Session F03: Shock-Boundary Layer Interaction

Investigating Near-Wall Structures in Swept SBLI with Fast-Response PSP

The global dynamic properties of swept Shock/Boundary-Layer Interactions (SBLI) have been historically difficult to study due to the highly three-dimensional flow topology and limitations of acquiring instantaneous pressure fluctuations using pointwise measurements with traditional sensors. A flowfield generated by the interaction of a Mach 2 turbulent boundary layer and an oblique shock induced by a sharp fin is investigated using polymer/ceramic Pressure-Sensitive Paint (PSP), where PSP is applied on the floor as well as the fin surface. This non-intrusive technique permits high-spatial-resolution pressure fields of the entire flowfield to be captured and is validated with in-situ unsteady pressure transducers at key locations in the interaction. These initial results show the presence of spatially coherent structures underneath the interaction that are strongly correlated with localized pressure fluctuations simultaneously recorded on the fin surface.   

Surface Pressure Field underneath a Sharp Fin Generated Shock Boundary Layer Interaction Using BinaryFIB Pressure Sensitive Paint

The single sharp fin-generated shockwave boundary layer interaction (SBLI) has been the subject of numerous fundamental aerodynamic studies based on its ubiquitous presence in the supersonic aircraft design. The information extracted from experimental studies of the complex three dimensional flowfield have traditionally had a limited spatial resolution, predominantly due in part to physical restrictions imposed by instrumentation hardware. The current investigation utilizes BinaryFIB Pressure Sensitive Paint (PSP), which has a significantly reduced temperature sensitivity, to acquire high fidelity, global steady pressure measurements on surfaces underneath this particular SBLI. The PSP, in conjunction with selectively placed traditional pressure ports connected to an electronic pressure scanner, is used to obtain the pressure field. Measurements were carried out to discern the effect of Reynolds number (Re/m ~ 15 x 10⁶ – 110 x 10⁶) and interaction strength over a range of supersonic Mach numbers (M_∞=2 – 4). Additionally, PSP will allow for the study of hard to instrument regions, particularly the inception region and the surface of the fin.   

Experimental Study of a Compliant Panel under a Mach 2 Compression Ramp Interaction

The dynamic response of a compliant panel under a shock wave/boundary layer interaction produced by a compression ramp is studied in a Mach 2 flow. The compliant panel, made of polycarbonate, is 2 mm thick and gives a fundamental-mode frequency of about 700 Hz. The compression ramp, located at the downstream end of the compliant panel, is double-sided with compression angles of 20 and 21 degrees. Side fences are used to maintain quasi-two dimensionality of the mean separated flow. The primary diagnostic techniques are high-speed stereo digital image correlation (DIC) and high-speed pressure sensitive paint (PSP). The DIC will give time resolved measurements of the displacement of the compliant panel in the vertical direction, and the PSP will give the surface pressure over the entire panel. The combination of pressure and displacement measurements will allow the characterization of the structural response of the panel due to the unsteady pressure loading caused by the shock-induced turbulent separated flow.   

Influence of air jet vortex generators on a supersonic shock wave / boundary layer interaction

Shock induced separation occurs in many air and space transportation applications. To alleviate its detrimental effects, we investigate the application of air-jet vortex generators for separation control. Small air jets induce longitudinal vortices in the boundary layer where they increase turbulent mixing. Consequently, the separation region decreases. We investigated a fully separated 24° compression ramp interaction at a flow Mach number of 2.5. Air jets are issued from circular orifices with a diameter of 0.1δ (incoming boundary layer thickness δ), inclined at an angle of 45° within the spanwise/wall-normal plane. The jet orifices are arranged in an equispaced line perpendicular to the main flow direction, and installed 8δ upstream of the ramp corner. The spanwise spacing was varied between 0.4 - 1.25δ and smaller spacings decrease the separation more effectively. The extent of the separation region was quantified from oil flow visualizations and we analyzed the control effect on the low-frequency shock unsteadiness based on sequences of focusing schlieren images. Mean and turbulent velocities are analyzed based on Particle Image Velocimetry measurements.
  

Direct Simulation of Fluid-Structure Interaction of Compression Ramp with an Embedded Compliant Panel

Sustained flight at hypersonic Mach numbers presents a unique challenge to robust vehicle design and control. Extreme aerothermal loading coupled with finite-rigidity, multifunctional structures can yield significant fluid-thermal-structural interaction (FTSI) in the skin panels and control surfaces of reusable hypersonic aircraft. Specifically, the excitation of a thermomechanically-compliant panel via external Shock-Wave Boundary-Later Interaction (SWBLI) presents an imperfectly-understood phenomenon with direct bearing on viable hypersonic vehicle design. The present talk outlines recent high-fidelity, multiphysics simulations of unsteady two-dimensional Mach 6 flow over a compression ramp featuring an embedded compliant panel. Unsteady surface-pressure loading generated by the SWBLI is compared between compliant and non-compliant configurations, and SWBLI-excited FSI is demonstrated. The low-frequency coupling between the corner separation region and the structure is investigated.