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Numerical simulation of a high-lift configuration with embedded fluidic actuators 32

Numerical simulation of a high-lift configuration with embed...

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32nd AIAA Applied Aerodynamics Conference 2014

作者： Vatsa, Veer N. Casalino, Damiano Lin, John C. Appelbaum, Jason NASA Langley Research Center Computational AeroSciences Branch Hampton VA 23681 United States Exa GmbH Curiestrasse 4 D-70563 Stuttgart Germany NASA Langley Research Center Flow Physics and Control Branch Hampton VA 23681 United States

ISBN: (数字)9781624102882

ISBN: (纸本)9781624102882

Numerical simulations have been performed for a vertical tail configuration with deflected rudder. The suction surface of the main element of this configuration is embedded with an array of 32 fluidic actuators that produce oscillating sweeping jets. Such oscillating jets have been found to be very effective for flow control applications in the past. In the current paper, a high-fidelity computational fluid dynamics (CFD) code known as the Powerflow® code is used to simulate the entire flow field associated with this configuration, including the flow inside the actuators. The computed results for the surface pressure and integrated forces compare favorably with measured data. In addition, numerical solutions predict the correct trends in forces with active flow control compared to the no control case. Effect of varying yaw and rudder deflection angles are also presented. In addition, computations have been performed at a higher Reynolds number to assess the performance of fluidic actuators at flight conditions.

关键词： flow control

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flow disturbance and surface roughness effects on cross-flow boundary-layer transition in supersonic flows 44

Flow disturbance and surface roughness effects on cross-flow...

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44th AIAA Fluid Dynamics Conference 2014

作者： Owens, Lewis R. Beeler, George B. Balakumar, P. McGuire, Patrick J. Flow Physics and Control Branch NASA Langley Research Center 1A East Reid St. M/S 170 Hampton VA 23681 United States Dept. of Mechanical and Aerospace Engineering University of California San Diego CA 92093 United States

ISBN: (数字)9781624102899

ISBN: (纸本)9781624102899

Fundamental experiments were performed to isolate and study cross-flow transition on a bi-convex, 35°-swept-wing model in Mach 2 flow conditions. Transition data acquired in this investigation were images of surface-based flow details resulting from the use of a sublimating chemical test technique. Freestream probes were used to acquire dynamic data to characterize the flow disturbance environment near the wing leading edge at each test condition. Surface profilometer measurements were acquired to characterize the wing model surface roughness near the leading edge for two different polish levels. Stationary cross-flow transition was observed for the higher total pressure test conditions for the polish level with a rougher finish. At the lower total pressure conditions, it was difficult to determine whether the cross-flow transition was dominated by either stationary or traveling disturbances. Traveling disturbances dominated the cross-flow transition at all test conditions for the polish level with a smoother finish. The impact of several distributed-roughness-element patterns on cross-flow transition was also explored. This study provides a broader perspective on what leading-edge wing surface roughness levels designers may want to consider for supersonic swept-wing laminar flow-control applications.

关键词： Swept wings

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Low-speed aerodynamic investigations of a hybrid wing body configuration 32

Low-speed aerodynamic investigations of a hybrid wing body c...

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32nd AIAA Applied Aerodynamics Conference 2014

作者： Vicroy, Dan D. Gatlin, Gregory M. Jenkins, Luther N. Murphy, Patrick C. Carter, Melissa B. Flight Dynamics Branch NASA Langley Research Center MS 308 Hampton VA 23681 United States Configuration Aerodynamics Branch NASA Langley Research Center MS 267 Hampton VA 23681 United States Flow Physics and Control Branch NASA Langley Research Center MS 170 Hampton VA 23681 United States Dynamic Systems and Control Branch NASA Langley Research Center MS 308 Hampton VA 23681 United States Configuration Aerodynamics Branch NASA Langley Research Center MS499 Hampton VA 23681 United States

ISBN: (数字)9781624102882

ISBN: (纸本)9781624102882

Two low-speed static wind tunnel tests and a water tunnel static and dynamic forced-motion test have been conducted on a hybrid wing-body (HWB) twinjet configuration. These tests, in addition to computational fluid dynamics (CFD) analysis, have provided a comprehensive dataset of the low-speed aerodynamic characteristics of this nonproprietary configuration. In addition to force and moment measurements, the tests included surface pressures, flow visualization, and off-body particle image velocimetry measurements. This paper will summarize the results of these tests and highlight the data that is available for code comparison or additional analysis.

