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Aeroacoustic study of a high-fidelity aircraft model: Part 2-unsteady surface pressures

Aeroacoustic study of a high-fidelity aircraft model: Part 2...

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18th AIAA/CEAS Aeroacoustics Conference 2012 (33rd AIAA Aeroacoustics Conference)

作者： Khorrami, Mehdi R. Neuhart, Dan H. NASA Langley Research Center Computational AeroSciences Branch Hampton VA 23681 United States NASA Langley Research Center Flow Physics and Control Branch Hampton VA 23681 United States

ISBN: (纸本)9781600869327

In this paper, we present unsteady surface pressure measurements for an 18%-scale, semi-span Gulfstream aircraft model. This high-fidelity model is being used to perform detailed studies of airframe noise associated with main landing gear, flap components, and gear-flap interaction noise, as well as to evaluate novel noise reduction concepts. The aerodynamic segment of the tests, conducted in the NASA Langley Research center 14- by 22-Foot Subsonic Tunnel, was completed in November 2010. To discern the characteristics of the surface pressure fluctuations in the vicinity of the prominent noise sources, unsteady sensors were installed on the inboard and outboard flap edges, and on the main gear wheels, struts, and door. Various configurations were tested, including flap deflections of 0°, 20°, and 39°, with and without the main landing gear. The majority of unsteady surface pressure measurements were acquired for the nominal landing configuration where the main gear was deployed and the flap was deflected 39°. To assess the Mach number variation of the surface pressure amplitudes, measurements were obtained at Mach numbers of 0.16, 0.20, and 0.24. Comparison of the unsteady surface pressures with the main gear on and off shows significant interaction between the gear wake and the inboard flap edge, resulting in higher amplitude fluctuations when the gear is present.

关键词： NASA

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Time-resolved measurements of the aero-optic effects of a compressible shear layer 43

Time-resolved measurements of the aero-optic effects of a co...

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43rd AIAA Plasmadynamics and Lasers Conference 2012

作者： Kelly, Ryan Rennie, R. Mark Jumper, Eric J. Center for Flow Physics and Control University of Notre Dame IN 46556 United States Department of Mechanical and Aerospace Engineering Hessert Laboratory for Aerospace Research Notre Dame IN 46556 United States

ISBN: (数字)9781624101878

ISBN: (纸本)9781624101878

Experimental data are presented of time-resolved optical wavefront measurements through a two-dimensional, weakly compressible, subsonic shear layer. The measurements were obtained in the University of Notre Dame's compressible shear layer wind tunnel (CSWLT), which is capable of generating high-speed shear-layer flows up to Mach 1.0 that can also be regularized using mechanical forcing applied at the trailing edge of the splitter plate between the high and low-speed flows. Experimental data are also presented demonstrating the use of a Phase-Lock-Loop (PLL) controller to detect and lock on to the phase of the shear-layer optical aberration;this kind of PLL control would be used to synchronize a deformable mirror with the shear-layer aberration, as part of a feedforward adaptive-optic correction for the optical effect of the shear layer. © 2012 by Ryan Kelly, R. Mark Rennie, and Eric J. Jumper. Published by the American Institute of Aeronautics and Astronautics, Inc.

关键词： Aberrations

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flow disturbance characterization measurements in the national transonic facility (invited) 50

Flow disturbance characterization measurements in the nation...

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50th AIAA Aerospace Sciences Meeting Including the New Horizons Forum and Aerospace Exposition

作者： King, Rudolph A. Andino, Marlyn Y. Melton, Latunia Eppink, Jenna Kegerise, Michael A. Tsoi, Andrew Flow Physics and Control Branch NASA Langley Research Center Hampton VA 23681 United States Computational Aerosciences Branch NASA Langley Research Center Hampton VA 23681 United States University of Colorado at Boulder United States

