Towards Effective Flow Control and Mitigation of Shock Effects in Aeronautical Applications

To design the next-generation aircraft and engines, efficient numerical tools need to be developed. Shock wave/boundary layer interactions are indeed putting the current industrial methods to the test: high-order methods lack robustness while modeling assumptions in low-fidelity methods make them not accurate enough. This work presents the first steps toward improving turbulence modeling in harmonic methods for shock-induced separated flows in turbomachinery applications, using high-fidelity data. A high-order solver based on the flux reconstruction framework is employed for performing high-fidelity simulations. Robustness is ensured by an enhanced artificial viscosity, allowing to capture shocks while not damping turbulence. A canonical oblique shock wave/boundary layer is first investigated to validate the solver and the results are in excellent agreement with the abundant existing literature. Then, the periodically forced transonic flow over a bump is considered. A study of the sensitivity to the perturbation frequency is carried out and highlights different flow regimes. The performance of the harmonic method for the bump case is finally shown to be inferior compared to the unsteady results. As a future step, the high-fidelity data generated will help to reduce the gap between the two.

Shock-wave/turbulent boundary layer interactions (SBLIs) are a typical hallmark of high-speed aerodynamics. Common examples of SBLIs can be found both in external flows, such as transonic/supersonic airfoils, wing-body junctions, aircraft control surfaces, and in internal flows such as engine supersonic inlets, compressors and turbines [1]. More broadly, SBLIs occur whenever a shock wave encounters a turbulent boundary layer developing on a solid surface. The impact of a shock on a boundary layer typically results in substantial flow separation, which can lead to significant decreases in performance. This chapter concentrates on the numerical aspects of SBLIs. In the first part, we elucidate the numerical framework utilized to simulate these interactions using Direct Numerical Simulations (DNS), both in Cartesian and curvilinear coordinates. Special emphasis has been placed on the handling of convective terms, which have been reformulated into a convenient split form ensuring the discrete preservation of total kinetic energy. Following this, the chapter presents qualitative and quantitative outcomes from a classical time-reversibility test. Subsequently, it delves into a more practical scenario, detailing the results of a fully turbulent supersonic compression corner.

An innovative, flexible wall-pressure sensor array for unsteady flow conditions has been developed and evaluated in a turbulent shockwave-boundary layer interaction (SBLI) setup at Mach 2. Compared to the previous version, the new sensor’s flexibility makes it easier to fit on different surfaces, while offering enhanced durability and improved sensitivity. The array comprises 18 circular sensors, each with a diameter of \(3\,{\text {mm}}\), fabricated using screen printing techniques from a thin piezoelectric PVDF film (thickness: \(110\,\upmu {\text {m}}\)). Remarkably, this sensor array achieves excellent spatial resolution while minimizing flow interference, all at a fraction of the cost associated with traditional dynamic pressure transducers.To validate its performance, the sensor array underwent dynamic calibration using a ball-drop impact test device. Subsequently, it was rigorously tested in a supersonic wind tunnel, demonstrating strong agreement with reference measurements obtained using a state-of-the-art Kulite pressure sensor. The resulting premultiplied power spectral density \(f\cdot PSD\) distributions align closely with findings reported in existing literature. Notably, the low-frequency unsteadiness region beneath the separation shock foot (\(X^*=0\)) exhibits a Strouhal range of \(St=0.03-0.05\).

