Session Q21: Compressible Flows II

Investigation of Dynamic Shock-Vortex Interactions in Compressible Low Reynolds Number Flows on 3D Wings

As the Martian environment combines very low atmospheric density with a lower than Earth speed of sound, rotorcrafts such as NASA’s Ingenuity craft operate in a unique compressible low Reynolds Number (Re) flow regime. Previous research conducted by the authors showcased a new phenomenon called dynamic shock-vortex interactions in which vortices created during dynamic stall events interacted with shocks formed during forced oscillation tests. The qualitative behavior of the phenomenon was characterized in the aforementioned study along with the characterization of the phenomenon with respect to Mach and reduced frequency variations for a NACA 0012 airfoil through the use of unsteady RANS simulations. The presented work will expand on the authors’ previous work be extending the analysis of these dynamic shock vortex interactions to 3D straight wings using LES methods. The span wise evolution of these dynamic shock-vortex interactions as well as the determination of the driving mechanisms involved in the formulation, transportation, and dissipation of these interactions will be of main focus.      

Investigation of Transitional SBLIs using Plasma Based Disturbances

Plasma-based disturbances are used to generate second-mode instability wave

packets in a Mach 5 laminar boundary layer (BL) and in the presence of a

cylinder induced shock boundary layer interaction (SBLI). The experiments

are performed in an indraft wind tunnel at fixed Re_u = 2.7 x 10⁶ m^-1. Local-

ized Arc Filament Plasma Actuators (LAFPAs) are first characterized in a

vacuum chamber to determine the effect of ambient pressure on the plasma

regime and the electrical energy deposition. Linear Stability Theory (LST)

predicts that second-mode frequency content varies between 20 kHz to 30 kHz

over the length of the wind tunnel test section. Synchronized wall pressure

(PCB) measurements and phase-locked Z-type schlieren imaging are used to

characterize the instability waves created by actuation. Frequency and wave-

length of the observed second mode wave packets match well with estimates

from LST. In the presence of an SBLI, a decrease in the frequency and an

increase in the associated wavelength is observed. Future work will address

the introduction of stronger forcing to study non-linear wave packets as well

as extend the scope of measurements and data analysis techniques relevant

to transitional SBLIs.   

Numerical study of the effect of fluid-structural coupling on shock turbulent boundary layer interaction

The effect of panel flexibility on shock turbulent-boundary-layer interaction (STBLI) is analyzed for an oblique shock impinging on an elastic fixed-fixed panel using high fidelity numerical simulations. The FSI solver couples a body-fitted, wall-modelled LES solver, a non linear dynamic finite element solid solver, and a spring-system analogy mesh deformation solver. The simulations follow previous experiments by Daub et al (2016) in which the Mach number M and oblique shock angle θ were changed to vary the strength of STBLI. Three cases of increasing shock strength are considered. A weakly coupled case of M=4 and θ=15, a moderately coupled case of M=4 and θ=20, and a stronger coupled case at M=3 and θ=17 for which the panel deflection significantly influences the flow field. Results will be presented for simulations accounting for the full spanwise width of the elastic panel, as well as reduced width spanwise periodic simulations.   

Influence of air-jet stability on the flow-control effectiveness of air-jet vortex generators

Shock-wave / turbulent boundary layer interaction are complex flow-fields, commonly encountered in many air and space transportation applications. A common flow control technique to alleviate its adverse effects and to mitigate separation is the application of air-jet vortex generators (AJVGs). AJVGs induce a system of streamwise vortices downstream of jet injection and enhance the distribution of momentum in the boundary layer. Consequently, the boundary layer is more resistant to separation and the shock-induced separation region decreases. However, the nature of the jet-in-crossflow interaction and the jet-induced vortical structures can be affected by the stability of the injected jet, which can have a direct impact on flow-control effectiveness; even under similar cross-flow and jet stagnation conditions. With a combined experimental-numerical effort, we study the influence of the stability of the flow in the jet-injection pipe: we compare a stable and fully developed injection-pipe flow with a partially-developed case and discuss their influence on flow-control effectiveness for an array of air-jet vortex generators on a fully-separated 24^o compression ramp-interaction.

LESs of single spanwise-inclined jets in to a supersonic cross-flow were performed for two injection-pipe lengths – (1) shorter pipe with a jet-development length of 4.24 d_jet and a longer pipe with a length of 18.3 d_jet, to assess the influence of jet-stability on the jet-induced vortical structures. To quantify the effects of a jet-array on shock-induced separation, experiments were conducted on a 24^o compression-ramp model under similar jet and cross-flow conditions. Oil-flow visualisations and PIV at multiple planes were utilized to quantify the separation behavior and turbulent statistics under the two different jet-stability conditions.   

STORT flight experiment - Fin-Induced Shock Boundary Layer Interactions.

Fin-induced Shock Wave Boundary-Layer Interaction (SBLI) experiments have been conducted in a Mach 5 laminar boundary layer. The tests have been conducted in the Indraft Supersonic Wind Tunnel (ISWT) at the University of Arizona at an Re_θ of approximately 1100. The purpose of this research is to characterize a fin-induced SBLI in support of the DLR STORT Program. Experiments are performed on the planar wall of the wind tunnel to investigate the influence of the fin's finite geometry (including moderate bluntness and expansion fans). In addition, these tests support other experiments conducted on a hollow cylinder to assess the influence of base curvature. The state of the boundary layer and key features of the flow topology have been identified and measured. Oil flow visualization and high-speed schlieren have been used to characterize fin-induced separation and vortex structures. The investigation is further supported with unsteady wall-pressure measurements in key regions of the fin-induced flow. This research improves the understanding of complex fin-induced SBLIs and will be used to augment studies on the larger hollow-cylinder model and the flight geometry.   

