Aerodynamic Redesign of a 5-Stage Axial Compressor by Re-Stagger and Bow

2015 ◽  
Author(s):  
Tao Ning ◽  
Chun-wei Gu ◽  
Xiao-tang Li ◽  
Tai-qiu Liu ◽  
Yao-bing Xiao
Keyword(s):  
Axial Compressor ◽  
Aerodynamic Optimization ◽  
Stall Margin ◽  
High Loss ◽  
Multi Stage ◽  
Pressure Profiles ◽  
Scale Test ◽  
Overall Performance ◽  
Operating Points

The paper is a preliminary methodology research of the aerodynamic redesign of a 5-stage axial compressor by re-stagger and bow with multistage CFD methods, with the purpose of developing the next upgrade version of this type of compressor. Prior to aerodynamic redesign, the validation study is carried out for overall performance, casing static pressure and spanwise total pressure profiles based on full-scale test data, proving that the multistage CFD applied is a relatively reliable tool for the analysis of the follow-up redesigning. Furthermore, at the near stall point, aerodynamic analysis demonstrates that significant separation exits in the last stator, which leads to the aerodynamic redesign focused on the last stator. To explore the potential of aerodynamic improvement, it is focused on the aerodynamic redesign by re-stagger and bow. Multi-stage CFD methods are applied in the whole redesign process. Two steps of redesigning are applied for aerodynamic optimization. Upgrade 1 by re-stagger is used to adjust the operating points to more reasonable region, which contributes about 9 percent stall margin increase. Based on upgrade 1, an unconventional asymmetric bow configure is employed in upgrade 2 to reduce the high loss region dominated by mainstream. The redesigning exploitation by re-stagger and bow in this research produces a total 16% increase in stall margin.

Assessment of a Novel Non-Axial Counter-Rotating Compressor Concept for Aero-Engines

2018 ◽  
Author(s):  
Quentin Dejour ◽  
Huu Duc Vo
Keyword(s):  
Pressure Rise ◽  
Axial Direction ◽  
Boundary Layer Separation ◽  
Low Flow ◽  
Compressor Stage ◽  
Mixed Flow ◽  
Stall Margin ◽  
High Loss ◽  
Multi Stage ◽  
Flow Compressor

This paper presents the first assessment of a new non-axial counter-rotating compressor concept. This concept consists of replacing the stator of a mixed-flow compressor stage or the diffuser of a centrifugal compressor stage with a counter-rotating rotor that will turn the flow back to the axial direction with much lower diffusion factor, while providing the equivalent in work of the upstream mixed-flow rotor or impeller. This concept has two advantages. First, the very high stage pressure rise means that only a single counter-rotating rotor may be required, making mechanical implementation simpler than for multi-stage axial counter-rotating compressors. Second, the replacement of the high flow turning (high loss) stator/diffuser in a non-axial stage with a low flow turning counter-rotating rotor gives the new concept potential for achieving higher efficiency than conventional non-axial compressors. As a first proof of concept, a subsonic counter-rotating mixed-flow compressor and its conventional (i.e. rotor-stator) equivalent have been designed with the intent of being implemented in a test rig. CFD simulations have been carried out for a comparative evaluation of both configurations. Results show that the counter-rotating mixed-flow compressor produces more than double the pressure rise of its conventional version with a slightly higher peak-efficiency while having a smaller axial length. Moreover, the counter-rotating configuration has a better stall margin than its conventional counterpart, for which the boundary layer separation from excessive flow turning in the stator causes early stall.

scholarly journals A Practical Compressor Casing Treatment

1997 ◽  
Author(s):  
Syed Khalid
Keyword(s):  
Suction Side ◽  
Velocity Variation ◽  
Cell Orientation ◽  
Stall Margin ◽  
Exit Velocity ◽  
Casing Treatment ◽  
Multi Stage ◽  
Pressure Profiles ◽  
Inclination Angles ◽  
Side Flow

A three stage compressor test incorporating casing inserts comprised of compound angled honeycomb cells demonstrated up to 10% higher stall margin than circumferential grooves casing treatment. This is attributed to effective tip flow energization resulting from the unsteady flow induced in and out of the cells as the blade tip sweeps by the cell openings. The rationale for selecting the cell inclination angles both relative to the normal and the tangential directions is discussed. The design intent of the cell orientation is to induce a high cell exit velocity as well as to impart a degree of flow alignment to the injected jets. A first order calculation of cell exit velocity variation based on the cell pressure/volume dynamics is indicative of unsteady blowing which is theorized to effectively mix the tip suction side flow and to enhance the tip flow streamwise momentum. This theory is partially substantiated by the presented compressor test results showing improved radial total pressure profiles, stage characteristics, and stall margin. Since a few unhealthy stages of a multi-stage compressor could make it stall prone, casing treatment of those weak stages could dramatically increase stall margin with negligible impact on overall adiabatic efficiency. In addition to the aerodynamic effectiveness, the mechanical suitability of this casing treatment to multistage compressors, based on its demonstrated abradability and packageability, is discussed.

