The Stacking of Compressor Stage Characteristics to Give an Overall Compressor Performance Map

1962 ◽  
Vol 13 (4) ◽  
pp. 349-367 ◽  
Author(s):  
M. D. C. Doyle ◽  
S. L. Dixon
Keyword(s):  
Axial Compressor ◽  
Compressor Stage ◽  
Individual Stage ◽  
Optimum Performance ◽  
Method Of Calculation ◽  
Multi Stage ◽  
Performance Map ◽  
Compressor Performance ◽  
Overall Performance ◽  
The Individual

SummaryA method of calculation is developed to compute the overall performance of a multi-stage axial compressor, from a knowledge of the individual stage characteristics, by a “stacking” technique. Compressor models are designed and their overall performance calculated. These results are compared to show, qualitatively, the effect of alterations in design and stage performance on overall performance and to find how compressors should be designed for optimum performance.

Compressor Fouling Modeling: Relationship Between Computational Roughness and Gas Turbine Operation Time

2012 ◽  
Vol 134 (5) ◽  
Author(s):  
Francesco Melino ◽  
Mirko Morini ◽  
Antonio Peretto ◽  
Michele Pinelli ◽  
Pier Ruggero Spina
Keyword(s):  
Fluid Dynamics ◽  
Surface Roughness ◽  
Gas Turbine ◽  
Gas Turbines ◽  
Axial Compressor ◽  
Operation Time ◽  
Compressor Stage ◽  
Performance Map ◽  
Computational Fluid Dynamics Cfd ◽  
Compressor Performance

Gas turbine axial compressor performance is heavily influenced by blade fouling. As a result, the gas turbines efficiency and producible power output decrease. Performance degradation of an axial compressor stage due to fouling can be analyzed by means of simulation through computational fluid dynamics (CFD) codes. Usually these methods reproduce the deteriorated blades by increasing their surface roughness and thickness. Another approach is the scaling of compressor stage performance maps. A model based on stage-by-stage techniques was presented in a previous work. This model is able to estimate the modifications of the overall compressor performance map as a function of the operating hours. The aim of the present study is to combine these two different approaches in order to relate the increase of blade computational surface roughness with compressor operating hours.

Compressor Fouling Modeling: Relationship Between Computational Roughness and Gas Turbine Operation Time

2011 ◽  
Author(s):  
Francesco Melino ◽  
Mirko Morini ◽  
Antonio Peretto ◽  
Michele Pinelli ◽  
Pier Ruggero Spina
Keyword(s):  
Fluid Dynamics ◽  
Surface Roughness ◽  
Gas Turbine ◽  
Gas Turbines ◽  
Axial Compressor ◽  
Operation Time ◽  
Compressor Stage ◽  
Performance Map ◽  
Computational Fluid Dynamics Cfd ◽  
Compressor Performance

Gas turbine axial compressor performance is heavily influenced by blade fouling; as a result, the gas turbines efficiency and producible power output decrease. Performance degradation of an axial compressor stage due to fouling can be analyzed by means of simulation through Computational Fluid Dynamics (CFD) codes. Usually these methods reproduce the deteriorated blades by increasing their surface roughness and/or thickness [1]. Another approach is the scaling of compressor stage performance maps. A model based on stage-by-stage techniques was presented in a previous work. This model is able to estimate the modifications of the overall compressor performance map as a function of the operating hours [2]. The aim of the present study is to combine these two different approaches in order to relate the increase of blade computational surface roughness with compressor operating hours.

Upgrade of a 16-Stage Industrial Compressor: Part II — Extension of the Analysis Method to the Design Function and Results

2006 ◽  
Author(s):  
Magdy S. Attia
Keyword(s):  
Axial Compressor ◽  
Design Environment ◽  
Axial Compressors ◽  
Performance Targets ◽  
Streamline Curvature ◽  
Multi Stage ◽  
Design Function ◽  
3D Effects ◽  
Compressor Performance ◽  
The Individual

A retrofit package that includes a slightly larger inlet and new, custom diffusion airfoils (CDA) was designed to replace the 16-stage axial compressor. The method used, and presented here, builds on earlier developments and is an extension of the scheme used to predict the compressor performance (Part I). The use of results from single-row 3D CFD, and their implementation into a streamline curvature (Throughflow) code lead to a better understanding of the compressor performance, which in turn lead to a better model of the compressor. This paper shows how the role of this newly developed model has been modified and adapted to the design environment. The 3D CFD results had previously provided a more accurate representation of deviation and losses, particularly at and near the end walls. The Throughflow code, when re-converged for design purposes, generated a much different solution for the individual streamlines than had been previously calculated using correlation or S1S2 analyses. Consequently, the newly generated boundary conditions for designing the individual stream sections, such as inlet and exit Mach numbers and air angles were also quite different. The designer then embarked on tailoring the individual sections to their respective duties under the guidelines of the newly developed method in true custom diffusion fashion. Iterations were conducted to optimize the section and airfoil shapes taking into consideration 3D effects. The end result was a systematic technique for designing multi-stage axial compressors and generating 3D airfoil shapes. The retrofit compressor upgrade package achieved its performance targets and delivered a measured polytropic efficiency of 93.4%.