The need for a better understanding of the low-frequency unsteadiness observed in shock wave/boundary layer interactions (SBLI) has driven research in this area for several decades. While numerous studies have been conducted on interactions with a turbulent boundary layer, in the context of transitional SBLI, research is still in its early stages and the low-frequency unsteadiness and the mechanisms underlying its origin remain poorly defined. In this study, large eddy simulations (LES) are performed in a \(M=1.7\) transitional shock reflection with separation. The objective is to examine any unsteadiness and the underlying mechanism. Beginning with a thorough assessment of stability theory, which suggests that transition in low supersonic compressible flows arises from the breakdown of oblique unstable boundary layer modes, the following question is posed. Do non-linear couplings between these oblique unstable modes and low-frequency unsteadiness emerge, and to what extent? To investigate quadratic couplings, high-order diagnostic is required. The results demonstrate that the unstable modes of the boundary layer interact non-linearly. High-frequency modes cascade non-linearly towards higher frequencies, initiating the turbulent cascade process, and towards lower frequencies. The low-frequency quadratic coupling with the flow characteristics at the separation point is responsible for the unsteadiness.

Shock-wave boundary layer interaction is a commonly found flow phenomenon in transonic and supersonic aerodynamics. However, interactions involving laminar boundary layers have received relatively less interest, due to the complexity of transition. As a result, experiments were performed to study laminar boundary layers and their interaction with shock-waves. The Mach number was 1.65 and the unit Reynolds number was varied between 5.6 and 11 million \(\text {m}^{-1}\). Pitot probes and hot-wire anemometry were employed for flow measurements. Experiments of the transition process of a natural laminar boundary layer captured the modal growth mechanisms of the primary instability, and a new time scale was found in the latter stages of the transition process. A new multi-sensor hot-wire probe was developed to study this new time scale, which revealed strange physical properties. Experiments of transitional SBLIs were performed on a \(6^{\circ }\) and a \(10^{\circ }\) compression ramp. A new non-dimensional parameter was developed for scaling the strength of the imposed shock, that was able to reconcile the large scatter in a diverse collection of length scales of transitional interactions. Measurements of the boundary layer transitional mechanisms over the interaction showed an accelerated growth over the separated shear layer, but surprisingly the growth of sub-harmonic instabilities was bypassed at reattachment. Finally, low-frequency unsteadiness at separation was found at Strouhal number of 0.05, similar to other studies on transitional interactions. A possible link between the presence of non-linearities over the separated shear layer and the low-frequency unsteadiness was found.

This chapter examines the influence of hole diameter, porosity, thickness-to-diameter ratio, and stagger angle on the performance of porous bleed control in mitigating the negative effects of shock-wave/boundary-layer interactions. A detailed numerical study focuses on the control of an irregular shock reflection, or Mach reflection, where a separation bubble below the shock foot is present in the uncontrolled case. Implementing bleed control modifies the flow field significantly, with variations in bleed rates upstream and downstream of the shock because of the external flow characteristics. The findings indicate that smaller hole diameters enhance bleed efficiency and control effectiveness, while porosity levels and thickness-to-diameter ratios exhibit complex trends, with a medium thickness-to-diameter ratio and a stagger angle of \(45^{\circ }\) emerging as optimal configurations for effective shock-wave/boundary-layer control.

Turbulent oblique Shock Wave Boundary Layer Interactions (SBLIs) were investigated experimentally in two rectangular test section blow down-type supersonic wind tunnels, to examine three-dimensionality induced by the presence of side-walls, as well as the low frequency separation bubble breathing oscillation. Testing was performed at Mach 2.5 and 2, with incident shock deflection angles of \(8^{\circ }\) and \(12^{\circ }\) at the Cambridge University (UCAM) and TU Delft (TUD) supersonic wind tunnel facilities respectively. In the UCAM facility, Conical shaped artificial corner separation bodies were used to generate corner waves, similar to those produced by corner separations, and vary their location with respect to the primary interaction. This resulted in a wide range of separation geometries underneath the primary interaction. Correlations between the separation length and pressure rise through interaction along streamwise strips revealed a quasi-2D relationship. The separation length was primarily correlated with the pressure rise from separation to reattachment. A secondary relationship was observed between the separation length and the pressure rise induced upstream of the interaction by corner waves. Corner waves modify the pressure rise in the interaction and this can lead to a significant reduction/elimination of separation in some regions. This strong control authority of pressure waves on the separation length informed the design of shock control bumps. Separation-bubble-shaped shock control bumps were tested in both test facilities with the goal of reducing separation, and dampening the low frequency bubble breathing oscillation. It was shown that these bumps are capable of significantly reducing and even eliminating flow separation. They also significantly dampened/eliminated the low frequency oscillation.

