scholarly journals Research on the Influence of Axisymmetric Endwall on EAT Performance

Energies ◽  
2021 ◽  
Vol 14 (8) ◽  
pp. 2215
Author(s):  
Han Teng ◽  
Wanyang Wu ◽  
Jingjun Zhong
Keyword(s):  
Total Pressure ◽  
Optimization Design ◽  
Pressure Ratio ◽  
Maximum Efficiency ◽  
Leakage Flow ◽  
Separation Flow ◽  
Suction Surface ◽  
Region Flow ◽  
Total Pressure Ratio ◽  
Isentropic Efficiency

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%.

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

2020 ◽  
Author(s):  
Botao Zhang ◽  
Bo Liu ◽  
Xin Sun ◽  
Hang Zhao
Keyword(s):  
Total Pressure ◽  
Pressure Ratio ◽  
Leading Edge ◽  
Tip Clearance ◽  
Leakage Flow ◽  
Compressor Cascade ◽  
Suction Surface ◽  
Blade Root ◽  
Flow Loss ◽  
Total Pressure Ratio

Abstract 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.

Experimental Investigation of Aspiration in a Multi-Stage High-Speed Axial-Compressor

2016 ◽  
Author(s):  
Jan Siemann ◽  
Ingolf Krenz ◽  
Joerg R. Seume
Keyword(s):  
Flow Field ◽  
Total Pressure ◽  
High Speed ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Reference Configuration ◽  
Corner Separation ◽  
Part Load ◽  
Total Pressure Ratio ◽  
Isentropic Efficiency

Reducing the fuel consumption is a main objective in the development of modern aircraft engines. Focusing on aircraft for mid-range flight distances, a significant potential to increase the engines overall efficiency at off-design conditions exists in reducing secondary flow losses of the compressor. For this purpose, Active Flow Control (AFC) by aspiration or injection of fluid at near wall regions is a promising approach. To experimentally investigate the aerodynamic benefits of AFC by aspiration, a 4½-stage high-speed axial-compressor at the Leibniz Universitaet Hannover was equipped with one AFC stator row. The numerical design of the AFC-stator showed significant hub corner separations in the first and second stator for the reference configuration at the 80% part-load speed-line near stall. Through the application of aspiration at the first stator, the numerical simulations predict the complete suppression of the corner separation not only in the first, but also in the second stator. This leads to a relative increase in overall isentropic efficiency of 1.47% and in overall total pressure ratio of 4.16% compared to the reference configuration. To put aspiration into practice, the high-speed axial-compressor was then equipped with a secondary air system and the AFC stator row in the first stage. All experiments with AFC were performed for a relative aspiration mass flow of less than 0.5% of the main flow. Besides the part-load speed-lines of 55% and 80%, the flow field downstream of each blade row was measured at the AFC design point. Experimental results are in good agreement with the numerical predictions. The use of AFC leads to an increase in operating range at the 55% part-load speed-line of at least 19%, whereas at the 80% part-load speed-line no extension of operating range occurs. Both speed-lines, however, do show a gain in total pressure ratio and isentropic efficiency for the AFC configuration compared to the reference configuration. Compared to the AFC design point, the isentropic efficiency ηis rises by 1.45%, whereas the total pressure ratio Πtot increases by 1.47%. The analysis of local flow field data shows that the hub corner separation in the first stator is reduced by aspiration, whereas in the second stator the hub corner separation slightly increases. The application of AFC in the first stage further changes the stage loading in all downstream stages. While the first and third stage become unloaded by application of AFC, the loading in terms of the De-Haller number increases in the second and especially in the fourth stage. Furthermore, in the reference as well as in the AFC configuration, the fourth stator performs significantly better than predicted by numerical results.

