Mechanism of compressor airfoil boundary layer flow control using nanosecond plasma actuation

https://doi.org/10.1016/j.ijheatfluidflow.2019.108502Get rights and content

Highlights

  • In this manuscript, the nanosecond plasma actuation is used to control the boundary layer flow of a typical CDA compressor airfoil.

  • The corresponding flow control mechanism has been uncovered using large eddy simulations.

  • It has been found that the nanosecond plasma actuation can influence the boundary layer flow through inducing two types of spanwise vortexes.

  • Especially, the plasma actuations located within the laminar boundary layer flow can act as a high-speed wall jet.

  • This finding further improves the ability of plasma actuation in controlling high-speed flows.

Abstract

The flow control effects of nanosecond plasma actuation on the boundary layer flow of a typical compressor controlled diffusion airfoil are investigated using large eddy simulation method. Three types of plasma actuation are designed to control the boundary layer flow, and two mechanisms of compressor airfoil boundary layer flow control using nanosecond plasma actuation have been found. The plasma actuations located within the laminar boundary layer flow can induce a small vortex structure through influencing on the density and pressure of the flow field. As the small vortex structure moves downstream along the blade surface with the main flow, it can suppress the turbulent flow mixing and reduce the total pressure loss. The flow control effect of the small vortex structure is summarized as wall jet effect. Differently, the plasma actuation located within the turbulent boundary layer flow can act on the shear layer flow and induce a large vortex structure. While moving downstream, this large vortex structure can suppress the turbulent flow mixing too.

Introduction

Compressor internal flow is characterized by strong unsteadiness and adverse pressure gradient (Cumpsty, 1989; Gao et al., 2015). As a result, flow transitions and separations are almost unavoidable for the boundary layer flow of compressor airfoils (Steinert and Starken, 1996; Lardeau et al., 2011; Scillitoe et al., 2019). Early transitions will increase the friction loss of blade surfaces, while too late transitions will increase the possibility of large-scale laminar boundary layer flow separations. For this reason, it is crucial to control the boundary layer flow effectively to improve the aerodynamics of compressor airfoils. Decades ago, researchers developed the compressor controlled diffusion airfoil (CDA) and obtained the optimized boundary layer flow state through controlling the adverse pressure gradient of compressor blade passage (Hobbs and Weingold, 1984; Behlke, 1986). With the CDA airfoil, the transition and separation of compressor blade boundary layer flow can be controlled, and resultantly, the flow loss is kept in a low level. As a matter of fact, the application of CDA airfoil was a landmark progress of aero-engine compressor developments. Later, researchers found the geometry of blade leading edge is rather influential to the aerodynamic performances of compressor airfoils. For traditional compressor airfoils, there often exists a small “spike” in the surface pressure distribution at the leading edge. These spikes may trigger boundary layer transitions or separations which will increase the compressor airfoil flow loss (E.Walraevens and Cumpsty, 1995; Goodhand and Miller, 2011; Wheeler et al., 2009). To reduce the negative influences of these spikes on the compressor blade boundary layer flow, the spikeless compressor airfoil is developed through optimizing the geometry of blade leading edge and the working efficiency improvement of aero-engine compressors is achieved resultantly.

Although the aerodynamic performances of compressor airfoil can be improved through optimizing the blade geometry, the boundary layer flow may experience unwanted transition or separation at off-design working conditions which will increase the flow loss dramatically (Goodhand and Miller, 2012). So it seems essential to develop flow control technique to control the compressor airfoil boundary layer flow actively to further improve the aerodynamic performances of aero-engine compressors whose working conditions will change dramatically with the variation of flying state. Plasma actuation is a typical active flow control technique and characterized by wide actuation frequency band and simple structure (Corke et al., 2010; Zheng et al., 2018). With these characteristics, plasma actuation has been acknowledged as an effective technique in controlling complex internal flows, such as compressor and turbine tip leakage flow (Zhang et al., 2019; Yu et al., 2019; Wang et al., 2019). Nanosecond plasma actuation, which is driven by a high impulse voltage and the corresponding pulse width is no bigger than 1 microsecond, is a typical unsteady flow control technique. It has been found that the nanosecond plasma actuation can suppress flow separations through imposing fast heating in the flow field (Correale et al., 2014). When the nanosecond plasma actuation is turned on, energy will be injected into the flow field within a very short time period (less than 1 microsecond) through gas discharge. As a result, the residual heat as well as shock wave can be observed in the discharge volume. Both the residual heat and shock wave can interact with the flow field, and achieve the flow control effect. It has been well acknowledged that the nanosecond plasma actuation can effectively control the boundary layer flow separations (Zheng et al., 2019; Little et al., 2012) and influence the boundary layer flow transitions (Little, 2019; Ullmer et al., 2015). So being an effective active flow control technique, the nanosecond plasma actuation has the potential to control the compressor airfoil boundary layer flow actively. Since no academic publications on this research topic can be found at present, exploratory investigations on the flow control of compressor airfoil using nanosecond plasma actuation are essential to uncover the flow control effects and mechanisms. With these exploratory investigations, it is expected that the plasma actuation based effective active flow control technique to control compressor airfoil boundary layer flow can be developed.