In this experimental study the effect of three-dimensional shock control bumps (SCB) on oblique shock wave/boundary layer interactions is investigated as a passive control method. It aims to develop an understanding of the effect of such devices on the interaction structure by means of studying the influence of the position of the bump with respect of the shock-impingement location, as well as the effect of the bump geometry (more in particular, the ramp section of the bump). The experiments were conducted in the ST-15 wind-tunnel at the Delft University of Technology for fully developed turbulent boundary layer conditions with \(Re_{\theta }\) of \(21.8 \cdot 10^3\) and freestream Mach number of 2.0. The control effectiveness is assessed from the size of the separated flow region, as well as the downstream boundary layer velocity profile. For this, PIV is employed as the main diagnostic method to characterise the flow field. In addition to this, high-speed Schlieren and oil flow measurements were performed to asses the effect of the SCB on the overall interaction structure.

The present work investigates the ability of the Partially-Averaged Navier-Stokes (PANS) method to reproduce transonic buffet, occurring on airfoils and wings at transonic regime under specific flow conditions. The designed test case for this analysis is the OAT15A unswept wing at Mach number \(\text {M}_{\infty }=0.73\) and Reynolds number \(\text {Re}_c=3\times 10^6\). The three-dimensional flow is studied by accounting for the wind tunnel walls in the experiments of Jacquin et al. [1]. The computations on a large-span, confined configuration revealed a strong three-dimensionality of the flow both before and after the buffet onset. The comparison with unsteady Reynolds-averaged Navier Stokes (URANS) results showed the benefits of PANS in resolving flow unsteadiness at different flow resolutions, especially on affordable CFD grids, at limited additional cost.

This study investigates the Shock-Wave Boundary Layer Interaction (SBLI) around supercritical aerofoils and wings in the Reynolds number range of \((2, 4.5) \times 10^6\). Physical analysis of the transonic buffet and its interaction with the shear-layer and near wake unsteadiness has been carried out in detail, showing the strong inter-dependence of the SBLI with the downstream unsteadiness. The numerical simulations have been performed with the NSMB—Navier Stokes Multi-Block code, using Delayed Detached Eddy Simulation and Organised Eddy Simulation. Improvement in the aerodynamic forces prediction has been discussed by means of a stochastic forcing approach, based on POD low-energy modes, applied in the DDES simulations around the V2C wing of Dassault Aviation, considering a constant spanwise section. A significant improvement in the aerodynamic performances has been shown and analysed thanks to specific electroactive morphing concepts based on trailing-edge vibration and on travelling waves along the suction side of an A320 morphing prototype. A decrease in the drag reduction in the order of 7% and in lift increase in the order of 3% have been obtained.

The complex flow physics of transonic compressors is linked with strong adverse pressure gradients, shock-wave boundary layer interactions (SBLI) and high level of unsteadiness. These are exacerbated by transitional effects coming along with altitude excitation. This chapter presents implications towards designing more efficient and compact turbomachineries for aeronautics, putting the mechanisms shock wave-boundary layer interactions in the center. Within this scope, a new design solution is proposed, which aims to mitigate limiting effects of transonic speeds at altitude, still maintaining the performance across the operating envelope with a holistic approach. The laminar boundary layer at altitude incurs a multi-shock pattern in the flow with a larger separated region accompanied with high losses and low-frequency unsteadiness. High fidelity numerical approaches are employed to capture the physics of the phenomenon accurately. It is followed by a numerical validation study carried out on a canonical test case, namely the Sandia axisymmetric transonic bump. The captured physics over the transonic, spherical bump is applicable to the flow in transonic compressor passages. Therefore, it is aimed to evaluate the strength and weakness of various numerical methods and different modelling approaches in respect to investigations of shock-boundary layer interaction.

