Experimental Studies on Stall Behavior in a Single Stage Transonic Axial Flow Compressor

2013 ◽  
Author(s):  
Dilipkumar Bhanudasji Alone ◽  
Subramani Satish Kumar ◽  
Shobhavathy M. Thimmaiah ◽  
Janaki Rami Reddy Mudipalli ◽  
A. M. Pradeep ◽  
...  
Keyword(s):  
Mach Number ◽  
Total Pressure ◽  
Pressure Ratio ◽  
Tangential Velocity ◽  
Axial Flow ◽  
Single Stage ◽  
Flow Angle ◽  
Axial Flow Compressor ◽  
Flow Compressor ◽  
Total Pressure Ratio

This paper describes the study of flow behavior of the transonic compressor stage in un-stalled and stalled conditions. Experiments were carried out in an open circuit single stage transonic axial flow compressor test rig. The test compressor was designed for 1.35 total to total pressure ratio at corrected mass flow rate of 22 kg/s. Both steady and unsteady measurements were carried out. The operating envelop of the compressor was experimentally determined to demark the stable and unstable operating range of the compressor at different operating speeds. Variations in the rotor inlet axial and tangential velocity in the tip region were studied using a calibrated single component hot wire probe. The compressor blade element performance was obtained at full flow and near stall conditions using a three hole aerodynamic probe. The variation in flow parameters like absolute flow angle, axial Mach number, absolute Mach number, tangential Mach number, static and total pressure ratio profiles at the rotor exit were obtained and their variations along the blade height were studied at full flow and near stall conditions. Static pressure variation in the tip region along the rotor chord was studied which showed reduction in slope as stall approached. Hotwire measurements showed abrupt variation in the axial velocity as compared to tangential velocity at stalled condition. It was observed that the flow turned in tangential direction at stall, as tangential component of velocity shows more fluctuations at stall in comparison with unstalled condition. The FFT analysis of the raw signals was performed and it was observed that the nature of the rotating stall was abrupt and stall cell travels nearly at half the rotor speed.

An Experimental Investigation About the Effect of the Partial Casing Treatment on a Transonic Axial-Flow Compressor Performance Under Inlet Distortion

2012 ◽  
Author(s):  
Maoyi Li ◽  
Wei Yuan ◽  
Xizhen Song ◽  
Yajun Lu ◽  
Zhiping Li ◽  
...  
Keyword(s):  
Total Pressure ◽  
Pressure Ratio ◽  
Axial Flow ◽  
Peak Efficiency ◽  
Stall Margin ◽  
Casing Treatment ◽  
Axial Flow Compressor ◽  
Inlet Distortion ◽  
Flow Compressor ◽  
Total Pressure Ratio

The traditional annulus casing treatment often pays the price of lowered efficiency for improving the stall margin of a compressor under inlet distortion. In view of the unsymmetry of the inlet flow-field of compressors, partial casing treatment was used to control the flow in a transonic axial-flow compressor with arc-skewed-slots deployed at different circumferential positions under inlet distortion. The experimental results indicate that when the partial casing treatment is arranged on the undistorted and distorted sectors, the stall margin is enhanced by 8.02%, with the relative peak efficiency improved simultaneously by 2.143%, compared with the case of solid casing at 98% rotating speed. By contrast, the traditional casing treatment increases the stall-margin by 23.13%, but decreases the relative peak efficiency by 0.752%. By analyzing dynamic and static experimental data, the mechanism underlying the partial casing treatment was also studied in detail here. The disturbances of inlet flow were restrained by annulus casing treatment, nevertheless the total pressure ratio was decreased obviously in the distorted sector. As a result, the stall-margin is improved, but the relative peak efficiency is decreased too. When the partial casing treatment was arranged on the undistortded and distorted sectors, the stall disturbances was thereby restrained. So the stall margin was enhanced. In addition, the total pressure ratio was improved by the partial casing treatment in the distorted and transition sectors, and thus the relative peak efficiency was also increased markedly.

