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Time-frequency analysis of boundary-layer instabilities generated by freestream laser perturbations 45

Time-frequency analysis of boundary-layer instabilities gene...

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45th AIAA Fluid Dynamics Conference, 2015

作者： Chou, Amanda Schneider, Steven P. Flow Physics and Control Branch NASA Langley Research Center HamptonVA23681 United States School of Aeronautics and Astronautics Purdue University West LafayetteIN47907 United States

ISBN: (数字)9781624103629

ISBN: (纸本)9781624103629

A controlled disturbance is generated in the freestream of the Boeing/AFOSR Mach-6 Quiet Tunnel (BAM6QT) by focusing a high-powered Nd:YAG laser to create a laser-induced breakdown plasma. The plasma then cools, creating a freestream thermal distur-bance that can be used to study receptivity. The freestream disturbance convects down-stream in the Mach-6 wind tunnel to interact with a flared cone model. The adverse pressure gradient created by the flare of the model is capable of generating second-mode instability waves that grow large and become nonlinear before experiencing natural transition in quiet flow. The freestream laser perturbation generates a wave packet in the boundary layer at the same frequency as the natural second mode, complicating time-independent analyses of the effect of the laser perturbation. The data show that the laser perturbation creates an instability wave packet that is larger than the natural waves on the sharp flared cone. The wave packet is still difficult to distinguish from the natural instabilities on the blunt flared cone. © 2015, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.

关键词： Wave packets

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Boundary layer instabilities generated by freestream laser perturbations 53

Boundary layer instabilities generated by freestream laser p...

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53rd AIAA Aerospace Sciences Meeting, 2015

作者： Chou, Amanda Schneider, Steven P. NASA Langley Research Center Flow Physics and Control Branch HamptonVA23681 United States School of Aeronautics & Astronautics Purdue University West LafayetteIN47907 United States

ISBN: (数字)9781624103438

ISBN: (纸本)9781624103438

A controlled, laser-generated, freestream perturbation was created in the freestream of the Boeing/AFOSR Mach-6 Quiet Tunnel (BAM6QT). The freestream perturbation convected downstream in the Mach-6 wind tunnel to interact with a flared cone model. The geometry of the flared cone is a body of revolution bounded by a circular arc with a 3-meter radius. Fourteen PCB 132A31 pressure transducers were used to measure a wave packet generated in the cone boundary layer by the freestream perturbation. This wave packet grew large and became nonlinear before experiencing natural transition in quiet flow. Breakdown of this wave packet occurred when the amplitude of the pressure fluctuations was approximately 10% of the surface pressure for a nominally sharp nosetip. The initial amplitude of the second mode instability on the blunt flared cone is estimated to be on the order of 10-6 times the freestream static pressure. The freestream laser-generated perturbation was positioned upstream of the model in three different configurations: on the centerline, offset from the centerline by 1.5 mm, and offset from the centerline by 3.0 mm. When the perturbation was offset from the centerline of a blunt flared cone, a larger wave packet was generated on the side toward which the perturbation was offset. The offset perturbation did not show as much of an effect on the wave packet on a sharp flared cone as it did on a blunt flared cone. © 2015, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.

关键词： Boundary layers

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A comparative study of simulated and measured gear-flap flow interaction 21

A comparative study of simulated and measured gear-flap flow...

引用

21st AIAA/CEAS Aeroacoustics Conference, 2015

作者： Khorrami, Mehdi R. Mineck, Raymond E. Yao, Chungsheng Jenkins, Luther N. NASA Langley Research Center Computational AeroSciences Branch HamptonVA23681 United States Yorktown VA23692 United States NASA Langley Research Center Flow Physics and Control Branch HamptonVA23681 United States

ISBN: (数字)9781624103674

ISBN: (纸本)9781624103674

The ability of two CFD solvers to accurately characterize the transient, complex, interacting flowfield asso-ciated with a realistic gear-flap configuration is assessed via comparison of simulated flow with experimental measurements. The simulated results, obtained with NASA’s FUN3D and Exa’s Powerflow® for a high-fidelity, 18% scale semi-span model of a Gulfstream aircraft in landing configuration (39° flap deflection, main landing gear on and off) are compared to two-dimensional and stereo particle image velocimetry measurements taken within the gear-flap flow interaction region during wind tunnel tests of the model. As part of the bench-marking process, direct comparisons of the mean and fluctuating velocity fields are presented in the form of planar contour plots and extracted line profiles at measurement planes in various orientations stationed in the main gear wake. The measurement planes in the vicinity of the flap side edge and downstream of the flap trailing edge are used to highlight the effects of gear presence on tip vortex development and the ability of the computational tools to accurately capture such effects. The present study indicates that both computed datasets contain enough detail to construct a relatively accurate depiction of gear-flap flow interaction. Such a finding increases confidence in using the simulated volumetric flow solutions to examine the behavior of pertinent aer-odynamic mechanisms within the gear-flap interaction zone. © 2015, American Institute of Aeronautics and Astronautics Inc, AIAA. All Rights Reserved.

