Boundary Layer Interaction Control Research Articles

Moving wall effect on normal shock wave–turbulent boundary layer interaction on an airfoil

Purpose This paper aims to describe a proposal for an innovative method of normal shock wave–turbulent boundary layer interaction (SBLI) and shock-induced separation control. Design/methodology/approach The concept is based on the introduction of a tangentially moving wall upstream of the shock wave and in the interaction region. The SBLI control mechanism may be implemented as a closed belt floating on an air cushion, sliding over two cylinders and forming the outer skin of the suction side of the airfoil. The presented exploratory numerical study is conducted with SPARC solver (steady 2D RANS). The effect of the moving wall is presented for the NACA 0012 airfoil operating in transonic conditions. Findings To assess the accuracy of obtained solutions, validation of the computational model is demonstrated against the experimental data of Harris, Ladson & Hill and Mineck & Hartwich (NASA Langley). The comparison is conducted not only for the reference (impermeable) but also for the perforated (permeable) surface NACA 0012 airfoils. Subsequent numerical analysis of SBLI control by moving wall confirms that for the selected velocity ratios, the method is able to improve the shock-upstream boundary layer and counteract flow separation, significantly increasing the airfoil aerodynamic performance. Originality/value The moving wall concept as a means of normal shock wave–turbulent boundary layer interaction and shock-induced separation control has been investigated in detail for the first time. The study quantified the necessary operational requirements of such a system and practicable aerodynamic efficiency gains and simultaneously revealed the considerable potential of this promising idea, stimulating a new direction for future investigations regarding SBLI control.

Study on the mechanism of shock wave and boundary layer interaction control using high-frequency pulsed arc discharge plasma

This paper studies the response characteristics of shock wave and boundary layer interaction (SWBLI) controlled by high-frequency pulsed arc discharge (PAD) in a Mach 2.5 flow. The dynamic evolution of SWBLI disturbed by arc plasma energy deposition was captured, and the controlling mechanism under different exciting power and frequency was explored. The results showed that the blast wave induced by PADs had a strong impact on SWBLI structures and distorted the separation shock wave. During the downstream propagation, the controlling gas bubbles (CGBs) delivered a continuous thermal excitation to the boundary layer and reached the maximum penetration depth near the semi-cylinder. The arc discharge in the SWBLI region induced larger energy deposition, which made the heating zone obtain the highest initial temperature and longest heating duration. Under the plasma condition of 1 × 1011 W/m3/15 kHz, both the upstream part of the shear layer and the foot portion of the reattachment shock wave were removed. When setting the excitation to 2.5 × 1010 W/m3/60 kHz, a thermal exciting surface of merged CGBs was formed and the separation shock wave was completely replaced by an equivalent compression-wave system. A better drag-reduction effect on the flow field would be produced by the actuator with an increased operating power or frequency, and a drag reduction rate of nearly 25.5% was achieved under the 2.5 × 1010 W/m3/60 kHz control condition.

Hypersonic flow control of shock wave/turbulent boundary layer interactions using magnetohydrodynamic plasma actuators

The effect of magnetohydrodynamic (MHD) plasma actuators on the control of hypersonic shock wave/turbulent boundary layer interactions is investigated here using Reynolds-averaged Navier-Stokes calculations with low magnetic Reynolds number approximation. A Mach 5 oblique shock/turbulent boundary layer interaction was adopted as the basic configuration in this numerical study in order to assess the effects of flow control using different combinations of magnetic field and plasma. Results show that just the thermal effect of plasma under experimental actuator parameters has no significant impact on the flow field and can therefore be neglected. On the basis of the relative position of control area and separation point, MHD control can be divided into four types and so effects and mechanisms might be different. Amongst these, D-type control leads to the largest reduction in separation length using magnetically-accelerated plasma inside an isobaric dead-air region. A novel parameter for predicting the shock wave/turbulent boundary layer interaction control based on Lorentz force acceleration is then proposed and the controllability of MHD plasma actuators under different MHD interaction parameters is studied. The results of this study will be insightful for the further design of MHD control in hypersonic vehicle inlets.

