Robust Optimization Design of Single-Stage Transonic Axial Compressor Considering the Manufacturing Uncertainties

2018 ◽  
Author(s):  
Zhihui Li ◽  
Yanming Liu ◽  
Ramesh K. Agarwal
Keyword(s):  
Total Pressure ◽  
Optimization Design ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Tip Clearance ◽  
Robust Design Optimization ◽  
Compressor Performance ◽  
Overall Performance ◽  
Transonic Axial Compressor ◽  
Total Pressure Ratio

Manufacturing uncertainties always lead to significant variability in compressor performance. In this work, the tip clearance uncertainties inherent in a transonic axial compressor are quantified to determine their effect on performance. The validated tip clearance losses model in conjunction with the 3D reynolds averaged navier-stokes (RANS) solver are utilized to simulate these uncertainties and quantify their effect on the adiabatic efficiency, total pressure ratio and choked mass flow. The sensitivity analysis method is employed to figure out which parameters play the most significant roles in determining the overall performance of compressor. To propagate these uncertainty factors, the non-intrusive polynomial chaos expansion (PCE) algorithm is used in this paper and the probability distributions of compressor performance are successfully predicted. A robust design optimization has been carried out based on the combination of the genetic algorithm (GA) and the uncertainty quantification (UQ) method, leading to a robust compressor rotor design for which the overall performance is relatively insensitive to variability in tip clearance without reducing the sources of the manufacturing noise. The optimization results show that the mean value of the adiabatic rotor efficiency is improved by 1.4 points with the overall variation of that reduced by 64.1%, while the total pressure ratio is slightly improved when compared to the prototype.

scholarly journals Centrifugal Compressor Performance Improvement through Multi Splitter Impeller

2019 ◽  
Vol 26 (2) ◽  
pp. 6-14
Author(s):  
Adil Malik ◽  
Qun Zheng ◽  
Shafiq R. Qureshi ◽  
Salman A. Ahmed ◽  
D. KB Gambo
Keyword(s):  
Performance Improvement ◽  
Centrifugal Compressor ◽  
Total Pressure ◽  
Pressure Ratio ◽  
Suction Surface ◽  
Pressure Surface ◽  
Percent Increase ◽  
Compressor Performance ◽  
Overall Performance ◽  
Total Pressure Ratio

Abstract In the paper, a back swept impeller of centrifugal compressor is experimentally studied and numerically validated and modified to increase its pressure ratio and improve efficiency, as well as to analyse the effect of splitter blade location between two main blades. The back swept multi splitter blade impeller was designed with a big splitter positioned close to the main blade suction surface and a smaller splitter close to the pressure surface. Adding this multi splitter improves the overall performance of the modified impeller due to less intensive flow separation and smaller pressure loss. In particular, the total pressure ratio was observed to increase from 4.1 to 4.4, with one percent increase in efficiency.

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.

Effects of Non-axisymmetric Casing Grooves Combined with Airflow Injection on Stability Enhancement of an Axial Compressor

2017 ◽  
Vol 0 (0) ◽  
Author(s):  
C. T. Dinh ◽  
K. Y. Kim
Keyword(s):  
Total Pressure ◽  
Stokes Equations ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Range Extension ◽  
Stable Range ◽  
Stall Margin ◽  
Injection Angle ◽  
Transonic Axial Compressor ◽  
Total Pressure Ratio

AbstractThis paper presents a performance evaluation of non-axisymmetric casing grooves combined with airflow injection in a transonic axial compressor with NASA Rotor 37, using three-dimensional Reynolds-averaged Navier-Stokes equations with the k-ε turbulence model. An axisymmetric casing groove was divided circumferentially into 36 non-axisymmetric grooves. The numerical results for adiabatic efficiency and total pressure ratio were validated with experimental data. A parametric study for stall margin, stable range extension, peak adiabatic efficiency, and total pressure ratio at peak adiabatic efficiency of the compressor was performed using five parameters: the front and rear lengths, the height of the casing groove, the injection mass flow rate, and the injection angle. The non-axisymmetric casing grooves combined with injection improve greatly the stall margin and stable range extension of the transonic axial compressor, but reduce only slightly the peak adiabatic efficiency in all cases, compared to the results for a smooth casing.