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.

scholarly journals A Navier-Stokes Analysis of the Effect of Tip Clearance on Compressor Stall Margin

1995 ◽  
Author(s):  
Robert P. Dring ◽  
William D. Sprout ◽  
Harris D. Weingold
Keyword(s):  
Three Dimensional ◽  
Axial Compressor ◽  
Tip Clearance ◽  
Navier Stokes ◽  
Stall Margin ◽  
Multi Stage ◽  
High Pressure Compressor ◽  
Practical Tool ◽  
The Impact ◽  
Pressure Compressor

A three-dimensional Navier-Stokes calculation was used to analyze the impact of rotor tip clearance on the stall margin of a multi-stage axial compressor. This paper presents a summary of: (1) a study of the sensitivity of the results to grid refinement, (2) an assessment of the calculation’s ability to predict stall margin when the stalling row was the first rotor in a multi-stage rig environment, (3) an analysis of the impact of including the effects of the downstream stator through body force effects on the upstream rotor, and (4) the ability of the calculation to predict the impact of tip clearance on stall margin through a calculation of the rear seven airfoil rows of an eleven stage high pressure compressor rig. The result of these studies was that a practical tool is available which can predict stall margin, and the impact of tip clearance, with reasonable accuracy.

Squealer Tip Treatment Design for Axial Compressors

2020 ◽  
Author(s):  
Giuseppe Bruni ◽  
James Taylor ◽  
Senthil Krishnababu ◽  
Robert Miller ◽  
Roger Wells
Keyword(s):  
Thickness Distribution ◽  
Axial Compressor ◽  
Leading Edge ◽  
Experimental Tests ◽  
Suction Side ◽  
Rotor Blades ◽  
Superior Performance ◽  
Stall Margin ◽  
Wall Flows ◽  
Multi Stage

Abstract End-wall flows are amongst the main sources of losses in the rear stages of a typical multi-stage axial compressor. Reducing the tip leakage losses in the rotor blades and vanes can provide an increased efficiency and stall margin of a given axial compressor stage. One approach is to use squealer tips, which are traditionally designed to minimize the effect of tip rubbing. However, squealers can also provide a significant performance benefit, when designed considering aerodynamics from the beginning, as shown in this paper. A CFD based methodology, in which the blade or vane thickness distribution is varied in a controlled manner was developed. This design methodology was used to create different types of squealer tip geometry for a representative stage in a low speed compressor rig. Three different tip concepts were designed, based on a Suction Side Squealer, on a Pressure Side Squealer and on the combination of the two being merged between the leading edge and trailing edge, this new design is called the SuPr Tip. Subsequent experimental tests carried out agreed with the predicted relative ranking of the different squealer designs and on the superior performance of the SuPr tip design over the others, thus validating the methodology and the design process.

Some effects of non-axisymmetric tip clearance layout on axial compressor aerodynamics

2021 ◽  
pp. 095765092110470
Author(s):  
Haohao Zhang ◽  
Haowan Zhuang ◽  
Jinfang Teng ◽  
Mingmin Zhu ◽  
Xiaoqing Qiang
Keyword(s):  
Transient Flow ◽  
Axial Compressor ◽  
Flow Fields ◽  
Tip Clearance ◽  
Inlet Guide Vane ◽  
Leakage Flow ◽  
Tip Leakage Flow ◽  
Stall Margin ◽  
Tip Leakage ◽  
Overall Performance

A steady and unsteady numerical research is carried out to explore some effects of a specific non-axisymmetric tip clearance layout on the overall performance and stability of an axial compressor stage. For a 4-stage low-speed research compressor (LSRC) in Shanghai Jiao Tong University (SJTU), one-eighth annulus of the inlet guide vane and the first stage rotor was modeled for this study. After the validation for the uniform tip clearance case, a specific non-axisymmetric tip clearance layout is chosen from several random cases generated by the Gaussian Probabilistic Density Function method. Unsteady time-averaged results at the near stall condition show that the chosen non-axisymmetric layout can improve the isentropic efficiency by 1.3% and extend the stall margin by 4%. Detailed analyses on flow fields are carried out to interpret the performance improvement. Due to the circumferential layout of clearance sizes, the inlet mass flow and incidence are redistributed in both the radial and circumferential directions. It leads to blade loading and tip leakage flow varying with the tip clearance size. The quantification of blockage manifests that the blockage arising from the tip leakage flow is significantly alleviated in the non-axisymmetric layout, which leads to improvements in overall performance and stall margin. Transient flow fields at the rotor tip are also analyzed at the near stall condition. For the non-axisymmetric layout, low-momentum regions originating from larger clearance sizes oscillate and develop downstream in one blade passage period.

