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

2021 ◽  
pp. 095441002110490
Author(s):  
Guang Wang ◽  
Wuli Chu ◽  
Haoguang Zhang ◽  
Zhentao Guo
Keyword(s):  
Total Pressure ◽  
Peak Velocity ◽  
Active Flow Control ◽  
Pressure Ratio ◽  
Excitation Frequency ◽  
Axial Compressor ◽  
Synthetic Jet ◽  
Optimal Excitation ◽  
Active Flow ◽  
Total Pressure Ratio

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.

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

2020 ◽  
Author(s):  
Zijing Chen ◽  
Bo Liu ◽  
Xiaoxiong Wu
Keyword(s):  
Genetic Algorithm ◽  
Total Pressure ◽  
Fitness Function ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Local Data ◽  
Real Coded Genetic Algorithm ◽  
Overall Efficiency ◽  
Total Pressure Ratio ◽  
Peak Value

Abstract 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.

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.

Numerical Investigation of Active Flow Control Using Tip Synthetic Jet on the Performance of a High-Speed Axial Compressor Rotor

2020 ◽  
Author(s):  
Guang Wang ◽  
Wuli Chu
Keyword(s):  
Flow Control ◽  
High Speed ◽  
Active Flow Control ◽  
Axial Compressor ◽  
Synthetic Jet ◽  
Tip Leakage Flow ◽  
Tip Leakage ◽  
Compressor Rotor ◽  
The Stability ◽  
Active Flow

Abstract In order to weaken the negative effect of tip leakage flow and improve the tip flow condition, this paper introduces synthetic jet into the flow control field of axial compressor, and proposes a method of active flow control by arranging synthetic jet at the tip. A high-speed axial compressor rotor of the author’s research group is taken as the numerical simulation object. On the basis of keeping geometric parameters of the synthetic jet actuator unchanged, this paper studies the influence of applying tip synthetic jet on aerodynamic performance of the compressor rotor at three axial positions of −10%Ca, 0%Ca and 21.35%Ca respectively. The results show that when tip synthetic jet is in the above three positions, comprehensive stability margin of the compressor rotor increases by 2.62%, 3.77% and 12.46% respectively, and efficiency near stall point increases by 0.22%, 0.25 and 0.47% respectively. This shows that when tip synthetic jet is far away from blade, the aerodynamic performance improvement of the compressor rotor is limited, and when tip synthetic jet is just above the leading edge, the effect of expanding stability is the best and the efficiency is the most improved. The mechanism of tip synthetic jet can increase the stability of the compressor rotor is that when the actuator is in the blowing stage, it can blow the low-speed air flow of blade top to downstream, and when the actuator is in the suction stage, it can suck the low-speed air flow of blade top into slot, so as to alleviate the top blockage and realize the stability expansion. The mechanism of tip synthetic jet can improve the efficiency of compressor rotor is that the blowing and suction of actuator weaken the intensity of tip leakage flow, reduce the size of vortex core and also reduce the flow loss of the compressor rotor correspondingly.

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.

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.

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.

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

PLoS ONE ◽  
2021 ◽  
Vol 16 (1) ◽  
pp. e0245711
Author(s):  
Asad Islam ◽  
Hongwei Ma
Keyword(s):  
Total Pressure ◽  
Numerical Study ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Aerodynamic Performance ◽  
Ansys Cfx ◽  
Transonic Axial Compressor ◽  
Total Pressure Ratio ◽  
Flow Case ◽  
Probe Block

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.

Time Resolved Simulation and Experimental Validation of the Flow in Axial Slot Casing Treatments for Transonic Axial Compressors

2008 ◽  
Author(s):  
Giovanni A. Brignole ◽  
Florian C. T. Danner ◽  
Hans-Peter Kau
Keyword(s):  
Mass Flow ◽  
Total Pressure ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Compressor Stage ◽  
Time Resolved ◽  
Transonic Compressor ◽  
Design Speed ◽  
Casing Treatments ◽  
Total Pressure Ratio

Building on the experience of previous investigations, a casing treatment was developed and applied to an axial transonic compressor stage, in literature referred to as Darmstadt Rotor 1. The aerodynamics of the experimental compressor stage was improved by applying axially orientated semicircular slots to the original plain casing, which both enhanced the operating range and design point efficiency. A gain in total pressure ratio along the entire design speed line was also observed. Within the scope of this study four different axial casing treatments were designed. Their effect on the flow in a transonic axial compressor stage was investigated parametrically using time-resolved 3D-FANS simulations with a mesh of approximately 4.8 · 106 grid points. This research aims to identify correlations between the geometrical cavity design and the changed channel flow. The findings help to formulate parameters for evaluating the performance of casing treatments. These criteria can further be used as target functions in the design optimisation process. The predicted behaviour of the transonic compressor was validated against experiments as well as an alternative numerical model, the non-linear harmonic method. Both confirmed the effect of the slots in raising efficiency as well as moving the design speed line towards higher pressure ratios. In the experiments, the addition of the slots increased the total pressure ratio at stall conditions by more than 5% and reduced mass flow from 87.5% of the design mass flow to less than 77.5% compared to the original geometry.

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