Marlyn Y. Andino, John C. H. Lin, S. Roman, E. Graff, M. Gharib, E. Whalen, I. Wygnanski
Active flow control (AFC) subscale experiments were conducted at the Lucas Wind Tunnel of the California Institute of Technology. Tests were performed on a generic vertical tail model at low speeds. Fluidic oscillators were used at the trailing edge of the main element (vertical stabilizer) to redirect the flow over the rudder and delay or prevent flow separation. Side force increases in excess of 50% were achieved with a 2% momentum coefficient (C μ ) input. The results indicated that a collective C μ of about 1% could increase the side force by 30-50%. This result is achieved by reducing the spanwise flow on the swept back wings that contributes to early flow separation near their tips. These experiments provided the technical backdrop to test the full-scale Boeing 757 vertical tail model equipped with a fluidic oscillator system at the National Full-scale Aerodynamics Complex 40-by 80-foot Wind Tunnel, NASA Ames Research Center. The C μ is shown to be an important parameter for scaling a fluidic oscillator AFC system from subscale to full-scale wind tunnel tests. The results of these tests provided the required rationale to use a fluidic oscillator AFC configuration for a follow-on flight test on the Boeing 757 ecoDemonstrator.
{"title":"Active Flow Control on Vertical Tail Models.","authors":"Marlyn Y. Andino, John C. H. Lin, S. Roman, E. Graff, M. Gharib, E. Whalen, I. Wygnanski","doi":"10.2514/1.J057876","DOIUrl":"https://doi.org/10.2514/1.J057876","url":null,"abstract":"Active flow control (AFC) subscale experiments were conducted at the Lucas Wind Tunnel of the California Institute of Technology. Tests were performed on a generic vertical tail model at low speeds. Fluidic oscillators were used at the trailing edge of the main element (vertical stabilizer) to redirect the flow over the rudder and delay or prevent flow separation. Side force increases in excess of 50% were achieved with a 2% momentum coefficient (C μ ) input. The results indicated that a collective C μ of about 1% could increase the side force by 30-50%. This result is achieved by reducing the spanwise flow on the swept back wings that contributes to early flow separation near their tips. These experiments provided the technical backdrop to test the full-scale Boeing 757 vertical tail model equipped with a fluidic oscillator system at the National Full-scale Aerodynamics Complex 40-by 80-foot Wind Tunnel, NASA Ames Research Center. The C μ is shown to be an important parameter for scaling a fluidic oscillator AFC system from subscale to full-scale wind tunnel tests. The results of these tests provided the required rationale to use a fluidic oscillator AFC configuration for a follow-on flight test on the Boeing 757 ecoDemonstrator.","PeriodicalId":80384,"journal":{"name":"AIAA student journal. American Institute of Aeronautics and Astronautics","volume":"87 1","pages":"3322-3338"},"PeriodicalIF":0.0,"publicationDate":"2019-07-17","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"88114151","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Jenna L. Eppink, R. Wlezien, R. King, Meelan Choudhari
Experimental measurements were performed on a swept flat-plate model with an airfoil leading edge and imposed chordwise pressure gradient to determine the effects of a backward-facing step on transition in a low-speed stationary crossflow-dominated boundary layer. Detailed hot-wire measurements were performed for three step heights ranging from 36 to 49% of the boundary-layer thickness at the step and corresponding to subcritical, nearly critical, and critical cases. In general, the step had a small localized effect on the growth of the stationary crossflow vortex, whereas the unsteady disturbance amplitudes increased with increasing step height. Intermittent spikes in instantaneous velocity began to appear for the two larger step heights. A physical explanation was provided for the mechanism leading to transition and the sudden movement in the transition front due to the critical steps. The large localized velocity spikes, which ultimately led to an intermittent breakdown of the boundary layer, were the result of nonlinear interactions of the different types of unsteady instabilities with each other and with the stationary crossflow vortices. Thus, the unsteady disturbances played the most important role in transition, but the stationary crossflow vortices also had a significant role via the modulation and the increased amplitude of the unsteady disturbances.
