Boundary Layer Transition Control Research Articles

Experiments for control of boundary layer transition by plasma actuators

Plasma flow control technology is a new type of active flow control technology, and this paper carries out certain research on the basic principle of plasma flow control technology and its application on the wing model. The paper introduces the theory of plasma discharge and analyzes how plasma can control the flow field. A parallel dielectric barrier discharge plasma exciter is then designed and tested by using the Particle Image Velocimetry (PIV) technique to investigate the spatial flow field distributions of single-stage and three-stage actuators, and the mechanism of jet generation by the actuator is analyzed. Finally, the plasma actuator was utilized to control flow separation on a super-zero-boundary airfoil, and the oil droplet interferometric method was used to study the actuator’s role in controlling boundary layer transition. The experiment revealed that increasing the actuator discharge voltage can further reduce wall drag and more effectively delay the transition. Under the given conditions of a head-on angle of 6° and the incoming wind speed of 20 m/s, a two-stage actuator’s continuous discharge parameter was set to an output voltage of 12 kV and an output frequency of 4.7 kHz. The results showed that the transition position can be delayed by 13%, indicating the effectiveness of the actuator’s jet effect in enhancing the stability of the boundary layer flow.

The stabilizing effect of grooves on Görtler instability-induced boundary layer transition in hypersonic flow

Görtler vortex-induced hypersonic boundary layer transition controlled by grooves is investigated using direct numerical simulations and spatial bi-global stability analysis. In the simulations, Görtler vortices are excited by wall steady blowing and suction with spanwise wavelengths of 3 mm. It is found that when the wall is covered with grooves, the Görtler streaks keep more regular even at the end of the model. In addition, the skin friction coefficient is reduced efficiently. Furthermore, the wall-normal and spanwise velocity shear are both reduced, suppressing growths of secondary instabilities. In conclusion, grooves can delay Görtler vortex-induced transition by modifying the Görtler streaks structure and instability, which would shed light on hypersonic boundary layer transition control.

Development of a Micro Plasma Actuator Array for Near-Wall Body-Force Generation

In order to apply plasma actuator to boundary layer transition control, it is necessary to introduce a body force into the appropriate part of the boundary layer, especially near the wall surface. To develop a device that can introduce body force only near the wall surface, we fabricated micro plasma actuators arrays using photolithography, and investigated the generated body force and jet characteristics. Velocity distribution measurements using a hot-wire anemometer were used to analyze the jet characteristics, and the body force distribution of the conventional and micro plasma actuators was clarified.

Blind disturbance separation and identification in a transitional boundary layer using minimal sensing

A novel approach is presented for identifying disturbance sources in wall-bounded shear flows. The underlying approach models the flow state, as measured by sensors embedded in the flow, as a mixture of disturbance sources. The degenerate unmixing estimation technique is adopted as a blind source separation technique to recover the separate sources and their unknown mixing process. The efficiency of this approach stems from its ability to isolate any, a priori unknown, number of sources, using two sensors only. Furthermore, by adding a single additional sensor, the method is expanded to also determine the propagation velocity vector of each of the isolated sources, based on sensor readings from three sensors appropriately located in the flow field. Theoretical guidelines for locating the sensors are provided. The power of the method is demonstrated via computer simulations and wind-tunnel experiments. The numerical study considers disturbances comprising discrete Tollmien–Schlichting waves and wave packets. Linear stability theory is used to model source mixtures acquired by sensors placed in a Blasius boundary layer. The experimental study investigates the flow over a flat plate, with hot wires as sensors, and a loudspeaker and plasma actuators as source generators. Based on numerical and experimental demonstrations, it is believed that the new approach should prove useful in various applications, including active control of boundary layer transition from laminar to turbulent flow.

