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Stereo particle image velocimetry measurements of transition downstream of a forward-facing step in a swept-wing boundary layer 55

Stereo particle image velocimetry measurements of transition...

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55th AIAA Aerospace Sciences Meeting

作者： Eppink, Jenna L. NASA Langley Research Center Flow Physics and Control Branch M.S. 170 HamptonVA23681 United States

ISBN: (纸本)9781624104473

Stereo particle image velocimetry measurements were performed downstream of a forward-facing step in a stationary-crossow dominated flow. Three different step heights were studied with the same leading-edge roughness configuration to determine the effect of the step on the evolution of the stationary-crossow. Above the critical step height, which is approximately 68% of the boundary-layer thickness at the step, the step caused a significant increase in the growth of the stationary crossow. For the largest step height studied (68%), premature transition occurred shortly downstream of the step. The stationary crossow amplitude only reached approximately 7% of Ue in this case, which suggests that transition does not occur via the high-frequency secondary instabilities typically associated with stationary crossow transition. The next largest step of 60% δ still caused a significant impact on the growth of the stationary crossow downstream of the step, but the amplitude eventually returned to that of the baseline case, and the transition front remained the same. The smallest step height (56%) only caused a small increase in the stationary crossow amplitude and no change in the transition front. A final case was studied in which the roughness on the leading edge of the model was enhanced for the lowest step height case to determine the impact of the stationary crossow amplitude on transition. The stationary crossow amplitude was increased by approximately four times, which resulted in premature transition for this step height. However, some notable differences were observed in the behavior of the stationary crossow mode, which indicate that the interaction mechanism which results in the increased growth of the stationary crossow downstream of the step may be different in this case compared to the larger step heights. © 2017, American Institute of Aeronautics and Astronautics Inc. All rights reserved.

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Time resolved stereo particle image velocimetry measurements of the instabilities downstream of a backward-facing step in a swept-wing boundary layer 47

Time resolved stereo particle image velocimetry measurements...

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47th AIAA Fluid Dynamics Conference, 2017

作者： Eppink, Jenna L. Yao, Chung-Sheng NASA Langley Research Center Flow Physics and Control Branch M.S. 170 HamptonVA23681 United States

ISBN: (数字)9781624105005

ISBN: (纸本)9781624105005

Time-resolved particle image velocimetry (TRPIV) measurements are performed down-stream of a swept backward-facing step, with a height of 49% of the boundary-layer thickness. The results agree well qualitatively with previously reported hotwire measurements, though the amplitudes of the fluctuating components measured using TRPIV are higher. Nonetheless, the low-amplitude instabilities in the flow are fairly well resolved using TR-PIV. Proper orthogonal decomposition is used to study the development of the traveling crossflow and Tollmien-Schlichting (TS) instabilities downstream of the step and to study how they interact to form the large velocity spikes that ultimately lead to transition. A secondary mode within the traveling crossflow frequency band develops with a wavelength close to that of the stationary crossflow instability, so that at a certain point in the phase, it causes an increase in the spanwise modulation initially caused by the stationary crossflow mode. This increased modulation leads to an increase in the amplitude of the TS mode, which, itself, is highly modulated through interactions with the stationary crossflow. When the traveling crossflow and TS modes align in time and space, the large velocity spikes occur. Thus, these three instabilities, which are individually of low amplitude when the spikes start to occur (u′rms=Ue © 2017, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.

关键词： Velocity measurement

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Three-dimensional ice-accretion measurement methodology for experimental aerodynamic simulation

Three-dimensional ice-accretion measurement methodology for ...

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作者： Broeren, Andy P. Addy, Harold E. Lee, Sam Monastero, Marianne C. McClain, Stephen T. NASA John H. Glenn Research Center Icing Branch 21000 Brookpark Rd. MS 11-2 ClevelandOH44135 United States Vantage Partners LLC ClevelandOH44135 United States University of Illinois at Urbana-Champaign UrbanaIL61801 United States Baylor University Mechanical Engineering Dept. One Bear Place No. 97356 WacoTX76798 United States Rensselaer Polytechnic Institute Department of Mechanical Aerospace and Nuclear Engineering Center for Flow Physics and Control Watervliet Facility 110 8th St. TroyNY12180 United States

