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Improved Unsteady Far-Field Drag Breakdown Method and Application to Complex Cases

AIAA JOURNAL 2016年第6期54卷 1907-1921页

作者： Toubin, Helene Bailly, Didier Costes, Michel Off Natl Etud & Rech Aerosp Appl Aerodynam Dept F-92190 Meudon France

Far-field drag extraction has the advantage over near-field integration of providing a phenomenological breakdown of drag. A decomposition into components linked to shock waves, viscous interactions, and lift-induced vortices is straightforward for steady flows. A formulation based on thermodynamic considerations is used at ONERA-The French Aerospace Lab and by its partners, but it is restricted to steady cases. A generalization to unsteady flows of the Van der Vooren formulation has been developed and tested on three unsteady cases previously. The proposed method allowed the breakdown of drag into the three usual components only;however, the induced drag coefficient remained ill-defined. This unsteady formulation is here modified to better express the induced drag. A new drag component is identified as a propagation and acoustics contribution. The new formulation is then applied to complex cases: two-dimensional and three-dimensional pitching cases, and an OAT15A profile subject to buffet simulated by zonal detached-eddy simulation computations.

关键词： Freestream Velocity Shock Waves One Dimensional flow Wingtip Vortices Acoustics Detached Eddy Simulation Aerodynamic Efficiency flow visualization techniques Supercritical Airfoils Power Spectral Density

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Comparison of Large-Scale Three-Dimensional Features in Zero- and Adverse-Pressure-Gradient Turbulent Boundary Layers

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AIAA JOURNAL 2015年第12期53卷 3686-3699页

作者： Melnick, M. Blake Thurow, Brian S. Auburn Univ Dept Aerosp Engn Adv Flow Diagnost Lab Auburn AL 36849 USA

Large-scale features in a zero-pressure-gradient turbulent boundary layer are compared with those in two adverse-pressure-gradient turbulent boundary layers using a scanning three-dimensional flow-visualization technique. The boundary-layer flow was formed on the wall of a wind tunnel and seeded with smoke introduced through a slit. A ramp was mounted on the opposite wall to create an adverse pressure gradient (beta = 8.1 and 9.7). A high-repetition-rate laser formed into a light sheet was scanned in the wall-normal direction, and the two-dimensional images were reconstructed into a nearly instantaneous three-dimensional flow-visualization volume. Observations about the boundary-layer structures for Reynolds numbers Re-theta. between 7400 and 8400 are made by investigating the instantaneous flow-visualization volumes, the features from binary-image slices, and conditional averaged volumes around a structure. In general, structures in a zero-pressure-gradient boundary layer are larger and spaced farther apart compared to the adverse-pressure-gradient cases. flow features identified through conditional averaging extend higher into the freestream flow for the adverse-pressure-gradient case, and are more distinct and separated from the wall than in the zero-pressure-gradient case. There were no noticeable differences in the spanwise direction for the three cases.

关键词： flow visualization techniques Boundary Layer Separation Reynolds Numbers Blowdown Wind Tunnels Particle Image Velocimetry Accelerating flow Vortices Shear Stress Image Processing Static Pressure

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Ethylene Transverse Jets in Supersonic Crossflows

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JOURNAL OF PROPULSION AND POWER 2015年第3期31卷 773-788页

作者： Park, Gisu Park, Chul Jin, Yuin Choi, Hojin Byun, Jongryul Hwang, Kiyoung Korea Adv Inst Sci & Technol Div Aerosp Engn Taejon 305701 South Korea Korea Adv Inst Sci & Technol Taejon 305701 South Korea Agcy Def Dev Adv Prop Technol Ctr Taejon 305600 South Korea

flowfield and flame-holding characteristics of ethylene jets were examined experimentally. flow visualization and CH * chemiluminescence images were obtained in a shock tunnel. Model configurations with and without a cowl or a cavity were tested for jet-to-freestream momentum flux ratios varying from 1.35 to 3.07. Flame-holding characteristics were studied using a cavity configuration with a cowl. Four flow conditions with the total temperature varying from 1610 to 2320K were studied keeping the overall fuel-air equivalence ratio at 0.2. The data showed that shock impingement due to the cowl influenced the jet to penetrate shallower into the freestream than for the case without the cowl. The shock impingement reduced the fluctuating level of the jet's periodical motion. At 2080 K, combustion was maintained inside the cavity, and this condition provided a continuous ignition downstream of the cavity (i.e., flaming holding is achieved at a much lower enthalpy than seen in the existing data). At 2320 K, the trend was similar. At temperatures lower than 2080 K, the flame holding was not maintained.

