Parametric study on active flow control of a transonic axial compressor rotor using endwall synthetic jet

High-load axial compressor is the mainstream of current compressor design and development. In order to improve the aerodynamic performance of high-load axial compressor, an active flow control method in which a synthetic jet is applied to the endwall is proposed. Taking the transonic axial compressor NASA Rotor 35 as the research object, using a single factor analysis method, the influence of five different excitation positions, three different excitation frequencies, and three different jet peak velocities on the aerodynamic performance of the compressor was studied in turn, and obtained the influence law of the endwall synthetic jet excitation parameters. The results show that all three parameters have important effects on the performance of the compressor. Among the excitation parameters studied in this paper, there is an optimal excitation position of 25% Ca. When excited at this position, the flow margin of the compressor is expanded the most. On the basis of maintaining the optimal excitation position and the maximum jet peak velocity, the calculation results found that the jet frequency has little effect on the compressor’s near stall flow rate, but has a great impact on the total pressure ratio and efficiency. The pressure ratio and efficiency increase with the increase of the excitation frequency. However, there seems to be a threshold of the excitation frequency. Only when the excitation frequency is greater than the threshold can the total pressure ratio and efficiency be higher than the prototype compressor. The jet peak velocity has the smallest impact on the compressor performance. Based on the optimal excitation position and the excitation frequency exceeding the threshold, even if the jet peak velocity is small, the compressor can obtain a higher flow margin, total pressure ratio, and efficiency than the prototype compressor. As the jet peak velocity increases, the performance of compressor can be further improved.

Simulation of the effect of nonuniform fouling thickness on an axial compressor stage performance

This paper focuses on the numerical simulation of the flow status in the compressor with the condition of fouling. NASA stage 35 was considered as the object, and the commercial code ANSYS CFX was used. The deposition rate of contaminants on the surface was considered to be different along the height of the blade. A data from related study shows that the deposition rate of the contaminant on the side close to the hub is higher than the side near the shroud part. Based on the deposition law, this paper simulated the fouling of the compressor blades by changing the thickness on the blade surface. This subject only changed the thickness of the stator blade surface because of a data showing that the fouling on the stator blade surface is almost double that on the rotor blade surface. In the condition that the roughness value of the blade surface is constant, only the stable working range of the compressor is effected by the change of the surface thickness of the stator blade. There is a positive relationship between the value of compressor minimum flow rate and the value of thickness increment. After fouling the total pressure ratio and isentropic efficiency degenerated 1.59% and 3.76%, respectively.

Preliminary Design Tool for Centrifugal Compressors

Commonly, compressor designs rely on previous machines that can be slightly modified to achieve new operating requirements. In some cases, however, a completely new design is needed because no previous designs are available for the specific operating range of interest. Without a previous design, it is difficult to make initial trade studies of an appropriate impeller diameter, speed, and number of compression stages. While new compressor designs are a common occurrence in applied research applications, conceptual design typically require a point-by-point process to balance the requirements with acceptable design parameters. This can be done manually or through automation to optimize for a specific operating parameter, such as efficiency. The authors are unaware of any tool available that bounds the range of design parameters for a centrifugal compressor stage without applying a point-by-point method. In this work, two common references for conceptual compressor design were cross-checked to develop an Excel-based tool to quickly determine the design space for a given set of compressor requirements. The tool relies on design experience presented by Aungier and Baljé as well as other experience drawn from available literature [1],[2]. The sheet functions from a series of assumptions based within the design experience and requires inputs regarding the desired power, fluid flow rate, and total-to-total pressure ratio, as well as inlet conditions. While the tool currently assumes an ideal gas, future revisions can include calls to REFPROP for a real gas.

Research on the Influence of Axisymmetric Endwall on EAT Performance

To improve the performance of electrically assisted turbochargers (EATs), the influences of the hub profile and the casing profile on EAT performance were numerically studied by controlling the upper and lower endwall profiles. An artificial neural network and a genetic algorithm were used to optimize the endwall profile, considering the total pressure ratio and the isentropic efficiency at the peak efficiency point. Different performances of the prototype EAT and the optimized EAT under variable clearance sizes were discussed. The endwall profile affects an EAT by making the main flow structure in the endwall area decelerate and then accelerate due to the expansion and contraction of the meridional surface, which weakens the secondary leakage flow of the prototype EAT and changes the momentum ratio of the clearance leakage flow and the separation flow in the suction surface corner area. Because the tip region flow has a more significant influence on EAT performance, the optimal casing scheme has a better effect than the hub scheme. The optimization design can increase the isentropic efficiency of the maximum efficiency point by 1.5%, the total pressure ratio by 0.67%, the mass flow rate by 1.2%, and the general margin by 6.4%.