ISBN: (数字)9781600869365

flow physics and control Branch, NASA Langley Research center, Hampton, VA 23681 Recent flow measurements have been acquired in the National Transonic Facility (NTF) to assess the unsteady flow environment in the test section. The primary purpose of the test is to determine the feasibility of the NTF to conduct laminar-flow-control testing and boundary-layer transition sensitive testing. The NTF can operate in two modes, warm (air) and cold/cryogenic (nitrogen) test conditions for testing full and semispan scaled models. The warm-air mode enables low to moderately high Reynolds numbers through the use of high tunnel pressure, and the nitrogen mode enables high Reynolds numbers up to flight conditions, depending on aircraft type and size, utilizing high tunnel pressure and cryogenic temperatures. NASA's Environmentally Responsible Aviation (ERA) project is interested in demonstrating different laminar-flow technologies at flight-relevant operating conditions throughout the transonic Mach number range and the NTF is well suited for the initial ground-based demonstrations. In the current test, we acquired data for Mach and unit Reynolds numbers ranging from 0.2 ≤ M ≤ 0.95 and 3.3 × 106 s/q m/( m), etc. Based on the current measurements and previous data, an assessment was made that the NTF is a suitable facility for ground-based demonstrations of laminar-flow technologies at flight-relevant conditions in the cryogenic mode.

关键词： Laminar flow

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High-reynolds number circulation control testing in the national transonic facility (invited)

High-reynolds number circulation control testing in the nati...

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50th AIAA Aerospace Sciences Meeting Including the New Horizons Forum and Aerospace Exposition

作者： Milholen II, William E. Jones, Gregory S. Chan, David T. Goodliff, Scott L. NASA Langley Research Center Hampton VA 23681 United States Configuration Aerodynamics Branch Mail Stop 499 United States Flow Physics and Control Branch Mail Stop 170 United States

A new capability to test active flow control concepts and propulsion simulations at high Reynolds numbers in the National Transonic Facility at the NASA Langley Research center is being developed. The first active flow control experiment was completed using the new FAST-MAC semi-span model to study Reynolds number scaling effects for several circulation control concepts. Testing was conducted over a wide range of Mach numbers, up to chord Reynolds numbers of 30 million. The model was equipped with four onboard flow control valves allowing independent control of the circulation control plenums, which were directed over a 15% chord simple-hinged flap. Preliminary analysis of the uncorrected lift data showed that the circulation control increased the low-speed maximum lift coefficient by 33%. At transonic speeds, the circulation control was capable of positively altering the shockwave pattern on the upper wing surface and reducing flow separation. Furthermore, application of the technique to only the outboard portion of the wing demonstrated the feasibility of a pneumatic based roll control capability.

关键词： Reynolds number

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Vortex formation of a finite span synthetic jet

Vortex formation of a finite span synthetic jet

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50th AIAA Aerospace Sciences Meeting Including the New Horizons Forum and Aerospace Exposition

作者： Van Buren, Tyler Amitay, Michel Whalen, Edward Rensselaer Polytechnic Institute Mechanical Aerospace and Nuclear Engineering 110 8th street Troy NY 12180 United States Boeing Company Boeing Research and Technology Active Flow Control 325 James S McDonnell Blvd Hazelwood MO 63042 United States Center for Flow Physics and Control 110 8th street Troy NY 12180 United States

The effects of different geometries and input parameters on the flow structures of a finite span synthetic jet are explored in quiescent flow experiments using stereoscopic PIV. Common geometrical parameters, such as neck height, and orifice aspect ratio are varied along with the jet performance characteristics such as Strouhal numbers and Reynolds numbers. Orifice neck heights of 2, 4 and 6mm were tested at aspect ratios of 6, 12, and 18. Jet Strouhal numbers of 0.115 to 0.157 were tested at Reynolds numbers of 615. It is found that as the aspect ratio was increased the size of the vortical structures decreases, the vorticity dissipation rate increases, and axis switching occurred farther away from the orifice. The neck length had a large effect on the structures' size and strength, their dissipation rate, and the jet's spreading. Moreover, the driving frequency was found to affect the vortical structures spreading due to the streamwise spacing between structures, as well as drastically affecting dissipation of the structures in both planes. © 2012 by the American Institute of Aeronautics and Astronautics, Inc.

关键词： Aspect ratio

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Investigation and active control of bluff body vortex shedding using plasma actuators 6

Investigation and active control of bluff body vortex sheddi...

