A Parametric Study on the Effect of Casing Treatment Slots on Performance Enhancement of a Single Stage Axial Flow Compressor

This paper numerically investigates the effect of slots on the performance of a highly subsonic axial flow compressor. The axial flow compressor consisting of 21 rotor blades with NACA 65 series blade profile was used for the simulations. The present results were obtained using different turbulence models and shear stress transport model was found to be the best one. Studies were conducted to determine the influence of length, depth and skew angle of the slot on the compressor performance. The slot width and centre-to-centre distance between the successive slots were maintained at 6.3% Ca and 11% Ca, respectively. The present study was carried out at different slot lengths (0 to 100% of axial chord, 20 to 100% of axial chord and 40 to 100% of axial chord), slot depths (7.9, 11 and 15.7% Ca) and skew angles (0°, 30°, 45° and 60°). The slot length of 20 to 100% of Ca, depth of 15.7% Ca and skew angle of 60° resulted in the best compressor performance leading to 22.1% stall margin improvement. Subsequently, flow characteristics were studied without and with slots.

Stall Warning Strategy Based On Fast Wavelet Analysis in a Multi-Stage Axial Flow Compressor

As a reliable stall warning strategy, the fast wavelet method was introduced to successfully predict the aerodynamic instability of a multi-stage axial flow compressor. One single sensor installed at each stage is proved to be sufficient to predict the stability status in a three-stage axial flow compressor. The whole prediction strategy includes the dynamic pressure signal capture, disturbance extraction using decomposition and reconstruction via fast wavelet transform, and stall warning index calculation based on statistical probability distribution. On this premise, the first occurrence of the stall in this three-stage axial flow compressor is predicted to be within the first stage, which is consistent with the stall route captured by the eight transducers around the casing wall. Thereafter, the stall warning index is used to monitor the stability status during the continuous throttling process. Furthermore, the validation using tip air injection and inlet radial distortion indicated that the stall warning index decreases as the compressor's stability improves. Conversely, the deterioration of stability causes the increase of the stall warning index. Thus, experimental results demonstrate that the stall warning method based on fast wavelet analysis can predict the aerodynamic instability in actual application.

Effect of Circumferential Single Casing Groove Location on the Flow Stability under Tip-Clearance Effect in a Transonic Axial Flow Compressor Rotor

Numerical simulations have been performed to study the effect of the circumferential single-grooved casing treatment (CT) at multiple locations on the tip-flow stability and the corresponding control mechanism at three tip-clearance-size (TCS) schemes in a transonic axial flow compressor rotor. The results show that the CT is more efficient when its groove is located from 10% to 40% tip axial chord, and G2 (located at near 20% tip axial chord) is the best CT scheme in terms of stall-margin improvement for the three TCS schemes. For effective CTs, the tip-leakage-flow (TLF) intensity, entropy generation and tip-flow blockage are reduced, which makes the interface between TLF and mainstream move downstream. A quantitative analysis of the relative inlet flow angle indicates that the reduction of flow incidence angle is not necessary to improve the flow stability for this transonic rotor. The control mechanism may be different for different TCS schemes due to the distinction of the stall inception process. For a better application of CT, the blade tip profile should be further modified by using an optimization method to adjust the shock position and strength during the design of a more efficient CT.

Numerical study on the effect of blade tip clearance change on stall margin of the transonic axial flow compressor rotor

The change of the blade tip clearance size has an important impact on the performance of the compressor. Considering that the performance curve of the compressor is often limited by surge and stall boundaries, this paper used the numerical simulation method to investigate the influence mechanism of the blade tip clearance size change on the stall margin of transonic axial flow compressor rotor. By mathematically decomposing the calculation formula of the stall margin of rotor, the approximate calculation formula of the change of rotor’s stall margin was obtained. Then, the detailed quantitative analysis of the factors that affect the rotor’s stall margin was carried out, the influence weights of various factors on the rotor’s stall margin was also obtained. Finally, the physical mechanism of the change of the rotor’s performance parameters was obtained by the analysis of rotor tip flow field after the blade tip clearance size change.

Design Optimization of a Multistage Axial Flow Compressor Based on Full-Blade Surface Parameterization and Phased Strategy

A new design optimization method is proposed for the problem of high-precision aerodynamic design of multistage axial compressors. The method mainly contains three aspects: full-blade surface parametrization can significantly reduce the number of control variables per blade row and increase the degrees of freedom of the leading edge blade angle compared with the traditional semiblade parametric method; secondly, the artificial bee colony algorithm improved initialization and food source exploration and exploitation mechanism to enhance the global optimization ability and convergence speed, and a distributed optimization system is built on the supercomputing platform based on this method; finally, a phased optimization strategy based on the “synchronous change in multirow blades” is proposed, and expert experience is introduced to achieve a better balance between exploration and exploitation. The optimization method is applied to the AL-31F four-stage low-pressure compressor. As a result, the adiabatic efficiency is improved by 0.67% and the surge margin is improved by 3.1% under the premise that the total pressure ratio and mass flow rate satisfy the constraints, which verifies the effectiveness and engineering practicality of the proposed optimization method in the field of multistage axial flow compressor aerodynamic optimization.