关键词： flow visualization

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flow-field investigation of gear-flap interaction on a gulfstream aircraft model 20

Flow-field investigation of gear-flap interaction on a gulfs...

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20th AIAA/CEAS Aeroacoustics Conference 2014

作者： Yao, Chungsheng Jenkins, Luther N. Bartram, Scott M. Harris, Jerome Khorrami, Mehdi R. Derry Mace, W. NASA Langley Research Center Hampton VA 23681 United States Sierra-Lobo Hampton VA 23681 United States Flow Physics and Control Branch NASA Langley Research Center MS 170 United States Advanced Sensing and Optical Measurements Branch NASA Langley Research Center MS 493 United States Computational AeroSciences Branch NASA Langley Research Center MS 128 United States Technical Applications Branch NASA Langley Research Center MS 289 United States

ISBN: (数字)9781624102851

ISBN: (纸本)9781624102851

Off-surface flow measurements of a high-fidelity 18% scale Gulfstream aircraft model in landing configuration with the main landing gear deployed are presented. Particle Image Velocimetry (PIV) and Laser Velocimetry (LV) were used to measure instantaneous velocities in the immediate vicinity of the main landing gear and its wake and near the inboard tip of the flap. These measurements were made during the third entry of a series of tests conducted in the NASA Langley Research center (LaRC) 14- by 22-Foot Subsonic Tunnel (14x22) to obtain a comprehensive set of aeroacoustic measurements consisting of both aerodynamic and acoustic data. The majority of the off-body measurements were obtained at a freestream Mach number of 0.2, angle of attack of 3 degrees, and flap deflection angle of 39 degrees with the landing gear on. A limited amount of data was acquired with the landing gear off. LV was used to measure the velocity field in two planes upstream of the landing gear and to measure two velocity profiles in the landing gear wake. Stereo and 2-D PIV were used to measure the velocity field over a region extending from upstream of the landing gear to downstream of the flap trailing edge. Using a special traverse system installed under the tunnel floor, the velocity field was measured at 92 locations to obtain a comprehensive picture of the pertinent flow features and characteristics. The results clearly show distinct structures in the wake that can be associated with specific components on the landing gear and give insight into how the wake is entrained by the vortex at the inboard tip of the flap.

关键词： Velocity measurement

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Attached and separated flow simulations with a hybrid RANS - Dynamic LES method

Attached and separated flow simulations with a hybrid RANS -...

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44th AIAA Fluid Dynamics Conference 2014

作者： Stoellinger, Michael Heinz, Stefan Balakumar, Ponnampalam Department of Mechanical Engineering University of Wyoming 1000 E. University Avenue Laramie WY 82071 United States Department of Mathematics University of Wyoming 1000 E. University Avenue Laramie WY 82071 United States NASA Langley Research Center Flow Physics and Control Branch Hampton VA 23681 United States

ISBN: (纸本)9781624102899

The feasibility of using a dynamic subgrid-scale (SGS) model in a hybrid RANS-LES method is studied in this work. If the transition between RANS and LES is located very close to the wall (such as in low to moderate Reynolds number flows) both the RANS and LES models require the use of a damping function if they are based on an eddy viscosity model. Most damping functions for SGS models are based on the geometric wall distance (such as the van Driest damping function) and thus not ideal in flows with complex geometries. As an alternative, dynamic SGS models have been shown to provide the required damping in wall resolved LES. Therefore, dynamic SGS models seem to be a good choice to be used in hybrid RANS-LES methods. In this work, a realizable dynamic one-equation SGS model is combined with a RANS model that is based on a novel damping function to a hybrid RANS - dynamic LES method. The new RANS damping function is based on the ratio of the Kolmogorov time scale to the mean shear time scale and thus does not rely on the geometric wall distance. The damping function is validated with pure RANS simulations of plane channel flows and external aerodynamic simulations of the NACA 0012 airfoil over a wide range of angle of attacks. The performance of the new hybrid RANS - dynamic LES method is demonstrated for the same flows and the encouraging results warrant the small computational overhead of the dynamic procedure.