Morphing of an adaptive shock control bump using pressurized chambers

The dawn of research on shock and boundary layer interaction control dates back to the 1970s, when humped transonic aerofoils were first studied as a means to improve the performance of supercritical aerofoil technology at off-design conditions. Since then, shock control bumps have been found to be promising devices for such kind of flow control. They have a smearing effect on the shock wave structure achieved through isentropic pre-compression of the flow upstream of the main shock and can significantly lower wave drag without incurring unacceptable viscous losses. However, their performance is strongly dependent on a set of geometrical parameters which must be adjusted according to the ever-changing flight conditions. A concept for an adaptive shock control bump is therefore presented. The proposed actuation mechanism aims at a compact, lightweight and simple structure which could be integrated into the spoiler region of near-future aircraft without major design changes required. Numerical optimization of a simplified analytical model of the structure is used to investigate the shock control bump adaptation to various aerodynamic target shapes. Compromises between geometrical conformity and both structural and actuation related requirements are studied. Furthermore, an outlook is given on design issues related to three-dimensional effects on a finite span shock control bump.

Shock wave/turbulence boundary layer interaction control with the secondary recirculation jet in a supersonic flow

The separation induced by the shock wave/boundary layer interaction is not beneficial for the engine operation, and it should be mitigated by the active or passive control approaches. In the current study, the flow control induced by the secondary recirculation jet has been evaluated numerically, and this approach combines the boundary layer bleed and the boundary layer suction together. At the same time, the influences of the height and length of the recirculation channel, as well as the dimensions and positions of the bleed and suction slots, on the flow control efficiency have been analyzed in detail. The obtained results show that the potions of the bleed and suction slots have a great impact on the flow control efficiency, and the separation length would decrease when the suction slot is placed on the windward side of the separation zone, as well as the peak pressure along the bottom wall. The secondary shock wave system induced by the secondary recirculation jet is beneficial for the mixing augmentation in the supersonic flow. However, the dimensions of the bleed and suction slots only have a slight impact on the flow control efficiency, and these factors can be omitted in the design process of the shock wave/boundary layer interaction model.

Plasma Synthetic Jet Actuators for Active Flow Control

The plasma synthetic jet actuator (PSJA), also named as sparkjet actuator, is a special type of zero-net mass flux actuator, driven thermodynamically by pulsed arc/spark discharge. Compared to widely investigated mechanical synthetic jet actuators driven by vibrating diaphragms or oscillating pistons, PSJAs exhibit the unique capability of producing high-velocity (>300 m/s) pulsed jets at high frequency (>5 kHz), thus tailored for high-Reynolds-number high-speed flow control in aerospace engineering. This paper reviews the development of PSJA in the last 15 years, covering the major achievements in the actuator working physics (i.e., characterization in quiescent air) as well as flow control applications (i.e., interaction with external crossflow). Based on the extensive non-dimensional laws obtained in characterization studies, it becomes feasible to design an actuator under several performance constraints, based on first-principles. The peak jet velocity produced by this type of actuator scales approximately with the cubic root of the non-dimensional energy deposition, and the scaling factor is determined by the electro-mechanical efficiency of the actuator (O(0.1%–1%)). To boost the electro-mechanical efficiency, the energy losses in the gas heating phase and thermodynamic cycle process should be minimized by careful design of the discharge circuitry as well as the actuator geometry. Moreover, the limit working frequency of the actuator is set by the Helmholtz natural resonance frequency of the actuator cavity, which can be tuned by the cavity volume, exit orifice area and exit nozzle length. In contrast to the fruitful characterization studies, the application studies of PSJAs have progressed relatively slower, not only due to the inherent difficulties of performing advanced numerical simulations/measurements in high-Reynolds-number high-speed flow, but also related to the complexity of designing a reliable discharge circuit that can feed multiple actuators at high repetition rate. Notwithstanding these limitations, results from existing investigations are already sufficient to demonstrate the authority of plasma synthetic jets in shock wave boundary layer interaction control, jet noise mitigation and airfoil trailing-edge flow separation.

Shockwave—boundary layer interaction control by plasma aerodynamic actuation: An experimental investigation

The potential of controlling shockwave—boundary layer interactions (SWBLIs) in air by plasma aerodynamic actuation is demonstrated. Experiments are conducted in a Mach 3 in-draft air tunnel. The separation-inducing shock is generated with a diamond-shaped shockwave generator located on the wall opposite to the surface electrodes, and the flow properties are studied with schlieren imaging and static wall pressure probes. The measurements show that the separation phenomenon is weakened with the plasma aerodynamic actuation, which is observed to have significant control authority over the interaction. The main effect is the displacement of the reflected shock. Perturbations of incident and reflected oblique shocks interacting with the separation bubble in a rectangular cross section supersonic test section are produced by the plasma actuation. This interaction results in a reduction of the separation bubble size, as detected by phase-lock schlieren images. The measured static wall pressure also shows that the separation-inducing shock is restrained. Our results suggest that the boundary layer separation control through heating is the primary control mechanism.

