Effects of recessed blade tips on stall margin in a transonic axial compressor

This paper deals with the numerical investigations on the effect of trailing edge crenulation on the performance of a transonic axial compressor rotor. Crenulation is broadly considered as a series of small notches or slots at the edge of a thin object, like a plate. Incorporating such notches at the trailing edge of a compressor cascade has shown beneficial effect in terms of reduction in total pressure loss due to enhanced mixing in the wake region. These notches act as vortex generators to produce counter rotating vortices, which increase intermixing between the free stream flow and the low momentum wake fluid. Considering the positive effects of crenulation in a cascade, it was hypothesized that the same technique would work in a rotating compressor to enhance its performance and stall margin. However, the present CFD simulations on a transonic compressor rotor have given mixed results. Whereas the peak total pressure ratio in the presence of trailing edge crenulation reduced, the stall margin improved by 2.97% compared to the rotor with straight edge blades. The vortex generation at the crenulated trailing edge was not as strong as reported in case of linear compressor cascade, but it was able to influence the flow field in the rotor tip region so as to energize the low momentum end-wall flow in the aft part of the blade passage. This beneficial effect delayed flow separation and allowed the mass flow rate to be reduced to still lower levels resulting in improved stall margin. The reduction in pressure ratio with crenulation was surprising and might be due to increased mixing losses downstream of the blade.

A CFD Study of Circumferential Groove Casing Treatments in a Transonic Axial Compressor

Volume 7: Turbomachinery, Parts A, B, and C ◽

10.1115/gt2010-23606 ◽

2010 ◽

Cited By ~ 2

Author(s):

Haixin Chen ◽

Xudong Huang ◽

Ke Shi ◽

Song Fu ◽

Matthew A. Bennington ◽

...

Keyword(s):

Axial Compressor ◽

Data Sampling ◽

Stall Margin ◽

Casing Treatment ◽

Circumferential Groove ◽

Compressor Rotor ◽

Radial Profiles ◽

Flow Mechanisms ◽

Transonic Axial Compressor ◽

And Performance

Numerical investigations were conducted to predict the performance of a transonic axial compressor rotor with circumferential groove casing treatment. The Notre Dame Transonic Axial Compressor (ND-TAC) was simulated by Tsinghua University with an in-house CFD code (NSAWET) for this work. Experimental data from the ND-TAC were used to define the geometry, boundary conditions and data sampling method for the numerical simulation. These efforts, combined with several unique simulation approaches, such as non-matched grid boundary technology to treat the periodic boundaries and interfaces between groove grids and the passage grid, resulted in good agreement between the numerical and experimental results for overall compressor performance and radial profiles of exit total pressure. Efforts were made to study blade level flow mechanisms to determine how the casing treatment impacts the compressor’s stall margin and performance. The flow structures in the passage, the tip gap and the grooves as well as their mutual interactions were plotted and analyzed. The flow and momentum transport across the tip gap in the smooth wall and the casing treatment configurations were quantitatively compared.

AERODYNAMIC PERFORMANCE OF AN AXIAL COMPRESSOR WITH A CASING GROOVE COMBINED WITH INJECTION

Transactions of the Canadian Society for Mechanical Engineering ◽

10.1139/tcsme-2013-0018 ◽

2013 ◽

Vol 37 (3) ◽

pp. 283-292 ◽

Cited By ~ 5

Author(s):

Dae-Woong Kim ◽

Jin-Hyuk Kim ◽

Kwang-Yong Kim

Keyword(s):

Turbulence Model ◽

Parametric Study ◽

Stokes Equations ◽

Three Dimensional ◽

Axial Compressor ◽

Aerodynamic Performance ◽

Navier Stokes ◽

Navier Stokes Equations ◽

Stall Margin ◽

Transonic Axial Compressor

Aerodynamic performance of a transonic axial compressor with a casing groove combined with injection has been investigated in this work. Three-dimensional Reynolds-averaged Navier–Stokes equations with k-ε turbulence model are discretized by finite volume approximations and solved on hexahedral grids for the flow analyses. For parametric study, the front and rear lengths and height of the casing groove are selected as the geometric parameters and are changed with constant injection to investigate their effects on the stall margin and peak adiabatic efficiency. As a result of the parametric study, the maximum stall margin and peak adiabatic efficiency are found to be obtained in the axial compressor having 70% height of the reference groove. The results show that the application of the casing groove combined with injection to an axial compressor is effective for the simultaneous improvement of both the stall margin and peak adiabatic efficiency of the compressor.