Aerodynamic Design and Analysis of a High Pressure Ratio Aspirated Compressor Stage

2000 ◽  
Author(s):  
Ali A. Merchant ◽  
Mark Drela ◽  
Jack L. Kerrebrock ◽  
John J. Adamczyk ◽  
Mark Celestina
Keyword(s):  
Boundary Layer ◽  
Total Pressure ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Adverse Pressure Gradient ◽  
Compressor Stage ◽  
High Pressure Ratio ◽  
Total Pressure Ratio ◽  
Isentropic Efficiency

The pressure ratio of axial compressor stages can be significantly increased by controlling the development of blade and endwall boundary layers in regions of adverse pressure gradient by means of boundary layer suction. This concept is validated and demonstrated through the design and analysis of a unique aspirated compressor stage which achieves a total pressure ratio of 3.5 at a tip speed of 1500 ft/s. The aspirated stage was designed using an axisymmetric through-flow code coupled with a quasi three-dimensional cascade plane code with inverse design capability. Validation of the completed design was carried out with three-dimensional Navier-Stokes calculations. Spanwise slots were used on the rotor and stator suction surfaces to bleed the boundary layer with a total suction requirement of 4% of the inlet mass flow. Additional bleed of 3% was also required on the hub and shroud near shock impingement locations. A three-dimensional viscous evaluation of the design showed good agreement with the quasi three-dimensional design intent, except in the endwall regions. The three-dimensional viscous analysis predicted a mass averaged total pressure ratio of 3.7 at an isentropic efficiency of 93% for the rotor, and a mass averaged total pressure ratio of 3.4 at an isentropic efficiency of 86% for the stage.

Non-Axisymmetric End Wall Profiling in Transonic Compressors—Part II: Design Study of a Transonic Compressor Rotor Using Non-Axisymmetric End Walls—Optimization Strategies and Performance

2009 ◽  
Author(s):  
Steffen Reising ◽  
Heinz-Peter Schiffer
Keyword(s):  
Secondary Flow ◽  
Total Pressure ◽  
Cfd Simulation ◽  
Pressure Ratio ◽  
Cross Flow ◽  
Secondary Flows ◽  
Transonic Compressor ◽  
End Wall ◽  
Total Pressure Ratio ◽  
Isentropic Efficiency

Secondary flows involving cross flow at high stage loading in modern axial compressors contribute significantly to efficiency limits. This paper summarizes an approach to control end wall flow using non-axisymmetric end walls. The challenge is to find the optimal non-axisymmetric end wall shape that results in the largest gain in performance. An automated multi-objective optimizer connected to a 3-D RANS flow solver was used to design the end wall contour. The process chain was applied to the rotor hub end wall of Configuration I of the Darmstadt Transonic Compressor. Several optimization strategies involving different objective functions to be minimized and the corresponding performances were compared. The parameters considered within the optimization process were isentropic stage efficiency, pressure loss in the rotor, throat area and secondary kinetic energy (SKE). A parameter variation was undertaken, leading to the following observations: Strong penalties on SKE at the rotor outlet and moderate penalties on isentropic efficiency, throat area and pressure ratio led to the best design. Isentropic efficiency could be raised by 0.12%, SKE at the rotor exit was reduced while the total pressure ratio of the stage remained constant. Strong penalties on efficiency and pressure ratio, a moderate one on throat area and a small one on SKE at the rotor outlet all led to a smaller increase in efficiency: 0.06%. On the other hand, a slight raise in the total pressure ratio could be achieved. A third optimization, eliminating the restriction on the throat area, was carried out to see which benefit in performance could be achieved without this geometrical restriction. Since the throat areas of all optimized geometries differ slightly from the datum value, an estimation was derived to see the extent to which the end wall profiling and cross section enlargement contribute to the improvements. Finally, a method to display secondary flows in turbomachinery is introduced. A second CFD simulation is used to calculate the primary flow where the hub end wall is defined as an Euler wall to avoid the end wall boundary layer and so eliminate the cause for some of the secondary flow mechanisms. This method clearly shows how the characteristics of secondary flow can be positively influenced by using non-axisymmetric end walls.

Robust Optimization Design of Single-Stage Transonic Axial Compressor Considering the Manufacturing Uncertainties

2018 ◽  
Author(s):  
Zhihui Li ◽  
Yanming Liu ◽  
Ramesh K. Agarwal
Keyword(s):  
Total Pressure ◽  
Optimization Design ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Tip Clearance ◽  
Robust Design Optimization ◽  
Compressor Performance ◽  
Overall Performance ◽  
Transonic Axial Compressor ◽  
Total Pressure Ratio

