database. The configuration investigated in this paper had a 30° flap deflection, and in addition the pilot’s right aileron were deflected and undeflected in different cases. The solutions were computed at an airspeed of 58 KEAS, for an altitude of 6000 feet, and a flight Reynolds number of 0.588e+06 per foot. To evaluate the high-lift distributed electric propulsion system without aileron deflections, the solutions were computed for an angle-of-attack sweep from −2° to 20°, with the high-lift propellers blowing and the cruise propellers excluded from the simulation. To investigate the aileron effectiveness, the solutions were computed at angles of attack of −2° and 14°, for aileron deflections from −25° to 18° with the high-lift propellers blowing and the cruise propellers windmilling at idle-power. The high-lift propellers and the idle-power cruise propellers were modeled with an actuator disk. Results show negligible differences in lift, drag and pitching moment whether the idle-powered cruise propellers were included or excluded from the simulation. In general, the Kestrel and USM3D codes compared well for lift, drag and pitching moment for the landing configuration with no aileron control. The Kestrel code, using the Spalart-Allmaras turbulence model with the rotation corrections terms, predicted an increasing lift coefficient up to maximum lift coefficient of 4.65 at a 15° angle of attack. The USM3D code, using the Spalart-Allmaras turbulence model turbulence model with the Quadratic Constitutive Relation predicted an increasing lift coefficient up to maximum lift coefficient of 4.5 at a 12° angle of attack, with the lift remaining constant through a 14◦ angle of attack. A possible difference in lift coefficient between the codes for high angles of attack may result from the different available options used with the standard Spalart-Allmaras turbulence model. The Kestrel code predicted more drag across the range of angle of attack than the USM3D code. The codes compared well for pitching moment coefficient across the range of angle of attack. The Kestrel and USM3D codes compared well for aileron effectiveness at a 2° angle of attack. The Kestrel code predicts better aileron effectiveness than USM3D at a 14° angle of attack. There is more flow separation in the region outboard of the last high-lift nacelle for the USM3D solutions at a 14° angle of attack, which diminishes the ability of the aileron to be effective.

X-57 Maxwell Airplane↗

NASA NTRS1996

Prediction of Airfoil Characteristics With Higher Order Turbulence Models

This study focuses on the prediction of airfoil characteristics, including lift and drag over a range of Reynolds numbers. Two different turbulence models, which represent two different types of models, are tested. The first is a standard isotropic eddy-viscosity two-equation model, and the second is an explicit algebraic stress model (EASM). The turbulent flow field over a general-aviation airfoil (GA(W)-2) at three Reynolds numbers is studied. At each Reynolds number, predicted lift and drag values at different angles of attack are compared with experimental results, and predicted variations of stall locations with Reynolds number are compared with experimental data. Finally, the size of the separation zone predicted by each model is analyzed, and correlated with the behavior of the lift coefficient near stall. In summary, the EASM model is able to predict the lift and drag coefficients over a wider range of angles of attack than the two-equation model for the three Reynolds numbers studied. However, both models are unable to predict the correct lift and drag behavior near the stall angle, and for the lowest Reynolds number case, the two-equation model did not predict separation on the airfoil near stall.

Gatski, Thomas B.↗

NASA NTRS1993

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

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.

Moes, Timothy R.↗

NASA NTRS1993

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

Moes, Timothy R.↗

NASA NTRS1994

Rarefield-Flow Shuttle Aerodynamics Flight Model

A model of the Shuttle Orbiter rarefied-flow aerodynamic force coefficients has been derived from the ratio of flight acceleration measurements. The in-situ, low-frequency (less than 1Hz), low-level (approximately 1 x 10(exp -6) g) acceleration measurements are made during atmospheric re-entry. The experiment equipment designed and used for this task is the High Resolution Accelerometer Package (HiRAP), one of the sensor packages in the Orbiter Experiments Program. To date, 12 HiRAP re-entry mission data sets spanning a period of about 10 years have been processed. The HiRAP-derived aerodynamics model is described in detail. The model includes normal and axial hypersonic continuum coefficient equations as function of angle of attack, body-flap deflection, and elevon deflection. Normal and axial free molecule flow coefficient equations as a function of angle of attack are also presented, along with flight-derived rarefied-flow transition bridging formulae. Comparisons are made between the aerodynamics model, data from the latest Orbiter Operational Aerodynamic Design Data Book, applicable computer simulations, and wind-tunnel data.

