Multi Disciplinary Design Optimization and Performance Evaluation of a Single-Stage Transonic Axial Compressor

2012 ◽  
Author(s):  
Young-Seok Kang ◽  
Tae-Choon Park ◽  
Soo-Seok Yang ◽  
Sae-Il Lee ◽  
Dong-Ho Lee
Keyword(s):  
Safety Factor ◽  
Incidence Angle ◽  
Axial Compressor ◽  
Aerodynamic Performance ◽  
Design Constraints ◽  
Flow Angle ◽  
Blade Loading ◽  
Final Design ◽  
Outlet Flow ◽  
Transonic Axial Compressor

The multi-disciplinary optimization (MDO) method, which integrates aerodynamic performance and structural stability, was utilized in the development of a single-stage transonic axial compressor. Numerical simulations and compressor tests were also carried out to evaluate the aerodynamic performance and safety factor of the optimized compressor. The rotor has 60 design parameters with twelve most sensitive design variables selected for design optimization. The stator was redesigned according to the rotor outlet flow angle variation to match the stator incidence angle by −1∼0 degrees, while maintaining the stage outlet flow angle. The design goal is to maximize both the stage efficiency and the safety factor from the baseline scratch compressor design. The object function is composed of the normalized efficiency and safety factor with weight factors. Initially, an approximation model was created to search for the global optimization within given ranges of variables and considering several design constraints. The genetic algorithm was used to explore the Pareto front of the optimization to find the maximum objective function values. The final design was chosen after a second stage gradient-based optimization process to improve the accuracy of the optimization. The CFD results showed that more blade loading is burden to the hub region by increasing the incidence angle. The fore part blade loading gradually decreases along the span-wise direction. In addition, normal shock, which spreads along the hub to the blade tip, is confined in the rotor flow passage and pressure surface shock coincidence point moves to be closer to the blade leading edge, indicating an increase in the amount of blade loading. FEA analyses showed that the blade root stress has been drastically relieved, because the optimized blade has trapezoid-shaped hub design relative to the baseline design. The final design achieved efficiency gain of 3.69% and showed a higher safety factor by 2.3 times relative to the baseline model, while maintaining its stage mass flow rate and total-to-total pressure within the design constraints. The compressor performance test data showed good agreement with the optimized design and CFD results. However, there is room for improvement in the optimization process to reflect off-design performance so as to secure more stable compressor operation ranges.

Design Optimization of a Single-Stage Transonic Axial Compressor and Test Evaluation of Its Aerodynamic Performance

2012 ◽  
Vol 15 (6) ◽  
pp. 77-84 ◽  
Author(s):  
Tae Choon Park ◽  
Young-Seok Kang ◽  
Oh-Sik Hwang ◽  
Ji-Han Song ◽  
Byeung Jun Lim
Keyword(s):  
Design Optimization ◽  
Axial Compressor ◽  
Aerodynamic Performance ◽  
Single Stage ◽  
Test Evaluation ◽  
Transonic Axial Compressor

Application of Smith Chart for Non-Repeating Stages in Axial Compressors

2013 ◽  
Author(s):  
Daniel Hernández ◽  
Antonio Antoranz ◽  
Raúl Vázquez
Keyword(s):  
Early Stage ◽  
Axial Compressor ◽  
Preliminary Design ◽  
The Novel ◽  
Flow Angle ◽  
Axial Compressors ◽  
Absolute Flow ◽  
Final Design ◽  
Semi Empirical ◽  
Smith Chart