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6th AIAA flow control Conference 2012

作者： Koc, Ilteris Britcher, Colin P. Wilkinson, Stephen P. Aselsan Inc. Guidance and Electro-Optics Division Ankara 06011 Turkey Old Dominion University Mechanical and Aerospace Engineering Department Norfolk Va 23529 United States NASA Langley Research Center Flow Physics and Control Branch MS 170 Hampton VA 23681 United States

ISBN: (数字)9781624101885

ISBN: (纸本)9781624101885

Phase-averaging methods and Particle Image Velocimetry (PIV) are used for the investigation of active control of bluff body vortex shedding using plasma actuators. Aerodynamic characteristics of a circular cylinder and an axisymmetric body (ogive cylinder) are first analyzed using a proposed phase averaging technique for the PIV data. Later, effect of plasma actuators on the aerodynamic characteristics of these bodies is investigated. Modification of the streamwise pressure gradient, leading to the establishment of a thinner boundary layer downstream of the circular cylinder was accomplished. Phase synchronization of vortex shedding was also obtained for the circular cylinder. Effects of different configurations of plasma actuators on the circumference, on the base, and in a streamwise direction were investigated for the ogive cylinder. It was observed that direct alteration of the mean flow about a body was not as effective as the boundary layer flow control where the flow instabilities are exploited. Also the three dimensionalities in this flow made it significantly more complex to analyze. © 2012 by Ilteris Koc, Colin P. Britcher and Stephen P. Wilkinson. Published by the American Institute of Aeronautics and Astronautics, Inc.

关键词： Circular cylinders

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Crossflow instability on a wedge-cone at mach 3.5

Crossflow instability on a wedge-cone at mach 3.5

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42nd AIAA Fluid Dynamics Conference and Exhibit 2012

作者： Beeler, George B. Wilkinson, Stephen P. Balakumar, P. Mcdaniel, Keith S. NASA Langley Research Center Hampton VA 23681 United States Analytical Mechanics Associates Inc Hampton VA 23681 United States Flow Physics and Control Branch M/S 170 United States Hypersonic Air Breathing Propulsion Branch M/S 168 United States

ISBN: (纸本)9781600869334

As a follow-on activity to the HyBoLT flight experiment, a six degree half angle wedge-cone model at zero angle of attack has been employed to experimentally and computationally study the boundary layer crossflow instability at Mach 3.5 under low disturbance freestream conditions. Computed meanflow and linear stability analysis results are presented along with corresponding experimental Pitot probe data. Using a model-mounted probe survey apparatus, data acquired to date show a well defined stationary crossflow vortex pattern on the flat wedge surface. This effort paves the way for additional detailed, calibrated flow field measurements of the crossflow instability, both stationary and traveling modes, and transition-to-turbulence under quiet flow conditions as a means of validating existing stability theory and providing a foundation for dynamic flight instrumentation development.

关键词： Probes

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Aeroacoustic study of a high-fidelity aircraft model: Part 1- steady aerodynamic measurements

Aeroacoustic study of a high-fidelity aircraft model: Part 1...

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18th AIAA/CEAS Aeroacoustics Conference 2012 (33rd AIAA Aeroacoustics Conference)

作者： Khorrami, Mehdi R. Hannon, Judith A. Neuhart, Dan H. Markowski, Gregory A. van de Ven, Thomas NASA Langley Research Center Computational AeroSciences Branch Hampton VA 23681 United States NASA Langley Research Center Flow Physics and Control Branch Hampton VA 23681 United States NASA Langley Research Center Jacobs Technology ROME Group Hampton VA 23681 United States NASA Langley Research Center Jacobs Technology ROME Group Hampton VA 23681 United States Gulfstream Aerospace Corporation Acoustics and Vibration Group Savannah GA 31402 United States

ISBN: (纸本)9781600869327

In this paper, we present steady aerodynamic measurements for an 18% scale model of a Gulfstream air-craft. The high fidelity and highly-instrumented semi-span model was developed to perform detailed aeroacoustic studies of airframe noise associated with main landing gear/flap components and gear-flap in-teraction noise, as well as to evaluate novel noise reduction concepts. The aeroacoustic tests, being conducted in the NASA Langley Research center 14- by 22-Foot Subsonic Tunnel, are split into two entries. The first entry, completed November 2010, was entirely devoted to the detailed mapping of the aerodynamic charac-teristics of the fabricated model. Flap deflections of 39°, 20°, and 0° with the main landing gear on and off were tested at Mach numbers of 0.16, 0.20, and 0.24. Additionally, for each flap deflection, the model was tested with the tunnel both in the closed-wall and open-wall (jet) modes. During this first entry, global forces (lift and drag) and extensive steady and unsteady surface pressure measurements were obtained. Preliminary analysis of the measured forces indicates that lift, drag, and stall characteristics compare favorably with Gulfstream's high Reynolds number flight data. The favorable comparison between wind-tunnel and flight data allows the semi-span model to be used as a test bed for developing/evaluating airframe noise reduction concepts under a relevant environment. Moreover, initial comparison of the aerodynamic measurements ob-tained with the tunnel in the closed- and open-wall configurations shows similar aerodynamic behavior. This permits the acoustic and off-surface flow measurements, planned for the second entry, to be conducted with the tunnel in the open-jet mode. © copyright protection in the United States.

