Session X24: Flow Control: Separation and High-Speed Flows

A Parametric Study of Active Flow Control on a Rectangular Multi-Stream Supersonic Jet Nozzle

Active control techniques are applied to a multi-stream rectangular supersonic nozzle. The nozzles unique geometry and operating conditions produce a highly three-dimensional turbulent field comprised of several canonical flows such as shear layers, shock-shear interactions, and supersonic mixing layers. The flow features a core (M = 1.6) flow that coalesces with a bypass (M = 1.0) stream behind a splitter plate and exits onto a deck plate on one side meant to simulate airframe integration. This interaction results in the emission of a high frequency tone that propagates throughout the flow field. Alleviation of this tone, as well as the near-field unsteadiness in the aft-deck region, and modification of the shock train are the areas sought to be addressed by the active control techniques. Active control to the system is implemented via steady blowing in the near-splitter plate region through an array of span-wise microjets. Several microjet configurations are tested experimentally varying the hole size, spacing, angle, and injection location. All the experiments were conducted within an anechoic chamber, providing the acquisition of simultaneous and synchronized pressure measurements in the far-field (via an array of microphones), in the near-field (via pressure transducers embedded within the aft deck plate), and time-resolved schlieren imaging at two downstream locations (providing measurements up to 3.5 diameters downstream). Spectral-Proper Orthogonal Decomposition techniques are applied to the schlieren image sets to quantify the spatio-temporal evolution of coherent structures for the different control configurations. This parametric study of active control configurations seeks to progress towards optimal control and understanding of the system as part of an extensive experimental (Syracuse University) and simulated (via LES at The Ohio State University) campaign.   

Evolution of Flow Control Techniques on Supersonic Multi-Stream Jet Flow

Flow control is performed on a Multi-Aperture Rectangular Single Expansion Ramp Nozzle (MARS) using an array of microjets perpendicular to the flow. This campaign focuses on how the microjet’s configuration, location, and characteristics influence the highly complex flow problem presented by a multi-stream rectangular jet nozzle. The jet nozzle mixes two main flows: a core flow at Mach 1.6, and a bypass flow at Mach 1. This effectively simulates a modern multi-stream jet nozzle. The campaign is the culmination of two prior campaigns, passive control built into the splitter plate separating the two main flows of the nozzle, and active control via a microjet array just downstream from the splitter plate. The goal of this campaign being to replicate, with active control, the passive control’s successes as effective mixing and noise mitigation. The experimental rig is housed within an anechoic chamber allowing for clear acoustic analysis. Near field pressure measurements (embedded into the aft-deck plate) will study the local unsteadiness at the nozzle exit. Far field pressure measurements will study the general directivity of the acoustic emissions. Streamwise and cross-stream PIV, in addition to streamwise Schlieren imaging will study each configurations impact on the shock structures.   

Fundamental study on plasma-assisted flow control at external nozzle of high-speed transport vehicle

This study explores the effect of dc plasma-assisted rapid flow control on the flow structure over 15-degree external nozzle surface connected at rear side of supersonic linear spike nozzles. An external nozzle is a part of rear fuselage surface structure that navigates exhaust gas from the engine. Typically, around a quarter of the total thrust is produced at the external nozzle surface and thereby flow control at the external nozzle area can have significant impact on vehicle's aerodynamic control and can be a feasible solution for rapid and effective aerodynamic control method. A major attention is focused on transient phenomena related to plasma-flow interaction. Experiments were conducted in supersonic wind tunnel at Mach numbers of 2 and 5. 0.2-second steady pulse plasma was generated at pairs of electrodes installed and flush-mounted in a crossflow direction in front of the expansion ramp model that represents the external nozzle. The transient flow structure and plasma behavior were visualized with schlieren method and high speed camera to clarify the flow and plasma interaction. The surface pressure distribution on the wall, especially at the front and rear side of electrodes and at the ramp, was measured with fast pressure transducers. DC plasma generated in the front, upstream side, of the external nozzle ramp shifts the shock position from the ranp to the electrode location, subsequently forming a small separation zone and resulting in significant pressure change behind electrodes. Pressure change was characterized as a function of flow/plasma parameters.   