Not Participating

An idealized shock/boundary-layer interaction problem that is two-dimensional in the mean but with a three-component mean velocity field is studied using large-eddy simulations. The problem isolates some aspects of three-dimensionality while avoiding others, and can thus offer insights into the differences between two- and three-dimensional problems. The addition of a crossflow increases the size of the separation bubble quite substantially when normalized by the boundary layer thickness or the momentum thickness, but less substantially when normalized by the displacement thickness. All cases studied are found to have a laminar separation bubble before the point of true separation that extends for at least one boundary layer thickness but remains fully within the viscous sublayer. To within the expected uncertainty, the frequency spectra of the wall pressure fluctuations do not change substantially with the addition of the crossflow, despite the increased size of the separation bubble.   

Mean Flow Similarity Behavior in 3D Shock Boundary Layer Interactions

Shock boundary layer interactions (SBLIs) are prevalent in 3D configurations associated with high-speed inlets and control surfaces. In comparison to canonical 2D interactions, SBLIs generated from swept geometries are poorly understood, particularly regarding the demarcation between various mean flow topologies. Historical studies on 3D SBLIs mean flow have identified cylindrical and conical similarities, with the transition between these states hypothesized to coincide with inviscid shock detachment at an adjusted Mach number, M_slip.  In the current study, 3D compressible N-S equations are used to investigate the mean flow similarity of laminar SBLIs induced by swept impinging oblique shocks and swept compression ramps on flat plates. Results show cases that exhibit quasi-conical features despite being well below the inviscid shock detachment limit, seemingly in contrast to the proposed hypothesis. Quasi-cylindrical features are observed only when the leading edge (LE) of the flat plate is swept. For turbulent flows, preliminary RANS calculations show behavior that is more consistent with past research. The role of the shock generator sweep, deflection, aspect ratio, M_slip, and LE sweep on the mean similarity are studied for both laminar and turbulent cases.    

Fin-Induced Shock Boundary Layer Interaction on Mach 5 cylinder.

Fin-induced Shock Boundary Layer Interactions (SBLIs) have been investigated on a hollow cylinder in the Mach 5 Ludwieg tube (LT5) at the University of Arizona. The study has been conducted in support of a DLR flight test as part of the STORT program and complements parallel experiments of similar SBLIs on planar surfaces. A half-scale model of the fin-base diameter flight geometry has been incorporated into the hollow cylinder model with variable total length between 400 and 1000 mm, such that the boundary layer state can be controlled between laminar/transitional/turbulent. Effects of surface curvature and the finite fin geometry have been assessed using high-speed schlieren, surface pressure transducers, and IR thermography. The inferred flow topology indicates separation is present around the base of the fin, and show the SBLI remains largely in Zone 1 (also referred to as the viscous inception region) due to the relatively large boundary layer that the fin experiences. Expansion waves from the fin further influence the nature of radial SBLI development. Results are contrasted to flat plate experiments of a similarly proportioned fin geometry and provide vital insights into understanding future data to be gathered from the flight experiment.   

Large-eddy simulations of turbulent compressible channels with a traveling wavy wall

Large-eddy simulations of turbulent compressible flow over a wavy channel and a channel with a backward traveling wave deforming wall are performed at a subsonic Mach number, M=0.5, and a supersonic Mach number, M=1.5. The Reynolds number (Re) based on the bulk flow velocity (U) and channel half-width (H) is 3000 and the channel walls are adiabatic. The wave steepness and wavelength are 0.07H and 0.5πH, respectively. At both Mach numbers, flow over the stationary wavy wall separates from the downhill of the wave and creates reverse flow near the trough of the wave. Due to the fluid compression in the supersonic channel, shock waves are generated over the peaks of the wave. When the wall undergoes a traveling wave motion with a wave speed of C=0.25πU, flow reattaches and the reverse flow zone is removed, which results in the decrease of the turbulent kinetic energy of the flow near the waves. Furthermore, the shockwaves of the supersonic wavy channel are removed by the traveling waves. The mechanism of shock removal is further investigated by performing simulations of a channel with wave speeds C=0.1πU and C=-0.1πU.   

Automated extraction of interface perturbations from explosively driven gas clouds in varying confinement

The gas cloud produced by a Winchester 209 primer is characterized by an explosive expansion of gases forming a jet outward from the primer. The gas cloud mixing produced by impinging this jet on an acrylic plate with a confined expansion volume is studied using high speed refractive imaging axially along the primer. Schlieren imaging allows the identification of the interface between the explosive product gases and the surrounding air. The interface location is extracted using automated morphological image processing techniques. The automated techniques are validated against a manual tracking of the interface, and good agreement is found. Using quarter symmetry, a dominant frequency and wavelength for the perturbations on the gas cloud interface is extracted using FFTs. The confinement of the expanding gas cloud is varied by increasing the space between acrylic plates. The evolution of the dominant wavelengths with confinement variation and time is tracked. Comparisons are drawn to existing experimental and theoretical literature on the evolution of the mixing region at the interface for both radius and wavelength with respect to time.