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AIAA Science and Technology Forum and Exposition (AIAA SciTech Forum)

作者： Jana, Amartya Sucheendran, Mahesh M. Indian Inst Technol Hyderabad 502285 Telangana India IIT Hyderabad Dept Mech & Aerosp Engn Hyderabad Telangana India

ISBN: (数字)9781624106095

ISBN: (纸本)9781624106095

Underwater supersonic gas jets are widely utilized in submerged propulsion systems such as missiles launched from submarines. Unlike the single phase supersonic jets (gas jet in air), supersonic jets which are submerged underwater can give rise to high pressure pulses moving back into the nozzle assembly. This study is focused on the numerical simulation of the initial transients of underwater supersonic air jets. The results from the numerical simulations are used to visualize the oscillation of pressure, Mach number and jet interface under varying operating parameters to envisage the physics of the flow field. Several aspects of the submerged jets are compared with those of the single phase jets to understand the major differences in the multiphase environment. Subsequently, the submerged jet will be studied more elaborately under varying operating parameters. Thereby, this work attempts to investigate and understand the different physical aspects of the development of the flow structure of a supersonic jet submerged in water.

关键词： Mach Numbers Total Variation Diminishing Propulsion System Submarines Incompressible flow Supersonic Nozzle Hydrostatic Pressure Navier Stokes Solver flow visualization techniques Coefficient of Viscosity

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Numerical Investigation of Aerodynamic Noise Generation by Porous-Coated Staggered Cylinders 28

Numerical Investigation of Aerodynamic Noise Generation by P...

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28th AIAA/CEAS Aeroacoustics Conference

作者： Sharma, Sparsh Geyer, Thomas F. BTU Cottbus Senftenberg Chair Numer Flow & Gas Dynam D-03046 Cottbus Germany Brandenburg Univ Technol Cottbus Senftenberg Tech Acoust Grp D-03050 Cottbus Germany German Aerosp Ctr DLR Inst Electrified Aero Engines D-03050 Cottbus Germany

ISBN: (数字)9781624106644

ISBN: (纸本)9781624106644

The flow field around two staggered cylinders has been studied extensively due to its importance in various engineering applications. Despite the great interest attracted by both numerical and experimental research, few studies have been published on the generation of noise by this configuration. In the present work, we investigate the flow field and noise generation in the far field of a staggered cylinder configuration in which the downstream cylinder is oriented at different angles to the horizontal axis. The angles are varied between 0 and 90 degrees. Numerical simulations are performed using the large-eddy simulation approach with the Smagorinsky subgrid length scale model. Ffowcs-Williams and Hawkings techniques are used to post-process the pressure fluctuations to predict the far-field acoustics. A corresponding detailed experimental analysis, carried out in an aeroacoustic free-jet wind tunnel, is used to validate the numerical method. The resulting sound pressure level spectra show the characteristic tonal features.

关键词： Aerodynamic Force Sound Pressure Level Large Eddy Simulation Wind Tunnel Tests Acoustic Measurement Karman Vortex Street flow Separation flow visualization techniques Porous Materials Tandem Cylinders

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Control of Shock-Induced Separation of a Turbulent Boundary Layer Using Air-Jet Vortex Generators

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AIAA JOURNAL 2021年第3期59卷 927-939页

作者： Ramaswamy, Deepak Prem Schreyer, Anne-Marie Rhein Westfal TH Aachen Chair Fluid Mech Wullnerstr 5a D-52062 Aachen Germany Rhein Westfal TH Aachen Inst Aerodynam Wullnerstr 5a D-52062 Aachen Germany

An experimental analysis was carried out to study the control effectiveness and turbulence behavior of an array of steady air-jet vortex generators on a shock-induced boundary-layer separation. Control was applied to a 24 degrees compression-ramp interaction at Mach 2.52 and Re-theta c = 8225. A parametric study varying the jet spacing revealed that an adequate interaction between the vortices is essential for effective control, and the most favorable control effect was achieved with a jet spacing of 0.76 delta. This case was analyzed in more detail using particle image velocimetry. The results reveal a strong corrugation of the separation zone due to the injection of the jets with a substantial reduction in its area along the entire spanwise extent of the air-jet array, compared with the uncontrolled case. Measurements of velocity fluctuations along the interaction region also show that the jet injection results in a reduction in turbulent intensity after reattachment, contributing to an overall reduction in amplification factor across the interaction region. The expected increase in turbulent mixing, which is essentially responsible for the control effect, was confirmed in the region in-between jet injection and separation shock. Farther away from the surface, the flow is not affected, which reduces potential drag penalties.