Experimental investigation of the micro-ramp based shock wave and turbulent boundary layer interaction control

Using a nanoparticle-based planar laser-scattering technique and supersonic particle image velocimetry, we investigated the effects of micro-ramp control on incident shockwave and boundary-layer interaction (SWBLI) in a low-noise supersonic wind-tunnel with Mach number 2.7 and Reynolds number Rθ = 5845. High spatiotemporal resolution wake structures downstream of the micro-ramps were detected, while a complex evolution process containing a streamwise counter-rotating vortex pair and large-scale hairpin-like vortices with Strouhal number Stδ of about 0.5–0.65 was revealed. The large-scale structures could survive while passing through the SWBLI region. Reflected shockwaves are clearly seen to be distorted accompanied by high-frequency fluctuations. Micro-ramp applications have a distinct influence on flow patterns of the SWBLI field that vary depending on spanwise locations. Both the shock foot and separation line exhibit undulations corresponding with modifications of the velocity distribution of the incoming boundary layer. Moreover, by energizing parts of the boundary flow, the micro-ramp is able to dampen the separation.

Shockwave–turbulent boundary layer interaction control using magnetically driven surface discharges

This study demonstrates the potential for shockwave–turbulent boundary layer interaction control in air using low current DC constricted surface discharges forced by moderate strength magnetic fields. An analytical model describing the physics of magnetic field forced discharge interaction with boundary layer flow is developed and compared to experiments. Experiments are conducted in a Mach 2.6 indraft air tunnel with discharge currents up to 300 mA and magnetic field strengths up to 5 Tesla. Separation- and non-separation-inducing shocks are generated with diamond-shaped shockwave generators located on the wall opposite to the surface electrodes, and flow properties are measured with schlieren imaging, static wall pressure probes and acetone flow visualization. The effect of plasma control on boundary layer separation depends on the direction of the Lorentz force (j × B). It is observed that by using a Lorentz force that pushes the discharge upstream, separation can be induced or further strengthened even with discharge currents as low as 30 mA in a 3-Tesla magnetic field. If shock-induced separation is present, it is observed that by using Lorentz force that pushes the discharge downstream, separation can be suppressed, but this required higher currents, greater than 80 mA. Acetone planar laser scattering is used to image the flow structure in the test section and the reduction in the size of recirculation bubble and its elimination are observed experimentally as a function of actuation current and magnetic field strength.

Design of "Smart" flap actuators for swept shock wave/tubulent boundary layer interaction control

Design of "Smart" flap actuators for swept shock wave/tubulent boundary layer interaction control

Shock wave – boundary layer interaction control by wall ventilation

Negative effects of shock wave – boundary layer interaction are restricting the velocity of modern commercial airplanes. Therefore, a lot of research was focused on methods of its control in the last decade. The effectiveness of passive, hybrid and active methods have been investigated and compared with one another. The main results are presented in this paper. These three methods were investigated experimentally. Static pressure distribution along the wall and the cavity were measured. At one location downstream of interaction the stagnation pressure was measured with a Pitot probe above the wall on the whole height of the λ-foot. Comparison of these results allowed us to judge the effectiveness of the method. The Schlieren system and the Mach–Zehnder interferometer were used to visualise the flow field structure.

A Passive Control of Interaction of Condensation Shock Wave anc Boundary Layer(II)

There were appreciable progresses on the study of shock wave / boundary layer interaction control in the transonic flow without nonequilibrium condensation. But in general, the actual flows associated with those of the airfoil of high speed flight body, the cascade of steam turbine and so on accompany the nonequilibrium condensation, and under a certain circumstance condensation shock wave occurs. Condensation shock wave / boundary layer interaction control is quite different from that of case without condensation, because the droplets generated by the result of nonequilibrium condensation may clog the holes of the porous wall for control and the flow interaction mechanism between the droplets and the porous system is concerned in the flow with nonequilibrium condensation. In these connections, it is necessary to study the condensation shock wave / boundary layer interaction control by cavity in the flow accompanying nonequilibrium condensation with condensation shock wave. In the present study, experiments were made on a roof mounted half circular arc in an indraft type supersonic wind tunnel to evaluate the effects of the porosity, the porous wall area and the depth of cavity on the pressure distribution around condensation shock wave. It was found that the porosity of 12% which was larger than the case of without nonequilibrium condensation produced the largest reduction of pressure fluctuations in the vicinity of condensation shock wave. The results also showed that wider porous area, deeper cavity for the same porosity of 12% are more favourable passive effect than the cases of its opposite. opposite.