The configuration of an axial compressor, including the mean radius, the annulus lines, stage loading or number of stages, flow parameter, work split and stage reactions, are all of them selected in the preliminary design phase. For the success of the final design, to attain the proper selection is mandatory. A representative geometry of the airfoils is not available at this early stage of the design process. Therefore the former parameters use to be selected based only on the designer prior experience and/or empirical correlations. Under these circumstances, the so called Smith Chart is a valuable tool that can provide simple guidelines to the designer and a preliminary assessment of the compressor efficiency. The use of this chart can be also extended to get the main features of the airfoils, like flow angles, turning, Mach and Reynolds number, diffusion factor, aspect ratio, etc. as well as to compare different design candidates. Several authors have produced their own diagrams by analytical or semi-empirical approach. The repeating stage hypothesis, which has been usually assumed, implies no change in inlet and outlet absolute flow angles and constant axial velocity throughout the stage. The density rise through the stage is compensated by reducing the annulus height and so the annulus wall slope along the compressor is directly obtained from the continuity equation, being in most of the cases not representative of real compressors. In order to have a more representative annulus, in the present work, the repeating stage hypothesis has not been assumed. The annulus shape (height and slope wall angle) is therefore defined by the designer and in order to close the equations of the problem, the absolute exit flow angle of every stage is required. The optimization of the compressor by the novel proposed method is more complicate because of the higher number of variables. However the method has the advantage to reduce the design iterations due to its more reliable results. The aim of this paper is to introduce the novel method of non-repeating stages and to show how this approach can be used in the preliminary design of an axial compressor.

AERODYNAMIC PERFORMANCE OF AN AXIAL COMPRESSOR WITH A CASING GROOVE COMBINED WITH INJECTION

2013 ◽  
Vol 37 (3) ◽  
pp. 283-292 ◽  
Author(s):  
Dae-Woong Kim ◽  
Jin-Hyuk Kim ◽  
Kwang-Yong Kim
Keyword(s):  
Turbulence Model ◽  
Parametric Study ◽  
Stokes Equations ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Aerodynamic Performance ◽  
Navier Stokes ◽  
Navier Stokes Equations ◽  
Stall Margin ◽  
Transonic Axial Compressor

Aerodynamic performance of a transonic axial compressor with a casing groove combined with injection has been investigated in this work. Three-dimensional Reynolds-averaged Navier–Stokes equations with k-ε turbulence model are discretized by finite volume approximations and solved on hexahedral grids for the flow analyses. For parametric study, the front and rear lengths and height of the casing groove are selected as the geometric parameters and are changed with constant injection to investigate their effects on the stall margin and peak adiabatic efficiency. As a result of the parametric study, the maximum stall margin and peak adiabatic efficiency are found to be obtained in the axial compressor having 70% height of the reference groove. The results show that the application of the casing groove combined with injection to an axial compressor is effective for the simultaneous improvement of both the stall margin and peak adiabatic efficiency of the compressor.

scholarly journals Development of mathematical model and computer program for primary design of transonic axial compressor

2020 ◽  
Vol 4 (4) ◽  
pp. 16-27
Author(s):  
A. I. Borovkov ◽  
◽  
Yu. B. Galerkin ◽  
O. A. Solovieva ◽  
A. A. Drozdov ◽  
...  
Keyword(s):  
Mathematical Model ◽  
Pressure Loss ◽  
Incidence Angle ◽  
Axial Compressor ◽  
Compressor Stage ◽  
Empirical Coefficient ◽  
Axial Compressors ◽  
Profile Losses ◽  
Transonic Axial Compressor ◽  
The Mathematical Model

The mathematical model underlying the program for calculating and designing axial compressors is presented. The process of calculating the pressure loss in the elements of the axial compressor stage flow path is described. The loss coefficient consists of losses on the limiting surfaces, secondary losses and profile losses. The effect of roughness on the pressure loss is taken into account by introducing the corresponding empirical coefficient. An algorithm for calculating the blades and vanes angles of the impeller and the guide apparatus is presented by calculating the incidence angle and the lag angle of the flow. The flow lag angle is the sum of the lag angle of the flow on the profile and the lag angle due to viscous flow on the limiting surfaces

scholarly journals Design Optimization of a Dual-Bleeding Recirculation Channel to Enhance Operating Stability of a Transonic Axial Compressor

Energies ◽  
2021 ◽  
Vol 15 (1) ◽  
pp. 159
Author(s):  
Tien-Dung Vuong ◽  
Kwang-Yong Kim
Keyword(s):  
Design Optimization ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Aerodynamic Performance ◽  
Maximum Pressure ◽  
Stall Margin ◽  
Design Variables ◽  
Operating Stability ◽  
Transonic Axial Compressor