{"title":"Influence of a Backward-Facing Step on Swept-Wing Boundary-Layer Transition.","authors":"Jenna L. Eppink, R. Wlezien, R. King, Meelan Choudhari","doi":"10.2514/1.J057437","DOIUrl":"https://doi.org/10.2514/1.J057437","url":null,"abstract":"Experimental measurements were performed on a swept flat-plate model with an airfoil leading edge and imposed chordwise pressure gradient to determine the effects of a backward-facing step on transition in a low-speed stationary crossflow-dominated boundary layer. Detailed hot-wire measurements were performed for three step heights ranging from 36 to 49% of the boundary-layer thickness at the step and corresponding to subcritical, nearly critical, and critical cases. In general, the step had a small localized effect on the growth of the stationary crossflow vortex, whereas the unsteady disturbance amplitudes increased with increasing step height. Intermittent spikes in instantaneous velocity began to appear for the two larger step heights. A physical explanation was provided for the mechanism leading to transition and the sudden movement in the transition front due to the critical steps. The large localized velocity spikes, which ultimately led to an intermittent breakdown of the boundary layer, were the result of nonlinear interactions of the different types of unsteady instabilities with each other and with the stationary crossflow vortices. Thus, the unsteady disturbances played the most important role in transition, but the stationary crossflow vortices also had a significant role via the modulation and the increased amplitude of the unsteady disturbances.","PeriodicalId":80384,"journal":{"name":"AIAA student journal. American Institute of Aeronautics and Astronautics","volume":"7 1","pages":"267-278"},"PeriodicalIF":0.0,"publicationDate":"2019-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"75322441","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Reduced-order modeling using the Volterra series approach has been successfully applied in the past decades to weakly nonlinear aerodynamic and aeroelastic systems. However, aspects regarding the identification of the kernels associated with the convolution integrals of Volterra series can profoundly affect the quality of the resulting reduced-order model (ROM). An alternative method for their identification based on artificial neural networks is evaluated in this work. This relation between the Volterra kernels and the internal parameters of a time-delay neural network is explored for the application in the reduced-order modeling of nonlinear unsteady aerodynamic loads. An impulse-type Volterra-based ROM is also under consideration for comparison. All aerodynamic data used for the construction of the reduced-order models are obtained from computational fluid dynamics (CFD) simulations of the NACA 0012 airfoil using the Euler equations. Prescribed inputs in pitch and in plunge degrees of freedom at different free-stream Mach numbers are used to evaluate the range of applicability of the obtained models. For weakly nonlinear test cases, the modeling performance of the neural network Volterra ROM was comparable to the impulse-type ROM. Additional accuracy and adequate modeling of stronger nonlinearities, however, could only be attained with the inclusion of the neural network kernels of higher-order in the Volterra ROM. A generic expression is derived for the kernel function of p th -order from the internal parameters of a time-delay neural network.
{"title":"Volterra Kernels Assessment via Time-Delay Neural Networks for Nonlinear Unsteady Aerodynamic Loading Identification.","authors":"N. C. G. de Paula, F. Marques, W. Silva","doi":"10.2514/1.J057229","DOIUrl":"https://doi.org/10.2514/1.J057229","url":null,"abstract":"Reduced-order modeling using the Volterra series approach has been successfully applied in the past decades to weakly nonlinear aerodynamic and aeroelastic systems. However, aspects regarding the identification of the kernels associated with the convolution integrals of Volterra series can profoundly affect the quality of the resulting reduced-order model (ROM). An alternative method for their identification based on artificial neural networks is evaluated in this work. This relation between the Volterra kernels and the internal parameters of a time-delay neural network is explored for the application in the reduced-order modeling of nonlinear unsteady aerodynamic loads. An impulse-type Volterra-based ROM is also under consideration for comparison. All aerodynamic data used for the construction of the reduced-order models are obtained from computational fluid dynamics (CFD) simulations of the NACA 0012 airfoil using the Euler equations. Prescribed inputs in pitch and in plunge degrees of freedom at different free-stream Mach numbers are used to evaluate the range of applicability of the obtained models. For weakly nonlinear test cases, the modeling performance of the neural network Volterra ROM was comparable to the impulse-type ROM. Additional accuracy and adequate modeling of stronger nonlinearities, however, could only be attained with the inclusion of the neural network kernels of higher-order in the Volterra ROM. A generic expression is derived for the kernel function of p th -order from the internal parameters of a time-delay neural network.","PeriodicalId":80384,"journal":{"name":"AIAA student journal. American Institute of Aeronautics and Astronautics","volume":"69 1","pages":"1725-1735"},"PeriodicalIF":0.0,"publicationDate":"2018-12-27","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"78186562","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