Aerodynamic Optimization Design of a Supersonic Compressor Rotor with High Pressure Ratio

Supersonic compressors have a high wheel speed and operational capability, which facilitate a high stage pressure ratio. However, the strong shock waves in the passage of a supersonic rotor and the interference between shock waves and boundary layers can lead to large flow loss and low efficiency. Moreover, the existing design of a high-load supersonic compressor has the problem of small stall margin. In this study, an automatic optimization method including 2D profile optimization and 3D blade optimization is proposed to achieve a high efficiency at the design point of a supersonic compressor rotor under the premise of reaching the desired mass flow rate and total pressure ratio. According to the analysis of flow near the stall point of the supersonic compressor rotor, the mechanism responsible for rotor tip stall is established, that is, the aerodynamic throat appeared inside the flow passage, reducing the ability of the blade tip to withstand back pressure, and the low-speed areas caused by the tip-leakage-vortex breakage and boundary layer separation reduced the flow capacity of the blade tip. Based on the reasons for rotor stall, three methods are proposed to improve the stall margin, which include increasing the exit radius of the upper meridian, forward sweep of the blade tip, and increasing the chord length of the blade tip. The above method is used to design a supersonic rotor with a total pressure ratio of 2.8, which exhibits an efficiency of 0.902 at the design point and a stall margin of 18.11%.

Numerical study of probe parameters on performance of a transonic axial compressor

The paper shows the effect of the probe on the performance of a transonic axial speed compressor. The unobstructed flow case with the experimental data was validated and used as a guide for all subsequent study cases. The aerodynamic performance for different probe parameters were calculated numerically using ANSYS-CFX. This covered the results on compressor output from changing probe axial positions, the radial immersion depths, the size of the probe, and the total number of probes. The findings were evaluated in relation to the total pressure ratio, performance, margin of deflation and stability. The velocity part distributions further showed that the probe block and raises the flow Mach value, which is the explanation why the compressor rotor’s total pressure ratio is lost. In fact, the parameters of the sample will significantly influence the calculation outcomes and affect the standard margin. The range of stability was also affected, which changes the performance trend from the choke to the stall. Consequently, the collection of correct probe parameters with fewer impact on compressor output is addressed.

Investigation of Recirculating Casing Treatment for Low-Pressure Ratio Mixed-Flow Micro-Compressor

This paper investigates the recirculating casing treatment (RCT) of a low total pressure ratio micro-compressor to achieve stall margin enhancement while minimizing the design point efficiency penalty. Three RCT injection and extraction configurations are studied, including combined slot-duct, ducts only, and slot only. The numerical approach is validated with a tested micro-compressor using RCT. A very good agreement is achieved between the predicted speedlines and the measured results. To minimize the design point efficiency loss, it is observed that the optimal location of extraction and injection is where the recirculated flow rate can be minimized at the design point. To maximize stall margin, extraction location should favor minimizing the tip blockage such as at the location where the tip flow separation of the baseline blade is fully developed. In addition, the slot configuration that generates pre-swirl to the upstream flow is beneficial to improve stall margin due to reduced incidence. The highest stall margin enhancement achieved is 9.49% with the slot geometry that has the extraction at the 62%C chordwise location, but has a design point efficiency loss of 1.9%. Overall, a small efficiency penalty of 0.6% at the design point is achieved for the final design with the stall margin increased by 6.2%.

Investigation Into the Effects of Hub Rotation on the Hub Leakage Flow of Cantilever Stator in a Transonic Axial Compressor