Numerical Studies on Effect of Stepped Tip Clearance Height on the Performance of Single Stage Transonic Axial Flow Compressor

2013 ◽  
Author(s):  
Pritam Batabyal ◽  
Dilipkumar B. Alone ◽  
S. K. Maharana
Keyword(s):  
Numerical Study ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Axial Flow ◽  
Single Stage ◽  
Tip Clearance ◽  
Baseline Model ◽  
Axial Flow Compressor ◽  
Design Speed ◽  
Flow Compressor

This paper presents a numerical case study of various stepped tip clearances and their effect on the performance of a single stage transonic axial flow compressor, using commercially available software ANSYS FLUENT 14.0. A steady state, implicit, three dimensional, pressure based flow solver with SST k-Ω turbulence model has been selected for the numerical study. The stepped tip clearances have been compared with the baseline model of zero tip clearance at 70% and 100 % design speed. It has been observed that the compressor peak stage efficiency and maximum stage pressure ratio decreases as the tip clearances in the rear part are increased. The stall margin also increases with increase in tip clearance compared to the baseline model. An ‘optimum’ value of stepped tip clearance has been obtained giving peak stage compressor performance. The CFD results have been validated with the earlier published experimental data on the same compressor at 70% design speed.

Numerical Efforts of Aerodynamic Re-Design in a Single-Stage Transonic Axial Compressor: Part 1 — Stator Design

2017 ◽  
Author(s):  
Justin (Jongsik) Oh
Keyword(s):  
Pressure Ratio ◽  
Axial Compressor ◽  
Axial Flow ◽  
Single Stage ◽  
Design Flow ◽  
Design Parameters ◽  
Original Design ◽  
Circular Arc ◽  
Surge Margin ◽  
Flow Compressor

In many aerodynamic design parameters for the axial-flow compressor, three variables of tailored blading, blade lean and sweep were considered in the re-design efforts of a transonic single stage which had been designed in 1960’s NASA public domains. As Part 1, the re-design was limited to the stator vane only. For the original MCA (Multiple Circular Arc) blading, which had been applied at all radii, the CDA (Controlled Diffusion Airfoil) blading was introduced at midspan as the first variant, and the endwalls of hub and casing (or tip) were replaced with the DCA (Double Circular Arc) blading for the second variant. Aerodynamic performance was predicted through a series of CFD analysis at design speed, and the best aerodynamic improvement, in terms of pressure ratio/efficiency and operability, was found in the first variant of tailored blading. It was selected as a baseline for the next design efforts with blade lean, sweep and both combined. Among 12 variants, a case of positively and mildly leaned blades was found the most attractive one, relative to the original design, providing benefits of an 1.0% increase of pressure ratio at design flow, an 1.7% increase of efficiency at design flow, a 10.5% increase of the surge margin and a 32.3% increase of the choke margin.

The Development of an Aerodynamic Performance Prediction Tool for Modern Axial Flow Compressor Profiles

2006 ◽  
Author(s):  
Dario Bruna ◽  
Carlo Cravero ◽  
Mark G. Turner
Keyword(s):  
Performance Prediction ◽  
Parametric Design ◽  
Flow Analysis ◽  
Axial Flow ◽  
Navier Stokes ◽  
Flow Angle ◽  
Flow Solver ◽  
Aerodynamic Loss ◽  
Axial Flow Compressor ◽  
Flow Compressor

The development of a computational tool (MP-LOS) for the aerodynamic loss modeling and prediction for axial-flow compressor blade sections is presented in this paper. A state-of-the-art quasi 3-D flow solver, MISES, has been used for the flow analysis on existing airfoil geometries in many working conditions. Different values of inlet flow angle, inlet Mach number, AVDR, Reynolds number and solidity have been chosen to investigate a possible working range. The target is a loss prediction formulation that will be introduced into throughflow or axisymmetric Navier-Stokes codes for the performance prediction of multistage axial flow compressors. The loss coefficient has been correlated to the flow parameters that have shown an influence on the profile loss for the blades under study. The proposed correlation, using the described computational approach, can be extended to any profile family with the aid of any code for the parametric design of blade profiles.