Transition Effect on the Flow Dynamics in a Compressor Blade Passage

An experimental investigation was carried out to study the effect of the boundary layer transition on the flow dynamics in the blade passage of a compressor cascade. A model of a turbine compressor passage was designed and assembled in a transonic wind tunnel for this purpose. Two different flow control methods were used in the experiment to induce the transition upstream of the shock wave, one concerning the microstep application and the other using distributed roughness strips. Two locations of spanwise microsteps for the transition trigger were chosen, one at the leading edge and the other closer to the shock wave position. The distributed roughness in the form of standard sandpaper strips with different heights was applied in three various locations on the blade upstream of the shock. The main objective of investigations is to present the influence of the laminar and turbulent shock wave-boundary layer interaction on the flow dynamics in a compressor fan passage, and the specific objective is to show how the parameters of particular transition control methods affect the flow dynamics in the investigated channel. The major challenge for this research was to minimize the disturbance caused by the microstep or roughness to the laminar boundary layer, while still ensuring a successful transition. Very interesting results were obtained in the flow control application for the boundary layer transition control, demonstrating a positive effect on the flow unsteadiness in changing the nature of the interaction.

Numerical study of the suction flow control of the supersonic boundary layer transition in a framework of gas-kinetic scheme

In this paper, we numerically study the supersonic boundary layer transition process and the suction flow control method by employing the DNS technology. By adopting the high resolution GKS method, the mechanism of the boundary transition and the statistical characteristics of the compressible turbulent boundary layer are investigated. Results indicate that the streamwise vortex and the three-dimensional shear layer with high instability are the main causes of the boundary layer transition. Suction control cases show that the laminar boundary layer suction can weaken the development of the unstable disturbance waves in the boundary layer, and delay the streamwise position of the local disturbance. Also, the space distributions of the vortex structure are weakened to delay the process of the transition.

Review of progress in some issues and engineering application of hypersonic boundary layer transition

A variety of complicated flow phenomena such as shock wave, boundary layer, flow separation, rarefied gas effects and real gas effects exist in the flows over hypersonic flight vehicles. The hypersonic boundary layer transition is not only a basic aerodynamics issue, but also a hot topic and challenging subject in the research of hypersonic flows. The accurate prediction and the effective control of boundary layer transition can achieve meticulous design of aerodynamic force and heating, thus improving the vehicle performance and enhancing the mission capability. In this paper, the research progress in typical instability characteristics commonly existed in the flight vehicles with complex configurations in engineering is reviewed, and the complicated issues of boundary layer transition in engineering practice which are urgent to be solved are proposed. Meanwhile, the engineering application direction of the research on hypersonic boundary layer transition is specified. Moreover, the investigations on flow control of hypersonic boundary layer are reviewed in order to realize the boundary layer control in the design of hypersonic vehicle in future to enhance the flight performance.

Transition effect on aerodynamic performance of compressor profile

PurposeThe purpose of this paper is to study experimentally the effect of transition and also the roughness height on the flow structure of the shock wave boundary layer interaction in the blades passage of a compressor cascade.Design/methodology/approachA model of a turbine compressor passage was designed and assembled in a transonic wind tunnel. In the experiment, the distributed roughness with different heights and locations was used to induce transition upstream of the shock wave.FindingsRecommendation regarding the roughness parameters for the application depends on what is more important as goal, whether the reduction of losses or unsteadiness. In case if more important are the losses reduction, a good choice for the roughness location seems to be the one close to the shock wave position.Research limitations/implicationsThe knowledge gained by this paper will enable the implementation of an effective laminar flow technology for engines in which the interaction of a laminar boundary layer with a shock wave takes place in the propulsion system and causes severe problems.Originality/valueThe paper focuses on the influence of the boundary layer transition induced by different roughness values and locations on aerodynamic performance of a compressor cascade. Very valuable results were obtained in the roughness application for the boundary layer transition control, demonstrating a positive effect in changing the nature of the interaction and also some negative influence in case of oversized roughness height, which cannot be found in the existing literature.