This paper is aimed at uncovering the mechanism of compressor airfoil boundary layer flow control using nanosecond plasma actuation. The nanosecond plasma actuation is imposed on the suction surface of a typical compressor CDA airfoil, and the influences of nanosecond plasma actuation on the suction surface boundary layer flow are investigated using LES (large eddy simulation) method. It has been found that the nanosecond plasma actuation can control the compressor airfoil boundary layer flow through inducing two types of spanwise vortexes. Results of this paper provide the basis of developing plasma actuation based flow control technique to control compressor airfoil boundary layer flow actively.

Section snippets

Airfoil model and simulation method

A typical compressor CDA airfoil (Steinert et al., 1990) is investigated in this paper. The airfoil model as well as the corresponding main aerodynamic parameters are shown in Fig. 1. The design inlet Mach number of the airfoil is 0.62, and the chord length is set as 70 mm during the simulations to keep in line with previous experimental investigations (Steinert et al., 1990). Correspondingly, the Reynolds number of the inlet flow is 9.8 × 105. The design inlet and outlet flow angles are 137°

Characteristics of the compressor airfoil boundary layer flow

The time-averaged static pressure and axial wall shear stress coefficients on the blade suction surface are shown in Fig. 3. The static pressure coefficient Cp is defined asCp=(p¯p¯1)/12ρU12where p¯ is the time-averaged local static pressure and p1¯ is the time-averaged static pressure of inlet flow. U1 is the time-averaged velocity of inlet flow. As for the simulation cases without plasma actuation, the time-averaged results here and in the following parts are obtained from arithmetical

Plasma actuation layouts

In this paper, the nanosecond plasma actuation is used to control the compressor airfoil boundary layer flow. Based on the characteristics of the compressor airfoil boundary layer flow, the dielectric barrier discharge plasma actuators are imposed on the blade suction surface at three locations. As shown in Fig. 6, the Actu1 plasma actuation is located at the blade leading edge, which is expected to influence the laminar boundary layer flow directly. The Actu2 plasma actuation is located

Flow control effects of Actu1 and Actu2 plasma actuations

The contours of spanwise vorticity (ωs) on the mid-span S1 surface at different moments after Actu1 plasma actuation is imposed are shown in Fig. 9. According to Fig. 9(a), the compressive wave can be observed 2 μs after the Actu1 plasma actuation is imposed. Then the compressive wave will spread outward and increase the pressure of the flow field. According to Fig. 9(b), a small vortex structure (SVS) with concentrated spanwise vorticity can be observed on the blade suction surface 22.5 μs

Flow control effects of Actu3 plasma actuation

Different from the Actu1 and Actu2 plasma actuation, the Actu3 plasma actuation is located within the turbulent boundary layer flow region. The contours of spanwise vorticity on the midspan S1 surface at different moments after Actu3 plasma actuation is imposed are shown in Fig. 14. It can be found that the disturbance imparted into the turbulent flow by the Actu3 plasma actuation can impose dramatic influence on the flow field.

According to Fig. 14(a), the compressive wave induced by nanosecond

Discussion

The distributions of time-averaged total pressure loss coefficient along the pitchwise direction with and without plasma actuations are shown in Fig. 17. The results are obtained at the axial location 30% chordlength from the blade trailing edge. As shown in Fig. 4, the blade wake with strong total pressure loss is aroused from the interactions of boundary layer flow on the blade surfaces. Correspondingly, the high total pressure loss coefficient can be observed within 0.4–0.8 pitch in Fig. 17.

Conclusions

The flow control effects of nanosecond plasma actuation on the boundary layer flow of a typical CDA compressor airfoil are investigated in this paper using LES simulation method, and main conclusions obtained are listed as below.

The boundary layer flow remains as laminar at the front part of the blade, and under the influences of laminar separation bubble, the boundary layer flow transits to turbulent at the rear part of the blade.

Based on the characteristics of the boundary layer flow, three

Declaration of Competing Interest

None.

Acknowledgments

This paper is supported by the National Natural Science Foundation of China (nos. 51906254, 51790511).

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