scholarly journals Effect of Circumferential Single Casing Groove Location on the Flow Stability under Tip-Clearance Effect in a Transonic Axial Flow Compressor Rotor

Energies ◽  
2021 ◽  
Vol 14 (19) ◽  
pp. 6143
Author(s):  
Xiaoxiong Wu ◽  
Bo Liu ◽  
Botao Zhang ◽  
Xiaochen Mao
Keyword(s):  
Control Mechanism ◽  
Incidence Angle ◽  
Axial Flow ◽  
Flow Stability ◽  
Tip Clearance ◽  
Tip Leakage Flow ◽  
Flow Angle ◽  
Axial Flow Compressor ◽  
Compressor Rotor ◽  
Flow Compressor

Numerical simulations have been performed to study the effect of the circumferential single-grooved casing treatment (CT) at multiple locations on the tip-flow stability and the corresponding control mechanism at three tip-clearance-size (TCS) schemes in a transonic axial flow compressor rotor. The results show that the CT is more efficient when its groove is located from 10% to 40% tip axial chord, and G2 (located at near 20% tip axial chord) is the best CT scheme in terms of stall-margin improvement for the three TCS schemes. For effective CTs, the tip-leakage-flow (TLF) intensity, entropy generation and tip-flow blockage are reduced, which makes the interface between TLF and mainstream move downstream. A quantitative analysis of the relative inlet flow angle indicates that the reduction of flow incidence angle is not necessary to improve the flow stability for this transonic rotor. The control mechanism may be different for different TCS schemes due to the distinction of the stall inception process. For a better application of CT, the blade tip profile should be further modified by using an optimization method to adjust the shock position and strength during the design of a more efficient CT.

Some Blade Designs for an Axial Flow Compressor Stage

1954 ◽  
Vol 58 (517) ◽  
pp. 61-64
Author(s):  
R. G. Taylor
Keyword(s):  
Mach Number ◽  
Axial Flow ◽  
Compressor Stage ◽  
Axial Flow Compressor ◽  
Design Specifications ◽  
Reaction Condition ◽  
Forced Vortex ◽  
Vortex Solutions ◽  
Flow Compressor ◽  
Constant Mach Number

Two design conditions for an axial flow compressor stage are proposed and examined. These are, the constant reaction condition (incorporating I “ radial equilibrium ”), and the condition that the Mach number at inlet to the rotor shall be invariant with radius. In addition, the combination of these two properties in one stage is considered. It is found, with further assumptions regarding the nature of the flow, that a forced vortex type of flow will satisfy both design specifications. The forced vortex solutions for the various cases are presented, and for constant Mach number at inlet to the rotor, more general solutions are given.

Flow Studies on a Single Stage Transonic Axial Flow Compressor Retrofitted With Circumferential Grooves and Varied Rotor-Stator Axial Gap

2017 ◽  
Author(s):  
Anand P. Darji ◽  
Dilipkumar Bhanudasji Alone ◽  
Chetan S. Mistry
Keyword(s):  
Numerical Study ◽  
Computational Study ◽  
Axial Flow ◽  
Single Stage ◽  
Experimental Results ◽  
Stall Margin ◽  
Axial Flow Compressor ◽  
Transonic Compressor ◽  
Stall Margin Improvement ◽  
Flow Compressor

A transonic axial flow compressor undergoes severe vibrations due to instabilities like stall and surge when it operates at lower mass flow rate in the absence of any control devices. In present study, the attempt was made to understand the combine impact of circumferential casing grooves (CCG) of constant aspect ratio and different axial spacing between rotor and stator on the operating stability of single stage transonic axial compressor and that of rotor alone using numerical simulation. The optimum rotor-stator gap in the presence of grooved casing treatment was identified. The steady state numerical analysis was performed by using three-dimensional Reynolds Average Navier-Stokes equation adapting shear stress transport (SST) k-ω turbulence model. The study is reported in two sections. First section includes the detailed numerical study on baseline case having smooth casing wall (SCW). The computational results were validated with the experimental results available at Propulsion Division of CSIR-NAL, Bangalore. The computational study shows good agreement with experimental results. The second section comprises the effects of optimum designs of CCG and various axial spacing on the stall margin improvement of transonic compressor. Current computational study shows that the axial spacing between rotor and stator is an important parameter for improvement in stall margin not only for SCW but also for CCG. Therefore, the highest stall margin improvement of 9% has achieved for 75% axial spacing.

