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%.

Numerical Studies on the Effect of Gurney Flap on Aerodynamic Performance and Stall Margin of a Transonic Axial Compressor Rotor

2014 ◽  
Author(s):  
Mudassir Ahmed M. Rafeeq ◽  
Quamber H. Nagpurwala ◽  
Subbaramu Shivaramaiah
Keyword(s):  
Total Pressure ◽  
Rotor Blade ◽  
Peak Pressure ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Stall Margin ◽  
Compressor Rotor ◽  
Numerical Studies ◽  
Gurney Flap ◽  
Total Pressure Ratio

Numerical studies have been carried out on the effectiveness of trailing edge Gurney flap on a transonic axial compressor rotor. The baseline geometry of the rotor blade was modified at the trailing edge by introducing Gurney flaps of varying depth and span-wise length, viz. 1 mm, 2 mm and 3 mm depth with 20% span length of Gurney flap from tip (designated as GF1-20, GF2-20 and GF3-20 respectively), and 1 mm depth with 50% and 100% span length (designated as GF1-50 and GF1-100 respectively). Geometric models of the compressor rotor without and with Gurney flaps were generated using CATIA V5 software and CFD simulations at 100% design rotor speed were carried out using ANSYS CFX software. Results have shown that the compressor total pressure ratio increased with increase in both depth and spanwise length of Gurney flap. Peak pressure ratio increased from 1.51 for baseline case to 1.58 for rotor GF1-100. However, the peak isentropic efficiency remained almost constant for various Gurney flap configurations, except for GF1-100 which showed a tendency for improvement in efficiency. The stall margin reduced with the introduction of Gurney flap and was lowest for configuration GF1-100 which gave highest peak pressure ratio. Higher blade loading with Gurney flap was responsible for lowering the stall margin. Analysis of the flow through the blade passages has shown clear formation of trailing end vortex structure in the presence of Gurney flap that resulted in bending of the streamlines towards suction surface of the rotor blade, with consequent reduction in flow deviation and increased flow deflection, and hence increased total pressure ratio.

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

2020 ◽  
Author(s):  
Kewei Xu ◽  
Gecheng Zha
Keyword(s):  
Pressure Ratio ◽  
Numerical Approach ◽  
Efficiency Loss ◽  
Stall Margin ◽  
Design Point ◽  
Casing Treatment ◽  
Final Design ◽  
Upstream Flow ◽  
Total Pressure Ratio ◽  
Good Agreement

Abstract 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%.

scholarly journals Design and Optimization of Multiple Circumferential Casing Grooves Distribution Considering Sweep and Lean Variations on the Blade Tip

Energies ◽  
2018 ◽  
Vol 11 (9) ◽  
pp. 2401
Author(s):  
Weimin Song ◽  
Yufei Zhang ◽  
Haixin Chen
Keyword(s):  
Pressure Ratio ◽  
Three Dimensional ◽  
Axial Distribution ◽  
Stall Margin ◽  
Design Point ◽  
Casing Treatment ◽  
Design And Optimization ◽  
Practical Engineering ◽  
Engineering Strategy ◽  
Blade Tip

This paper focuses on the design and optimization of the axial distribution of the circumferential groove casing treatment (CGCT). Effects of the axial location of multiple casing grooves on the flow structures are numerically studied. Sweep and lean variations are then introduced to the blade tip, and their influences on the grooves are discussed. The results show that the ability of the CGCT to relieve the blockage varies with the distribution of grooves, and the three-dimensional blading affects the performance of both the blade and the CGCT. Accordingly, a multi-objective optimization combining the CGCT design with the sweep and lean design is conducted. Objectives, including the total pressure ratio and the adiabatic efficiency, are set at the design point; meanwhile, the choking mass flow and the near-stall performance are constrained. The coupling between the CGCT and the blade is improved, which contributes to an optimal design point performance and a sufficient stall margin. The sweep and lean in the tip redistribute the spanwise and chordwise loading, which enhances the ability of the CGCT to improve the blade’s performance. This work shows that the present CGCT-blade integrated optimization is a practical engineering strategy to develop the working capacity and efficiency of a compressor blade while achieving the stall margin extension.

Numerical Investigation on the Effects of Circumferential Coverage of Injection in a Transonic Compressor With Discrete Tip Injection

2014 ◽  
Author(s):  
Wei Wang ◽  
Wuli Chu ◽  
Haoguang Zhang ◽  
Yanhui Wu
Keyword(s):  
Total Pressure ◽  
Pressure Ratio ◽  
Time Lag ◽  
Three Dimensional ◽  
Stability Margin ◽  
Transonic Compressor ◽  
Blade Tip ◽  
Tip Injection ◽  
The Stability ◽  
Total Pressure Ratio

Discrete tip injection upstream of the rotor tip is an effective technique to extend stability margin for a compressor system in an aeroengine. The current study investigates the effects of injectors’ circumferential coverage on compressor performance and stability using time-accurate three-dimensional numerical simulations for multi passages in a transonic compressor. The percentage of circumferential coverage for all the six injectors ranges from 6% to 87% for the five investigated configurations. Results indicate that circumferential coverage of tip injection can greatly affect compressor stability and total pressure ratio, but has little influence on adiabatic efficiency. The improvement of compressor total pressure ratio is linearly related with the increasing circumferential coverage. The unsteady flow fields show that there exists a non-ignorable time lag of the injection effects between the passage inlet and outlet, and blade tip loading will not decline until the injected flow reaches the passage outlet. Stability improves sharply with the increasing circumferential coverage when the coverage is less than 27%, but increases flatly for the rest. It is proven that the injection efficiency which is a measurement of averaged blockage decrement in the injected region is an effective guideline to predict the stability improvement.

