Comparative Study of Tip Injection in a Transonic and Subsonic Compressor

Discrete tip injection is an effective method to enhance stability of compressors. This study compares the effects of injection parameters on compressor performance and underlying mechanisms in two different compressors. The transonic compressor is studied using unsteady simulations and the subsonic compressor is mainly investigated with experiment. Results show that tip injection improves stable operating range by 35.6% and 77.9% for the transonic compressor and subsonic compressor, respectively, without decreasing compressor efficiency. The effects of circumferential coverage percentage and injector throat height on compressor stability are similar in the two compressors when the injection velocity is double the velocity of main flow. The optimal injector throat height which is normalized by the tip clearance size is the same for the two compressors, and the best circumferential coverage percentage for the subsonic compressor is lower than that in the transonic compressor. For the two compressors, the adaption of the main flow to the discrete tip injection is unsteady, and the hysteresis effect that the recovery of tip blockage lags behind the recovery of tip leakage vortex accounts for the improved stability using partial coverage of injection. The injection efficiency, which is defined to quantify the improved quality of the flow field in the injection domain, is proven to determine the stall limits by studying the effects of several injection parameters. The guidelines built in the subsonic compressor can be used in the transonic compressor to design tip injection, but the optimal values of some injection parameters should be reconfirmed.

Shapley Additive Explanations of Multi-Geometrical Variable Coupling Effect in Transonic Compressor

Optimization algorithms in the compressor detailed design stage generate big data of geometries and corresponding performances, but these data are often not exploited efficiently to unveil hidden compressor design guidance. In this work, the Shapley Additive Explanations (SHAP) method from game theory is proposed as an efficient methodology to extract design guidelines from databases. A database was generated when optimizing the blade features (sweep, lean, end-bend) of Rotor 37. Based on this, a neural network is trained to predict compressor efficiency. The SHAP method is then applied to explain the neural network behavior, which provides information on the sensitivity of single geometrical variables and the coupling effect between multiple geometrical variables. Results show that the near-tip sweep and mid-span lean angles are most influential on efficiency. Within the same group of variables, the adjacent variables tend to present strong positive coupling effects on efficiency. Among different groups, evident coupling effects are observed between sweep and lean and between lean and end-bend, but the coupling effect between sweep and end-bend is negligible. Flow mechanisms behind the coupling effects are discussed. For near-tip lean angles L3 and L4, the positive coupling effect is due to the change of the passage shock. For near-tip lean angle L4 and sweep angle S4, the change of detached shock leads to a negative coupling effect. The proposed data mining method based on the neural network and SHAP is promising and transferable to other turbomachinery optimization databases in the future.

Rotating tip clearance asymmetry and role of endwall injection in stability desensitization of a transonic fan

The current study is aimed at understanding the effect of rotating tip clearance asymmetry on the operability and performance of a transonic compressor. Another objective of this investigation is to determine the influence of tip injection on reducing the detrimental effects of clearance asymmetry. Three dimensional unsteady Reynolds-averaged Navier–stokes simulations have been performed from choke to stall for different arrangements of non-uniform blade heights in a transonic fan. Furthermore, numerical computations have been conducted with endwall injection of air. The numerical results have been validated against experimental data. Results show that having the same mean tip clearance, the asymmetric compressor is less stable than the axisymmetric configuration. However, the peak pressure rise is found to be almost linearly correlated to the mean tip clearance for both the axisymmetric and asymmetric compressors. It is found that tip injection can desensitize the compressor to the tip clearance asymmetry. Results further reveal that tip clearance asymmetry does not change the compressor path to instability. However, endwall injection is found to be able to change the compressor stalling mode. Investigations concerning rotating non-uniformity (caused by non-uniform blade heights) are very few in open literature. The obtained results can assist in predicting the effect of rotating tip clearance asymmetry on the stability and performance of high-speed compressor rotors. Furthermore, the results uncover how tip injection can desensitize the compressor stability and affect its path into instability, which is one of the most important questions in the turbomachinery world.

Effect of Airfoil Vortex Generator on the Performance and Stability of a Transonic Axial Compressor

In the present research work, effect of airfoil vortex generator on performance and stability of transonic compressor stage is investigated through CFD simulations. In turbomachines vortex generators are used to energize boundary and generated vortex is made to interact with tip leakage flow and secondary flow vortices formed in rotor and stator blade passage. In the present numerical investigation symmetrical airfoil vortex generator is placed on rotor casing surface close to leading edge, anticipating that vortex generated will be able to disturb tip leakage flow and its interaction with rotor passage core flow. Six different vortex generator configuration are investigated by varying distance between vortex generator trailing edge and rotor leading edge. Particular vortex generator configuration shows maximum improvement of stall margin and operating range by 5.5% and 76.75% respectively. Presence of vortex generator alters flow blockage by modifying flow field in rotor tip region and hence contributes to enhancement of stall margin. As a negative effect, interaction of vortex generator vortices and casing causes surface friction and high entropy generation. As a result compressor stage pressure ratio and efficiency decreases.

Effect of Rotor Tip Casing Treatment on the Performance and Stability of a Transonic Axial Compressor

In a transonic compressor rotor, tip leakage flow interacts with passage shock, casing boundary layer and secondary flow. This leads to increase in total pressure loss and reduction of compressor stability margin. Casing treatment is one of the passive endwall geometry modification technique to control tip leakage flow interaction. In the present investigation effect of rotor tip casing treatment is investigated on performance and stability of a NASA 37 transonic compressor stage. Existing literature reveals, that endwall casing treatment slots i.e., porous casing treatment, axial slots axially skewed slots, circumferential grooves, recirculating casing treatment etc. are able to improve compressor stability margin with penalty on stage efficiency. Turbomachinery engineers and scientists are still focusing their research work to identify an endwall casing treatment configuration with improves both compressor stall margin as well as stage efficiency. Hence in the current work, as an innovative idea, effect of casing treatment slot along rotor tip mean camber line is investigated on NASA 37 compressor stage. Casing treatment slot with rectangular cross-section was created along the rotor tip mean camber line. Four different casing treatment configurations were created by changing number of slots on rotor casing surface. In all four configurations casing treatment slot width and height remains same. Flow simulation of NASA 37 compressor stage was performed with all these four casing treatment configurations. A maximum stall margin improvement of 3% was achieved with a particular slot configuration, but without any increase in compressor stage efficiency.

Effect of three non-axisymmetric stator schemes on compressor performance with distorted inlet

In the boundary layer ingesting propulsion system, the compressor suffers from a non-uniform flow field. The compressor operating with distorted inflow continuously results in the loss of aerodynamic performance and stability margin. In this paper, three non-axisymmetric configurations are described for the stator of a transonic compressor to match the non-uniform flow field. The flow fields with distorted inflow at near stall condition are obtained and analyzed, the effects of the prototype stator and the three non-axisymmetric stators on aerodynamic performance are compared in detail. Results show that the non-axisymmetric stator schemes can effectively improve the stability margin of the transonic compressor and the maximum stability margin is relatively increased by 22.3% in all the three non-axisymmetric stators. The non-axisymmetric stator design is effective on decreasing the aerodynamic losses and improving the performance of the compressor operating with distorted inflow. Overall, the results show that in the design of the non-axisymmetric stator, the adoption of a curved-twisted blade and the increase of cascade solidity have the potential to reduce loss sources caused by distorted inflow.