Numerical Simulation of Steady Air Injection Flow Control Effects on a Transonic Axial Flow Compressor

2012 ◽  
Vol 224 ◽  
pp. 352-357
Author(s):  
Islem Benhegouga ◽  
Ce Yang
Keyword(s):  
Mass Flow Rate ◽  
Mass Flow ◽  
Flow Rate ◽  
Pressure Ratio ◽  
Axial Flow ◽  
Leading Edge ◽  
Air Injection ◽  
Axial Flow Compressor ◽  
Blade Tip ◽  
Flow Compressor

In this work, steady air injection upstream of the blade leading edge was used in a transonic axial flow compressor, NASA rotor 37. The injectors were placed at 27 % upstream of the axial chord length at blade tip, the injection mass flow rate is 3% of the chock mass flow rate, and 3 yaw angles were used, respectively -20°, -30°, and -40°. Negative yaw angles were measured relative to the compressor face in opposite direction of rotational speeds. To reveal the mechanism, steady numerical simulations were performed using FINE/TURBO software package. The results show that the stall mass flow can be decreased about 2.5 %, and an increase in the total pressure ratio up to 0.5%.

Oxy-Fuel Turbine Through-Flow Design

2014 ◽  
Author(s):  
Majed Sammak ◽  
Srikanth Deshpande ◽  
Magnus Genrup
Keyword(s):  
Mach Number ◽  
Total Pressure ◽  
Tangential Velocity ◽  
Flow Angle ◽  
Power Turbine ◽  
Through Flow ◽  
The Mean ◽  
Line Design ◽  
Reaction Degree ◽  
Relative Flow

The objective of the paper is to present the through-flow design of a twin-shaft oxy-fuel turbine. The through-flow design is the subsequent step after the turbine mean-line design. The through-flow phase analyses the flow in both axial and radial directions, where the flow is computed from hub to tip and along streamlines. The parameterization of the through-flow is based on the mean-line results, so principal features such as blade angles at the mean-line into the through-flow phase should be retained. Parameters such as total inlet pressure and temperature, mass flow, rotation speed and turbine geometries are required for the through-flow modelling. The through-flow study was performed using commercial software — AxCent(™) from Concepts NREC. The rotation speed of the twin-shaft power turbine was set to 7200 rpm, while the power turbine was set to 4800 rpm. The mean-line design determined that the twin-shaft turbine should be designed with two compressor turbine stages and three power turbine stages. The through-flow objective was to study the variations in the thermodynamic parameters along the blade. The power turbine last-stage design was studied because of the importance of determining exit Mach number distribution of the rotor tip. The last stage was designed with damped forced condition. The term ‘damped’ is used because the opening from the tip to the hub is limited to a certain value rather than maintaining the full concept of forced vortex. The study showed the parameter distribution of relative Mach number, total pressure and temperature, relative flow angle and tangential velocity. Through-flow results at 50% span and mean-line results showed reasonable agreement between static pressure, total pressure, reaction degree and total efficiency. Other parameters such as total temperature and relative Mach number showed some difference which can be attributed to working fluid in AxCent being pure CO2. The relative tip Mach number at rotor exit was 1.03, which is lower than the maximum typically allowed value of 1.2. The total pressure distribution was smooth from hub to tip which minimizes the spanwise gradient of total pressure and thus reduces the strength of secondary vortices. The reaction degree distribution was presented in the paper and no problems were revealed in the reaction degree at the hub. Rotor blades were designed to produce a smooth exit relative flow angle distribution. The relative flow angle varied by approximately 5° from hub to tip. The tangential velocity distribution was proportional to blade radius, which coincided with forced vortex design. Through-flow design showed that the mean-line design of a twin-shaft oxy-fuel turbine was suitable.

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