To study the azimuthal development of boundary-layer instabilities, a controlled, laser-generated perturbation was created in the freestream of the Boeing/U.S. Air Force Office of Scientific Research Mach 6 Quiet Tunnel. The freestream perturbation convected downstream in the wind tunnel to interact with a flared-cone model. The flared cone is a body of revolution bounded by a circular arc with a 3 m radius. Pressure transducers were used to measure a wave packet generated in the cone boundary layer by the freestream perturbation. Nine of these sensors formed three stations of azimuthal arrays and were used to determine the azimuthal variation of the wave packets in the boundary layer. The freestream laser-generated perturbation was positioned upstream of the model in three different configurations: along the centerline axis, offset from the centerline axis by 1.5 mm, and offset from the centerline axis by 3.0 mm. When the freestream perturbation was offset from the centerline of a flared cone with a 1.0 mm nose radius, a larger wave packet was generated on the side toward which the perturbation was offset. As a result, transition occurred earlier on that side. The offset perturbation did not have as large of an effect on the boundary layer of a nominally sharp flared cone.
{"title":"Azimuthal Variation of Instabilities Generated on a Flared Cone by Laser Perturbations.","authors":"A. Chou, S. Schneider","doi":"10.2514/1.J056419","DOIUrl":"https://doi.org/10.2514/1.J056419","url":null,"abstract":"To study the azimuthal development of boundary-layer instabilities, a controlled, laser-generated perturbation was created in the freestream of the Boeing/U.S. Air Force Office of Scientific Research Mach 6 Quiet Tunnel. The freestream perturbation convected downstream in the wind tunnel to interact with a flared-cone model. The flared cone is a body of revolution bounded by a circular arc with a 3 m radius. Pressure transducers were used to measure a wave packet generated in the cone boundary layer by the freestream perturbation. Nine of these sensors formed three stations of azimuthal arrays and were used to determine the azimuthal variation of the wave packets in the boundary layer. The freestream laser-generated perturbation was positioned upstream of the model in three different configurations: along the centerline axis, offset from the centerline axis by 1.5 mm, and offset from the centerline axis by 3.0 mm. When the freestream perturbation was offset from the centerline of a flared cone with a 1.0 mm nose radius, a larger wave packet was generated on the side toward which the perturbation was offset. As a result, transition occurred earlier on that side. The offset perturbation did not have as large of an effect on the boundary layer of a nominally sharp flared cone.","PeriodicalId":80384,"journal":{"name":"AIAA student journal. American Institute of Aeronautics and Astronautics","volume":"23 1","pages":"1867-1877"},"PeriodicalIF":0.0,"publicationDate":"2018-02-22","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"82948920","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
A parametric experimental study was performed with sweeping jet actuators (fluidic oscillators) to determine their effectiveness in controlling flow separation on an adverse pressure gradient ramp. Actuator parameters that were investigated include blowing coefficients, operation mode, pitch and spreading angles, streamwise location, and size. Surface pressure measurements and surface oilflow visualization were used to characterize the effects of these parameters on the actuator performance. 2D Particle Image Velocimetry measurements of the flow field over the ramp and hot-wire measurements of the actuator's jet flow were also obtained for selective cases. In addition, the sweeping jet actuators were compared to other well-known flow control techniques such as micro-vortex generators, steady blowing, and steady vortex-generating jets. The results confirm that the sweeping jet actuators are more effective than steady blowing and steady vortex-generating jets for this ramp configuration. The results also suggest that an actuator with a wider jet spreading (110 vs. 70 degrees) placed closer (2.3 vs. 7 boundary layer thickness upstream) to the flow separation location provides better performance. Different actuator sizes obtained by scaling down the actuator geometry produced different jet spreading. Scaling down the actuator (based on the throat dimensions) from 6.35 × 3.18 mm to 3.81 × 1.9 mm resulted in similar flow control performance; however, scaling down the actuator further to 1.9 × 0.95 mm reduced the actuator efficiency by reducing the jet spreading considerably. The results of this study provide insight that can be used to design and select the optimal sweeping jet actuator configuration for flow control applications.
