scholarly journals Simulation of the effect of nonuniform fouling thickness on an axial compressor stage performance

2021 ◽  
Vol 13 (7) ◽  
pp. 168781402110304
Author(s):  
Hai-Ou Sun ◽  
Li-Song Wang ◽  
Zhong-Yi Wang ◽  
Meng Wang ◽  
Yan-Hua Wang ◽  
...  
Keyword(s):  
Deposition Rate ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Blade Surface ◽  
Compressor Blades ◽  
Ansys Cfx ◽  
Commercial Code ◽  
Stator Blade ◽  
Total Pressure Ratio ◽  
Isentropic Efficiency

This paper focuses on the numerical simulation of the flow status in the compressor with the condition of fouling. NASA stage 35 was considered as the object, and the commercial code ANSYS CFX was used. The deposition rate of contaminants on the surface was considered to be different along the height of the blade. A data from related study shows that the deposition rate of the contaminant on the side close to the hub is higher than the side near the shroud part. Based on the deposition law, this paper simulated the fouling of the compressor blades by changing the thickness on the blade surface. This subject only changed the thickness of the stator blade surface because of a data showing that the fouling on the stator blade surface is almost double that on the rotor blade surface. In the condition that the roughness value of the blade surface is constant, only the stable working range of the compressor is effected by the change of the surface thickness of the stator blade. There is a positive relationship between the value of compressor minimum flow rate and the value of thickness increment. After fouling the total pressure ratio and isentropic efficiency degenerated 1.59% and 3.76%, respectively.

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.

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.

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.

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.

Some Experiments With a Supersonic Axial Compressor Stage

1987 ◽  
Vol 109 (3) ◽  
pp. 388-397 ◽  
Author(s):  
A. J. Wennerstrom
Keyword(s):  
Pressure Ratio ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Design Flow ◽  
Vortex Generators ◽  
Boundary Layer Control ◽  
Compressor Stage ◽  
Design Speed ◽  
Isentropic Efficiency ◽  
Layer Control

Between 1970 and 1974, ten variants of a supersonic axial compressor stage were designed and tested. These included two rotor configurations, three rotor tip clearances, addition of boundary-layer control consisting of vortex generators on both the outer casing and the rotor, and the introduction of slots in the stator vanes. Design performance objectives were a stage total pressure ratio of 3.0 with an isentropic efficiency of 0.82 at a tip speed of 1600 ft/s (488 m/s). The first configuration passed only 70 percent of design flow at design speed, achieving a stage pressure ratio of 2.25 at a peak stage isentropic efficiency of 0.61. The rotor was grossly separated. The tenth variant passed 91.4 percent of design flow at design speed, producing a stage pressure ratio of 3.03 with an isentropic efficiency of 0.75. The rotor achieved a pressure ratio of 3.59 at an efficiency of 0.87 under the same conditions. Major conclusions were that design tools available today would undoubtedly permit the original goals to be met or exceeded. However, the application for such a design is currently questionable because efficiency goals considered acceptable for most current programs have risen considerably from the level considered acceptable at the inception of this effort. Splitter vanes placed in the rotor permitted very high diffusion levels to be achieved without stalling. However, viscous effects causing three-dimensional flows violating the assumption of flow confined to concentric stream tubes were so strong that a geometry optimization does not appear practical without a three-dimensional, viscous analysis. Passive boundary-layer control in the form of vortex generators and slots does appear to offer some benefit under certain circumstances.

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.

Adjoint Optimization of Multistage Axial Compressor Blades with Static Pressure Constraint at Blade Row Interface

2016 ◽  
Vol 33 (2) ◽  
Author(s):  
Jia Yu ◽  
Lucheng Ji ◽  
Weiwei Li ◽  
Weilin Yi
Keyword(s):  
Pressure Distribution ◽  
Adjoint Method ◽  
Static Pressure ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Optimal Solution ◽  
Optimization Procedure ◽  
Compressor Blades ◽  
Static Pressure Distribution ◽  
Practical Engineering

AbstractAdjoint method is an important tool for design refinement of multistage compressors. However, the radial static pressure distribution deviates during the optimization procedure and deteriorates the overall performance, producing final designs that are not well suited for realistic engineering applications. In previous development work on multistage turbomachinery blade optimization using adjoint method and thin shear-layer N-S equations, the entropy production is selected as the objective function with given mass flow rate and total pressure ratio as imposed constraints. The radial static pressure distribution at the interfaces between rows is introduced as a new constraint in the present paper. The approach is applied to the redesign of a five-stage axial compressor, and the results obtained with and without the constraint on the radial static pressure distribution at the interfaces between rows are discussed in detail. The results show that the redesign without the radial static pressure distribution constraint (RSPDC) gives an optimal solution that shows deviations on radial static pressure distribution, especially at rotor exit tip region. On the other hand, the redesign with the RSPDC successfully keeps the radial static pressure distribution at the interfaces between rows and make sure that the optimization results are applicable in a practical engineering design.