Determining the adverse aerodynamic effects due to ice accretion often relies on dry-air wind-tunnel testing of artificial, or simulated, ice shapes. Recent developmentsinice-accretion documentation methods have yieldeda laserscanning capability that can measure highly three-dimensional features of ice accreted in icing wind tunnels. The objective of the current study was to evaluate the aerodynamic accuracy of ice-accretion simulations generated from laser-scan data. Tests were conductedinthe NASA Glenn Icing Research Tunnel, where ice accretions were measured with laser scanning before the molding process. Aerodynamic performance testing was conducted at the University of Illinois low-speed wind tunnel at a Reynolds number of 1.8 × 106 and a Mach number of 0.18 with an airfoil model designed to accommodate artificial ice shapes. The ice-accretion molds were used to fabricate one set of artificial ice shapes from polyurethane castings. The laser-scan data were usedto fabricate another setof artificial ice shapes using rapid-prototype manufacturing. The iced-airfoil results with both sets of artificial ice shapes were compared to evaluate the aerodynamic simulation accuracy. The data showed that the laser-scan and rapid-prototype manufacturing approach was capableof replicating small ice features within the reported accuracies ofthe laser-scan measurement and rapid-prototyping methods, thus providing a new capability for high-fidelity ice-accretion documentation and artificial ice-shape fabrication for icing research. © 2017 by the authors.

关键词： Ice

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Design of an axisymmetric afterbody test case for CFD validation 23

Design of an axisymmetric afterbody test case for CFD valida...

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23rd AIAA Computational Fluid Dynamics Conference, 2017

作者： Disotell, Kevin J. Rumsey, Christopher L. NASA Langley Research Center HamptonVA23681 United States NASA Langley Research Center Flow Physics and Control Branch MS 170 HamptonVA23681 United States NASA Langley Research Center Computational AeroSciences Branch MS 128 HamptonVA23681 United States

ISBN: (数字)9781624105067

ISBN: (纸本)9781624105067

As identified in the CFD Vision 2030 Study commissioned by NASA, validation of advanced RANS models and scale-resolving methods for computing turbulent flow fields must be supported by continuous improvements in fundamental, high-fidelity experiments designed specifically for CFD implementation. In accordance with this effort, the underpinnings of a new test platform referred to herein as the NASA Axisymmetric Afterbody are presented. The devised body-of-revolution is a modular platform consisting of a forebody section and afterbody section, allowing for a range of flow behaviors to be studied on interchangeable afterbody geometries. A body-of-revolution offers advantages in shape definition and fabrication, in avoiding direct contact with wind tunnel sidewalls, and in tail-sting integration to facilitate access to higher Reynolds number tunnels. The current work is focused on validation of smooth-body turbulent flow separation, for which a six-parameter body has been developed. A priori RANS computations are reported for a risk-reduction test configuration in order to demonstrate critical variation among turbulence model results for a given afterbody, ranging from barely-attached to mild separated flow. RANS studies of the effects of forebody nose (with/without) and wind tunnel boundary (slip/no-slip) on the selected afterbody are presented. Representative modeling issues that can be explored with this configuration are the effect of higher Reynolds number on separation behavior, flow physics of the progression from attached to increasingly-separated afterbody flows, and the effect of embedded longitudinal vortices on turbulence structure. © 2017, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.

关键词： Computational fluid dynamics

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High lift common research model for wind tunnel testing: An active flow control perspective 55

High lift common research model for wind tunnel testing: An ...

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55th AIAA Aerospace Sciences Meeting

作者： Lin, John C. Melton, Latunia P. Viken, Sally A. Andino, Marlyn Y. Koklu, Mehti Hannon, Judith A. Vatsa, Veer N. Flow Physics and Control Branch MS 170 NASA Langley Research Center HamptonVA23681 United States Configuration Aerodynamics Branch MS 499 NASA Langley Research Center HamptonVA23681 United States Computation AeroSciences Branch MS 128 NASA Langley Research Center HamptonVA23681 United States

ISBN: (数字)9781624104473

ISBN: (纸本)9781624104473

This paper provides an overview of a research and development effort sponsored by the NASA Advanced Air Transport Technology Project to achieve the required high-lift performance using active flow control (AFC) on simple hinged flaps while reducing the cruise drag associated with the external mechanisms on slotted flaps of a generic modern transport aircraft. The removal of the external fairings for the Fowler flap mechanism could help to reduce drag by 3.3 counts. The main challenge is to develop an AFC system that can provide the necessary lift recovery on a simple hinged flap high-lift system while using the limited pneumatic power available on the aircraft. Innovative low-power AFC concepts will be investigated in the flap shoulder region. The AFC concepts being explored include steady blowing and unsteady blowing operating in the spatial and/or temporal domain. Both conventional and AFC-enabled high-lift configurations were designed for the current effort. The high-lift configurations share the cruise geometry that is based on the NASA Common Research Model, and therefore, are also open geometries. A 10%-scale High Lift Common Research Model (HL-CRM) is being designed for testing at the NASA Langley Research center 14- by 22-Foot Subsonic Tunnel during fiscal year 2018. The overall project plan, status, HL-CRM configurations, and AFC objectives for the wind tunnel test are described. © 2017, American Institute of Aeronautics and Astronautics. All Rights reserved.