The present work performed a comprehensive investigation to find the effects of a dual-bleeding port recirculation channel on the aerodynamic performance of a single-stage transonic axial compressor, NASA Stage 37, and optimized the channel’s configuration to enhance the operating stability of the compressor. The compressor’s performance was examined using three parameters: The stall margin, adiabatic efficiency, and pressure ratio. Steady-state three-dimensional Reynolds-averaged Navier–Stokes analyses were performed to find the flow field and aerodynamic performance. The results showed that the addition of a bleeding channel increased the recirculation channel’s stabilizing effect compared to the single-bleeding channel. Three design variables were selected for optimization through a parametric study, which was carried out to examine the influences of six geometric parameters on the channel’s effectiveness. Surrogate-based design optimization was performed using the particle swarm optimization algorithm coupled with a surrogate model based on the radial basis neural network. The optimal design was found to increase the stall margin by 51.36% compared to the case without the recirculation channel with only 0.55% and 0.28% reductions in the peak adiabatic efficiency and maximum pressure ratio, respectively.

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.

Investigations Into Aerodynamic Performance of Turbine Stages With Flexible Shroud Seals

2020 ◽  
Author(s):  
Xin Yan ◽  
Yabo Wang ◽  
Kang Zhang ◽  
Xinbo Dai ◽  
Kun He
Keyword(s):  
Steam Turbine ◽  
Flow Patterns ◽  
Aerodynamic Performance ◽  
Original Design ◽  
Flow Angle ◽  
Large Power ◽  
Aerodynamic Efficiency ◽  
Forward Bending ◽  
Outlet Flow ◽  
Reaction Degree

Abstract The present paper utilizes a numerical method to investigate the effect of flexible shroud seals, including the forward bending flexible seals and backward bending flexible seals, on aerodynamic performance of high pressure steam turbine stages. At first, the wear performance of flexible seal is analyzed with the Finite Element Analysis method. It shows that wear in flexible strip is so small that only the installation clearance needs to be considered in operation process. Then, by replacing the labyrinth shroud seals with flexible shroud seals, the aerodynamic efficiency, outlet flow angle distributions, and reaction degree distributions in two-stages are obtained. At three installation clearances, interactions between leakage flow and main flow, as well as the flow patterns in flexible shroud seals, are visualized and also compared with the original design case. The numerical results indicate that turbine stages configured with forward bending flexible shroud seals have a very close aerodynamic performance to those configured with conventional labyrinth shroud seals at the same clearance, whereas the turbine stages configured with backward bending flexible seals have lower total-total isentropic efficiency than those with conventional labyrinth shroud seals. By replacing the conventional labyrinth shroud seals with forward bending flexible shroud seals (at the same clearance), the aerodynamic efficiency, outlet flow angles, limiting streamlines, secondary flow patterns in shroud region and reaction degree distributions in stages are almost not affected. Since the forward bending flexible seal allows relatively smaller installation clearance than the conventional labyrinth seal, application of this kind of seal in rotor blade tip gap is much beneficial to achieve lower leakage rate and higher aerodynamic performance in large power steam turbine stages.

Parametric Study on Aerodynamic Performance of a Transonic Axial Compressor with a Casing Groove and Tip Injection

2013 ◽  
Vol 284-287 ◽  
pp. 872-877 ◽  
Author(s):  
Dae Woong Kim ◽  
Jin Hyuk Kim ◽  
Kwang Yong Kim
Keyword(s):  
Parametric Study ◽  
Stokes Equations ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Aerodynamic Performance ◽  
Navier Stokes ◽  
Stall Margin ◽  
Transonic Axial Compressor ◽  
Tip Injection

This paper presents a parametric study on aerodynamic performance of a transonic axial compressor combined with a casing groove and tip injection using three-dimensional Reynolds-average Navier-Stokes equations. The front and rear lengths and height of the groove are selected as the geometric parameters to investigate their effects on the stall margin and peak adiabatic efficiency. These parameters are changed with constant injection. The validation of the numerical results is performed in comparison with experimental data for the total pressure ratio and adiabatic efficiency. As the results of the parametric study, the maximum stall margin and peak adiabatic efficiency are obtained in the axial compressor having 70% groove height of the reference groove. The stall margin and peak adiabatic efficiency in other cases are also improved in comparison with the axial compressors with the smooth casing and reference groove. The results show that both the stall margin and the peak adiabatic efficiency are considerably improved by the application of the casing groove combined with tip injection in an axial compressor.

Sign in / Sign up

Export Citation Format

Share Document