Effect of inlet distortion on the performance of axial transonic contra-rotating compressor

2016 ◽  
Vol 232 (1) ◽  
pp. 42-54 ◽  
Author(s):  
Hanru Liu ◽  
Yangang Wang ◽  
Songchuan Xian ◽  
Wenbin Hu
Keyword(s):  
Total Pressure ◽  
Pressure Ratio ◽  
Leading Edge ◽  
Tip Leakage Flow ◽  
Stall Margin ◽  
Tip Leakage ◽  
Inlet Distortion ◽  
Flow Mixing ◽  
Inflow Conditions ◽  
Total Pressure Ratio

The present paper numerically conducted full-annulus investigation on the effects of circumferential total pressure inlet distortion on the performance and flow field of the axial transonic counter-rotating compressor. Results reveal that the inlet distortion both deteriorates the performance of the upstream and downstream rotors resulting in reduction of total pressure ratio, efficiency and stall margin of the transonic contra-rotating compressor. Regarding the development of distortion inside compressor, the downstream rotor reinforces the air-flow mixing effects and, thus, attenuates the distortion intensity significantly. Under the distorted inflow conditions, the detached shockwave at the leading edge of downstream rotor interacts with the tip leakage flow and causes the blockage of the blades passage, which is one important reason for the transonic contra-rotating compressor stall.

Effect of Trailing Edge Crenulation on the Performance and Stall Margin of a Transonic Axial Compressor

2013 ◽  
Author(s):  
S. Subbaramu ◽  
Quamber H. Nagpurwala ◽  
A. T. Sriram
Keyword(s):  
Beneficial Effect ◽  
Total Pressure ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Compressor Cascade ◽  
Trailing Edge ◽  
Stall Margin ◽  
Positive Effects ◽  
Compressor Rotor ◽  
Transonic Axial Compressor

This paper deals with the numerical investigations on the effect of trailing edge crenulation on the performance of a transonic axial compressor rotor. Crenulation is broadly considered as a series of small notches or slots at the edge of a thin object, like a plate. Incorporating such notches at the trailing edge of a compressor cascade has shown beneficial effect in terms of reduction in total pressure loss due to enhanced mixing in the wake region. These notches act as vortex generators to produce counter rotating vortices, which increase intermixing between the free stream flow and the low momentum wake fluid. Considering the positive effects of crenulation in a cascade, it was hypothesized that the same technique would work in a rotating compressor to enhance its performance and stall margin. However, the present CFD simulations on a transonic compressor rotor have given mixed results. Whereas the peak total pressure ratio in the presence of trailing edge crenulation reduced, the stall margin improved by 2.97% compared to the rotor with straight edge blades. The vortex generation at the crenulated trailing edge was not as strong as reported in case of linear compressor cascade, but it was able to influence the flow field in the rotor tip region so as to energize the low momentum end-wall flow in the aft part of the blade passage. This beneficial effect delayed flow separation and allowed the mass flow rate to be reduced to still lower levels resulting in improved stall margin. The reduction in pressure ratio with crenulation was surprising and might be due to increased mixing losses downstream of the blade.

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.

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

2020 ◽  
Author(s):  
Song Huang ◽  
Chuangxin Zhou ◽  
Chengwu Yang ◽  
Shengfeng Zhao ◽  
Mingyang Wang ◽  
...  
Keyword(s):  
Total Pressure ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Aerodynamic Performance ◽  
Blade Loading ◽  
Stall Margin ◽  
Axial Compressors ◽  
Design Speed ◽  
Total Pressure Ratio ◽  
Highly Loaded

Abstract 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.

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.

Design of a Highly Loaded Transonic Two-Stage Fan Using Swept and Bowed Blading

2011 ◽  
Author(s):  
Hailiang Jin ◽  
Donghai Jin ◽  
Fang Zhu ◽  
Ke Wan ◽  
Xingmin Gui
Keyword(s):  
Total Pressure ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Leading Edge ◽  
Navier Stokes ◽  
Two Stage ◽  
Stall Margin ◽  
Stator Vane ◽  
Total Pressure Ratio ◽  
Highly Loaded

This paper presents the design of a highly loaded transonic two-stage fan using several advanced three-dimensional blading techniques including forward sweep and “hub bending” in rotors and several bowed configurations in stators. The effects of these blading techniques on the performance of the highly loaded transonic two-stage fan were investigated on the basis of three-dimensional Navier-Stokes predictions. The results indicate that forward sweep has insignificant impact on the total pressure ratio and adiabatic efficiency of the fan. The throttling range of the fan is found to be improved by forward sweep because the shock in the forward swept rotor is expelled later upstream to the leading edge than that in the unswept one. Hub bending design technique increases the efficiency in the hub region of R1 due to the reduction of the low momentum zone in the hub region near the trailing edge. The stator vane design has a pronounced impact on the performance of the fan. The total pressure ratio, adiabatic efficiency, and stall margin of the schemes with the bowed vanes are increased significantly compared to the scheme with the straight vanes. The large corner stall in the straight S1 vane is reduced effectively by the bowed S1 vanes. Moreover, the strong corner stall in the straight S2 vane is fully eliminated by the bowed S2 vanes. Among the bowed vane schemes, the scheme with positive bowed (P. B.) hub and negative bowed (N. B.) tip vanes has the best efficiency and stall margin performances thanks to the superiority of the performance over the midspan regions of the bowed vanes.

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