{"title":"Effects of Sweeping Jet Actuator Parameters on Flow Separation Control.","authors":"M. Koklu","doi":"10.2514/1.J055796","DOIUrl":"https://doi.org/10.2514/1.J055796","url":null,"abstract":"A parametric experimental study was performed with sweeping jet actuators (fluidic oscillators) to determine their effectiveness in controlling flow separation on an adverse pressure gradient ramp. Actuator parameters that were investigated include blowing coefficients, operation mode, pitch and spreading angles, streamwise location, and size. Surface pressure measurements and surface oilflow visualization were used to characterize the effects of these parameters on the actuator performance. 2D Particle Image Velocimetry measurements of the flow field over the ramp and hot-wire measurements of the actuator's jet flow were also obtained for selective cases. In addition, the sweeping jet actuators were compared to other well-known flow control techniques such as micro-vortex generators, steady blowing, and steady vortex-generating jets. The results confirm that the sweeping jet actuators are more effective than steady blowing and steady vortex-generating jets for this ramp configuration. The results also suggest that an actuator with a wider jet spreading (110 vs. 70 degrees) placed closer (2.3 vs. 7 boundary layer thickness upstream) to the flow separation location provides better performance. Different actuator sizes obtained by scaling down the actuator geometry produced different jet spreading. Scaling down the actuator (based on the throat dimensions) from 6.35 × 3.18 mm to 3.81 × 1.9 mm resulted in similar flow control performance; however, scaling down the actuator further to 1.9 × 0.95 mm reduced the actuator efficiency by reducing the jet spreading considerably. The results of this study provide insight that can be used to design and select the optimal sweeping jet actuator configuration for flow control applications.","PeriodicalId":80384,"journal":{"name":"AIAA student journal. American Institute of Aeronautics and Astronautics","volume":"18 1","pages":"100-110"},"PeriodicalIF":0.0,"publicationDate":"2018-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"90484957","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
This paper presents mechanisms to explain, as well as mathematics to model, time-averaged spatially resolved amplitude observations of number density and number density unsteadiness in a Mach 10 flow as it transitions from the freestream, through a bow-shock wave, and into the gas cap created by a blunt-body model. The primary driver for bow-shock unsteadiness is freestream unsteadiness or "tunnel noise." Primary unsteadiness is bow-shock oscillation. It scales spatially with the number density first derivative and is modeled using a sech2(z) term. Secondary weaker unsteadiness begins as freestream unsteadiness and increases linearly in direct proportion to the gas number density across the bow shock and into the gas cap. This is the well-known amplification of the freestream turbulent kinetic energy mechanism and is modeled using a tanh(z) term. Total unsteadiness [fit using tanh(z) term + sech2(z) term] is expressed as the number density standard deviation and modeled as a linear combination of these two independent, simultaneous, and nonlinear unsteadiness mechanisms. Relationships between mechanism coefficients and various flowfield and wind-tunnel parameters are discussed. For example, bow-shock and gas cap oscillation amplitudes are linearly correlated with stagnation pressure and, by deduction, freestream unsteadiness.