Manufacturing uncertainties always lead to significant variability in compressor performance. In this work, the tip clearance uncertainties inherent in a transonic axial compressor are quantified to determine their effect on performance. The validated tip clearance losses model in conjunction with the 3D reynolds averaged navier-stokes (RANS) solver are utilized to simulate these uncertainties and quantify their effect on the adiabatic efficiency, total pressure ratio and choked mass flow. The sensitivity analysis method is employed to figure out which parameters play the most significant roles in determining the overall performance of compressor. To propagate these uncertainty factors, the non-intrusive polynomial chaos expansion (PCE) algorithm is used in this paper and the probability distributions of compressor performance are successfully predicted. A robust design optimization has been carried out based on the combination of the genetic algorithm (GA) and the uncertainty quantification (UQ) method, leading to a robust compressor rotor design for which the overall performance is relatively insensitive to variability in tip clearance without reducing the sources of the manufacturing noise. The optimization results show that the mean value of the adiabatic rotor efficiency is improved by 1.4 points with the overall variation of that reduced by 64.1%, while the total pressure ratio is slightly improved when compared to the prototype.

Use of Inlet Guide Vanes for a Miniature Centrifugal Compressor

2005 ◽  
Author(s):  
Xiaoyi Li ◽  
Lei Zhou ◽  
Jay Kapat ◽  
Louis C. Chow
Keyword(s):  
Centrifugal Compressor ◽  
Total Pressure ◽  
High Speed ◽  
Cooling System ◽  
Pressure Ratio ◽  
Working Fluid ◽  
Guide Vanes ◽  
Inlet Guide Vanes ◽  
Total Pressure Ratio ◽  
Isentropic Efficiency

A novel design for a high-speed, miniature centrifugal compressor for a miniature RTBC (reverse turbo Brayton cycle) cryogenic cooling system is the focus of this paper. Due to the geometrical restriction imposed by the cryocooling system, the outer radius of the compressor is limited to 2.5 cm. Such a small compressor must rotate at a high speed in order to provide an acceptable pressure ratio. Miniature design precludes the use of inducers with large angles. In order to compensate for the absence of conventional inducers, the proposed design uses inlet guide vanes (IGV) that produce preswirl at the impeller inlet. IGV is followed by a radial impeller and an axial diffuser. The design speed for this compressor is 313,000 rpm for an overall static-to-total pressure ratio of 1.7 with helium as the working fluid for the compressor and the cryocooling system. The analysis undertaken in this paper is for an aerodynamically similar design with air as the working fluid. The rotational speed is 108,000 rpm and the overall static-to-total pressure ratio of 1.55. This paper concentrates on computational prediction of the performance of the compressor. The three-dimensional transient simulation is performed by using sliding mesh model (SMM). Blade tip leakage is not considered in the computation presented here. The unsteady solution predicts the interaction between IGV and the impeller, and between the impeller and the diffuser. The isentropic efficiency of impeller is found to be 81% at the design point. Based on the results obtained in this study, the inlet angle of diffuser vanes is modified to match the gas flow at the impeller exit, resulting in an increase of the isentropic efficiency of diffuser from 8.6% to 74.8%. It is also found that the performance of upstream components — IGV and impeller, are not affected by the performance of the diffuser.

scholarly journals Centrifugal Compressor Performance Improvement through Multi Splitter Impeller

2019 ◽  
Vol 26 (2) ◽  
pp. 6-14
Author(s):  
Adil Malik ◽  
Qun Zheng ◽  
Shafiq R. Qureshi ◽  
Salman A. Ahmed ◽  
D. KB Gambo
Keyword(s):  
Performance Improvement ◽  
Centrifugal Compressor ◽  
Total Pressure ◽  
Pressure Ratio ◽  
Suction Surface ◽  
Pressure Surface ◽  
Percent Increase ◽  
Compressor Performance ◽  
Overall Performance ◽  
Total Pressure Ratio

Abstract In the paper, a back swept impeller of centrifugal compressor is experimentally studied and numerically validated and modified to increase its pressure ratio and improve efficiency, as well as to analyse the effect of splitter blade location between two main blades. The back swept multi splitter blade impeller was designed with a big splitter positioned close to the main blade suction surface and a smaller splitter close to the pressure surface. Adding this multi splitter improves the overall performance of the modified impeller due to less intensive flow separation and smaller pressure loss. In particular, the total pressure ratio was observed to increase from 4.1 to 4.4, with one percent increase in efficiency.