关键词： flow Conditions Freestream Conditions Chemiluminescence Enthalpy flow visualization techniques Curved Shocks Shear Layers Oblique Shock Combustors Scramjet

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Swept-Wing Boundary-Layer Transition and Turbulent flow Physics from Multipoint Measurements

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AIAA JOURNAL 2014年第2期52卷 338-347页

作者： Glauser, Mark N. Saric, William S. Chapman, Keith L. Reibert, Mark S. Syracuse Univ Dept Mech & Aerosp Engn Syracuse NY 13244 USA Texas A&M Univ Dept Aerosp Engn College Stn TX 77843 USA Clarkson Univ Dept Mech & Aeronaut Engn Potsdam NY 13699 USA Arizona State Univ Dept Mech & Aerosp Engn Phoenix AZ 85004 USA

Multipoint surface shear stress and hot-wire velocity measurements were obtained within a three-dimensional, swept-wing boundary layer. Proper orthogonal decomposition, spectra, and spatial correlations were used to identify the presence of flow structures and their spatial evolution. The first proper orthogonal decomposition crossflow spatial eigenmode from the shear stress obtained in the laminar region has a wavelength of 12 mm, which is consistent with the crossflow vortex spanwise periodicity. The first proper orthogonal decomposition crossflow spatial eigenmode obtained in the turbulent region is not nearly as distinct, but a large spatial structure approximately 24 mm in span is observed, suggesting that, through the transition process, two neighboring crossflow vortices are interacting with a stronger one, canceling a close weaker one. The proper orthogonal decomposition was also applied in span to hot-wire velocity measurements at 3 mm above the wing surface in both the laminar and turbulent regimes. flow structure in the turbulent flow regime has been identified that is very similar to that obtained in the laminar transition zone. This suggests that the initial instability and transition physics continue to play a key role even after the flow has transitioned to turbulence.

关键词： Boundary Layer Transition Natural Laminar flow Swept Wing Shear Stress Vortex Structure flow visualization techniques Freestream Velocity Surface Roughness Velocity Profiles Hot Wire Anemometry

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Tunnel Background Noise on Compressible Convex-Corner flows

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JOURNAL OF AIRCRAFT 2013年第4期50卷 1011-1015页

作者： Chung, Kung-Ming Chang, Po-Hsiung Chang, Keh-Chin Natl Cheng Kung Univ Aerosp Sci & Technol Res Ctr Tainan 711 Taiwan Natl Cheng Kung Univ Inst Aeronaut & Astronaut Tainan 711 Taiwan

The perforated test section walls of a transonic wind tunnel are responsible for generating edgetones, thus contributing significantly to the overall noise level and possibly invalidating the measurements of surface pressure fluctuations. In the present study, the wind tunnel was assembled with perforated top/bottom walls, solid side walls, and four perforated walls. The focus was to investigate the effects of tunnel background noise on compressible convex-corner flows, which correspond to the upper control surface of an aircraft wing. External acoustic disturbance has a minor effect on the mean surface pressure distributions, including flow expansion and recompression near the corner. Higher peak pressure fluctuations, which are associated with shock excursion phenomena, are observed with more intense tunnel background noise, particularly for the flow with initial separated boundary layers.

关键词： Boundary Layer Separation Transonic Wind Tunnel Lift to Drag Ratio Freestream Mach Number flow visualization techniques Separated flows Pressure Transducers Supersonic Speed Control Surfaces Acoustic Excitation

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AIAA JOURNAL 2012年第4期50卷 985-988页

作者： Chung, Kung-Ming Chang, Po-Hsiung Chang, Keh-Chin Natl Cheng Kung Univ Aerosp Sci & Technol Res Ctr Tainan 701 Taiwan Natl Cheng Kung Univ Inst Aeronaut & Astronaut Tainan 701 Taiwan

Boundary-layer flow in the vicinity of a sharp corner was investigated. A scaling parameter to characterize the transition from subsonic to transonic expansion flows, incipient shock-induced boundary-layer separation, and separation length was proposed. It was found that the flow expansion near the corner apex and peak pressure fluctuations in the recompression process could be scaled linearly with similarity parameter β(<13) for subsonic expansion flows and initial transonic expansion flows. Once the local Mach number exceeded the critical Mach number, the increment of (mean surface static pressure = freestream static pressure) _min with similarity parameter (≊13-20) became less significant. For the test cases of extensive shock-induced boundary-layer separation, the highly intermittent wall-pressure signals due to shock excursion phenomena resulted in intense pressure fluctuations in the recompression process.

关键词： Boundary Layer Separation Freestream Mach Number flow visualization techniques Static Pressure Standard Deviation Compressible flow Supersonic Speed Transonic Wind Tunnel Pressure Transducers flow Characteristics

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flow Topology Around Gas, Liquid, and Three-Dimensional Obstacles in Hypersonic flow

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AIAA JOURNAL 2012年第1期50卷 100-108页

作者： Asma, C. O. Tirtey, S. Schloegel, F. von Karman Inst Fluid Dynam Turbomachinery Dept B-1640 Rhode St Genese Belgium Univ Queensland St Lucia Qld 4072 Australia

The flow topologies of gas and liquid injections into hypersonic crossflow are compared with the flow topology around three-dimensional rigid obstacles in hypersonic flow. Experimental techniques such as oil flow visualization, infrared thermography, a sublimation technique, and high-speed schlieren photography are applied for a Mach 6 crossing flow on a flat plate. The shock system and the vortex system are visualized and discussed for all cases. Pressure ratios for gas injection, momentum flux ratio for liquid injection, and Reynolds number for three-dimensional obstacles are taken into consideration. The effects of nondimensional numbers on the flow topologies are analyzed. Similarities and differences among different topologies are explained by studying the vortical structures, making use of computational fluid dynamics simulations.

