Shock wave/boundary layer interactions occur in a wide range of supersonic internal and external flows, and often these interactions are associated with turbulent boundary layer separation. The resulting separated flow is associated with large-scale, low-frequency unsteadiness whose cause has been the subject of much attention and debate. In particular, some researchers have concluded that the source of low-frequency motions is in the upstream boundary layer, whereas others have argued for a downstream instability as the driving mechanism. Owing to substantial recent activity, we are close to developing a comprehensive understanding, albeit only in simplified flow configurations. A plausible model is that the interaction responds as a dynamical system that is forced by external disturbances. The low-frequency dynamics seem to be adequately described by a recently proposed shear layer entrainment-recharge mechanism. Upstream boundary layer fluctuations seem to be an important source of disturbances, but the evidence suggests that their impact is reduced with increasing size of the separated flow.

Low-Frequency Unsteadiness of Shock Wave/Turbulent Boundary Layer Interactions

Progress in shock wave/boundary layer interactions

We present wall-modelled large-eddy simulations (WLES) of oblique shock waves interacting with the turbulent boundary layers (TBLs) (nominal and ) developed inside a duct with an almost-square cross-section ( ) to investigate three-dimensional effects imposed by the lateral confinement of the flow. Three increasing strengths of the incident shock are considered, for a constant Mach number of the incoming air stream , by varying the height (1.1, 3 and 5 mm) of a compression wedge located at a constant streamwise location that spans the top wall of the duct at a 20° angle. Simulation results are first validated with particle image velocimetry (PIV) experimental data obtained at several vertical planes (one near the centre of the duct and three near one of the sidewalls) for the 1.1 and 3 mm-high wedge cases. The instantaneous and time-averaged structure of the flow for the stronger-interaction case (5 mm-high wedge), which shows mean flow reversal, is then investigated. Additional spanwise-periodic simulations are performed to elucidate the influence of the sidewalls, and it is found that the structure and location of the shock system, as well as the size of the separation bubble, are significantly modified by the lateral confinement. A Mach stem at the first reflected interaction is present in the simulation with sidewalls, whereas a regular shock intersection results for the spanwise-periodic case. Low-frequency unsteadiness is observed in all interactions, being stronger for the secondary shock reflections of the shock train developed inside the duct. The downstream evolution of secondary turbulent flows developed near the corners of the duct as they traverse the shock system is also studied.

https://www.cambridge.org/core/services/aop-cambridge-core/content/view/S0022112014005059

Confinement effects in shock wave/turbulent boundary layer interactions through wall-modelled large-eddy simulations

This paper presents a description of the physical principles of aerodynamic power savings from boundary layer ingestion propulsion and a quantitative evaluation of the boundary layer ingestion bene...

Boundary layer ingestion propulsion benefit for transport aircraft

An investigation into parameters affecting separation in normal shock wave/boundary layer interactions (SBLIs) has been conducted. It has been shown that the effective aspect ratio of an experimental facility (defined as δ * /tunnel width) is a critical factor in determining when shock-induced separation will occur. Experiments examining M ∞ =1.4 and 1.5 normal shock waves in a wind tunnel with a small rectangular cross-section have been performed and show that a link exists between the extent of shock-induced separation on the tunnel centre-line and the size of corner-flow separations. In tests where the corner-flows were modified ahead of the shock (through suction and vortex generators), the extent of separation around the tunnel centre-line was seen to vary significantly. These observations are attributed to the way corner flows modify the three-dimensional shock-structure and the impact this has on the magnitude of the adverse pressure gradient experienced by the tunnel wall boundary layers.

/pdf/corner-effect-and-separation-in-transonic-channel-flows-1resijub2u.pdf

Corner effect and separation in transonic channel flows

Wind tunnel experiments were performed to quantify the aerodynamic benefit of boundary layer ingestion (BLI) for the D8 transport aircraft concept. Two powered 1:11 scale, 13.4 ft span models, in BLI and non-BLI versions, were tested at the NASA Langley 14⇥22 Foot Subsonic Wind Tunnel to directly compare their performance. The models share the same basic airframe and propulsor units for the most direct comparison. They are also fully tripped to make the measured BLI benefit results scalable to the full-size aircraft. The comparison metric is the propulsor power required to produce a given net stream-wise force on the entire aircraft. The results show that the model BLI propulsors require 6% less electrical power at the simulated cruise point. These experiments provide the first back-to-back assessment quantifying the aerodynamic benefits of BLI for a civil aircraft. The BLI benefit quoted is preliminary in nature because it is defined in terms of electrical power, but we are in the process of obtaining a value in terms of flow power and there is indication that the BLI saving will remain essentially the same.

/pdf/preliminary-experimental-assessment-of-the-boundary-layer-2aos67c3pb.pdf

Preliminary Experimental Assessment of the Boundary Layer Ingestion Benefit for the D8 Aircraft

This paper presents the experimental assessment of the aerodynamic benefit of boundary layer ingestion for an advanced design civil transport aircraft, the D8 “double bubble,” carried out from 2010...

https://arc.aiaa.org/doi/pdf/10.2514/1.J055755

Boundary Layer Ingestion Benefit of the D8 Transport Aircraft

To investigate whether vortex generators can be an effective form of passive flow control an experimental investigation has been conducted in a small-scale wind tunnel. With specific emphasis on supersonic inlet applications flow separation was initiated using a combined terminal shock wave and subsonic diffuser: a configuration that has been developed as a part of a program to produce a more inlet-relevant flowfield in a small-scale wind tunnel than previous studies. When flow control was initially introduced little overall flow improvement was obtained as the losses tended to be redistributed instead of removed. It became apparent that there existed a strong coupling between the center-span flow and the corner flows. As a consequence, only when flow control was applied to both the corner flows and center-span flow was a significant flow improvement obtained. When corner suction and center-span vortex generators were employed in tandem separation was much reduced and wall-pressure and stagnation pressure...

Shock wave/boundary-layer interaction control using a combination of vortex generators and bleed

This article reviews research into the potential of vortex generators to mitigate shock-induced separation. Studies ranging from those conducted in the early post-war era to those performed recently are discussed. On the basis of the investigations described in this report, it is clear that vortex generators can alleviate shock-induced boundary layer separation. Yet, it will be shown that their potential and efficiency varies considerably in practical applications. Much more success is reported in transonic test cases compared to separation induced in purely supersonic interactions. Under a variety of flow conditions, the best performance is achieved with vortex generators with a height of roughly half the boundary layer thickness and a shape similar to a swept vane. Notwithstanding this, vortex generator performance is not as consistent as it is in low-speed applications. Further work is required before vortex generators can be implemented into the design process for eliminating shock-induced separation on transonic wings and in supersonic inlets.

/pdf/a-review-of-the-use-of-vortex-generators-for-mitigating-3eb2zlxx8d.pdf

Neil Titchener

Papers

Corner effect and separation in transonic channel flows

Preliminary Experimental Assessment of the Boundary Layer Ingestion Benefit for the D8 Aircraft

Boundary Layer Ingestion Benefit of the D8 Transport Aircraft

Shock wave/boundary-layer interaction control using a combination of vortex generators and bleed

A review of the use of vortex generators for mitigating shock-induced separation