关键词： Aeroacoustics

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Receptivity to roughness, acoustic, and vortical disturbances in supersonic boundary layers over swept wings

Receptivity to roughness, acoustic, and vortical disturbance...

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41st AIAA Fluid Dynamics Conference and Exhibit 2011

作者： Balakumar, P. King, Rudolph A. Flow Physics and Control Branch NASA Langley Research Center Hampton VA 23681 United States

ISBN: (纸本)9781600869471

The receptivity and interaction of stationary and traveling crossflow instability of three-dimensional supersonic boundary layers over a swept biconvex wing with a blunt leading edge are numerically investigated for a freestream Mach number of 3. The steady and unsteady flow fields are obtained by solving the full Navier-Stokes equations. The receptivity of the boundary layer to surface roughness, freestream acoustic waves, and freestream vorticity waves are numerically investigated. The initial amplitudes of the stationary vortices generated by 1 μm roughness elements is about 2000 times larger than the initial amplitudes of the traveling disturbances generated by vortical disturbances with an amplitude level of ũv/U∞ = 1*10-5. The traveling crossflow vortices generated by acoustic disturbances of pac/p∞ = 1*10-5 are about 8 times weaker than the vortices generated by the vorticity waves. The interaction of stationary and traveling disturbances was investigated by solving the equations with both surface roughness and vortical disturbances. When the initial amplitudes of the stationary disturbances are large compared to the traveling disturbances, the stationary vortex dominates the perturbation field. When the amplitudes arc comparable, the traveling vortex prevails and the stationary vortex is suppressed.

关键词： Vortex flow

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Recent AAOL in-flight wavefront measurements of aero-optics and implications for aero-optics beam control in tactical laser weapon systems

Recent AAOL in-flight wavefront measurements of aero-optics ...

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42nd AIAA Plasmadynamics and Lasers Conference 2011

作者： Goorskey, David J. Whiteley, Matthew R. Gordeyev, Stanislav Jumper, Eric J. MZA Associates Dayton OH 45430 United States Center for Flow Physics and Control University of Notre Dame Department of Aerospace and Mech. Eng. Notre Dame IN 46556 United States

ISBN: (纸本)9781624101472

Aero-optics disturbances were measured in- flight with the Airborne Aero-Optics Laboratory (AAOL) 1 ft (4 in. clear aperture) flat-windowed turret at an altitude of 15 kft altitude and at a Mach number of 0.5 for 47o az., 42o el. (forward-looking) and 139o az., 70o el. (rear- looking) turret angles. Wavefronts were collected using a FasCam CCD array camera with an AMO lenslet array in a Shack-Hartman configuration at a 20 kHz frame-rate. 2,000 consecutive wavefronts of each of the two turret pointing angles measured were scaled to a tactical high energy laser (HEL) laser weapon scenario and their wavefront error statistics analyzed. The two wavefront sequences were decomposed using proper orthogonal decomposition (POD) modes and their spatial frequency content evaluated. From this, the minimum number of deformable mirror (DM) actuators needed to compensate these disturbances was estimated to be approximately 18 actuators across the aperture diameter. Finally, a the performance of a simple integrator-type adaptive optics (AO) control system against the two scaled wavefront sequences was investigated. It was found that the forward-looking aero-optics aberrations were about three-times less in magnitude than the rear-looking aero-optics aberrations. Consequently, the open-loop Strehl ratio of the forward-looking case was considerably greater than the open-loop Strehl ratio of the rear-looking case. Nevertheless, in for both the forward- looking and rear-looking cases, in order to achieve compensation better than open-loop, it was found necessary to have AO sample rates greater than 10 kHz with one frame of latency. These results suggest that significant AO compensation improvements will be achieved mainly by reducing the latency as much as possible, perhaps with adaptive/predictive AO controllers. © 2011 by the American Institute of Aeronautics and Astronautics, Inc.

关键词： Wavefronts

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