关键词： NASA

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Flight test of an AFC enhanced vertical tail 8th

Flight test of an AFC enhanced vertical tail

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8th AIAA flow control Conference, 2016

作者： Whalen, Edward A. Shmilovich, Arvin Spoor, Marc Tran, John Vijgen, Paul Lin, John C. Andino, Marlyn The Boeing Company Boeing Research and Technology 325 James S. McDonnell Blvd MS S306-4030 HazelwoodMO63042 United States The Boeing Company Boeing Research and Technology 14900 Bolsa Chica Rd H017-D335 Huntington BeachCA92647 United States The Boeing Company Boeing Commercial Airplanes SeattleWA98124 United States Flow Physics and Control Branch NASA Langley Research Center MS 170 HamptonVA23666 United States Boeing Commercial Airplanes MukilteoWA98275 United States

ISBN: (纸本)9781624104329

The size of the vertical tail and rudder of a commercial aircraft is influenced by the requirement that the airplane must be controllable in the event of an engine failure. As a result, on twin-engine airplanes it is typically larger than is required for stability and control under normal flight conditions. Despite the rarity of the need for this level of control, the airplane must carry this size vertical tail and rudder with it at all times. The primary impact of this is additional drag and weight that reduces the fuel efficiency of the airplane. Active flow control can be used to enhance the control authority of the rudder by mitigating flow separation on it at high rudder deflection and side slip angles. This effect can be utilized to provide the required level of rudder control authority from a physically smaller vertical tail. As a result, fuel efficiency of the airplane may be increased due to the reduced size and drag of the smaller vertical tail throughout the operational envelope of the airplane. In April 2015, Boeing and NASA tested a pneumatic sweeping jet based active flow control system on the vertical tail of the modified Boeing 757 ecoDemonstrator. This test was the culmination of several years of development that included small scale wind tunnel testing, CFD simulations, a system integration study, full-scale wind tunnel testing and the flight demonstration. The data from the flight test indicate that the flow control actuation may be able to provide an approximately 14% increase in side force at the maximum tested rudder deflection and at critical sideslip angles. © 2016, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.

关键词： Wind tunnels

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The interaction of a backward-facing step and cross flow instabilities in boundary-layer transition 53

The interaction of a backward-facing step and cross flow ins...

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53rd AIAA Aerospace Sciences Meeting, 2015

作者： Eppink, Jenna L. Wlezien, Richard W. King, Rudolph A. Choudhari, Meelan NASA Langley Research Center Flow Physics and Control Branch HamptonVA23681 United States Iowa State University 2271 Howe Hall Room 1200A AmesIA50011 United States

ISBN: (数字)9781624103438

ISBN: (纸本)9781624103438

A low-speed experiment was performed on a swept at plate model with an imposed pressure gradient to determine the effect of a backward-facing step on transition in a stationary-cross flow dominated ow. Detailed hot-wire boundary-layer measurements were performed for a step height of approximately 50% of the boundary layer thickness at the step. Three leading-edge roughness configurations were tested to determine the ef- fect of stationary-cross flow amplitude on transition, and it was found that increasing the stationary-cross flow amplitude caused an upstream movement of the transition front. The step caused a local increase in the growth of the stationary cross flow instabilities, but the stationary cross flow amplitude at transition was suffciently low ( © 2015 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

关键词： Modulation

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Plasma flow control on a landing gear model 53

Plasma flow control on a landing gear model

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53rd AIAA Aerospace Sciences Meeting, 2015

作者： Wicks, Michael Thomas, Flint O. Corke, Thomas C. Nelson, Chris Patel, Mehul Cain, Alan B. Department of Aerospace and Mechanical Engineering Center for Flow Physics and Control University of Notre Dame Notre DameIN46556 United States Innovative Technology Applications Company LLC 14929 Royalbrook Dr ChesterfieldMO63017-7722 United States