关键词： Damping

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An experimental study of roughness-induced instabilities in a supersonic boundary layer 7

An experimental study of roughness-induced instabilities in ...

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AIAA AVIATION 2014 -7th AIAA Theoretical Fluid Mechanics Conference 2014

作者： Kegerise, Michael A. King, Rudolph A. Choudhari, Meelan Li, Fei Norris, Andrew NASA Langley Research Center Flow Physics and Control Branch M. S. 170 Hampton VA 23681 United States NASA Langley Research Center Computational Aerosciences Branch M. S. 128 Hampton VA 23681 United States NASA Langley Research Center Hypersonic Air Breathing Propulsion Branch M. S. 168 Hampton VA 23681 United States

ISBN: (数字)9781624102936

ISBN: (纸本)9781624102936

Progress on an experimental study of laminar-to-turbulent transition induced by an isolated roughness element in a supersonic laminar boundary layer is reported in this paper. Here, the primary focus is on the effects of roughness planform shape on the instability and transition characteristics. Four different roughness planform shapes were considered (a diamond, a circle, a right triangle, and a 45° fence) and the height and width of each one was held fixed so that a consistent frontal area was presented to the oncoming boundary layer. The nominal roughness Reynolds number was 462 and the ratio of the roughness height to the boundary layer thickness was 0.48. Detailed ow-field surveys in the wake of each geometry were performed via hot-wire anemometry. High- and low-speed streaks were observed in the wake of each roughness geometry, and the modified mean flow associated with these streak structures was found to support a single dominant convective instability mode. For the symmetric planform shapes|the diamond and circular planforms|the instability characteristics (mode shapes, growth rates, and frequencies) were found to be similar. For the asymmetric planform shapes|the right-triangle and 45° fence planforms| the mode shapes were asymmetrically distributed about the roughness-wake centerline. The instability growth rates for the asymmetric planforms were lower than those for the symmetric planforms and therefore, transition onset was delayed relative to the symmetric planforms.

关键词： Stability

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Preliminary experimental assessment of the boundary layer ingestion benefit for the D8 aircraft 52

Preliminary experimental assessment of the boundary layer in...

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52nd AIAA Aerospace Sciences Meeting - AIAA Science and Technology Forum and Exposition, SciTech 2014

作者： Uranga, Alejandra Drela, Mark Greitzer, Edward M. Titchener, Neil A. Lieu, Michael K. Siu, Nina M. Huang, Arthur C. Gatlin, Gregory M. Hannon, Judith A. Massachusetts Institute of Technology Cambridge MA 02139 United States NASA Langley Research Center Hampton VA 23681 United States Dept. of Aeronautics and Astronautics 77 Massachusetts Avenue Cambridge MA 02139 United States Dept. of Aeronautics and Astronautics United States Configuration Aerodynamics Branch Mail Stop 267 United States Flow Physics and Control Branch Mail Stop 170 United States

ISBN: (数字)9781624102561

ISBN: (纸本)9781624102561

Wind tunnel experiments were performed to quantify the aerodynamic benefit of boundary layer ingestion (BLI) for the D8 transport aircraft concept. Two powered 1:11 scale, 13.4 ft span models, in BLI and non-BLI versions, were tested at the NASA Langley 14×22 Foot Subsonic Wind Tunnel to directly compare their performance. The models share the same basic airframe and propulsor units for the most direct comparison. They are also fully tripped to make the measured BLI benefit results scalable to the full-size aircraft. The comparison metric is the propulsor power required to produce a given net stream-wise force on the entire aircraft. The results show that the model BLI propulsors require 6% less electrical power at the simulated cruise point. These experiments provide the first back-to-back assessment quantifying the aerodynamic benefits of BLI for a civil aircraft. The BLI benefit quoted is preliminary in nature because it is defined in terms of electrical power, but we are in the process of obtaining a value in terms of flow power and there is indication that the BLI saving will remain essentially the same.

关键词： Wind tunnels

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Evaluation of dielectric-barrier-discharge actuator substrate materials 45

Evaluation of dielectric-barrier-discharge actuator substrat...

