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.

Off-Design Prediction of Transonic Axial Compressors: Part 2 — Generalized Mean-Line Loss Modelling Methodology

2018 ◽  
Author(s):  
John Kidikian ◽  
Marcelo Reggio
Keyword(s):  
Total Pressure ◽  
Pressure Losses ◽  
Axial Compressors ◽  
Design Performance ◽  
Blade Row ◽  
Performance Map ◽  
Modelling Methodology ◽  
Generalized Mean ◽  
Line Loss ◽  
Loss Modelling

In Part 1, of this two-part paper, an off-design mean-line code and a generalized methodology to obtain “tuning” factors were presented. It was shown that the modified factors were capable of predicting the off-design performance of four well documented NASA transonic axial compressors. In this paper, Part 2, a generalized methodology to create correlations for the rotor and stator total pressure losses, deviation angles, and blade row inlet and exit blockage factors is presented. The generalized mean-line loss modelling methodology will allow the compressor designer to decommission the use of the performance map scaling techniques. In its place, the generalized predictive methodology will accurately estimate the off-design performance of transonic axial compressors and can be used to fill the gaps of missing data.

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.

scholarly journals Performance Improvement of Axial Compressors and Fans with Plasma Actuation

2009 ◽  
Vol 2009 ◽  
pp. 1-13 ◽  
Author(s):  
Sebastien Lemire ◽  
Huu Duc Vo ◽  
Michael W. Benner
Keyword(s):  
Computational Study ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Suction Side ◽  
Boundary Layer Separation ◽  
Plasma Actuator ◽  
Control Device ◽  
Axial Compressors ◽  
Compressor Rotor ◽  
Flow Control Device

This paper proposes the use of plasma actuator to suppress boundary layer separation on a compressor blade suction side to increase axial compressor performance. Plasma actuators are a new type of electrical flow control device that imparts momentum to the air when submitted to a high AC voltage at high frequency. The concept presented in this paper consists in the positioning of a plasma actuator near the separation point on a compressor rotor suction side to increase flow turning. In this computational study, three parameters have been studied to evaluate the effectiveness of plasma actuator: actuator strength, position and actuation method (steady versus unsteady). Results show that plasma actuator operated in steady mode can increase the pressure ratio, efficiency, and power imparted by the rotor to the air and that the pressure ratio, efficiency and rotor power increase almost linearly with actuator strength. On the other hand, the actuator's position has limited effect on the performance increase. Finally, the results from unsteady simulations show a limited performance increase but are not fully conclusive, due possibly to the chosen pulsing frequencies of the actuator and/or to limitations of the CFD code.

scholarly journals Influence of Axial Compressor Fouling on Gas Turbine Unit Perfomance Based on Different Schemes and With Different Initial Parameters

1998 ◽  
Author(s):  
A. P. Tarabrin ◽  
V. A. Schurovsky ◽  
A. I. Bodrov ◽  
J.-P. Stalder
Keyword(s):  
Gas Turbine ◽  
Quantitative Evaluation ◽  
Turbine Unit ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Gas Turbine Unit ◽  
Small Deviations ◽  
Axial Compressors ◽  
Temperature And Pressure

This paper presents an attempt of a quantitative evaluation of deterioration, due to fouling, in the performance of gas turbine units of different schemes (one-shaft, two-shaft and three-shaft) but with axial compressors of the same type and dimensions. This paper also examines the influence of the gas turbine unit initial parameters (inlet turbine temperature and pressure ratio) on the gas turbine unit sensitivity to axial compressor fouling. The evaluation of the gas turbine unit sensitivity to fouling is performed based on the small deviations method. The index of sensitivity to fouling (ISF) proposed in (Tarabrin et al, 1996) is used. It is proposed that not only the ISF characterizing the axial compressor sensitivity to fouling but also the scheme and the initial gas turbine unit parameters influencing the gas turbine unit sensitivity to axial compressor fouling, should be taken into consideration.

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.

Maximum Loading Capacity of Tandem Blades in Axial Compressors

2018 ◽  
Author(s):  
Baojie Liu ◽  
Du Fu ◽  
Xianjun Yu
Keyword(s):  
Axial Compressor ◽  
Pressure Rise ◽  
Design Tool ◽  
Stable Operation ◽  
Loading Capacity ◽  
Axial Compressors ◽  
Maximum Loading ◽  
Work Input ◽  
Dimensional Parameters ◽  
Highly Loaded

Tandem blades are widely reported to be superior to a single-blade configuration under the aerodynamic circumstance with a large flow turning in a stator or a high work input in a rotor. Aiming at the design of a highly loaded rear stage of a high pressure compressor with the advanced concept, the maximum loading capacity of a tandem-blade configuration, which is rarely described in open literature, is fundamentally necessary to be explicit in order to determine a stable operation range. A diffuser analogy is carefully carried out between the tandem-blade geometry and the diffuser passage using a reliable and robust numerical method. The analysis approach to effectively predicting the maximum static pressure rise is verified by the limited results of computational fluid dynamics (CFD) and experiments. In addition, the maximum loading capacity of the tandem-blade configuration is compared with that of the single-blade configuration to define a more favorable design range of meanline parameters. The results indicate that the tandem blade outperforms the conventional blade in a specific design space and the approach can be a potential design tool to guide the selection of one-dimensional parameters of tandem blades in a highly loaded axial compressor.