关键词： Vortex Generators Supersonic Boundary Layers Particle Image Velocimetry Vortices Von Karman Constant Freestream Velocity flow visualization techniques Trisonic Wind Tunnels Stress Distribution flow Conditions

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Influences of Steps on the Hypersonic Boundary-Layer Transition on a Cone

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AIAA JOURNAL 2021年第2期59卷 439-446页

作者： Xu, Xiwang Yi, Shihe Zhang, Feng Zhang, Bo Liu, Xiaolin Natl Univ Def Technol Coll Aerosp Sci & Engn Deya Rd 109 Changsha 410073 Peoples R China

The influences of steps on the hypersonic boundary-layer transition on a cone are investigated in a hypersonic quiet wind tunnel, using high-frequency pressure sensors and nano-tracer-based planar laser scattering (NPLS) techniques. The high-frequency surface pressure fluctuations are analyzed using power spectral density analysis, band-pass filtering, and cross-correlation calculation. Quantitative analysis of the NPLS images is carried out based on the grayscale information and time interval. The results demonstrate that the second-mode wave amplitude increases initially and then drops along the streamwise, whereas the characteristic frequency decreases monotonously. However, the second-mode wave amplitude in the 0.4 mm (0.5 delta(s)) and 0.6 mm (0.75 delta(s)) forward-facing step models is constantly rising, and the obvious second-mode wave first appears closer to the downstream than the smooth cone, which suggests that the forward-facing steps could suppress the second-mode disturbances and shifts the transition downstream. Nonetheless, there is only a clear promotion effect when the height of the forward-facing step is 1.6 mm (2 delta(s)). In addition, the backward-facing step could destabilize the boundary layer and shift the transition upstream, and this effect is enhanced by increasing the step height.

关键词： Laminar Turbulent Transition Power Spectral Density Hypersonic Wind Tunnels Band Pass Filter Pressure Sensors Supersonic flow flow visualization techniques Angle of Attack Hot Wire Anemometry Charge Coupled Device

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Response Evaluation of cntTSP for Detection of Dynamic Boundary Layer Transition in Low-Speed Wind Tunnel

Response Evaluation of cntTSP for Detection of Dynamic Bound...

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AIAA Science and Technology Forum and Exposition (AIAA SciTech Forum)

作者： Ikami, Tsubasa Fujita, Koji Nagai, Hiroki Yorita, Daisuke Tohoku Univ Grad Sch Engn Dept Aerosp Engn Sendai Miyagi 9808577 Japan Tohoku Univ Inst Fluid Sci Sendai Miyagi 9808577 Japan German Aerosp Ctr DLR Inst Aerodynam & Flow Technol Dept Expt Methods D-37073 Gottingen Germany

ISBN: (数字)9781624106095

ISBN: (纸本)9781624106095

cntTSP is a flow visualization technique that uses the carbon nano-tube (CNT) as a thin-film heater for Temperature-Sensitive Paint (TSP). The cntTSP technique can measure the movement of the boundary-layer transition on a moving wing model dynamically. This study conducted a wind tunnel test applying cntTSP to a wing model with a pitch-sweep motion. The detected boundary-layer transition position had a constant difference compared to the position in a pitch-fixed condition. Furthermore, we investigated the effect of the parameters of TSP thickness (5 and 40 mu m), sweep speed (0.25, 0.50, 1.00 degrees/s), and the with or without of a screen layer, and found that the tested parameters were not related to the difference of the transition position.

关键词： Boundary Layer Transition Low Speed Wind Tunnel Carbon Nanotubes flow visualization techniques Thin Films Temperature Sensitive Paint Lift Coefficient Turbulent flow Heat Transfer Coefficients Fluorescence

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Evaluation of CFD for Simulation of HighSupersonic Control-Surface Effectiveness

Evaluation of CFD for Simulation of HighSupersonic Control-S...

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AIAA Aviation Forum and Exposition

作者： Ross, James C. Denison, Marie F. Childs, Robert E. Garcia, Joseph A. Stremel, Paul M. Hawke, Veronica M. Spooner, Hannah R. Reed, Matthew A. Kleb, Bil Watkins, Anthony N. Danehy, Paul M. Burns, Ross A. Fahringer, Timothy W. Borg, Stephen E. Robinson, Philip E. NASA Ames Res Ctr Moffett Field CA 94035 USA NASA Langley Res Ctr Hampton VA 23681 USA NASA Johnson Space Ctr Houston TX 77058 USA