Reconciling Compressor Performance Differences for Varying Ambient Inlet Conditions

2015 ◽  
Author(s):  
Natalie R. Smith ◽  
Reid A. Berdanier ◽  
John C. Fabian ◽  
Nicole L. Key
Keyword(s):  
Total Pressure ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Tip Clearance ◽  
Ambient Conditions ◽  
Midwest United States ◽  
Thermal Growth ◽  
Reynolds Number Effects ◽  
Inlet Conditions ◽  
Total Pressure Ratio

Careful experimental measurements can capture small changes in compressor total pressure ratio that arise with subtle changes in an experiment’s configuration. Research facilities that use unconditioned atmospheric air must account for changes in ambient compressor inlet conditions to establish repeatable performance maps. A unique dataset from a threestage axial compressor has been acquired over the duration of 12 months in the Midwest United States where ambient conditions change significantly. The trends show a difference in compressor total pressure ratio measured on a cold day versus a warm day despite correcting inlet conditions to sea level standard day. To reconcile these differences, this paper explores correcting the compressor exit thermodynamic state, Reynolds number effects, and variations in rotor tip clearance as a result of differences in thermal growth.

Experimental and Numerical Investigation of a Circumferential Groove Casing Treatment in a Low Speed Axial Research Compressor at Different Tip Clearances

2017 ◽  
Author(s):  
Matthias Rolfes ◽  
Martin Lange ◽  
Konrad Vogeler ◽  
Ronald Mailach
Keyword(s):  
Total Pressure ◽  
Solid Wall ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Chord Length ◽  
Single Stage ◽  
Tip Clearance ◽  
Low Speed ◽  
Operating Range ◽  
Circumferential Groove

The demand of increasing pressure ratios for modern high pressure compressors leads to decreasing blade heights in the last stages. As tip clearances cannot be reduced to any amount and minimum values might be necessary for safety reasons, the tip clearance ratios of the last stages can reach values notably higher than current norms. This can be intensified by a compressor running in transient operations where thermal differences can lead to further growing clearances. For decades, the detrimental effects of large clearances on an axial compressor’s operating range and efficiency are known and investigated. The ability of circumferential casing grooves in the rotor casing to improve the compressor’s operating range has also been in the focus of research for many years. Their simplicity and ease of installation are one reason for their continuing popularity nowadays, where advanced methods to increase the operating range of an axial compressor are known. In a previous paper [1], three different circumferential groove casing treatments were investigated in a single stage environment in the Low Speed Axial Research Compressor at TU Dresden. One of these grooves was able to notably improve the operating range and the efficiency of the single stage compressor at very large rotor tip clearances (5% of chord length). In this paper, the results of tests with this particular groove type in a three stage environment in the Low Speed Axial Research Compressor are presented. Two different rotor tip clearance sizes of 1.2% and 5% of tip chord length were investigated. At the small tip clearance, the grooves are almost neutral. Only small reductions in total pressure ratio and efficiency compared to the solid wall can be observed. If the compressor runs with large tip clearances it notably benefits from the casing grooves. Both, total pressure and efficiency can be improved by the grooves in a similar extent as in single stage tests. Five-hole probe measurements and unsteady wall pressure measurements show the influence of the groove on the flow field. With the help of numerical investigations the different behavior of the grooves at the two tip clearance sizes will be discussed.

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.

Design and Test of a Transonic Axial Splittered Rotor

2015 ◽  
Author(s):  
Garth V. Hobson ◽  
Anthony J. Gannon ◽  
Scott Drayton
Keyword(s):  
Mass Flow ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Design Procedure ◽  
Tip Clearance ◽  
Turbine Engine ◽  
Transonic Compressor ◽  
Compressor Rotor ◽  
Transonic Axial Compressor ◽  
Commercial Off The Shelf

A new design procedure was developed that uses commercial-off-the-shelf software (MATLAB, SolidWorks, and ANSYS-CFX) for the geometric rendering and analysis of a transonic axial compressor rotor with splitter blades. Predictive numerical simulations were conducted and experimental data were collected in a Transonic Compressor Rig. This study advanced the understanding of splitter blade geometry, placement, and performance benefits. In particular, it was determined that moving the splitter blade forward in the passage between the main blades, which was a departure from the trends demonstrated in the few available previous transonic axial compressor splitter blade studies, increased the mass flow range with no loss in overall performance. With a large 0.91 mm (0.036 in) tip clearance, to preserve the integrity of the rotor, the experimentally measured peak total-to-total pressure ratio was 1.69 and the peak total-to-total isentropic efficiency was 72 percent at 100 percent design speed. Additionally, a higher than predicted 7.5 percent mass flow rate range was experimentally measured, which would make for easier engine control if this concept were to be included in an actual gas turbine engine.

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.

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