关键词： Low Speed Wind Tunnel Freestream Mach Number High Performance Aircraft Trailing Edge Flap NASA Langley Research center Landing Gear Aerodynamic Performance Mean Aerodynamic Chord Lifting Wing Particle Image Velocimetry

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flow around low aspect ratio cylinders and their applications for flow control 55

Flow around low aspect ratio cylinders and their application...

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55th AIAA Aerospace Sciences Meeting

作者： Gildersleeve, Samantha Clingman, Dan Amitay, Michael Rensselaer Polytechnic Institute Department of Mechanical Aerospace and Nuclear Engineering 110 8th Street TroyNY12180 United States The Boeing Company MiltonWA98354 United States Center for Flow Physics and Control 110 8th Street TroyNY United States

ISBN: (纸本)9781624104473

The implementation of an array of finite span, low aspect ratio circular pins as a flow control means to mitigate separation over a flapped airfoil was investigated experimentally in a wind tunnel. As a first step, the flowfield associated with the interaction of a single finite span, low aspect ratio static/dynamic cylindrical pin with a laminar boundary layer over a flat plate was explored. The local and global flowfields were explored for the single pin, which showed that adding a dynamic motion resulted in an increased magnitude of vortical structures associated with a low aspect ratio cylinder, increased downwash over the pin and entrainment of fluid around the cylinder sides. Additionally, the pin’s mean height is explored in a parametric study to determine the effects that the mean height of a static pin with respect to the local boundary layer imposed on the nearfield region. Results indicated that the magnitude of downwash and flow acceleration around the cylinder is directly related to its mean height and blockage effect to the free stream. The pin erected to approximately the same height as the local boundary layer appears to provide the most significant results in terms of enhanced downwash and minimal wake deficit. Lastly, the effect of an array of static cylinders, placed just upstream of a deflected control surface, on the separated flow over the control surface was explored via surface pressure and Stereo Particle Image Velocimetry measurements. The array of pins was embedded in a turbulent boundary layer, where the height of the pin was approximately ½ of the local boundary layer thickness. The static array of pins demonstrated reduction (or complete mitigation) of the separated flow over the control surface, as well as a change in the global circulation;thus, an enhancement of the lift force and reduction of drag is anticipated. Overall, the features discussed have provided strong evidence of the potential that finite span, low aspect ratio cylindr

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Sweeping jet optimization studies 8

Sweeping jet optimization studies

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8th AIAA flow control Conference, 2016

作者： Melton, LaTunia Pack Koklu, Mehti Andino, Marlyn Lin, John C. Edelman, Louis Flow Physics and Control Branch NASA Langley Research Center HamptonVA United States

ISBN: (数字)9781624104329

ISBN: (纸本)9781624104329

Progress on experimental efforts to optimize sweeping jet actuators for active flow control (AFC) applications with large adverse pressure gradients is reported. Three sweeping jet actuator configurations, with the same orifice size but different internal geometries, were installed on the flap shoulder of an unswept, NACA 0015 semi-span wing to investigate how the output produced by a sweeping jet interacts with the separated flow and the mechanisms by which the flow separation is controlled. For this experiment, the flow separation was generated by deflecting the wing’s 30% chord trailing edge flap to produce an adverse pressure gradient. Steady and unsteady pressure data, Particle Image Velocimetry data, and force and moment data were acquired to assess the performance of the three actuator configurations. The actuator with the largest jet deflection angle, at the pressure ratios investigated, was the most efficient at controlling flow separation on the flap of the model. Oil flow visualization studies revealed that the flow field controlled by the sweeping jets was more three-dimensional than expected. The results presented also show that the actuator spacing was appropriate for the pressure ratios examined. © 2016, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.

关键词： Actuators

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Active flow control uwsing sweeping jet actuators on a semi-span wing model 54th

Active flow control uwsing sweeping jet actuators on a semi-...

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54th AIAA Aerospace Sciences Meeting, 2016

作者： Pack Melton, LaTunia Koklu, Mehti Flow Physics and Control Branch NASA Langley Research Center HamptonVA United States