Responses of a Two-Dimensional Multi-Stream Supersonic Jet Flow to External Perturbations

Resolvent and stability analyses are performed for a two-dimensional supersonic jet flow bounded by an aft-deck. The goal is to uncover the underlying physics of the base flow and to gain insights into flow responses of control-induced perturbations. The jet flow consists of a Mach 1.6 main stream and a Mach 1.0 bypass stream separated by a splitter plate (SP). A vortex shedding phenomenon induced by the mixing of the two streams is observed, and the shedding frequency corresponds to the far-field resonant tone observed in experiments. Using resolvent analysis, a range of frequencies with zero spanwise wavenumber is analyzed to examine the flow responses to external perturbations introduced by active control. At the shedding frequency, forcing modes are found near the SP trailing edge (SPTE) and response modes are in the SP shear layer. Therefore, steady sonic micro-jet actuation is placed around the SPTE at various injection angles with the control objective of suppressing the tone and also aft-deck surface loading. Actuation on the top surface of the SP can most effectively weaken shock strength, and actuation on the SPTE surface can significantly reduce surface loading. For optimal control cases, perturbation dynamics will be further examined to reveal the control mechanism.   

Bio-inspired Fractal Parapet to Mitigate Rooftop Suction over a Low-Rise Building in High Winds

The significant damage experienced by roofing components of low-rise buildings in high winds necessitates the enhancement of their aerodynamic performance. Roof failure is often initiated at the windward roof edges and corners, due to peak suctions induced by flow separation and rooftop vortices. Past studies have explored using parapets on the roof to mitigate such peak suctions. Among the various parapets, the continuous porous parapet shows the most effective mitigation effects by disrupting rooftop vortex structures.

This research used a biomimetic approach to identify wind resilient strategies in nature and apply them to design a novel parapet. Fractal patterns have been used to control turbulence scales, increase turbulence intensity and reduce wind velocity. A cross-grid fractal-based porous parapet was designed to attempt further reduction in rooftop suctions. Wind tunnel tests of a 1:6 scaled low-rise building model was conducted to examine  rooftop pressures of three configurations: no-parapet (control), porous parapet and fractal porous parapet. These configurations were tested at various wind directions to determine their effects on mitigating peak rooftop suctions. The results will provide insights into innovative flow control strategies by learning from nature.   

Experimental Investigation of Turbulent Boundary Layer Separation Control by 3D Printed Shark Skin Models

Turbulent boundary layer separation can be problematic in many engineering applications. However, nature may have a solution in the form of passively actuated micro-flaps inspired by shortfin mako shark scales, which have been proven to passively bristle under reversing flow conditions and control flow separation in DPIV experiments using real skin samples. An investigation of how these shark scales interact with reversing flow in the near-wall regions of the boundary layer is of interest to better understand the fluid-shark scale interactions. Previous studies investigated the passive bristling motion of the scales and determined that the responsive motion of the scales was vitally important to flow control. Moreover, the bristling height into the boundary layer appeared to be an important parameter for control, with scales reaching too far beyond the buffer layer having little to no effect on separation control compared to larger boundary layer thicknesses. Using a rotating cylinder above a flat plate in a water tunnel setup, an adverse pressure gradient was induced creating a separated region within a tripped turbulent boundary layer with approximate Reynolds numbers up to 8 x 10⁵. Three different 3D printed micro arrays for kinetic optimization (MAKO) were mounted into a plate to replicate low-resistance passive bristling angles of 50 degrees. The model scales were constructed with maximum bristling heights of 1.4 mm, 2 mm, and 2.8 mm (a y⁺ range of 11 – 31) to investigate how bristle height into the buffer layer affects separation control. This dynamically scaled low-speed flow study makes the boundary layer dynamics and shark scale motions more measurable while allowing for actuations heights of the scales to be within the buffer layer. Baseline studies document flow separation and reversing flow development in the presence of an adverse pressure gradient over a smooth plate for direct comparison. Results quantify the separation control and observe how bristle height affects flow control.   

