Comprehensive Application of a First Principles Based Methodology for Design of Axial Compressor Configurations

2011 ◽  
Author(s):  
Vishwas Iyengar ◽  
Lakshmi N. Sankar
Keyword(s):  
First Principles ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Leading Edge ◽  
Aerodynamic Characteristics ◽  
Design Parameters ◽  
Axial Compressors ◽  
Design Variables ◽  
Improved Accuracy ◽  
Total Pressure Ratio

Axial compressors are widely used in many aerodynamic applications. The design of an axial compressor configuration presents many challenges. It is necessary to retool the design methodologies to take advantage of the improved accuracy and physical fidelity of these advanced methods. Here, a first-principles based multi-objective technique for designing single stage compressors is described. The study accounts for stage aerodynamic characteristics and rotor-stator interactions. The proposed methodology provides a way to systematically screen through the plethora of design variables. This method has been applied to a rotor-stator stage similar to NASA Stage 35. By selecting the most influential design parameters and by optimizing the blade leading edge and trailing edge mean camber line angles, phenomena such as tip blockages, blade-to-blade shock structures and other loss mechanisms can be weakened or alleviated. It is found that these changes to the configuration can have a beneficial effect on total pressure ratio and stage adiabatic efficiency, thereby improving the performance of the axial compression system.

Comprehensive Application of a First Principles Based Methodology for Design of Axial Compressor Configurations

2012 ◽  
Vol 134 (6) ◽  
Author(s):  
Vishwas Iyengar ◽  
Lakshmi N. Sankar
Keyword(s):  
First Principles ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Leading Edge ◽  
Aerodynamic Characteristics ◽  
Design Parameters ◽  
Axial Compressors ◽  
Design Variables ◽  
Improved Accuracy ◽  
Total Pressure Ratio

Axial compressors are widely used in many aerodynamic applications. The design of an axial compressor configuration presents many challenges. It is necessary to retool the design methodologies to take advantage of the improved accuracy and physical fidelity of these advanced methods. Here, a first-principles based multiobjective technique for designing single stage compressors is described. The study accounts for stage aerodynamic characteristics and rotor-stator interactions. The proposed methodology provides a way to systematically screen through the plethora of design variables. This method has been applied to a rotor-stator stage similar to NASA Stage 35. By selecting the most influential design parameters and by optimizing the blade leading edge and trailing edge mean camber line angles, phenomena such as tip blockages, blade-to-blade shock structures and other loss mechanisms can be weakened or alleviated. It is found that these changes to the configuration can have a beneficial effect on total pressure ratio and stage adiabatic efficiency, thereby improving the performance of the axial compression system.

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.

A Study of Influences of Inlet Total Pressure Distortions On Clearance Flow in an Axial Compressor

2021 ◽  
Author(s):  
Guoming Zhu ◽  
Xiaolan Liu ◽  
Bo Yang ◽  
Moru Song
Keyword(s):  
Total Pressure ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Leading Edge ◽  
Stall Margin ◽  
Flow Cells ◽  
Static Pressure Distribution ◽  
Speed Flow ◽  
Clearance Flow ◽  
The Stability

Abstract The rotating distortion generated by upstream wakes or low speed flow cells is a kind of phenomenon in the inlet of middle and rear stages of an axial compressor. Highly complex inflow can obviously affect the performance and the stability of these stages, and is needed to be considered during compressor design. In this paper, a series of unsteady computational fluid dynamics (CFD) simulations is conducted based on a model of an 1-1/2 stage axial compressor to investigate the effects of the distorted inflows near the casing on the compressor performance and the clearance flow. Detailed analysis of the flow field has been performed and interesting results are concluded. The distortions, such as total pressure distortion in circumferential and radial directions, can block the tip region so that the separation loss and the mixing loss in this area are increased, and the efficiency and the total pressure ratio are dropped correspondingly. Besides, the distortions can change the static pressure distribution near the leading edge of the rotor, and make the clearance flow spill out of the rotor edge more easily under near stall condition, especially in the cases with co-rotating distortions. This phenomenon can be used to explain why the stall margin is deteriorated with nonuniform inflows.

Numerical Efforts of Aerodynamic Re-Design in a Single-Stage Transonic Axial Compressor: Part 1 — Stator Design

2017 ◽  
Author(s):  
Justin (Jongsik) Oh
Keyword(s):  
Pressure Ratio ◽  
Axial Compressor ◽  
Axial Flow ◽  
Single Stage ◽  
Design Flow ◽  
Design Parameters ◽  
Original Design ◽  
Circular Arc ◽  
Surge Margin ◽  
Flow Compressor

In many aerodynamic design parameters for the axial-flow compressor, three variables of tailored blading, blade lean and sweep were considered in the re-design efforts of a transonic single stage which had been designed in 1960’s NASA public domains. As Part 1, the re-design was limited to the stator vane only. For the original MCA (Multiple Circular Arc) blading, which had been applied at all radii, the CDA (Controlled Diffusion Airfoil) blading was introduced at midspan as the first variant, and the endwalls of hub and casing (or tip) were replaced with the DCA (Double Circular Arc) blading for the second variant. Aerodynamic performance was predicted through a series of CFD analysis at design speed, and the best aerodynamic improvement, in terms of pressure ratio/efficiency and operability, was found in the first variant of tailored blading. It was selected as a baseline for the next design efforts with blade lean, sweep and both combined. Among 12 variants, a case of positively and mildly leaned blades was found the most attractive one, relative to the original design, providing benefits of an 1.0% increase of pressure ratio at design flow, an 1.7% increase of efficiency at design flow, a 10.5% increase of the surge margin and a 32.3% increase of the choke margin.