ISBN: (数字)9781624106101

ISBN: (纸本)9781624106101

As part of studying the ability of Computational Fluid Dynamics (CFD) to accurately model important flow physics in the high-supersonic Mach number range, control-surface effectiveness on an entry vehicle for Mars exploration was examined. As with several other important flow regimes studied under the CFD as Surrogate for Wind Tunnel Testing at High Supersonic Speeds Project, a combined CFD and wind-tunnel study of a proposed Mars landing configuration was undertaken. The testing was performed in high-speed test section of the NASA Langley Research Center's Unitary Plan Wind Tunnel. The CFD team was an integral part of the overall evaluation team throughout the model development and test planning process and performed pre-test computations predicting the results of the testing. For the CFD predictions of the model in the wind tunnel, the flow into the test section was imposed as a boundary condition. The imposed inflow was based on a previous flow characterization study and companion CFD simulating the flow from the settling chamber through the test section. This paper presents a description of the control-surface effectiveness testing and the resulting aerodynamic data. A summary of comparisons of the experimental and CFD results will also be presented as well as a cost comparison of the two approaches to generating data on this previously untested vehicle shape.

关键词： Computational Fluid Dynamics Supersonic Wind Tunnels Aerodynamic Coefficients NASA Langley Research Center flow Conditions Mars Exploration Supersonic Speed Angle of Attack Flap Deflection flow visualization techniques

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Investigating Iced Swept-Wing flowfields Using a Swept Flat Plate and a Discontinuous Backward-Facing Step

Investigating Iced Swept-Wing Flowfields Using a Swept Flat ...

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AIAA Aviation Forum and Exposition

作者： Yoshida, William T. Bragg, Michael B. Woodard, Brian S. Univ Washington William E Boeing Dept Aeronaut & Astronaut Guggenheim Hall Room 307Box 352400 Seattle WA 98195 USA Univ Washington Coll Engn Room 211Guggenheim HallBox 352400 Seattle WA 98195 USA Univ Washington William E Boeing Dept Aeronaut & Astronaut Room 211Guggenheim HallBox 352400 Seattle WA 98195 USA Univ Illinois Dept Aerosp Engn 306 Talbot Lab104 S Wright St Urbana IL 61801 USA

ISBN: (数字)9781624106101

ISBN: (纸本)9781624106101

An experimental study of a swept flat plate has been performed in the University of Washington 3' by 3' Low-Speed Wind Tunnel to improve the understanding of iced swept-wing flowfields. The swept flat plate model included two interchangeable backward-facing steps that were mounted to the full span of the plate leading edge, as well as an adjustable trailing-edge flap. Each of the steps was designed to recreate one of two key flowfields characteristic of swept wings with artificial ice shapes while reducing the geometric complexity relative to the swept wing model. The first was a spanwise-running leading-edge vortex produced by certain low-fidelity artificial ice shapes. This type of flowfield is referred to as Type I, and was produced using the solid backward-facing step. Characteristic streamwise-running streaks of oil identify the second flowfield in surface flow visualization. This flow pattern was referred to as Type II, and was recreated with a modified backward-facing step that included spanwise-periodic gaps. Configurations of the flat plate were tested at four trailing-edge flap settings at a Reynolds number based on step height of 5.03 x 10(4), which corresponds to a Mach number of 0.176. The experimental techniques used included fluorescent-oil surface flow visualization, surface pressure measurements, and five-hole pressure probe measurements. Two tentative categories of Type II flowfield features were identified. In the first category, referred to as "vortex generatortype," each Type II flowfield feature was correlated to a single discontinuity in the step. flow visualization showed that these are likely the result of counter-rotating vorticity that forms in wakes of each solid feature in the discontinuous step. In the second category, referred to as "shear layer instability-type," Type II features did not correspond to a single discontinuity. This behavior may be attributed to the discontinuities exciting an instability in the separated shear layer, or

关键词： Flap Deflection flow visualization techniques Vortices flow Separation Aerodynamic Testing Low Speed Wind Tunnel Vortex Generators Trailing Edge Flap Pressure Probes Separated flows

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Influence of interaction strength on separation control with air-jet vortex generators

Influence of interaction strength on separation control with...