Design Optimization of a Dual-Bleeding Recirculation Channel to Enhance Operating Stability of a Transonic Axial Compressor

Energies ◽

10.3390/en15010159 ◽

2021 ◽

Vol 15 (1) ◽

pp. 159

Author(s):

Tien-Dung Vuong ◽

Kwang-Yong Kim

Keyword(s):

Design Optimization ◽

Pressure Ratio ◽

Three Dimensional ◽

Axial Compressor ◽

Aerodynamic Performance ◽

Maximum Pressure ◽

Stall Margin ◽

Design Variables ◽

Operating Stability ◽

Transonic Axial Compressor

The present work performed a comprehensive investigation to find the effects of a dual-bleeding port recirculation channel on the aerodynamic performance of a single-stage transonic axial compressor, NASA Stage 37, and optimized the channel’s configuration to enhance the operating stability of the compressor. The compressor’s performance was examined using three parameters: The stall margin, adiabatic efficiency, and pressure ratio. Steady-state three-dimensional Reynolds-averaged Navier–Stokes analyses were performed to find the flow field and aerodynamic performance. The results showed that the addition of a bleeding channel increased the recirculation channel’s stabilizing effect compared to the single-bleeding channel. Three design variables were selected for optimization through a parametric study, which was carried out to examine the influences of six geometric parameters on the channel’s effectiveness. Surrogate-based design optimization was performed using the particle swarm optimization algorithm coupled with a surrogate model based on the radial basis neural network. The optimal design was found to increase the stall margin by 51.36% compared to the case without the recirculation channel with only 0.55% and 0.28% reductions in the peak adiabatic efficiency and maximum pressure ratio, respectively.

Parametric Study on Aerodynamic Performance of a Transonic Axial Compressor with a Casing Groove and Tip Injection

Applied Mechanics and Materials ◽

10.4028/www.scientific.net/amm.284-287.872 ◽

2013 ◽

Vol 284-287 ◽

pp. 872-877 ◽

Cited By ~ 2

Author(s):

Dae Woong Kim ◽

Jin Hyuk Kim ◽

Kwang Yong Kim

Keyword(s):

Parametric Study ◽

Stokes Equations ◽

Pressure Ratio ◽

Three Dimensional ◽

Axial Compressor ◽

Aerodynamic Performance ◽

Navier Stokes ◽

Stall Margin ◽

Transonic Axial Compressor ◽

Tip Injection

This paper presents a parametric study on aerodynamic performance of a transonic axial compressor combined with a casing groove and tip injection using three-dimensional Reynolds-average Navier-Stokes equations. The front and rear lengths and height of the groove are selected as the geometric parameters to investigate their effects on the stall margin and peak adiabatic efficiency. These parameters are changed with constant injection. The validation of the numerical results is performed in comparison with experimental data for the total pressure ratio and adiabatic efficiency. As the results of the parametric study, the maximum stall margin and peak adiabatic efficiency are obtained in the axial compressor having 70% groove height of the reference groove. The stall margin and peak adiabatic efficiency in other cases are also improved in comparison with the axial compressors with the smooth casing and reference groove. The results show that both the stall margin and the peak adiabatic efficiency are considerably improved by the application of the casing groove combined with tip injection in an axial compressor.

Optimization of a Transonic Axial Compressor Considering Interaction of Blade and Casing Treatment to Improve Operating Stability

Volume 7: Turbomachinery, Parts A, B, and C ◽

10.1115/gt2011-45404 ◽

2011 ◽

Cited By ~ 1

Author(s):

Jin-Hyuk Kim ◽

Kwang-Jin Choi ◽

Kwang-Yong Kim

Keyword(s):