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.

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.

Off-Design Prediction of Transonic Axial Compressors: Part 1 — Mean-Line Code and Tuning Factors

2018 ◽  
Author(s):  
John Kidikian ◽  
Marcelo Reggio
Keyword(s):  
Pressure Ratio ◽  
Axial Compressor ◽  
Design Tool ◽  
Information Storage ◽  
Axial Compressors ◽  
Design Performance ◽  
Engineering Software ◽  
Line Loss ◽  
Loss Modelling ◽  
Line Analysis

With yearly advances in CFD techniques and methodologies, and the increased capacity and capabilities of computer CPU, GPU, and information storage, CFD has become a powerful design tool. However, despite its vast strengths, a CFD analysis is still based on the sound development of the 1D mean-line analysis methodology. This paper (part 1 of 2) describes an off-design axial compressor mean-line code, tested in a specialized engineering software for the development and analysis of a whole gas turbine engine, and the various tuning factors used to obtain an off-design performance match. It will be shown that, to obtain a proper match of the off-design performance of single-stage transonic axial compressors, both the rotor and stage pressure ratio, and the rotor temperature ratio are required to be converged upon. To do so, the off-design mean-line analysis requires the incorporation of a set of inlet & exit blockage factors and deviation angles that vary with the compressor performance conditions. This approach differs from the literature-based procedural assumptions (or rule-of-thumb) of fixed inlet and exit blockage factors of approximately “0.98”, and the use of a unique deviation angle based on Carter’s rule. The results obtained in this paper are then used to develop a generalized off-design mean-line loss modelling methodology (part 2 of 2) capable of predicting the off-design performance of four well documented NASA transonic axial compressors.

Analyses and Optimization of a Propulsion Cycle for Unmixed High Bypass Turbofan

2008 ◽  
Author(s):  
Adel Ghenaiet
Keyword(s):  
Fuel Consumption ◽  
Pressure Ratio ◽  
Specific Fuel Consumption ◽  
Design Parameters ◽  
Inlet Temperature ◽  
Turbine Inlet Temperature ◽  
High Pressure Ratio ◽  
Turbine Inlet ◽  
Design Variables ◽  
Bypass Ratio

This paper deals with a parametric study and an optimization for the design variables of a high bypass unmixed turbofan equipping commercial aircrafts. The objective of the first part of this study is to highlight the effects of the principal design parameters (bypass ratio, compression ratios, turbine inlet temperature etc..) on the uninstalled performance, in terms of specific thrust and specific fuel consumption. The second part concerns the optimization, aiming at finding the optimum design parameters concurrently minimizing the specific fuel consumption at cruise, and meeting the thrust requirement at takeoff. The cycle analyzer (on-design and off-design) as coupled to the optimization algorithm MMFD by adopting a random multi-starts search strategy is shown to be stable and converging. The predefined requirements and constraints have dictated utilizing an engine with a high-bypass ratio, high-pressure ratio and a moderate turbine inlet temperature. In general, the obtained results compare fairly well with typical data available for an equivalent ‘reference’ engine. This elaborated methodology is shown to be consistent with the conceptual design requirements and accuracy, because, it does not use components’ characteristics, and operates on simplifying assumptions. This present methodology can be readily adapted for other configurations of aero-engines as well, and easily integrated in a multi-disciplinary design approach.

Optimization of Highly Loaded Fan Rotor Based on Throughflow Model

2007 ◽  
Author(s):  
Hong Wu ◽  
Qiushi Li ◽  
Sheng Zhou
Keyword(s):  
Mass Flow ◽  
Total Pressure ◽  
Pressure Ratio ◽  
Optimization Method ◽  
Design Variables ◽  
Adaptive Simulated Annealing ◽  
Throughflow Model ◽  
Important Design ◽  
Total Pressure Ratio ◽  
Highly Loaded

This paper presents an optimization method for fan/compressor which couples throughflow model solving axisymmetric Euler equations with adaptive simulated annealing (ASA) algorithm. One of the advantages of this optimization method is that it spends much less time than 3D optimization due to the rapid solving of throughflow model. In addition, the optimization space is quite extensive because more design variables can be adjusted in throughflow phase, such as swirl distribution, hub curve and sweep. To validate this optimization method, a highly loaded fan rotor with pressure ratio of 3.06 as a baseline is optimized. During the optimization process, the objective function is total pressure ratio, moreover, mass flow and efficiency are selected as the constraint conditions. Three important design variables including swirl distribution, hub curve and sweep are parameterized using Bezier curve, and then optimized in throughflow model independently, finally the optimum designs are validated using 3D viscous CFD solver. It is shown that pressure ratio and rotor loading can be improved further through optimizing swirl distribution, however, hub and sweep curves take more effects on mass flow and efficiency respectively. The optimization results demonstrate the advantage and feasibility of this optimization method.

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.

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