When an aircraft engine operates at a transonic regime, the Shock Wave Boundary Layer Interaction (SBLI) plays a significant role in aerodynamic performance and its effects are detrimental while operating at low Reynolds conditions. To investigate SBLI effects numerically and experimentally, a linear cascade 3-profile test section has been designed for the IMP-PAN transonic blow-down wind tunnel facility. A novel technique of inlet valves with perforated plates setup has been installed upstream of the test section to reduce the Reynolds number in the test section. Three different Reynolds cases (\(7.4 \times 10^5\), \(5.8 \times 10^5\), and \(2.8 \times 10^5\)) have been chosen for detailed flow structure comparison. The main focus of the research is on the suction side of the middle profile where the shock wave interacts with the boundary layer resulting in boundary layer separation. A detailed boundary layer investigation and the location of the separation bubble beneath the shock foot have been compared. The total pressure losses have been compared based on wake measurements downstream of the middle profile.

The flow through a transonic compressor cascade is inherently unsteady due to the shock-boundary layer interactions (SBLI) on the blade. Despite decades of research, few details are known about the mechanisms that cause such behavior. This chapter presents a multidisciplinary study aiming to elucidate these mechanisms and optimize flow control methods to mitigate their effects. For this purpose, the Transonic Cascade TEAMAero was first optimized and its performance was validated experimentally. The cycle of shock oscillation was then compared using advanced experimental measurement techniques and high-fidelity LES. This comparison revealed a continuous propagation of pressure waves from a point upstream of the trailing edge. The interaction of these waves with the main bow shock at different points of the cycle was then linked to the frequencies observed in the oscillation spectrum. A configuration of this cascade with two roughness patches was finally optimized using a novel procedure developed. The optimal configurations obtained show how the targeted design of these devices can simultaneously mitigate shock oscillations and improve performance. This chapter demonstrates how the combined application of advanced numerical and experimental techniques needs to be intensified as researchers search for a global theory of SBLI in compressor blades.

High-pressure ratios in transonic compressor rotors and fan blades pose challenges such as supersonic speeds and shock waves, which can result in boundary layer separation and potential performance issues. This study presents experimental research exploring the effect of surface roughness on shockwave-boundary layer interaction (SBLI). By employing various measurement techniques, the study provides a comprehensive overview of how surface roughness affects SBLI, enriching the dataset for internal flows across four specimens with different roughness parameters. The findings indicate that lower roughness surfaces lead to larger separation bubbles and more typical shock structures, whereas increased roughness results in smaller separation bubbles but higher flow instability.

At cruise altitude, low Reynolds numbers result in a laminar boundary layer on the suction side of a transonic fan blade, extending to the shockwave/boundary-layer interaction. For transitional SBLIs with significant shock-induced separation, a shock oscillation mechanism occurs, characterized by the growth and natural suppression of the upstream laminar section of the separation bubble. The authors utilize a combination of numerical and experimental techniques across various cases, including a canonical case, cascades, and a 3D fan, to investigate the phenomenon. To validate the dynamic mechanism observed in large eddy simulations, experiments using high-speed Schlieren, spark light sh dowgraphy and PIV were conducted. The characteristic length scale for the oscillation mechanism, based on the travel distance of the laminar separation shock, is a key finding. The mechanism existence strongly depends on free stream turbulence and the boundary layer state. Oscillation frequencies are much lower for the turbulent oncoming boundary layer compared to the laminar case, which shows a strong link between the large scale movement of the laminar separation shock, the separation bubble, and reflected shock movement. In contrast, the turbulent interaction shows significantly less reflected shock travel distance. Preliminary full span LES simulations corroborate the link of findings to the application.