Assessment and parameterization of upstream shielding effect in quasi-roughness induced transition with direct numerical simulations

Boundary layer transition control is an essential technology for developing natural laminar flow airfoils, and one of the effective strategies is shielding effect that transition induced by large roughness can be delayed by low amplitude roughness around it. In this paper, we study the shielding effect brought by a two-dimensional strip placed upstream of a discrete quasi-roughness in transonic Blasius boundary layer flow with direct numerical simulations (DNS). With a comparison between cases with or without the shielding strip, we find that the strip can decrease the roughness Reynolds number Rek by lifting up the boundary layer, and reduce the stronger left vortex in wake structure which sheds from the top of the discrete quasi-roughness. Meanwhile, the strip enhances the weaker right vortex trailing from the horseshoe vortex. We also conduct a parametric study on strip height kup, and results show a trend that transition is gradually delayed in the left vortex, and meanwhile gets earlier in the right one with a kup increment in specific range. When kup exceeds a critical value, a reversal phenomenon is observed that transition happens much earlier, because the strip provides an excessively large disturbance which strengthens weak vortices that should have been dissipated rapidly. This work improves the understanding of shielding effect in transonic regime, and figures out the influence of shielding strip height to offer a useful reference for future studies and applications.

On the stochastic forcing of the boundary layer by plasma actuators

Turbulence generation by plasma actuator is one of the key barriers on the way of their application as virtual roughness elements in boundary layer transition control. Recent application of these devices for the crossflow instability control has shown that, along with the stationary disturbances creation, they perform a broadband unsteady forcing of the boundary layer, leading to the unwanted excitation of traveling modes. To study the physical mechanisms of this broadband forcing, characterization of the discharge-induced disturbances in a 2D subsonic boundary layer was performed. It has been shown that low-frequency velocity pulsations correlate with the discharge optical emission, i.e. they are caused by stochastic microdischarges dynamics in sDBD.

Experimental investigation of the natural and controlled disturbance development in a supersonic boundary layer on the swept wing

The results of investigation of the transverse flow inhomogeneity effect on the natural and controlled disturbance development in 3D supersonic boundary layer are considered. The transition curves were measured on a model of a smooth swept wing and on a swept wing with roughness elements of two configurations. The influence of periodic surface roughness on the growth of natural pulsations (without using a discharge) in the boundary layer is determined. In experiments with artificial disturbances it is found that the periodic modulation of mean flow can lead to stabilization of controlled disturbances in supersonic boundary layer on a swept wing. It was found out that the weak flow inhomogeneity can influence the effectiveness of the active/passive technology of supersonic boundary layer transition control, including the surface microroughness on a swept wing.

By-pass transition control with a DBD plasma actuator model coupled with a laminar kinetic energy turbulence model

The effect of a DBD plasma actuator on the by-pass boundary layer transition control is numerically investigated. A two-equation DBD plasma actuator model is coupled with a three-equation eddy-viscosity turbulence model, which adopts the laminar kinetic energy concept. The investigated test-cases concern zero and variable freestream pressure gradient transitional flows on a flat plate with a sharp leading edge, which belong to the ERCOFTAC experimental database. The experimental data describe the transition characteristics with no plasma actuation and hence, they are used to assess the turbulence model behaviour and as a reference point in order to quantify the coupling effect of the plasma actuator model with the laminar kinetic energy concept. The investigation is focused on the effect of the varying plasma actuator voltage on transitional and turbulent boundary layer characteristics, such as the Reynolds-stresses, the turbulent dissipation and the laminar kinetic energy distributions. The results show that with the DBD actuator activated, the turbulent quantities are suppressed, the transition onset location moves downstream and a boundary layer transition delay is observed.

By-pass transition control with a DBD plasma actuator model coupled with a laminar kinetic energy turbulence model

The effect of a DBD plasma actuator on the by-pass boundary layer transition control is numerically investigated. A two-equation DBD plasma actuator model is coupled with a three-equation eddy-viscosity turbulence model, which adopts the laminar kinetic energy concept. The investigated test-cases concern zero and variable freestream pressure gradient transitional flows on a flat plate with a sharp leading edge, which belong to the ERCOFTAC experimental database. The experimental data describe the transition characteristics with no plasma actuation and hence, they are used to assess the turbulence model behaviour and as a reference point in order to quantify the coupling effect of the plasma actuator model with the laminar kinetic energy concept. The investigation is focused on the effect of the varying plasma actuator voltage on transitional and turbulent boundary layer characteristics, such as the Reynolds-stresses, the turbulent dissipation and the laminar kinetic energy distributions. The results show that with the DBD actuator activated, the turbulent quantities are suppressed, the transition onset location moves downstream and a boundary layer transition delay is observed.