ISBN: (数字)9781624103438

ISBN: (纸本)9781624103438

This paper reports an effort to extend bluff-body plasma noise control technology demonstrated for a tandem cylinder configuration in [28,29] to a realistic landing gear geometry. Experiments and numerical simulations are presented which focus on plasma flow control of a generalized tandem geometry consisting of the interaction between the hydraulic strut and downstream torque arm of the Gulfstream G550 nose landing gear. Experiments and numerical simulations are preformed at a ReD=122,000 and M=0.081. This component work is considered prerequisite to the broader goal of demonstrating plasma noise control on a partially dressed, closed cavity 30% scale model of a Gulfstream G550 nose gear. © 2015 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

关键词： Numerical models

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flow separation control on a full-scale vertical tail model using sweeping jet actuators 53

Flow separation control on a full-scale vertical tail model ...

引用

53rd AIAA Aerospace Sciences Meeting, 2015

作者： Andino, Marlyn Y. Lin, John C. Washburn, Anthony E. Whalen, Edward A. Graff, Emilio C. Wygnanski, Israel J. NASA Langley Research Center Flow Physics and Control Branch MS 170 HamptonVA23681 United States NASA Langley Research Center Flow Physics and Control Branch MS 41 HamptonVA23681 United States The Boeing Company Boeing Research and Technology Mail Stop: S306-4030 HazelwoodMO63042 United States California Institute of Technology Graduate Aerospace Laboratories PasadenaCA91125 United States The University of Arizona Aerospace and Mechanical Engineering Department TucsonAZ85721 United States

ISBN: (数字)9781624103438

ISBN: (纸本)9781624103438

This paper describes test results of a joint NASA/Boeing research effort to advance Active flow control (AFC) technology to enhance aerodynamic efficiency. A full-scale Boeing 757 vertical tail model equipped with sweeping jets AFC was tested at the National Full-Scale Aerodynamics Complex 40- by 80-Foot Wind Tunnel at NASA Ames Research center. The flow separation control optimization was performed at 100 knots, a maximum rudder deflection of 30°, and sideslip angles of 0° and -7.5°. Greater than 20% increments in side force were achieved at the two sideslip angles with a 31-actuator AFC configuration. flow physics and flow separation control associated with the AFC are presented in detail. AFC caused significant increases in suction pressure on the actuator side and associated side force enhancement. The momentum coefficient (Cµ) is shown to be a useful parameter to use for scaling-up sweeping jet AFC from sub-scale tests to full-scale applications. Reducing the number of actuators at a constant total Cµ of approximately 0.5% and tripling the actuator spacing did not significantly affect the flow separation control effectiveness. © 2015, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.

关键词： Actuators

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Characterization of flap edge noise radiation from a high-fidelity airframe model 21

Characterization of flap edge noise radiation from a high-fi...

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21st AIAA/CEAS Aeroacoustics Conference, 2015

作者： Humphreys, William M. Khorrami, Mehdi R. Lockard, David P. Neuhart, Dan H. Bahr, Christopher J. NASA Langley Research Center Advanced Measurements and Data Systems Branch HamptonVA23681-2199 United States NASA Langley Research Center Computational AeroSciences Branch HamptonVA23681-2199 United States NASA Langley Research Center Flow Physics and Control Branch HamptonVA23681-2199 United States NASA Langley Research Center Aeroacoustics Branch HamptonVA23681-2199 United States

ISBN: (数字)9781624103674

ISBN: (纸本)9781624103674

The results of an experimental study of the noise generated by a baseline high-fidelity airframe model are presented. The test campaign was conducted in the open-jet test section of the NASA Langley 14- by 22-foot Subsonic Tunnel on an 18%-scale, semi-span Gulfstream airframe model incorporating a trailing edge flap and main landing gear. Unsteady surface pressure measurements were obtained from a series of sensors positioned along the two flap edges, and far field acoustic measurements were obtained using a 97-microphone phased array that viewed the pressure side of the airframe. The DAMAS array deconvolution method was employed to determine the locations and strengths of relevant noise sources in the vicinity of the flap edges and the landing gear. A Coherent Output Power (COP) spectral method was used to couple the unsteady surface pressures measured along the flap edges with the phased array output. The results indicate that outboard flap edge noise is dominated by the flap bulb seal cavity with very strong COP coherence over an approximate model-scale frequency range of 1 to 5 kHz observed between the array output and those unsteady pressure sensors nearest the aft end of the cavity. An examination of the COP spectra for the inboard flap proved inconclusive, most likely due to a combination of coherence loss caused by decorrelation of acoustic waves propagating through the thick wind tunnel shear layer and contamination of the spectra by tunnel background noise at lower frequencies. Directivity measurements obtained from integration of DAMAS pressure-squared values over defined geometric zones around the model show that the baseline flap and landing gear are only moderately directional as a function of polar emission angle. © 2015, American Institute of Aeronautics and Astronautics Inc, AIAA. All Rights Reserved.