ISBN: (纸本)9781624103933

Wind tunnel experiments were performed using active flow control on an unswept semispan wing model with a 30% chord trailing edge flap to aid in the selection of actuators for a planned high Reynolds number experiment. Two sweeping jet actuator sizes were investigated to determine the influence of actuator size on the active flow control system efficiency. Sweeping jet actuators with orifice sizes of 1 mm × 2 mm and 2 mm × 4 mm were selected because of the differences in actuator jet sweep angle. The parameters that were varied include actuator momentum, freestream velocity, and trailing edge flap deflection angle. Steady and unsteady pressure data, Particle Image Velocimetry data, and force and moment data were acquired to assess the performance of the two actuators. In addition to the wind tunnel experiments, benchtop studies of the actuators were performed to characterize the jets produced by each actuator. Benchtop investigations of the smaller actuator reveal that the jet exiting the actuator has a reduced sweep angle compared to published data for larger versions of this type of actuator. The larger actuator produces an oscillating jet that attaches to the external diffuser walls at low supply pressures and produces the expected sweep angles. The AFC results using the smaller actuators show that while the actuators can control flow separation, the selected spacing of 3.3 cm may be too large due to the reduced sweep angle. In comparison, the spacing for the larger actuators, 6.6 cm, appears to be optimal for the Mach numbers investigated. Particle Image Velocimetry results are presented and show how the wall jets produced by the actuators cause the flow to attach to the flap surface. © 2016, American Institute of Aeronautics and Astronautics Inc, AIAA. All Rights Reserved.

关键词： Actuators

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Measurements in a transitioning cone boundary layer at freestream mach 3.5 54

Measurements in a transitioning cone boundary layer at frees...

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54th AIAA Aerospace Sciences Meeting, 2016

作者： King, Rudolph A. Chou, Amanda Balakumar, Ponnampalam Owens, Lewis R. Kegerise, Michael A. Flow Physics and Control Branch NASA Langley Research Center HamptonVA23681 United States

ISBN: (数字)9781624103933

ISBN: (纸本)9781624103933

An experimental study was conducted in the Supersonic Low-Disturbance Tunnel to investigate naturally-occurring instabilities in a supersonic boundary layer on a 7° halfangle cone. All tests were conducted with a nominal freestream Mach number of M∞ = 3.5, total temperature of T0 = 299.8 K, and unit Reynolds numbers of Re∞ × 10−6 = 9.89, 13.85, 21.77, and 25.73 m−1. Instability measurements were acquired under noisy-flow and quiet-flow conditions. Measurements were made to document the freestream and the boundary-layer edge environment, to document the cone baseline flow, and to establish the stability characteristics of the transitioning flow. Pitot-pressure and hot-wire boundarylayer measurements were obtained using a model-integrated traverse system. All hotwire results were single-point measurements and were acquired with a sensor calibrated to mass flux. For the noisy-flow conditions, excellent agreement for the growth rates and mode shapes was achieved between the measured results and linear stability theory (LST). The corresponding N factor at transition from LST is N ~ 3.9. The stability measurements for the quiet-flow conditions were limited to the aft end of the cone. The most unstable first-mode instabilities as predicted by LST were successfully measured, but this unstable first mode was not the dominant instability measured in the boundary layer. Instead, the dominant instabilities were found to be the less-amplified, low-frequency disturbances predicted by linear stability theory, and these instabilities grew according to linear theory. These low-frequency unstable disturbances were initiated by freestream acoustic disturbances through a receptivity process that is believed to occur near the branch I locations of the cone. Under quiet-flow conditions, the boundary layer remained laminar up to the last measurement station for the largest Re∞, implying a transition N factor of N > 8.5. © 2016, American Institute of Aeronautics and Astronautics Inc, AIAA.

关键词： Stability

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Receptivity and forced response to acoustic disturbances in high-speed boundary layers 46

Receptivity and forced response to acoustic disturbances in ...

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46th AIAA Fluid Dynamics Conference, 2016

作者： Balakumar, P. King, Rudolph A. Chou, Amanda Owens, Lewis R. Kegerise, Michael A. Flow Physics and Control Branch NASA Langley Research Center HamptonVA23581 United States

ISBN: (数字)9781624104367

ISBN: (纸本)9781624104367

Supersonic boundary-layer receptivity to freestream acoustic disturbances is investigated by solving the Navier-Stokes equations for Mach 3.5 flow over a sharp flat plate and a 7-deg halfangle cone. The freestream disturbances are generated from a wavy wall placed at the nozzle wall. The freestream acoustic disturbances radiated by the wavy wall are obtained by solving the linearized Euler equations. The results for the flat plate show that instability modes are generated at all the incident angles ranging from zero to highly oblique. However, the receptivity coefficient decreases by about 20 times when the incident angle increases from zero to a highly oblique angle of 68 degrees. The results for the cone show that no instability modes are generated when the acoustic disturbances impinge the cone obliquely. The results show that the perturbations generated inside the boundary layer by the acoustic disturbances are the response of the boundary layer to the external forcing. The amplitude of the forced disturbances inside the boundary layer are about 2.5 times larger than the incoming field for zero azimuthal wavenumber and they are about 1.5 times for large azimuthal wavenumbers. © 2016, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.

关键词： Boundary layers

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