{"title":"Mach 10 Bow-Shock Unsteadiness Modeled by Linear Combination of Two Mechanisms.","authors":"R. Jeffrey Balla","doi":"10.2514/1.J055944","DOIUrl":"https://doi.org/10.2514/1.J055944","url":null,"abstract":"This paper presents mechanisms to explain, as well as mathematics to model, time-averaged spatially resolved amplitude observations of number density and number density unsteadiness in a Mach 10 flow as it transitions from the freestream, through a bow-shock wave, and into the gas cap created by a blunt-body model. The primary driver for bow-shock unsteadiness is freestream unsteadiness or \"tunnel noise.\" Primary unsteadiness is bow-shock oscillation. It scales spatially with the number density first derivative and is modeled using a sech2(z) term. Secondary weaker unsteadiness begins as freestream unsteadiness and increases linearly in direct proportion to the gas number density across the bow shock and into the gas cap. This is the well-known amplification of the freestream turbulent kinetic energy mechanism and is modeled using a tanh(z) term. Total unsteadiness [fit using tanh(z) term + sech2(z) term] is expressed as the number density standard deviation and modeled as a linear combination of these two independent, simultaneous, and nonlinear unsteadiness mechanisms. Relationships between mechanism coefficients and various flowfield and wind-tunnel parameters are discussed. For example, bow-shock and gas cap oscillation amplitudes are linearly correlated with stagnation pressure and, by deduction, freestream unsteadiness.","PeriodicalId":80384,"journal":{"name":"AIAA student journal. American Institute of Aeronautics and Astronautics","volume":"94 1","pages":"4274-4285"},"PeriodicalIF":0.0,"publicationDate":"2017-10-31","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"88918518","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
The performance of two wall models based on Reynolds-averaged Navier-Stokes is compared in large-eddy simulation of a high Reynolds number separating and reattaching flow over the NASA wall-mounted hump. Wall modeling significantly improves flow prediction on a coarse grid where the large-eddy simulation with the no-slip wall boundary condition fails. Low-order statistics from the wall-modeled large-eddy simulation are in good agreement with the experiment. Wall-pressure fluctuations from the resolved-scale solution are in good agreement with the experiment, whereas wall shear-stress fluctuations modeled entirely through the wall models appear to be significantly underpredicted. Although the two wall models produce comparable results in the upstream attached flow region, the nonequilibrium wall model outperforms the equilibrium wall model in the separation bubble and recovery region where the key assumptions in the equilibrium model are shown to be invalid.
{"title":"Wall-Modeled Large-Eddy Simulation of a High Reynolds Number Separating and Reattaching Flow.","authors":"G. Park","doi":"10.2514/1.J055745","DOIUrl":"https://doi.org/10.2514/1.J055745","url":null,"abstract":"The performance of two wall models based on Reynolds-averaged Navier-Stokes is compared in large-eddy simulation of a high Reynolds number separating and reattaching flow over the NASA wall-mounted hump. Wall modeling significantly improves flow prediction on a coarse grid where the large-eddy simulation with the no-slip wall boundary condition fails. Low-order statistics from the wall-modeled large-eddy simulation are in good agreement with the experiment. Wall-pressure fluctuations from the resolved-scale solution are in good agreement with the experiment, whereas wall shear-stress fluctuations modeled entirely through the wall models appear to be significantly underpredicted. Although the two wall models produce comparable results in the upstream attached flow region, the nonequilibrium wall model outperforms the equilibrium wall model in the separation bubble and recovery region where the key assumptions in the equilibrium model are shown to be invalid.","PeriodicalId":80384,"journal":{"name":"AIAA student journal. American Institute of Aeronautics and Astronautics","volume":"51 1","pages":""},"PeriodicalIF":0.0,"publicationDate":"2017-08-02","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"76757367","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
This paper describes a novel computational-fluid-dynamics-based numerical solution procedure for effective simulation of aerothermoacoustics problems with application to aerospace vehicles. A finite element idealization is employed for both fluid and structure domains, which fully accounts for thermal effects. The accuracies of both the fluid and structure capabilities are verified with flight- and ground-test data. A time integration of the structural equations of motion, with the governing flow equations, is conducted for the computation of the unsteady aerodynamic forces, which uses a transpiration boundary condition at the surface nodal points in lieu of the updating of the fluid mesh. Two example problems are presented herein to that effect. The first one relates to a cantilever wing with a NACA 0012 airfoil. The solution results demonstrate the effect of temperature loading that causes a significant increase in acoustic response. A second example, the hypersonic X-43 vehicle, is also analyzed; and relevant results are presented. The common finite element-based aerothermoelastic-acoustics simulation process, its applicability to the efficient and routine solution of complex practical problems, the employment of the effective transpiration boundary condition in the computational fluid dynamics solution, and the development and public domain distribution of an associated code are unique features of this paper.