A New Multistage Axial Compressor Designed With the Apnasa Multistage CFD Code: Part 1 — Code Calibration

2001 ◽  
Author(s):  
M. Mansour ◽  
S. Hingorani ◽  
Y. Dong
Keyword(s):  
Test Data ◽  
Flow Behavior ◽  
Measurement Data ◽  
Axial Compressor ◽  
Design Tool ◽  
Paper Test ◽  
Airfoil Optimization ◽  
Axial Compressors ◽  
Radial Profiles ◽  
Multistage Axial Compressor

The NASA Average-Passage multistage turbomachinery flow analysis code “APNASA” by J.J. Adamczyk (1985) has been validated, calibrated, and demonstrated at Honeywell Engines and Systems for the design of multistage axial compressors. APNASA was first calibrated against test data of two existing compressors and then used as a design tool in the design of a new modern multistage axial compressor. The results of the calibration, design effort, and the data measurements are presented in this two-part paper. In the present paper (Part 1) the results of the calibration for two multistage axial compressors are presented. The first compressor consists of four axial stages that were designed in the mid 1980s. The second compressor consists of three axial stages and was designed in the mid 1990s using viscous, three-dimensional CFD code, with airfoil optimization performed in single blade row fashion. The calibration work was aimed at developing meshing and modeling best practices and validating the code capability to simulate flow behavior in a multistage environment. Predictions are compared with test data for the axial compressor overall performance, individual stage performance, and detailed radial profiles at the stator vanes leading edge planes, throughout the compressor. Results show good agreement between APNASA predictions and measurement data. In particular, the results clearly demonstrate the ability of APNASA to capture the stage matching of multistage machines. As a result of this calibration/validation work, a new multistage axial compressor was subsequently designed, by using APNASA as the primary source of information for airfoil optimization (presented as Part 2 of this paper). Test results for the new compressor reveal that the design achieved its performance and operability goals in its first build. Details of the compressor design philosophy using APNASA and the comparison between APNASA simulation results and test data are presented in Part 2.

scholarly journals Development and Applications of a Stage Stacking Procedure

2012 ◽  
Author(s):  
Sameer Kulkarni ◽  
Mark L. Celestina ◽  
John J. Adamczyk
Keyword(s):  
Gas Turbine ◽  
Axial Compressor ◽  
Sensitivity Study ◽  
Flow Coefficient ◽  
Turbine Engines ◽  
Stable Operation ◽  
Gas Turbine Engines ◽  
Axial Compressors ◽  
Design Performance ◽  
Load Following

The preliminary design of multistage axial compressors in gas turbine engines is typically accomplished with mean-line methods. These methods, which rely on empirical correlations, estimate compressor performance well near the design point, but may become less reliable off-design. For land-based applications of gas turbine engines, off-design performance estimates are becoming increasingly important, as turbine plant operators desire peaking or load-following capabilities and hot-day operability. The current work develops a one-dimensional stage stacking procedure. This includes a newly-defined blockage term, which is used to estimate the off-design performance and operability range of a 13-stage axial compressor. The new blockage term is defined to give mathematical closure on static pressure, and values of blockage are shown to collapse to a curve as functions of stage inlet flow coefficient and corrected speed. Utility of the stage stacking procedure is demonstrated by estimation of the minimum corrected speed which allows stable operation of the compressor. Further utility of the stage stacking procedure is demonstrated with a bleed sensitivity study, which estimates a bleed schedule to expand the compressor’s operating range.

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.

Optimization of the Workflow of Multistage Axial Compressors Using Modern Gas Dynamics Computing Systems

2020 ◽  
Author(s):  
Grigorii Popov ◽  
Evgenii Goriachkin ◽  
Oleg Baturin ◽  
Valerii Matveev ◽  
Igor Egorov ◽  
...  
Keyword(s):  
Gas Dynamics ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Calculation Model ◽  
Cfd Modeling ◽  
Compressor Blades ◽  
Stability Margins ◽  
Axial Compressors ◽  
Cfd Models ◽  
Multi Stage

Abstract Current developmental level of computers and numerical methods of gas dynamics makes it possible to optimize compressors using 3D CFD models. Design variants for the compressor can be automatically generated that best suit all the design requirements and limitations. However, the methods and tool for optimizing compressors are not sufficiently developed for the successful application. The problems lie in the large size of the calculation model, the solution time and the requirements for computer resources. In present study, a method for finding the optimal configuration of the blades of multi-stage axial compressors using 3D CFD modeling and commercial optimization programs as the main tools was developed. The basic parameters of the compressor (efficiency, pressure ratio, mass flow rate, etc.) can be improved using the created method correcting the shape of the blade profiles and their relative position. The method considers presence of various constraints. When developing the method, special attention was paid to the creation of an algorithm for parameterizing the blade shape and a program based on it, which can automatically change the shape of the axial compressor blades. They were used by the authors during optimization as a tool that converts variable parameters into the “new” blade geometry. Recommendations were also found on the rational settings for the CFD models used in the optimization of axial compressors. The paper provides a brief overview of several works related to the optimization of multi-stage gas turbine axial compressors for various purposes (number of stages from 3 to 15), successfully performed using the developed method. As a result, an increase was achieved in efficiency, pressure ratio and stability margins.

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