A Criterion for Axial Compressor Hub-Corner Stall

2006 ◽  
Author(s):  
V. M. Lei ◽  
Z. S. Spakovszky ◽  
E. M. Greitzer
Keyword(s):  
Flow Control ◽  
Substantial Reduction ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Diffusion Parameter ◽  
Design Parameters ◽  
Diffusion Limit ◽  
Specific Flow ◽  
Multi Stage ◽  
Flow Variables

This paper presents a new criterion for estimating the size and strength of three-dimensional hub-corner stall in rotors and shrouded stators of multi-stage axial compressors. A simple, first-of-a-kind description for the formation of hub-corner stall is derived, consisting of (i) a stall indicator, which quantifies the extent of the reversed flow via the local blade loading and thus indicates whether corner stall occurs, and (ii) a diffusion parameter which defines the diffusion limit. The stall indicator can be cast in terms of a Zweifel loading coefficient. The diffusion parameter is based on preliminary design type flow variables and geometry. Computational simulations and single and multi-stage compressor data are used to show the applicability of the criterion over a range of blade design parameters. The criterion also enables determination of specific flow control actions needed to mitigate hub-corner stall. To illustrate the latter a flow control blade, designed using the ideas developed, is seen to achieve a substantial reduction in the flow non-uniformity associated with hub-corner stall.

A Criterion for Axial Compressor Hub-Corner Stall

2008 ◽  
Vol 130 (3) ◽  
Author(s):  
V.-M. Lei ◽  
Z. S. Spakovszky ◽  
E. M. Greitzer
Keyword(s):  
Flow Control ◽  
Substantial Reduction ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Design Flow ◽  
Diffusion Parameter ◽  
Design Parameters ◽  
Diffusion Limit ◽  
Specific Flow ◽  
Flow Variables

This paper presents a new criterion for estimating the onset of three-dimensional hub-corner stall in axial compressor rotors and shrouded stators. A simple first-of-a-kind description of hub-corner stall formation is developed which consists of (i) a stall indicator, which quantifies the extent of the separated region via the local blade loading and thus indicates whether hub-corner stall occurs, and (ii) a diffusion parameter, which defines the diffusion limit for unstalled operation. The stall indicator can be cast in terms of a Zweifel loading coefficient. The diffusion parameter is based on preliminary design flow variables and geometry. Computational simulations and single and multistage compressor data are used to show the applicability of the criterion over a range of blade design parameters. The criterion also enables determination of specific flow control actions to mitigate hub-corner stall. As an illustration, a flow control blade, designed using the ideas developed, is seen to produce a substantial reduction in the flow nonuniformity associated with hub-corner stall.

A Passive Flow Control to Mitigate the Corner Separation in an Axial Compressor by a Slotted Rotor Blade

2019 ◽  
Author(s):  
Sungho Yoon ◽  
Rao Ajay ◽  
Venkata Chaluvadi ◽  
Vittorio Michelassi ◽  
Ramakrishna Mallina
Keyword(s):  
Flow Control ◽  
Rotor Blade ◽  
Substantial Reduction ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Suction Side ◽  
Pressure Side ◽  
Stall Margin ◽  
Radial Migration ◽  
Corner Separation

Abstract The operability of the axial compressor is generally limited by endwall flows; either at the casing mainly due to the tip leakage flows or at the hub mainly due to three-dimensional corner separations. Therefore, it is crucial to improve flows near the endwalls to enhance the operability of the compressor. Based on a last-stage with cantilevered stator vanes, a small endwall slot was introduced to a rotor blade to mitigate the hub corner separation and maximize the aerodynamic operating range of axial compressors by natural aspiration. The developed flow control technology is numerically analyzed based on the in-house High-Speed Research Compressor (HSRC) which, in turn, represents the rear stage of a modern compressor. This compressor was predicted to stall due to hub corner separation on a rotor blade based on multistage CFD analysis. A small spanwise endwall slot, connecting the pressure side and the suction side of a compressor rotor blade, was introduced near the hub to provide the by-pass flows from the pressure side to the suction side (see Figure 1). This naturally-aspirated jet significantly reduced the three-dimensional corner separation which generally occurs where the suction side meets the hub. The substantial reduction of the three-dimensional corner separation, in turn, improved the aerodynamic stall margin of the compressor. The benefit is accomplished because the low momentum region near the hub was energized due to the naturally-aspirated jet through the endwall slot and the radial migration of the low momentum flow on the suction side was significantly reduced. A systematic parametric study was conducted to better understand the flow details and optimize the flow control without sacrificing aerodynamic efficiency. It was discovered that a very small slot, smaller than 10% of span, located near the endwall, was sufficient to have a more than 6% improvement of the stall margin with a negligible efficiency penalty (less than 0.1%). The naturally-aspirated flow through the small slot eliminates the source of the corner separation at the hub platform by strengthening the flow near the hub. This, in turn, reduces the overall aerodynamic blockage by decreasing the radial migration of the low momentum flow over a third of the span. Finally, evaluations of the mechanical strength and structural dynamics of slotted rotor blades, as well as the aerodynamic impact in a multi-stage environment were conducted and its results were discussed.