Stokes Equations ◽

Axial Compressor ◽

Tip Clearance ◽

Pareto Optimal ◽

Angle Distribution ◽

Pareto Optimal Solutions ◽

Optimal Solutions ◽

Stall Margin ◽

Multi Objective ◽

Transonic Axial Compressor

A multi-objective optimization of a transonic axial compressor with circumferential casing grooves has been carried out in the present study. A hybrid multi-objective evolutionary algorithm coupled with response surface approximation is used to optimize the stall margin and design speed efficiency of the transonic axial compressor. Three-dimensional Reynolds-averaged Navier-Stokes equations with the shear stress transport turbulence model are discretized by finite volume approximations and solved on hexahedral grids for the flow analysis. The stall margin and peak adiabatic efficiency are used as objective functions for the optimization. Tip clearance and angle distribution at blade tip are considered as design variables in addition to the depth of the circumferential casing grooves which was more sensitive variable than the width in the previous work (GT2010-22396). Latin-hypercube sampling as design-of-experiments is used to generate twenty five design points within the design space. A fast non-dominated sorting genetic algorithm with an ε–constraint strategy for the local search is applied to determine the global Pareto-optimal solutions. The trade-off between two objectives is determined and discussed with respect to the representative clusters in the Pareto-optimal solutions compared to the smooth casing.

Effects of Blade Damage on the Performance of a Transonic Axial Compressor Rotor

Volume 8: Turbomachinery, Parts A, B, and C ◽

10.1115/gt2012-68324 ◽

2012 ◽

Author(s):

Yanling Li ◽

Abdulnaser Sayma

Keyword(s):

Axial Compressor ◽

Rotating Stall ◽

Rotor Blades ◽

Design Stage ◽

Compressor Blades ◽

Stall Margin ◽

Compressor Rotor ◽

Surge Margin ◽

Compressor Performance ◽

Transonic Axial Compressor

Gas turbine axial compressor blades may encounter damage during service for various reasons. Debris from casing or foreign objects may impact blades causing damage near the rotor’s tip. This may result in deterioration of performance and reduction in the surge margin. Ability to assess the effect of damaged blades on the compressor performance and stability is important at both the design stage and in service. The damage to compressor blades breaks the cyclic symmetry of the compressor assembly. Thus computations have to be performed using the whole annulus. Moreover, if rotating stall or surge occurs, the downstream boundary conditions are not known and simulations become difficult. This paper presents an unsteady CFD analysis of compressor performance with tip curl damage. Tip curl damage typically occurs when rotor blades hit a loose casing liner. The computations were performed up to the stall boundary, predicting rotating stall patterns. The aim is to assess the effect of blade damage on stall margin and provide better understanding of the flow behaviour during rotating stall. Computations for the undamaged rotor are also performed for comparison. A transonic axial compressor rotor is used for the time-accurate numerical unsteady flow simulations, with a variable choked nozzle downstream simulating an experimental throttle. One damaged blade was introduced in the rotor assembly and computations were performed at 60% of the design rotational speed. It was found that there is no significant effect on the compressor stall margin due to one damaged blade despite the differences in rotating stall patterns between the undamaged and damaged assemblies.

Stability Enhancement of a Single-Stage Transonic Axial Compressor Using Inclined Oblique Slots

Energies ◽

10.3390/en14092346 ◽

2021 ◽

Vol 14 (9) ◽

pp. 2346

Author(s):

Tien-Dung Vuong ◽

Kwang-Yong Kim

Keyword(s):

Pressure Ratio ◽

Three Dimensional ◽

Axial Compressor ◽

Single Stage ◽

Rotor Blades ◽

Navier Stokes ◽

Optimized Design ◽

Stall Margin ◽

Casing Treatment ◽

Transonic Axial Compressor

A casing treatment using inclined oblique slots (INOS) is proposed to improve the stability of the single-stage transonic axial compressor, NASA Stage 37, during operation. The slots are installed on the casing of the rotor blades. The aerodynamic performance was estimated using three-dimensional steady Reynolds-Averaged Navier-Stokes analysis. The results showed that the slots effectively increased the stall margin of the compressor with slight reductions in the pressure ratio and adiabatic efficiency. Three geometric parameters were tested in a parametric study. A single-objective optimization to maximize the stall margin was carried out using a Genetic Algorithm coupled with a surrogate model created by a radial basis neural network. The optimized design increased the stall margin by 37.1% compared to that of the smooth casing with little impacts on the efficiency and pressure ratio.