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45th AIAA Plasmadynamics and Lasers Conference 2014

作者： Wilkinson, Stephen P. Siochi, Emilie J. Sauti, Godfrey Xu, Tian-Bing Meador, Mary Ann Guo, Haiquan Flow Physics and Control Branch Mail Stop 170 United States Advanced Materials and Processing Branch Mail Stop 397 United States Durability and Coatings Branch MS 49-3 United States NASA Langley Research Center Hampton VA 23681 United States National Institute of Aerospace 100 Exploration Way Hampton VA 23666 United States NASA Glenn Research Center Cleveland OH 44135 United States Ohio Aerospace Institute Brookpark OH 44142 United States

ISBN: (数字)9781624102905

ISBN: (纸本)9781624102905

A key, enabling element of a dielectric barrier discharge (DBD) actuator is the dielectric substrate material. While various investigators have studied the performance of different homogeneous materials, most often in the context of related DBD experiments, fundamental studies focused solely on the dielectric materials have received less attention. The purpose of this study was to conduct an experimental assessment of the body-force-generating performance of a wide range of dielectric materials in search of opportunities to improve DBD actuator performance. Materials studied included commonly available plastics and glasses as well as a custom-fabricated polyimide aerogel. Diagnostics included static induced thrust, electrical circuit parameters for 2D surface discharges and ID volume discharges, and dielectric material properties. Lumped-parameter circuit simulations for the ID case were conducted showing good correspondence to experimental data provided that stray capacitances are included. The effect of atmospheric humidity on DBD performance was studied showing a large influence on thrust. The main conclusion is that for homogeneous, dielectric materials at forcing voltages less than that required for streamer formation, the material chemical composition appears to have no effect on body force generation when actuator impedance is properly accounted for.

关键词： Dielectric barrier discharge

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Receptivity of hypersonic boundary layers to distributed roughness and acoustic disturbances 51

Receptivity of hypersonic boundary layers to distributed rou...

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51st AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition 2013

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

ISBN: (数字)9781624101816

ISBN: (纸本)9781624101816

Boundary-layer receptivity and stability of Mach 6 flows over smooth and rough sevendegree half-angle sharp-tipped cones are numerically investigated. The receptivity of the boundary layer to slow acoustic disturbances, fast acoustic disturbances, and vortical disturbances is considered. The effects of three-dimensional isolated roughness on the receptivity and stability are also simulated. The results for the smooth cone show that the instability waves are generated in the leading edge region and that the boundary layer is much more receptive to slow acoustic waves than to the fast acoustic waves. Vortical disturbances also generate unstable second modes, however the receptivity coefficients are smaller than that of the slow acoustic wave. Distributed roughness elements located near the nose region decreased the receptivity of the second mode generated by the slow acoustic wave by a small amount. Roughness elements distributed across the continuous spectrum increased the receptivity of the second mode generated by the slow and fast acoustic waves and the vorticity wave. The largest increase occurred for the vorticity wave. Roughness elements distributed across the synchronization point did not change the receptivity of the second modes generated by the acoustic waves. The receptivity of the second mode generated by the vorticity wave increased in this case, but the increase is lower than that occurred with the roughness elements located across the continuous spectrum. The simulations with an isolated roughness element showed that the second mode waves generated by the acoustic disturbances are not influenced by the small roughness element. Due to the interaction, a three-dimensional wave is generated. However, the amplitude is orders of magnitude smaller than the two-dimensional wave.

关键词： Acoustic waves

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Roughness induced transition in a supersonic boundary layer 43

Roughness induced transition in a supersonic boundary layer

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43rd AIAA Fluid Dynamics Conference

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

ISBN: (数字)9781624102141

ISBN: (纸本)9781624102141

Direct numerical simulation is used to investigate the transition induced by threedimensional isolated roughness elements in a supersonic boundary layer at a free stream Mach number of 3.5. Simulations are performed for two different configurations: one is a square planform roughness and the other is a diamond planform roughness. The mean-flow calculations show that the roughness induces counter rotating streamwise vortices downstream of the roughness. These vortices persist for a long distance downstream and lift the low momentum fluid from the near wall region and place it near the outer part of the boundary layer. This forms highly inflectional boundary layer profiles. These observations agree with recent experimental observations. The receptivity calculations showed that the amplitudes of the mass-flux fluctuations near the neutral point for the diamond shape roughness are the same as the amplitude of the acoustic disturbances. They are three times smaller for the square shape roughness.

关键词： Boundary layers

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