Roughness effect on shock wave boundary layer interaction area in compressor fan blades passage

The main objective of this research is to study experimentally the effect of transition on the flow structure of the shock wave boundary layer interaction in the blades passage of a compressor cascade. For this purpose a model of a turbine compressor passage was designed and assembled in a transonic wind tunnel. In the experiment the distributed roughness with different heights and locations was used to induce transition upstream of the shock wave. Three locations of distributed roughness in form of standard sandpaper strips were chosen to apply on the blade. The average roughness height comprised in a range of 8 to 16 μm. All together 8 different flow cases were investigated. The paper focuses on the influence of the boundary layer transition induced by different roughness values and locations on the flow pattern in the blades passage of a compressor cascade. Very promising results were obtained in the roughness application for the boundary layer transition control, demonstrating a positive effect in changing the nature of the interaction.

Measurement of corner separation zone response on a compression ramp to plasma actuation within the hypersonic boundary layer

This study was performed to characterize the dominant frequencies present in the boundary layer uptsream of and in the corner separation zone of a compression surface in Mach 4.5 flow and to determine a control effect of transient plasma actuation on the boundary layer. Schlieren imaging was used to distinguish the corner separation zone for 20°, 25°, and 30° compression ramps mounted on flat plates. Spectra of the natural disturbances present in the boundary layer and separation zone were gathered using a high-speed Shack-Hartmann wavefront sensor and surface mounted PCBTM pressure sensors while varying flow parameters by adjusting total pressure, temperature, and ramp angle. Shallow cavity discharge plasma actuators were used as a high-frequency localized thermal forcing mechanism of the boundary layer. The plasma effect was negligible for forcing frequencies (50 kHz) below the natural dominant frequency (~55-80 kHz). High frequency perturbations that can promote the transition to turbulence were amplified when the plasma forcing frequency (100 kHz) was higher than the natural dominant frequency (~55-80 kHz). This technique can potentially be used for active control of hypersonic boundary layer transition and the supersonic flow structure on the compression surface.

An overview of flow control activities at Dassault Aviation over the last 25 years

ABSTRACTIn France, the very first ideas on flow control were developed by Philippe Poisson-Quinton from the Office National d'Etudes et Recherches Aérospatiales (ONERA) in the 1950s. There was some renewal of this research topic in the early 1990s, first in the United States with scientists like Wygnanski and Gad-El Hak, and also in France at the initiative of Pierre Perrier from Dassault Aviation, who triggered a lot of research activities in this field both at ONERA and in the French National Centre for Scientific Research (CNRS) laboratories. The motivation was driven by the applications on Dassault Aviation military aircraft and Falcon business jets in order to contribute to the design, while facilitating performance optimisation and multi-disciplinary compromise. A few examples of flow control technologies, such as forebody vortex control, circulation control, flow separation control or boundary layer transition control using hybrid laminar flow control (HLFC), are presented to illustrate the applications and to explain the methodology used for the design of the flow control devices. The author also emphasises the current reaction of industry with respect to the integration of flow control technologies on an aircraft programme. The conclusion is related to the present status of the French research on this topic and to the next challenges to be addressed.