关键词： NASA

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Performance enhancement of a full-scale vertical tail model equipped with active flow control 53

Performance enhancement of a full-scale vertical tail model ...

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53rd AIAA Aerospace Sciences Meeting, 2015

作者： Whalen, Edward A. Lacy, Douglas Lin, John C. Andino, Marlyn Y. Washburn, Anthony E. Graff, Emilio Wygnanski, Israel J. The Boeing Company Boeing Research and Technology Mail Stop: S306-4030 HazelwoodMO63042 United States The Boeing Company Boeing Commercial Airplanes PO Box 3707 M/S 0R-MM SeattleWA98124 United States NASA Langley Research Center Flow Physics and Control Branch MS 170 HamptonVA23681 United States NASA Langley Research Center Flow Physics and Control Branch MS 41 HamptonVA23681 United States California Institute of Technology Graduate Aerospace Laboratories PasadenaCA91125 United States The University of Arizona Aerospace and Mechanical Engineering Department TucsonAZ85721 United States

ISBN: (数字)9781624103438

ISBN: (纸本)9781624103438

This paper describes wind tunnel test results from a joint NASA/Boeing research effort to advance active flow control (AFC) technology to enhance aerodynamic efficiency. A full-scale Boeing 757 vertical tail model equipped with sweeping jet actuators was tested at the National Full-Scale Aerodynamics Complex (NFAC) 40- by 80-Foot Wind Tunnel (40×80) at NASA Ames Research center. The model was tested at a nominal airspeed of 100 knots and across rudder deflections and sideslip angles that covered the vertical tail flight envelope. A successful demonstration of AFC-enhanced vertical tail technology was achieved. A 31-actuator configuration significantly increased side force (by greater than 20%) at a maximum rudder deflection of 30°. The successful demonstration of this application has cleared the way for a flight demonstration on the Boeing 757 ecoDemonstrator in 2015. © 2015 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

关键词： Demonstrations

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control of Tollmien-Schlichting waves using piezoelectrically driven oscillating surface 45

Control of Tollmien-Schlichting waves using piezoelectricall...

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45th AIAA Fluid Dynamics Conference, 2015

作者： Dell’Orso, Haley Tuna, Burak A. Amitay, Michael Department of Mechanical Aerospace and Nuclear Engineering Rensselaer Polytechnic Institute 110 8th Street TroyNY12180 United States Center for Flow Physics and Control Department of Mechanical Aerospace and Nuclear Engineering Rensselaer Polytechnic Institute 110 8th Street TroyNY12180 United States

ISBN: (数字)9781624103629

ISBN: (纸本)9781624103629

The feasibility of using dynamic surface modification (via discrete Piezoelectrically Driven Oscillating Surface, PDOS, actuators) to delay transition from laminar to turbulent flow on a flat plate was explored experimentally using Particle Image Velocimetry. Two circular PDOS actuators were used to excite and control the T-S waves on a flat plate. The actuators were placed at the mid-span of a flat plate at two streamwise locations downstream from the leading edge, where the free-stream velocity was U = 16 m/s. The upstream actuator was used to excite and phase-lock the T-S wave at an excitation frequency of 250 Hz (corresponding to the most amplified wave from stability analysis), and the downstream PDOS was used to mitigate the T-S wave by driving it at an anti-phase with the appropriate amplitude. Time-and phase-averaged flow fields were acquired at two downstream interrogation domains. The data from the first domain (located in between the two PDOS actuators) explored the effect of the excitation amplitude on the magnitude of the excited T-S wave. At the second window, located downstream of both PDOS actuators, activation of both actuators resulted in a significant reduction or augmentation of the T-S wave, depending on the phase shift between the two actuators. © 2015, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.

关键词： Actuators

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