Blanchard, Robert C.↗

NASA NTRS1985

Wind-Tunnel Results of Advanced High-Speed Propellers at Takeoff, Climb, and Landing Mach Numbers

Low-speed wind-tunnel performance tests of two advanced propellers have been completed at the NASA Lewis Research Center as part of the NASA Advanced Turboprop Program. The 62.2 cm (24.5 in.) diameter adjustable-pitch models were tested at Mach numbers typical of takeoff, initial climbout, and landing speeds (i.e., from Mach 0.10 to 0.34) at zero angle of attack in the NASA Lewis 10 by 10 Foot Supersonic Wind Tunnel. Both models had eight blades and a cruise-design-point operating condition of Mach 0.80, and 10.668 km (35,000 ft) I.S.A. altitude, a 243.8 m/s (800 ft/sec) tip speed, and a high power loading of 301 kW/sq m (37.5 shp/sq ft). Each model had its own integrally designed area-ruled spinner, but used the same specially contoured nacelle. These features reduced blade-section Mach numbers and relieved blade-root choking at the cruise condition. No adverse or unusual low-speed operating conditions were found during the test with either the straight blade SR-2 or the 45 deg swept SR-3 propeller. Typical efficiencies of the straight and 45 deg swept propellers were 50.2 and 54.9 percent, respectively, at a takeoff condition of Mach 0.20 and 53.7 and 59.1 percent, respectively, at a climb condition of Mach 0.34.

Stefko, George L.↗

NASA NTRS1984

Heat transfer and oil flow studies on a single-stage-to-orbit control-configured winged entry vehicle

Results are presented for oil flow and phase change paint heat transfer tests conducted on a 0.006 scale model of a proposed single stage to orbit control configured vehicle. The data were taken at angles of attack up to 40 deg at a free stream Mach number of 10 for Reynolds numbers based on model length of 0.5 x 10 to the 6th power, 1.0 x 10 to the 6th power and 2.0 x 10 to the 6th power. The magnitude and distribution of heating are characterized in terms of angle of attack and Reynolds number aided by an analysis of the flow data which are used to suggest the presence of various three dimensional flow structures that produce the observed heating patterns. Of particular interest are streak heating patterns that result in high localized heat transfer rates on the wing windward surface at low to moderate angles of attack. These streaks are caused by the bow-shock/wing-shock interaction and formation of the wing-shock. Embedded vorticity was found to be associated with these interactions.