scholarly journals Axial Compressor Mean-Line Analysis: Choking Modelling and Fully-Coupled Integration in Engine Performance Simulations

2021 ◽  
Vol 6 (1) ◽  
pp. 4
Author(s):  
Ioannis Kolias ◽  
Alexios Alexiou ◽  
Nikolaos Aretakis ◽  
Konstantinos Mathioudakis
Keyword(s):  
Axial Compressor ◽  
Engine Performance ◽  
Engine Model ◽  
Multi Stage ◽  
Transient Simulations ◽  
Full Knowledge ◽  
Compressor Performance ◽  
First Time ◽  
Performance Calculation ◽  
Fully Coupled

A mean-line compressor performance calculation method is presented that covers the entire operating range, including the choked region of the map. It can be directly integrated into overall engine performance models, as it is developed in the same simulation environment. The code materializing the model can inherit the same interfaces, fluid models, and solvers, as the engine cycle model, allowing consistent, transparent, and robust simulations. In order to deal with convergence problems when the compressor operates close to or within the choked operation region, an approach to model choking conditions at blade row and overall compressor level is proposed. The choked portion of the compressor characteristics map is thus numerically established, allowing full knowledge and handling of inter-stage flow conditions. Such choking modelling capabilities are illustrated, for the first time in the open literature, for the case of multi-stage compressors. Integration capabilities of the 1D code within an overall engine model are demonstrated through steady state and transient simulations of a contemporary turbofan layout. Advantages offered by this approach are discussed, while comparison of using alternative approaches for representing compressor performance in overall engine models is discussed.

scholarly journals Simulation of Compressor Performance Deterioration due to Erosion

1989 ◽  
Author(s):  
W. Tabakoff ◽  
A. N. Lakshminarasimha ◽  
M. Pasin
Keyword(s):  
Fault Model ◽  
Experimental Results ◽  
Performance Tests ◽  
Individual Stage ◽  
One Stage ◽  
Before And After ◽  
Performance Deterioration ◽  
Compressor Performance ◽  
Multistage Compressor ◽  
The Individual

Experimental results obtained from cascades and one stage compressor performance tests before and after erosion were used to test a fault model to represent erosion. This model was implemented on a stage stacking program developed to demonstrate the effect of erosion in a multistage compressor. The effect of the individual stage erosion on the overall compressor performance is also demonstrated.

Performance Prediction for Automotive Turbocharger Compressors

1975 ◽  
Vol 189 (1) ◽  
pp. 557-565 ◽  
Author(s):  
A. Whitfield ◽  
F. J. Wallace
Keyword(s):  
Performance Prediction ◽  
Flow Analysis ◽  
Compressor Stage ◽  
Thermodynamic Models ◽  
Flow Rates ◽  
Design Point ◽  
One Dimensional ◽  
Dimensional Flow ◽  
Performance Map ◽  
The Individual

A procedure to predict the complete performance map of turbocharger centrifugal compressors is presented. This is based on a one-dimensional flow analysis using existing published loss correlations that were available and thermodynamic models to describe the incidence loss and slip factor variation at flow rates which differ from the design point. To predict the losses within the complete compressor stage using a one-dimensional flow procedure, it is necessary to introduce a number of empirical parameters. The uncertainty associated with these empirical parameters is assessed by studying the effect of varying them upon the individual losses and upon the overall predicted performance.

Effect of Tip Gap Variation on the Performance of the Transonic Fan Stage With Tandem Stator

2019 ◽  
Author(s):  
K. Ananthakrishnan ◽  
Shyama Prasad Das ◽  
B. V. S. S. S. Prasad
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Axial Compressor ◽  
Compressor Stage ◽  
Tip Leakage ◽  
Pressure Loss Coefficient ◽  
Overall Performance ◽  
Transonic Fan ◽  
Total Pressure Loss Coefficient ◽  
Transonic Rotor

Abstract The main goal of modern axial compressor development is to increase the power to weight ratio with higher efficiency. In the present investigation, highly loaded single stage axial compressor with tandem stator vanes is used. Tandem vanes help in attaining the compact compressor stage along with high pressure loading. It is designed for a stage pressure ratio of 2, mass flow rate of 9.02 kg/s operating at 30800 rpm resulting in transonic flow field. The aerodynamic performance of this compressor detoriates due to the tip leakage and secondary flows. Steady-state numerical investigation is carried out to study the flow structures near the tip region of transonic rotor and how different tip gaps influence the overall performance of the compressor. Further the effects of tip leakage flow variation on the performance of tandem vanes are also highlighted. Transonic fan stage with baseline tip gap of 0.5mm is analyzed along with different tip clearance values ranging from 0 % to 3 % of axial chord. Three-dimensional viscous Reynolds Averaged Navier Stokes (RANS) equations are solved using SST k-ω turbulence model. Computational domain discretized with high quality hexahedral elements (Y+ < 2) in AUTOGRID, Numeca. The numerical procedure is verified against the experimental results of Rotor37 transonic rotor test case. Tip leakage losses contribute a substantial amount to the total loss of stage. Overall performance and the stall characteristics for the compressor stage has been evaluated for different tip gap variations.. Further, the topological properties are exploited to visualize the critical points and separation lines on rotor and tandem vanes. Increase in rotor total pressure loss coefficient is observed with increasing tip gap. In contrary, overall total pressure loss coefficient improves for smaller tip gap values and then detoriates. It is observed optimum tip gap height lies close to the 1.125mm, 2% of baseline design value.