Shape Optimization of Multi-Stage Axial Compressor Blades Using Adjoint Method With Static Pressure Constraint at Blade Row Interface

2015 ◽  
Author(s):  
Jia Yu ◽  
Lucheng Ji ◽  
Weiwei Li ◽  
Weilin Yi
Keyword(s):  
Pressure Distribution ◽  
Adjoint Method ◽  
Static Pressure ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Optimal Solution ◽  
Optimization Procedure ◽  
Compressor Blades ◽  
Static Pressure Distribution ◽  
Multi Stage

Adjoint method is an important tool for design refinement of multistage compressors. However, the radial static pressure distribution deviates during the optimization procedure and deteriorates the overall performance, producing final designs that are not well suited for realistic engineering applications. In previous development work on multistage turbomachinery blade optimization using adjoint method and thin shear-layer N-S equations, the entropy production is selected as the objective function with given mass flow rate and total pressure ratio as imposed constraints. The radial static pressure distribution at the interfaces between rows is introduced as a new constraint in the present paper. The approach is applied to the redesign of a five-stage axial compressor, and the results obtained with and without the constraint on the radial static pressure distribution at the interfaces between rows are discussed in detail. The results show that the redesign without radial static pressure distribution constraint (RSPDC) gives an optimal solution that shows deviations on radial static pressure distribution, especially at rotor exit tip region. On the other hand, the redesign with the RSPDC successfully keeps the radial static pressure distribution at the interfaces between rows and make sure that the optimization results are applicable in a practical engineering design.

scholarly journals Flow behavior in a contra-rotating transonic axial compressor

2021 ◽  
Vol 15 (3) ◽  
pp. 8440-8449
Author(s):  
Sarallah Abbasi ◽  
Maryam Alizadeh
Keyword(s):  
Flow Behavior ◽  
Pressure Ratio ◽  
Flow Analysis ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Navier Stokes ◽  
Tip Leakage Flow ◽  
Ansys Cfx ◽  
Energy Equations ◽  
Compressor Efficiency

This study investigated a three-dimensional flow analysis on a two-stage contra-rotating axial compressor using the Navier–Stokes, continuity, and energy equations with Ansys CFX commercial software. In order to validate the obtained results, the absolute and relative flow angles curves for each rotor in radial direction were extracted and compared with the other investigation results, indicating good agreement. The compressor efficiency curve also was extracted by varying the compressor pressure ratio and compressor efficiency against mass flow rate. The flow results revealed that further distortion of the flow structure in the second rotor imposed a greater increase in the amount of entropy, especially at near-stall conditions. The increase of entropy in the second rotor is due to the interference of the tip leakage flow with the main flow which consequently caused more drops in the second rotor, suggesting that more efficacy of flow control methods occurred in the second rotor than in the first rotor.

Modelling Unsteady Transition Effects in an Axial Compressor

2008 ◽  
Author(s):  
Adam D. Beevers ◽  
Joao Amaral-Teixeira ◽  
Roger Wells
Keyword(s):  
Boundary Layer ◽  
Axial Compressor ◽  
Leading Edge ◽  
Transition Process ◽  
Transition Model ◽  
Velocity Defect ◽  
Commercial Code ◽  
Stator Blade ◽  
Transition Effects ◽  
Unsteady Effects

Wake induced transition is simulated at mid-span on a C4 stator blade in a 1.5 stage low speed axial compressor using the CFX γ – θ transition model. IGV and rotor wake inputs were created from a succession of Fourier series produced from experimental data. The purpose of the study was to understand the effectiveness of the γ – θ model implemented in a commercial code to predict the unsteady effects of wake induced transition. The γ – θ transition model was found to predict wake-induced transition and the subsequent calmed region caused by the passing wakes. The wake velocity defect creates conditions within the boundary layer such that the high disturbance energy, which is diffused into the boundary layer at the leading edge from the wake, initiates the transition process. This high disturbance energy travels through the boundary layer directly behind the wake.

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