关键词： Transitional flow Horseshoe Vortex flow visualization techniques Aerodynamic Stability Skin Friction Coefficient Aspect Ratio Shock Layers Computational Fluid Dynamics Simulation Hypersonic Wind Tunnels Thermal Protection System

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Experimental Studies on a Rudimentary Four-Wheel Landing Gear

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AIAA JOURNAL 2012年第11期50卷 2435-2447页

作者： Venkatakrishnan, L. Karthikeyan, N. Mejia, Kevin Natl Aerosp Labs Expt Aerodynam Div Bangalore 560017 Karnataka India Boeing Co Boeing Engn Operat & Technol Aero Noise & Prop Labs Aero Testing Methods Seattle WA 98108 USA

Experiments were carried out on a rudimentary four-wheel landing-gear arrangement to understand the flow physics of this configuration. Measurements carried out included oil-flow visualization, steady surface pressures and two-dimensional particle image velocimetry measurements in the near wake at a Reynolds number based on wheel diameter of 1 x 10(6). The surface visualization demonstrated the significantly asymmetric behavior of the flow on both the front and back wheel. The surface-pressure data show that separation is earlier on the wing side as opposed to the ground side for both wheels, caused by the presence of the vertical strut. The data show that the truck and strut are dominant contributors to the total drag. The off-body measurements in the near wake document the confluent wakes from the wheel and the horizontal strut leading to a three-dimensional unsteady wake. The data presented here provide a comprehensive picture, both quantitatively and qualitatively, of the flow over a rudimentary landing gear and are useful for validation of computational fluid dynamics efforts in addition to understanding the flow physics of such configurations.

关键词： Landing Gear flow visualization techniques Freestream Velocity Particle Image Velocimetry Detached Eddy Simulation Landing Gear Arrangements Static Pressure Laminar Separation Bubble Vortices National Aerospace Laboratories

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Turbulent flows of an Acoustically Excited Elevated Transverse Jet

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AIAA JOURNAL 2012年第9期50卷 1964-1978页

作者： Huang, R. F. Hsu, C. M. Natl Taiwan Univ Sci & Technol Dept Mech Engn Taipei 106 Taiwan

The turbulent flow characteristics in the symmetry planes of unforced and excited elevated transverse jets were studied experimentally in a wind tunnel. Jet pulsations were generated by means of acoustic excitation. The jet Reynolds number and the jet-to-crossflow momentum-flux ratio were 1000 and 1.2, respectively. A high-speed particle image velocimeter was employed to measure the velocity field. Exciting the transverse jet at Strouhal numbers less than about 0.53 caused the synchronized flapping-jet mode. Exciting the transverse jet at Strouhal numbers larger than about 0.77 induced the synchronized shear-layer-vortices mode. As the transverse jet was excited in the synchronized flapping-jet mode, the jet shot to high altitude and was drastically less deflected than the unforced transverse jet. The jet-wake vortex ring disappeared and the bifurcation lines moved upstream. The fluctuation intensities, vorticities, and turbulent shear stress of the excited transverse jet exhibited drastically higher values than those of the unforced transverse jet. The Lagrangian integral time and length scales of the transverse jet excited at the resonance Strouhal number were drastically smaller than those of the unforced one.

关键词： Turbulent flow Strouhal Numbers Shear Layers flow visualization techniques Freestream Velocity Acoustic Excitation Particle Image Velocimetry Resonance Frequencies flow Characteristics Stress Distribution

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Separation-Control Mechanisms of Steady and Pulsed Vortex-Generator Jets

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JOURNAL OF PROPULSION AND POWER 2012年第6期28卷 1201-1213页

作者： Evans, Simon Hodson, Howard Worcester Polytech Inst Dept Mech Engn Worcester MA 01609 USA Univ Cambridge Whittle Lab Dept Engn Cambridge CB3 0DY England

Vortex-generator jets have been used to remove an open separation on the suction surface of blades within a linear compressor cascade at high incidence. Steady jets removed the separation with a jet velocity ratio of 70%, reducing the loss coefficient by 60% relative to the uncontrolled case. Pulsed jets tested over a range of reduced actuation frequencies also yielded a reduction in the loss coefficient. The mechanisms by which steady and pulsed vortex-generator jets delay separation and reduce loss were explored using particle-image velocimetry measurements. The results suggest that the mechanisms by which steady and pulsed vortex-generator jets achieve separation control are different. With steady blowing, streamwise vortices mix the separated boundary layer. In the pulsed case, however, at least some control authority is a result of spanwise structures, the roll up of which is encouraged by actuation at appropriate forcing frequencies.

关键词： Vortex Generator Jets Boundary Layer Separation Starting Vortex Particle Image Velocimetry Pulse Jet Compressor Cascade flow visualization techniques Mass flow Rate Freestream Velocity Separated flows

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