{"title":"Aerothermoelastic-Acoustics Simulation of Flight Vehicles.","authors":"K. Gupta, S. Choi, S. Lung, A. Ibrahim","doi":"10.2514/1.J055088","DOIUrl":"https://doi.org/10.2514/1.J055088","url":null,"abstract":"This paper describes a novel computational-fluid-dynamics-based numerical solution procedure for effective simulation of aerothermoacoustics problems with application to aerospace vehicles. A finite element idealization is employed for both fluid and structure domains, which fully accounts for thermal effects. The accuracies of both the fluid and structure capabilities are verified with flight- and ground-test data. A time integration of the structural equations of motion, with the governing flow equations, is conducted for the computation of the unsteady aerodynamic forces, which uses a transpiration boundary condition at the surface nodal points in lieu of the updating of the fluid mesh. Two example problems are presented herein to that effect. The first one relates to a cantilever wing with a NACA 0012 airfoil. The solution results demonstrate the effect of temperature loading that causes a significant increase in acoustic response. A second example, the hypersonic X-43 vehicle, is also analyzed; and relevant results are presented. The common finite element-based aerothermoelastic-acoustics simulation process, its applicability to the efficient and routine solution of complex practical problems, the employment of the effective transpiration boundary condition in the computational fluid dynamics solution, and the development and public domain distribution of an associated code are unique features of this paper.","PeriodicalId":80384,"journal":{"name":"AIAA student journal. American Institute of Aeronautics and Astronautics","volume":"36 1","pages":"49-56"},"PeriodicalIF":0.0,"publicationDate":"2016-06-16","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"86122062","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Jenna L. Eppink, R. Wlezien, R. King, Meelan Choudhari
A swept flat plate model with an imposed pressure gradient was experimentally investigated in a low-speed flow to determine the effect of a backward-facing step on transition in a stationary crossflow-dominated flow. Detailed hotwire measurements of boundary-layer flow were performed to investigate the upstream shift in transition due to a step height of 49% of the local unperturbed boundary-layer thickness. Increasing the initial stationary crossflow amplitude caused an upstream movement of the transition front for the backward-facing step case. The step caused a local increase in the growth of the stationary crossflow instabilities, but the stationary crossflow amplitude at transition was sufficiently low (<0.04Ue ) so that stationary crossflow was not solely responsible for transition. The unsteady velocity spectra downstream of the step were rich with unsteady disturbances in the 80- to 1500-Hz range. Three distinct families of disturbances were identified based on phase speed and wave angle, namely, a highly oblique disturbance (possibly traveling-crossflow-like), a Tollmien-Schlichting-wave-like disturbance, and a shear-layer instability. The stationary crossflow disturbances caused a modulation of the unsteady disturbances, resulting in spatially concentrated peaks in unsteady disturbance amplitude. This modulation of the unsteady disturbances is believed to be the reason for the upstream movement of the transition front with increasing stationary crossflow amplitude.
{"title":"Interaction of a Backward-Facing Step and Crossflow Instabilities in Boundary-Layer Transition.","authors":"Jenna L. Eppink, R. Wlezien, R. King, Meelan Choudhari","doi":"10.2514/1.J056267","DOIUrl":"https://doi.org/10.2514/1.J056267","url":null,"abstract":"A swept flat plate model with an imposed pressure gradient was experimentally investigated in a low-speed flow to determine the effect of a backward-facing step on transition in a stationary crossflow-dominated flow. Detailed hotwire measurements of boundary-layer flow were performed to investigate the upstream shift in transition due to a step height of 49% of the local unperturbed boundary-layer thickness. Increasing the initial stationary crossflow amplitude caused an upstream movement of the transition front for the backward-facing step case. The step caused a local increase in the growth of the stationary crossflow instabilities, but the stationary crossflow amplitude at transition was sufficiently low (<0.04Ue ) so that stationary crossflow was not solely responsible for transition. The unsteady velocity spectra downstream of the step were rich with unsteady disturbances in the 80- to 1500-Hz range. Three distinct families of disturbances were identified based on phase speed and wave angle, namely, a highly oblique disturbance (possibly traveling-crossflow-like), a Tollmien-Schlichting-wave-like disturbance, and a shear-layer instability. The stationary crossflow disturbances caused a modulation of the unsteady disturbances, resulting in spatially concentrated peaks in unsteady disturbance amplitude. This modulation of the unsteady disturbances is believed to be the reason for the upstream movement of the transition front with increasing stationary crossflow amplitude.","PeriodicalId":80384,"journal":{"name":"AIAA student journal. American Institute of Aeronautics and Astronautics","volume":"97 10 1","pages":"497-509"},"PeriodicalIF":0.0,"publicationDate":"2015-01-05","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"83462875","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
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