Aerodynamic optimization design of aspirated highly loaded fan stage

2018 ◽  
Vol 233 (9) ◽  
pp. 3426-3444
Author(s):  
Jinhuan Zhang ◽  
Zhenggui Zhou ◽  
Cui Cui
Keyword(s):  
Total Pressure ◽  
Rotor Blade ◽  
Optimization Design ◽  
Design Method ◽  
Pressure Ratio ◽  
Stream Surface ◽  
Design Goal ◽  
Suction Flow ◽  
Total Pressure Ratio ◽  
Highly Loaded

In this paper, an aerodynamic design method for an aspirated compressor/fan is developed. In the S2 through-flow design, the loss feedback is used to solve the inapplicability of the conventional loss model. In S1 profile design, an optimization design method is constructed in which the profile and the suction flow parameters are simultaneously handled as design parameters to couple the optimization design. A 3D optimization method is used to modify the profiles at the hub and tip of the rotor blade and the sweep and lean of the stator blade. An aspirated highly loaded fan stage (load coefficient of 0.69) was designed using the design method. A flow field simulation shows that at the design point, with a modest suction flow (4.84% of the inlet mass flow), very high isentropic efficiency (0.9213) is achieved, and the total pressure ratio (3.445) achieves its design goal (3.40), but the mass flow rate of the designed fan stage is 6.2% lower than the design goal. From the comparisons between the 2D flow fields on the S1 stream surfaces and the 3D flow fields at the corresponding blade spans, it is concluded that the flow presents nearly a form of a 2D S1 stream surface at most of the spans, and the 2D design method which is based on the S1/S2 stream surface in this paper is effective. Moreover, the flow is analyzed around the rotor root of the aspirated rotor, revealing a weak flow capacity in that area. This result suggests that desirable flow might not be set up when the designed profile has a large camber at the rotor blade root because the total pressure ratio cannot be improved without compromising the static pressure ratio.

Effect of swirl on the performance and stability of transonic axial compressor

2017 ◽  
Vol 232 (6) ◽  
pp. 608-621
Author(s):  
Baofeng Tu ◽  
Luyao Zhang ◽  
Jun Hu
Keyword(s):  
Mass Flow ◽  
Total Pressure ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Maximum Efficiency ◽  
Stable Boundary ◽  
Tip Leakage ◽  
Stable Operating ◽  
Swirl Intensity ◽  
Total Pressure Ratio

To investigate the effect of twin swirl and bulk swirl on the performance and stability of a transonic axial compressor, a blade swirl generator was designed and simulated with a transonic single rotor using steady and unsteady numerical calculation methods. The bulk swirl intensity was adjusted by replacing the blades with different camber angles. The twin swirl intensity was decreased by reducing the blade number. The counter-rotating bulk swirl generated a significant drop in both the efficiency and stall margin, and resulted in an increase in the choked mass flow, and total pressure ratio. The co-rotating bulk swirl generated a decrease in the mass flow, total pressure ratio and stable operating range. The counter-rotating bulk swirl resulted in suction surface boundary layer separation beyond 50% of the span-wise height as well as more serious tip leakage blockage. The twin swirl resulted in a decrease in the total pressure ratio, maximum efficiency and stable operating range. The steady and unsteady numerical calculation results were consistent, though some differences were observed in the values. For the steady calculation, the maximum efficiency and choked mass flow decreased by 0.88% and 1.74%, respectively, and the mass flow at the stable boundary increased by 3.92% as compared to the uniform inlet flow at twin swirl intensity of 24°. During the unsteady calculation, the mass flow exhibited an increase of only 2.2% at the stable boundary. Under twin swirl and co-rotating bulk swirl and uniform inlet flow, the leading edge spillage of the tip leakage flow resulted in compressor instability. The counter-rotating bulk swirl changed the mechanism of instability. The characterisation of the swirl distortion presented a difference between the steady and unsteady calculations near the stable boundary. The unsteady calculation exhibited a lower mass flow at the stable boundary point and a higher total pressure ratio.

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

2021 ◽  
Vol 2021 ◽  
pp. 1-24
Author(s):  
Cui Cui ◽  
Zhenggui Zhou ◽  
Endor Liu
Keyword(s):  
Shock Waves ◽  
Total Pressure ◽  
Optimization Design ◽  
Pressure Ratio ◽  
Strong Shock ◽  
Stall Margin ◽  
Design Point ◽  
Compressor Rotor ◽  
Blade Tip ◽  
Total 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%.

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