Numerical Analysis of the Effects of Non-Uniform Surface Roughness on Compressor Stage Performance

2010 ◽  
Author(s):  
Mirko Morini ◽  
Michele Pinelli ◽  
Pier Ruggero Spina ◽  
Mauro Venturini
Keyword(s):  
Surface Roughness ◽  
Gas Turbine ◽  
Traditional Approach ◽  
Fluid Dynamic ◽  
Turbine Blades ◽  
Compressor Stage ◽  
Turbine Performance ◽  
Performance Map ◽  
Overall Performance ◽  
Stage Performance

Gas turbine performance degradation over time is mainly due to the deterioration of compressor and turbine blades, which, in turn, causes a modification of the compressor and turbine performance maps. Since detailed information about the actual modification of the compressor and turbine performance maps is usually unavailable, component performance can be modeled and investigated (i) by scaling the overall performance map, or (ii) by using stage-by-stage models of the compressor and turbine and by scaling each single stage performance map to account for each stage deterioration, or (iii) by performing 3D numerical simulations, which allow to both highlight the fluid-dynamic phenomena occurring in the faulty component and grasp the effect on the overall performance of the component. In this paper, the authors address the most common and experienced source of loss for a gas turbine, i.e. compressor fouling. With respect to the traditional approach, which mainly aims at the identification of the overall effects of fouling, authors investigate a micro-scale representation of compressor fouling (e.g. blade surface deterioration and flow deviation). This allows (i) a more detailed investigation of the fouling effects (e.g. mechanism, location along blade height, etc.), (ii) a more extensive analysis of the causes of performance deterioration and (iii) the assessment of the effect of fouling on stage performance coefficients and on stage performance maps. The effects of a non-uniform surface roughness on both rotor and stator blades of an axial compressor stage are investigated by using a commercial CFD code. The NASA Stage 37 test case was used as the baseline geometry. The numerical model already validated against experimental data available in literature was used for the simulations. Different non-uniform combinations of surface roughness levels on rotor and stator blades were imposed. This makes it possible to highlight how the localization of fouling on compressor blades affects compressor performance, both at an overall and at a fluid-dynamic level.

Numerical Investigation of Axial Compressor Stages With Differing Degrees of Reaction

2014 ◽  
Author(s):  
Jens Ortmanns
Keyword(s):  
Numerical Investigation ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Pressure Rise ◽  
Wall Region ◽  
Compressor Stage ◽  
Tip Leakage Flow ◽  
Degree Of Reaction ◽  
Multi Stage ◽  
Flow Phenomena

In order to increase the efficiency of a compressor module, several loss sources such as aerofoil profile loss, secondary loss and clearance flow phenomena must be taken into account and balanced in the most efficient way. This current document presents the results of a numerical investigation based on a conventionally loaded high pressure compressor stage with different inlet and exit swirls. The effects of changing the degree of reaction on the compressor stage flow pattern is analysed in detail. In general, the correlation between the overall stage efficiency at constant pressure ratio and the degree of stage reaction is low. Nevertheless, the results show a direct impact on the rotor tip leakage flow and the secondary flow phenomena in the stator end-wall region when the degree of reaction is modified which is driven by the change in static pressure rise between the rotor and the stator passages. The balance of these two loss sources might have an impact on the efficiency and the stall behaviour of a multi-stage compressor.

Application of Fuzzy Logic to Axial Compressor Performance Map Prediction

2007 ◽  
Author(s):  
Kaveh Ghorbanian ◽  
Mohammad Gholamrezaei
Keyword(s):  
Fuzzy Logic ◽  
Axial Compressor ◽  
Optimization Procedure ◽  
Time Frame ◽  
Learning Capability ◽  
Design Variables ◽  
Performance Map ◽  
Compressor Performance ◽  
Short Time ◽  
Fuzzy Logic Technique

The present paper applies fuzzy logic technique to predict the performance map of an axial compressor. This technique relies on employing the information of a data curve in concert with the information at the design point. Further, the learning capability of ANN technique is integrated to the potential of fuzzy logic. A comparison of the predicted results with experimental data reveals a very good agreement. The proposed technique has not only the capability to model the nonlinear surge line as well as the kink in a classical compressor performance map but also it can be used as an alternative tool to foresee the effect of modification of design variables, as well as to guide the design optimization procedure in a short time frame.

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