Accelerated 3D Aerodynamic Optimization of Gas Turbine Blades

2014 ◽  
Author(s):  
Philipp Amtsfeld ◽  
Michael Lockan ◽  
Dieter Bestle ◽  
Marcus Meyer
Keyword(s):  
Turbine Blades ◽  
Three Dimensional ◽  
Optimization Process ◽  
Design Parameters ◽  
Blade Design ◽  
Aerodynamic Optimization ◽  
Multi Stage ◽  
Design Variables ◽  
Real Flow ◽  
Blade Model

State-of-the-art aerodynamic blade design processes mainly consist of two phases: optimal design of 2D blade sections and then stacking them optimally along a three-dimensional stacking line. Such a quasi-3D approach, however, misses the potential of finding optimal blade designs especially in the presence of strong 3D flow effects. Therefore, in this paper a blade optimization process is demonstrated which uses an integral 3D blade model and 3D CFD analysis to account for three-dimensional flow features. Special emphasis is put on shortening design iterations and reducing design costs in order to obtain a rapid automatic optimization process for fully 3D aerodynamic turbine blade design which can be applied in an early design phase already. The three-dimensional parametric blade model is determined by up to 80 design variables. At first, the most important design parameters are chosen based on a non-linear sensitivity analysis. The objective of the subsequent optimization process is to maximize isentropic efficiency while fulfilling a minimal set of constraints. The CFD model contains both important geometric features like tip gaps and fillets, and cooling and leakage flows to sufficiently represent real flow conditions. Two acceleration strategies are used to cut down the turn-around time from weeks to days. Firstly, the aerodynamic multi-stage design evaluation is significantly accelerated with a GPU-based RANS solver running on a multi-GPU workstation. Secondly, a response surface method is used to reduce the number of expensive function evaluations during the optimization process. The feasibility is demonstrated by an application to a blade which is a part of a research rig similar to the high pressure turbine of a small civil jet engine. The proposed approach enables an automatic aerodynamic design of this 3D blade on a single workstation within few days.

On the Coupling of Inverse Design and Optimization Techniques for Turbomachinery Blade Design

2006 ◽  
Author(s):  
Duccio Bonaiuti ◽  
Mehrdad Zangeneh
Keyword(s):  
Mechanical Design ◽  
Design Method ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Optimization Techniques ◽  
Operating Conditions ◽  
Inverse Design ◽  
Design Parameters ◽  
Optimization Strategy ◽  
Design And Optimization

Optimization strategies have been used in recent years for the aerodynamic and mechanical design of turbomachine components. One crucial aspect in the use of such methodologies is the choice of the geometrical parameterization, which determines the complexity of the objective function to be optimized. In the present paper, an optimization strategy for the aerodynamic design of turbomachines is presented, where the blade parameterization is based on the use of a three-dimensional inverse design method. The blade geometry is described by means of aerodynamic parameters, like the blade loading, which are closely related to the aerodynamic performance to be optimized, thus leading to a simple shape of the optimization function. On the basis of this consideration, it is possible to use simple approximation functions for describing the correlations between the input design parameters and the performance ones. The Response Surface Methodology coupled with the Design of Experiments (DOE) technique was used for this purpose. CFD analyses were run to evaluate the configurations required by the DOE to generate the database. Optimization algorithms were then applied to the approximated functions in order to determine the optimal configuration or the set of optimal ones (Pareto front). The method was applied for the aerodynamic redesign of two different turbomachine components: a centrifugal compressor stage and a single-stage axial compressor. In both cases, both design and off-design operating conditions were analyzed and optimized.