scholarly journals Complementary Use of CFD and Experimental Measurements to Assess the Impact of Shrouded and Cantilevered Stators in Axial Compressors

1999 ◽  
Author(s):  
M. S. Campobasso ◽  
A. Mattheiss ◽  
U. Wenger ◽  
A. Amone ◽  
P. Boncinelli
Keyword(s):  
Axial Compressor ◽  
Operating Conditions ◽  
Experimental Measurements ◽  
Stall Margin ◽  
Axial Compressors ◽  
Design Concepts ◽  
Overall Performance ◽  
The Impact

The performance of an axial compressor with either shrouded or cantilevered stators has been analyzed. The two configurations have been compared both at design and near stall operating conditions, with the aid of CFD and experimental measurements. Results show that shrouded or cantilevered stators impact differently on the overall performance of the tested compressor. A higher stall margin occurs with the cantilevered build, while the work coefficient and the efficiency of the shrouded build at design conditions are higher. An overall comparison of the shrouded and cantilevered design concepts has been carried out, not only in aerodynamic but also in economic terms.

Three-Dimensional Aerodynamic Optimization of a Multi-Stage Axial Compressor

2016 ◽  
Author(s):  
Tao Ning ◽  
Chun-wei Gu ◽  
Xiao-tang Li ◽  
Tai-qiu Liu
Keyword(s):  
Genetic Algorithm ◽  
Surrogate Model ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Optimization Method ◽  
Operating Point ◽  
Design Parameters ◽  
Aerodynamic Optimization ◽  
Stall Margin

An optimization method combined of a genetic algorithm, an artificial neural network, a CFD solver and a blade generator, is developed in this research and applied in the three-dimensional blading design of a newly designed highly-loaded 5-stage axial compressor. The adaptive probabilities of crossover and mutation, non-uniform mutation operator and elitism operator are employed to improve the convergence of the genetic algorithm. Considering both the optimization efficiency and effectiveness, a mixture of high-fidelity multistage CFD method and approximate surrogate model of the feed-forward ANN is used to evaluate the fitness. In particular, the database is updated dynamically and used to re-train the surrogate model of ANN for improving the accuracy for predicting. The last stator of the compressor is optimized at the near stall operating point. The tip bow with relative bow height Hb and bow angle αb are treated as design parameters. The adiabatic efficiency as well as the penalty of mass flow and total pressure ratio constitute the objective functions to be maximized. The optimum (Hb = 0.881, αb = 14.7°) obtains 0.4% adiabatic efficiency increase for the whole compressor at the optimized operating point. The detailed aerodynamic is compared between the baseline and optimized stator, and the mechanism is analyzed. The optimized version obtains 5.1% increase in stall margin and maintains the efficiency at the design point.

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.

Numerical Investigation on the Effect of Vortex Generator on Axial Compressor Performance

2014 ◽  
Author(s):  
Ruchika Agarwal ◽  
Anand Dhamarla ◽  
Sridharan R. Narayanan ◽  
Shraman N. Goswami ◽  
Balamurugan Srinivasan
Keyword(s):  
Cross Flow ◽  
Axial Compressor ◽  
Vortex Generator ◽  
Side Surface ◽  
Suction Side ◽  
Test Case ◽  
Stall Margin ◽  
Layer Flow ◽  
Flow Effects ◽  
End Wall

The performance of the compressor blade is considerably influenced by secondary flow effects, like the cross flow on the end wall as well as corner flow separation between the wall and the blade. The present work is focused on the studying the effects of Vortex Generator (VG) on NASA Rotor 37 test case using Computational Fluid Dynamics (CFD). VG helps in controlling the inception of the stall by generating vortices and energizes the low momentum boundary layer flow which enhances the rotor performance. Three design configuration namely, Counter-rotating, Co-rotating and Plow configuration VG are selected based on the improved aerodynamic performance discussed in reference [1]. These VG are located at 90% span and 42% chord on suction side surface of the blade. Among the three configurations, the first configuration has greater impact on the end wall cross flow and flow deflection which resulted in enhanced numerical stall margin of 5.4% from baseline. The reasons for this numerical stall margin improvement are discussed in detail.

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