Passive Hypersonic Boundary Layer Transition Control Using an Ultrasonically Absorptive Coating With Random Microstructure: Computational Analysis Based on the Ultrasonic Absorption Properties of Carbon-Carbon

Previous investigations in the High Enthalpy Shock Tunnel Göttingen of the German Aerospace Center (DLR) show that carbon fiber reinforced carbon ceramic (C/C) surfaces can be utilized to damp hypersonic boundary layer instabilities resulting in a delay of boundary layer transition onset. Linear stability analyses were performed using the DLR stability code NOLOT, NOnLocal Transition analysis code. To adapt the boundary condition to account for the characteristics of porous C/C material, the ultrasonic absorption properties of C/C were investigated experimentally and theoretically. Therefore, a test rig was set up to directly measure the reflection coefficient in the frequency and pressure range corresponding to the test conditions in HEG. In this frame, the reflection of ultrasonic waves from flat plate test samples with different porous layer thicknesses was investigated and compared to an ideally reflecting surface.The obtained results were used to improved the boundary condition used for stability analysis above porous surfaces. The numerical results, using the original as well as the improved boundary condition, were compared with wind tunnel tests. These experiments were performed at Mach 7.5 and different unit Reynolds numbers. A 7∘ half-angle cone model with a nose radius of 2.5mm and a total length of 1077mm was used. One-third of the metallic model surface in circumferential direction was replaced by C/C ceramics. The comparison between numeric and experiments includes the investigations of the second modes, the damping of the these modes and the resulting transition shift.

Ultrasonic absorption characteristics of porous carbon–carbon ceramics with random microstructure for passive hypersonic boundary layer transition control

Preceding studies in the high enthalpy shock tunnel Gottingen of the German Aerospace Center (DLR) revealed that carbon fibre reinforced carbon ceramic (C/C) surfaces can be utilized to damp hypersonic boundary layer instabilities leading to a delay of boundary layer transition onset. To assess the ultrasonic absorption properties of the material, a test rig was set up to measure the reflection coefficient at ambient pressures ranging from 0.1 × 105 to 1 × 105 Pa. For the first time, broadband ultrasonic sound transducers with resonance frequencies of up to 370 kHz were applied to directly cover the frequency range of interest with respect to the second-mode instabilities observed in previous experiments. The reflection of ultrasonic waves from three flat plate test samples with a porous layer thickness between 5 and 30 mm was investigated and compared to an ideally reflecting surface. C/C was found to absorb up to 19 % of the acoustic power transmitted towards the material. The absorption characteristics were investigated theoretically by means of the quasi-homogeneous absorber theory. The experimental results were found to be in good agreement with the theory.

Strategies for control of transitional and turbulent flows using plasma-based actuators

An exploratory numerical study of the control of transitional and turbulent separated flows by means of dielectric-barrier-discharge (DBD) actuators is presented. The flow fields are simulated employing a high-fidelity Navier–Stokes solver augmented with a phenomenological model representing the plasma-induced body forces imparted by the actuator on the fluid. Several applications are considered, including interaction of an actuator with a laminar boundary layer, suppression of wing stall, control of boundary layer transition on a plate, control of laminar separation over a ramp, and turbulent separation over a wall-mounted hump. Effective suppression of stall over a NACA 0015 airfoil at moderate Reynolds numbers is demonstrated using either co-flow or counter-flow actuators pulsed at a sufficiently high frequency. By contrast, continuous actuation is found to provide little control of separation. For a laminar boundary layer developing along a flat plate, a counter-flow DBD actuator is shown to provide an effective on-demand tripping device. This property is exploited for the suppression of laminar separation over a ramp. Control of turbulent boundary-layer separation over a wall-mounted hump suggests that once the flow is turbulent, control effectiveness is only achieved for higher actuator strengths with implications for the scalability of DBD devices to higher freestream velocities.

Global-optimization-based methods for use in boundary layer transition control

Turbulent flow has a significantly higher drag than the corresponding laminar flow at the same flow conditions, and therefore incurs a significant penalty of increased fuel consumption due to the extra thrust required. One possible way of decreasing the drag is to apply surface suction to delay the transition from laminar to turbulent flow. In this paper an aerofoil with three non-overlapping panels covering up to 20 per cent of chord for boundary layer transition control is considered. The problem is complicated by the fact that panels can change both their positions and lengths. The complexity of the optimization problem is such that it is not practical to perform the investigation using a single processor. A constrained global parallel algorithm based on a combination of deformed configuration methods and a controlled random search method is developed. It is shown that for the problem considered, good solutions can be found efficiently.