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.

Multi Stage Axial Compressor Design and Performance Evaluation

2011 ◽  
Author(s):  
Young Seok Kang ◽  
Tae Choon Park ◽  
Oh Sik Hwang ◽  
Soo Seok Yang
Keyword(s):  
Performance Test ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Pressure Rise ◽  
Multi Stage ◽  
Compressor Design ◽  
Test Result ◽  
Highly Loaded ◽  
Small Aircraft

Recently, needs for Unmanned Air Vehicle (UAV) and small aircraft are increasing and demands for small turbo jet or turbo fan engines are also increasing. Then, size and weight are the two main restrictions in UAV or small aircraft propulsion system applications. One method for resolving such a problem is to increase the pressure rise per stage and to reduce the number of stages. Nowadays, matured compressor aerodynamic design techniques enable us to design highly loaded axial compressors. This paper covers from the design step of a highly loaded transonic axial compressor to the performance test result and its analysis. At the fore part of the paper, aerodynamic process of a multi stage axial compressor is introduced. To satisfy both of the mass flow and pressure rise, the compressor should rotate at a high rotational speed. Therefore the transonic flow field forms in the rotor stages and it is designed with a relatively high pressure rise per stage to satisfy its design target. Basically, one dimensional and quasi three dimensional compressor design were carried with compressor design codes. The compressor stage consists of 3 stages, and the bulk pressure ratio is 2.5. The first stage is burdened with the highest pressure ratio and less pressure rises occur in the following stages. Also it is designed that tip Mach number of the first rotor row does not exceed 1.3. The final design was confirmed by iterating three dimensional CFD calculations to satisfy design target and some design intentions. In the latter part of the paper, its performance test processes are briefly introduced. The performance test result showed that the overall compressor performance targets; pressure ratio and efficiency are well achieved. From the test results, we found some clues for further improvement and optimization of the compressor aerodynamic performance.

A Prediction Model for Corner Separation/Stall in Axial Compressors

2010 ◽  
Author(s):  
Xianjun Yu ◽  
Baojie Liu
Keyword(s):  
Aspect Ratio ◽  
Pressure Gradient ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Experimental Results ◽  
Diffusion Parameter ◽  
Experimental Conditions ◽  
Corner Separation ◽  
Axial Compressors ◽  
Key Factor

Endwall corner stall can cause significant aerodynamic blockage and losses production. Hence, pre-prediction of it during the preliminary processes of compressors is important. However, Lieblein’s diffusion factor often fails near the endwall region for the strong three-dimensionality flow effects. A new model for predicting endwall corner stall phenomenon in axial compressors was developed based on the methodology used by Lei et al. [1] (J. Turbomach. 2008, 031006). At first, the influencing factors for the flows of endwall corner separation/stall were analyzed by numerical simulations. The results showed that, besides the parameters determining the loading of a two-dimensional blade profile, blade aspect ratio was also a key factor. Then, by using both the theoretical and empirical methods, a modified diffusion parameter, which can be used as a criterion for axial compressor corner stall, was defined to consider the combined effects of three factors: the streamwise pressure gradient, the circumferential pressure gradient and the passage mass-flow-rate redistribution effect (controlled mainly by blade aspect ratio). Finally, the stall criterion was validated by experimental results of various test facilities with different blade geometries and experimental conditions. The results showed that the modified diffusion parameter can predict the corner separation/stall flows in a good agreement with the experimental results in axial compressors without blade three-dimensional designs.