In order to explore the similarities and differences between the flow fields of cantilever stator and idealized compressor cascade with tip clearance, and to extend the cascade leakage model to compressors, the influence of stator hub rotation to represent cascade and cantilever stator on hub leakage flow was numerically studied. On this basis, the control strategy and mechanism of blade root suction were discussed. The results show that there is no obvious influence on stall margin of the compressor whether the stator hub is rotating or stationary. For rotating stator hub, the overall efficiency is decreased while the total pressure ratio is increased. At peak efficiency point and near stall point, the efficiency is reduced by about 0.43% and 0.34% individually, while the total pressure ratio is enlarged by about 0.23% and 0.27%, respectively. The gap leakage flow is promoted due to stator hub rotation, and the structure of the leakage vortex is weakened obviously. In addition, the hub leakage flow originating from the blade leading edge of rotating hub may contribute to double leakage near the trailing edge of the adjacent blade. However, the leakage flow directly out of the blade passage with stationary stator hub. The stator root loading and strength of the leakage flow increase with the rotation of the hub, and the leakage vortex is further away from the suction surface of the blade and is stretched to an ellipse closer to the endwall under the shear action. The rotating hub makes the flow loss near the stator gap increase, while the flow loss in the upper part of the blade root is decreased. Meanwhile, the total pressure ratio in the end area is increased. Blade root suction of cantilever stator can effectively control the hub leakage flow, inhibit the development of hub leakage vortex, and improve the flow capacity of the passage, thereby reducing the flow loss and modifying the flow field in the end zone.

Research on Optimization of the Design(Inverse) Issue of S2 Surface of Axial Compressor Based on Genetic Algorithm

In order to further improve the effectiveness of design(inverse) issue of S2 surface of axial compressor, a design method of optimization model based on real-coded genetic algorithm is instructed, with a detailed description of some important points such as the population setting, the fitness function design and the implementation of genetic operator. The method mainly takes the pressure ratio, the circulation as the optimization variables, the total pressure ratio and the overall efficiency of the compressor as the constraint condition and the decreasing of the diffusion factor of the compressor as the optimization target. In addition, for the propose of controlling the peak value of some local data after the optimization, a local optimization strategy is proposed to make the method achieve better results. In the optimization, the streamline curvature method is used to perform the iterative calculation of the aerodynamic parameters of the S2 flow surface, and the polynomial fitting method is used to optimize the dimensionality of the variables. The optimization result of a type of ten-stage axial compressor shows that the pressure ratio and circulation parameters have significant effect on the diffusion factor’s distribution, especially for the rotor pressure ratio. Through the optimization, the smoothness of the mass-average pressure ratio distribution curve of the rotors at all stages of the compressor is improved. The maximum diffusion factors in spanwise of rotor rows at the first, fifth and tenth stage of the compressor are reduced by 1.46%, 12.53% and 8.67%, respectively. Excluding the two calculation points at the root and tip of the blade because of the peak value, the average diffusion factors in spanwise are reduced by 1.28%, 3.46%, and 1.50%, respectively. For the two main constraints, the changes of the total pressure ratio and overall efficiency are less than 0.03% and 0.032%, respectively. In the end, a 3-d CFD numerical result is given to testify the effects of the optimization, which shows that the loss in the compressor is decreased by the optimization algorithm.

Effect of Backward Sweep on Aerodynamic Performance of a 1.5-Stage Highly Loaded Axial Compressor

As a degree of freedom in the three-dimensional blade design of axial compressors, the sweep technique significantly affects the aerodynamic performance of axial compressors. In this paper, the effects of backward sweep rotor configurations on the aerodynamic performance of a 1.5-stage highly loaded axial compressor at different rotational design speeds are studied by numerical simulation. The aim of this work is to improve understanding of the flow mechanism of backward sweep on the aerodynamic performance of a highly loaded axial compressor. A commercial CFD package is employed for flow simulations and analysis. The study found that at the design rotational speed, compared with baseline, backward sweep rotor configurations reduce the blade loading near the leading edge but slightly increases the blade loading near the trailing edge in the hub region. As the degree of backward sweep increases, the stall margin of the 1.5-stage axial compressor increase first and then decrease. Among different backward sweep rotor configurations, the 10% backward sweep rotor configuration has the highest stall margin, which is about 2.5% higher than that of baseline. This is due to the change of downstream stator incidence, which improves flow capacity near the hub region. At 80% rotational design speed, backward sweep rotor configurations improve stall margin and total pressure ratio of the compressor. It’s mainly due to the decreases of the rotor incidence near the middle span, which results in the decreases of separation on the suction surface. At 60% rotational design speed, detached shock disappears. Backward sweep rotor configurations deteriorate stall margin of the compressor, but increase total pressure ratio and adiabatic efficiency when the flow rate is lower than that at peak efficiency condition. Therefore, it’s necessary to consider the flow field structure of axial compressors at whole operating conditions in the design process and use the design freedom of sweep to improve the aerodynamic performance.