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AIAA Science and Technology Forum and Exposition (AIAA SciTech Forum)

作者： Ramaswamy, Deepak Prem Schreyer, Anne-Marie Rhein Westfal TH Aachen Inst Aerodynam & Chair Fluid Mech Wtillnerstr 5a D-52062 Aachen Germany

ISBN: (数字)9781624106095

ISBN: (纸本)9781624106095

A promising flow control strategy to mitigate shock-induced flow separation is the application of air-jet vortex generators. In the present study, experiments were conducted to investigate the influence of shock strength on the separation control effectiveness of air-jet vortex generators using flow visualisation tools and 2C particle image velocimetry. Two strongly separated compression ramp interactions with ramp angles of 24> and 33> were chosen for this investigation, and control is applied using a single row of steady air-jet vortex generators. The results show the presence of a large, closed separation bubble and a substantial reduction in separation for both the compression-ramp interactions. The streamwise separation length, scaled with the interaction length showa slightly larger reduction for the 33> compression ramp interaction, indicating a marginally improved control effectiveness of AJVGs for the stronger compression-ramp interaction. An assessment of the turbulent profiles showed qualitative similarity between the two cases and a reduction in turbulent amplification across the interaction due to jet-injection, with the stronger 33> interaction showing higher turbulent amplification reduction than the 24 > interaction.

关键词： Vortex Generators Boundary Layer Thickness Particle Image Velocimetry flow visualization techniques Freestream Mach Number Stress Distribution Vortices Trisonic Wind Tunnels Kinematic Viscosity flow Conditions

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Performance effects of winglets with leading edge protuberances on a rectangular wing

Performance effects of winglets with leading edge protuberan...

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AIAA Science and Technology Forum and Exposition (AIAA SciTech Forum)

作者： Flores-Mezarina, Juan A. Bravo-Mosquera, Pedro D. Ribeiro, Daniel G. Ceron-Munoz, Hernan D. Univ Sao Paulo Sao Carlos Sch Engn Dept Aeronaut Engn BR-13563120 Sao Carlos SP Brazil

ISBN: (数字)9781624106095

ISBN: (纸本)9781624106095

The leading edge protuberances are inspired on the humpback whale pectoral flippers. They act as control flow devices that reduce the flow three dimensional effects on wings and increase their aerodynamic performance at high angles of attack. Although there is plenty research about their dynamics, research on their possible applications is still in emerging phase. In this study, their effects on a winglet performance are assessed. A rectangular wing, a smooth leading edge winglet and three winglets with different leading edge protuberances sizes were designed and 3D printed to carry out CFD simulations, wind tunnel measurements and flow visualization tests. It was shown that the winglets with leading edge protuberances perform similarly to the base one in terms of lift at most angles of attack. In terms of drag, the winglet with the largest protuberances reduced parasite drag while the other two slightly increased it. On the other hand, all of them reduced the induced drag slope inversely proportional to their protuberances' size. Thus, the winglet with the largest protuberances had a significant aerodynamic efficiency improvement at low angles of attack. The CFD velocity curl contours downstream the wing showed that it also reduced the wing tip vortices' intensity. The leading edge protuberances showed the potential to enhance the winglet's performance on a rectangular wing.

关键词： Rectangular Wing Aircraft Configurations Aerodynamic Performance Aerodynamic Coefficients Wingtip Vortices Wind Tunnel Tests flow visualization techniques Numerical Simulation Wingtip Device Aspect Ratio

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Stability-Based Transition Transport Modeling for Unstructured Computational Fluid Dynamics at Transonic flow Conditions

Stability-Based Transition Transport Modeling for Unstructur...

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AIAA Science and Technology Forum and Exposition (AIAA SciTech)

作者： Stroer, Philip Krimmelbein, Normann Krumbein, Andreas Grabe, Cornelia German Aerosp Ctr DLR Inst Aerodynam & Flow Technol C2A2S2E Ctr Comp Applicat Aerosp Sci & Engn Bunsenstr 10 D-37073 Gottingen Germany German Aerosp Ctr DLR Inst Aerodynam & Flow Technol C2A2S2E Lilienthalpl 7 D-38108 Braunschweig Germany

ISBN: (数字)9781624105951

ISBN: (纸本)9781624105951

In the present work the extension of a recently published transition modeling approach is proposed. In the modified variant of the approach compressibility effects up to a Mach number of 1.6 are taken into account using the compressible formulation of the AHD transition criterion. Moreover, the Gleyzes criterion for the estimation of the transition onset location inside laminar separation bubbles was incorporated into the approach providing a higher accuracy for separation-induced transition points. The extensions were validated calculating the flat plate, the CAST10-2 airfoil and the Eppler 387 airfoil including grid resolution studies. Moreover, details with regard to the implementation and application of the extensions are given.

关键词： Compressible flow Computational Fluid Dynamics Supercritical Airfoils Transition Modeling Laminar Turbulent Transition Compressibility Effect Angle of Attack Kinematic Viscosity flow visualization techniques Convection

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