Viscous Throughflow Modelling for Multi-Stage Compressor Design

1992 ◽  
Author(s):  
M. A. Howard ◽  
S. J. Gallimore
Keyword(s):  
Shear Force ◽  
High Speed ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Design System ◽  
Flow Angle ◽  
Axial Compressors ◽  
Multi Stage ◽  
Viscous Shear ◽  
Compressor Design

An existing throughflow method for axial compressors, which accounts for the effects of spanwise mixing using a turbulent diffusion model, has been extended to include the viscous shear force on the endwall. The use of a shear force, consistent with a no-slip condition, on the annulus walls in the throughflow calculations allows realistic predictions of the velocity and flow angle profiles near the endwalls. The annulus wall boundary layers are therefore incorporated directly in the throughflow prediction. This eliminates the need for empirical blockage factors or independent annulus boundary layer calculations. The axisymmetric prediction can be further refined by specifying realistic spanwise variations of loss coefficient and deviation to model the three-dimensional endwall effects. The resulting throughflow calculation gives realistic predictions of flow properties across the whole span of a compressor. This is confirmed by comparison with measured data from both low and high speed multi-stage machines. The viscous throughflow method has been incorporated into an axial compressor design system. The method predicts the meridional velocity defects in the endwall region and consequently blading can be designed which allows for the increased incidence, and low dynamic head, near to the annulus walls.

scholarly journals Physics of Airfoil Clocking in a High-Speed Axial Compressor

1998 ◽  
Author(s):  
Daniel J. Dorney ◽  
Om P. Sharma ◽  
Karen L. Gundy-Burlet
Keyword(s):  
Relative Motion ◽  
High Speed ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Unsteady Aerodynamics ◽  
Navier Stokes ◽  
Jet Engines ◽  
Axial Compressors ◽  
Multi Stage ◽  
Significant Performance

Axial compressors have inherently unsteady flow fields because of relative motion between rotor and stator airfoils. This relative motion leads to viscous and inviscid (potential) interactions between blade rows. As the number of stages increases in a turbomachine, the buildup of convected wakes can lead to progressively more complex wake/wake and wake/airfoil interactions. Variations in the relative circumferential positions of stators or rotors can change these interactions, leading to different unsteady forcing functions on airfoils and different compressor efficiencies. In addition, as the Mach number increases the interaction between blade rows can be intensified due to potential effects. In the current study an unsteady, quasi-three-dimensional Navier-Stokes analysis has been used to investigate the unsteady aerodynamics of stator clocking in a 1-1/2 stage compressor, typical of back stages used in high-pressure compressors of advanced commercial jet engines. The effects of turbulence have been modeled with both algebraic and two-equation models. The results presented include steady and unsteady surface pressures, efficiencies, boundary layer quantities and turbulence quantities. The main contribution of the current work has been to show that airfoil clocking can produce significant performance variations at the Mach numbers associated with an engine operating environment. In addition, the growth of turbulence has been quantified to aid in the development of models for the multi-stage steady analyses used in design systems.

scholarly journals Quasi Three-Dimensional Design for a Novel Turbo-Vapor Compressor and the Last Stage of a Low-Pressure Steam Turbine

2021 ◽  
Vol 85 (2) ◽  
pp. 1-13
Author(s):  
Amin Mobarak ◽  
Mostafa Shawky Abdel Moez ◽  
Shady Ali
Keyword(s):  
Three Dimensional ◽  
Axial Compressor ◽  
Low Pressure ◽  
Turbine Rotor ◽  
Combined Cycle ◽  
Operating Conditions ◽  
3D Design ◽  
Compressor Rotor ◽  
Multi Stage ◽  
Last Stage

Turbo-vapor compressors (TVCs) are used to create a vacuum pressure in the evaporator of a novel combined cycle for electricity and freshwater production invented by Amin Mobarak. A novel design conceived of a TVC is introduced to increase the efficiency, allowable mass flow rate and reduce costs and losses. The system consists of a single axial compressor rotor followed by a single axial turbine rotor, which drives the upstream compressor, allowing high flow rates. A quasi-3D design is carried out for the TVC to calculate the flow velocity components and angles and ensure that the turbo-vapor turbine work is equal to the turbo-vapor compressor work. A preliminary design of the low-pressure power turbine (LPT) is done to examine the size and number of stages. The (LPT) size is twice the size of TVC at typical cycle operating conditions. A three-stage design is the most appropriate choice for the number of stages. It satisfies the accelerating relative flow condition at the last stage over a range of flow coefficients. A quasi-3D design is carried out for the LPT's last stage to ensure a multi-stage power turbine's safe design.

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