Analysis of receptivity and sensitivity of laminar flow separation over three-dimensional tapered wings

The structural sensitivity and receptivity of flows over tapered wings are studied numerically to understand the effect of Reynolds number and planform geometry variation and ultimately inform flow control. Low aspect ratio swept and tapered NACA 0015 wings are considered at 200 ≤ Re ≤ 600. Direct and adjoint TriGlobal linear stability analysis is carried out to identify the wavemaker of the leading unstable mode; this is found to have a compact structure located inside the laminar separation bubble (LSB) with regions associated with the suction and pressure side shear layers. The receptivity to momentum forcing is found to be highest near the separation line on the suction side. Both regions follow the spanwise location of maximum recirculation of the LSB when the wing is swept and tapered. The combination of leading edge (LE) and trailing edge (TE) sweep angles, rather than the taper ratio itself, affects the spanwise position of both the wavemaker and receptivity fields. An increase of LE sweep moves these regions towards the wing tip, whereas the forward sweep of the TE has the opposite effect and causes movement towards the root. These insights from stability theory are then used to inform the positioning of jet actuators in experimental flow control investigations, and numerical results are found to be consistent with those of experiments under the same boundary conditions. The present results establish a theoretical basis for future studies on tapered wings at higher Reynolds numbers.   

Control of Turbulent Boundary Layer Separation by 3D Printed Dolphin Skin Models

A problem in many flow applications is boundary layer separation; this study investigates the passive separation control mechanism that could result from the micro-grooves found on dolphin skin. A water tunnel was used to grow a turbulent boundary layer over rigid 3D printed plastic models, and an adverse pressure gradient was created with a rotating cylinder to induce flow separation in the region of study. Each model consists of dynamically similar sinusoidal grooves in the streamwise direction inspired by dolphin skin. The groove period studied is 5 mm, and for this study, the groove amplitude is varied from 0.9 mm, corresponding to the dolphin skin, to 1.5 mm. The hypothesis is that the dolphin-inspired case will lead to maximal flow separation control with minimal skin friction drag penalty. Time-resolve digital particle image velocimetry (TR-DPIV) is used to track the development of the flow separation within the boundary layer and compared to the smooth plate (non-grooved) wall cases. The grooved surfaces form embedded vortices within the sinusoidal cavities which can lead to a partial slip condition effect in the wall vicinity. DPIV results quantify momentum adjacent to the grooved surface and the corresponding flow separation.   

Physics-based flow control for swept and tapered wings using steady blowing

Abstract: An experimental investigation of flow control by steady blowing over swept and tapered wings was performed in a series of wind tunnel experiments with a mean-chord Reynolds number of 247,500. Two wings were explored each with a semi-aspect ratio of 2 and taper ratio of 0.269. One wing had an unswept leading edge and a 30 degree swept forward trailing edge and one wing had a 30 degree swept back leading edge and an unswept trailing edge. Both wings were instrumented for steady blowing from seven equally spaced jets across the leading edge. Aerodynamic loads were collected for multiple combinations of jets at a blowing ratio of 1. Volumetric mean flow fields were collected using SPIV to understand the effects of the jets on the large-scale wake structures at an angle of attack of 18 degrees. For the wing with the swept forward trailing edge, a single jet at the third-span location was the most effective at improving lift to drag ratio and alleviating pitch break. SPIV measurements show that this jet acted as a virtual fence, blocking the formation of the tip-to-root inverted ram’s horn structure that has been previously observed to form on this type of wing. Flow control on several planforms will be presented.   

Experimental Investigations of Separation and Reattachment Transient Flows Controlled by a Plasma Actuator

This study presents experimental findings on the separation and reattachment transient flow phenomena over a NACA0015 airfoil wing utilizing a plasma actuator for flow control. The investigation examines flow behavior during the activation and deactivation of the control device, offering valuable insights for future feedback-based active flow control strategies, with the potential to enhance aerodynamic capabilities. The experiments were performed at a Reynolds number of 66,000 and an angle of attack of 13 degrees, representing leading-edge separation without control. The plasma actuator, operated at 8 kV voltage, 30 kHz base frequency, and burst frequencies ranging from 100 Hz to 600 Hz, was mounted on the wing's leading edge. The study employed particle image velocimetry to measure flow field velocities and eight piezoelectric pressure sensors to capture surface pressure data. The primary focus of this paper is on the first proper orthogonal decomposition mode of the transient flow velocity field, which is extensively analyzed and discussed. The results provide detailed insights into the flow separation and reattachment transient processes, including their flow structures and temporal evolution. One significant finding is the observed time asymmetry between the separation and reattachment transient processes, indicating potential opportunities to enhance actuator efficiency.