Helms, V. T., III↗

NASA NTRS2015

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

Characterization of launch vehicle unsteady aerodynamics is a field best studied through experimentation, which is often carried out in the form of large scale wind tunnel testing. Measurement of the fluctuating pressures induced by the boundary layer noise is customarily made with miniature pressure transducers installed into a model of the vehicle of interest. Literature shows that noise level increases between two to five decibels (dB referenced to 20 micropascal) can be induced when the transducer surface is not mounted perfectly flush with the model outer surface. To reduce this artificially induced noise, special transducer holders have been used for aeroacoustic wind tunnel testing by NASA. This holder is a sleeve into which the transducer fits, with a cap that allows it to be mounted in a recessed hole in the model. A single hole in the holder allows the transport of the tunnel medium so the transducer can discriminate the fluctuating pressure due to the turbulent boundary layer noise. The holder is first dry fitted into the model and any difference in height between the holder and the model surface can be sanded flush. The holder is then removed from the model, the transducer glued inside the holder, and the holder replaced in the model, secured also with glue, thus eliminating the problem of noise level increases due to lack of flushness. In order to work with this holder design, special transducers have been ordered with their standard screen removed and the diaphragm moved as close to the top of the casing as possible to minimize any cavity volume. Although this greatly reduces induced noise due to the transducers being out of flush, the holders can also induce a cavity resonance that is usually at a very high frequency. This noise is termed transducer mount induced noise (XMIN). The peak of the mode can vary with the cavity depth, boundary layer noise that can excite the mode, tunnel flow medium, and the build of the transducers. Because the boundary layer flow interaction with the microphone holder seems to have some effect on the cavity resonance response, the NASA/Marshall Space Flight Center (MSFC) aeroacoustic team postulated that changing the design of the holder might reduce the cavity resonance response. An experiment was performed at the MSFC Trisonic Wind Tunnel comparing a new holder design with the traditional MSFC holder. The new holder design is intended to mimic the Kulite Semiconductors, Inc. B-screen, where the single hole of the traditional MSFC holder is replaced by a series of ten 0.007 inch diameter holes. This particular design serves to protect the transducer diaphragm as in the manufacturer's B-screen, but also increases the frequency of the XMIN such that it is not apparent in the data collected from aeroacoustic wind tunnel tests. MSFC compared the two holder designs by installing two transducers in standard holders on one side of a ten inch long ten degree cone, and two transducers in B-screen holders on the opposite side for a series of runs at different conditions. Over 11-13 February, 2013, the model was run at Mach numbers between 0.80 and 1.96 at angles of attack at 0 deg, plus or minus 1 deg, plus or minus 2 deg, plus or minus 4 deg, and plus or minus 8 deg. Model roll attitudes at 0 deg, 90 deg, and 180 deg were tested to isolate potential tunnel bias. This paper presents results from the test showing that the proposed holder design significantly reduces the influence of XMIN on measured fluctuating pressure levels.

Herron, A. J.↗

NASA NTRS1992

Nonlinear stability and control study of highly maneuverable high performance aircraft, phase 2

Research leading to the development of new nonlinear methodologies for the adaptive control and stability analysis of high angle of attack aircraft such as the F-18 is discussed. The emphasis has been on nonlinear adaptive control, but associated model development, system identification, stability analysis, and simulation were studied in some detail as well. Studies indicated that nonlinear adaptive control can outperform linear adaptive control for rapid maneuvers with large changes in angle of attack. Included here are studies on nonlinear model algorithmic controller design and an analysis of nonlinear system stability using robust stability analysis for linear systems.

Mohler, R. R.↗

NASA NTRS1978

Evaluation of interference in the OSU 6 in. by 22 in. transonic airfoil tunnel

Interference in a 6 in. by 22 in. two-dimensional wind tunnel was evaluated at Mach numbers up to 1.06 by comparing pressure distributions from airfoil models of differing size. Models of the NACA 0012 profile, having chords of 76, 152, and 305 mm, were used in one phase of the evaluation program, and models of a supercritical profile, having chords of 76 and 152 mm, were used in another. The confinement interference, was documented i.e., blockage, downwash and streamline curvature, all of which are quite small on a model having a chord of 152 mm and which can, for most applications, be ignored. Specifically, the corrections were lumped into an attack angle adjustment of -0.16 degrees per unit lift coefficient on a 152 mm model.

Lee, J. D.↗

NASA NTRS1990

Numerical simulation of the effects of variation of angle of attack and sweep angle on vortex breakdown over delta wings

In the present investigation of the vortical flowfield structure over delta wings at high angles of attack, three-dimensional Navier-Stokes numerical simulations were conducted to predict the complex leeward flowfield characteristics; these encompass leading-edge separation, secondary separation, and vortex breakdown. Attention is given to the effect on solution accuracy of circumferential grid-resolution variations in the vicinity of the wing leading edge, and well as to the effect of turbulence modeling on the solutions. When a critical angle-of-attack was reached, bubble-type vortex breakdown was found. With further angle-of-attack increase, a change from bubble-type to spiral-type vortex breakdown was predicted by the numerical solution.

Ekaterinaris, J. A.↗