International Journal of Emerging Technology and Advanced Engineering Website: www.ijetae.com (ISSN 2250-2459, ISO 9001:2008 Certified Journal, Volume 3, Issue 8, August 2013)

Effect of Variations in Aspect Ratio on Single Stage Axial Flow Compressor Using Numerical Analysis Kumbhar Anil H.1, Aashish Agarwal2 1

2

PG Student, Asso. Professor, Technocrats Institute of Technology, Bhopal, M.P, India. At the stall point aerodynamic instabilities would be initiated. Beyond this point local flow re-circulation would be induced due to flow separation. The present day research and developmental efforts in the area of axial flow compressors for aero engine application are aimed to improve the efficiency. In experimental, numerical and theoretical investigations have been performed in the past, studying the effects of aspect ratio on compressor performance, but yielded mixed results Meherwan P. Boyce [2005],G.J. Fohmi[1971], J.H. Horlock [1967], Ronuld J. Steinke und James E. Crozcse [1967], observed a Effects of Aspect Ratio on blade loading, blade excitation, and the pre-twist blade angles. At high aspect ratios the blades had to be designed with mid span shrouds, and tip shrouds. This decreases the efficiency of the stage; however, without the shrouds the pre-twist blade angle had to be increased and the blade excitation resulted in blade failure. Ronuld J. Steinke und James E. Crozcse study under the Lewis Reseurch Center Cleveland, Ohio. Results of thier studies do not indicate any basic limit on aspect ratio other than possible supersonic meridional velocities for acceptable design point performance. They note that off-design performance and mechanical design problems may limit the maximum usable aspect ratio. The high axial pressure gradients associated with higher aspect ratio blading could require refinements in the profile loss correlation that was used in order to predict performance accurately. Lars Sommer , Dieter Bestle [2011], Introduces a new curvature based design parameterization of twodimensional high pressure compressor blade sections to be used in a multi-criteria aerodynamic design optimization process. Lingen Chen, Jun Luo , Fengrui Sun & Chih Wu.[2005], This paper focuses on the efficiency optimization of an axial-flow compressor stage using onedimensional flow-theory. The universal characteristic relation a axial-flow compressor stage is obtained. The influences of various parameters on the relations are also analyzed. This paper describes the effect of aspect ratio by using commercial code AxStream to investigate the influence of aspect ratio on a single stage subsonic axial flow compressor.

Abstract— The performance of single-stage axial flow compressors has been observed to be adversely affected by an aspect ratio (the ratio of blade height to axial chord length). The current investigation deals with a numerical based analysis for the effects of aspect ratio for single stage subsonic axial flow compressor through CFD analysis using commercial code AxStream. The investigation aims to identify the effect of varying aspect ratio for single stage axial flow compressor with fixed aspect ratio through the stage and change the AR for IGV and stator blade while AR for rotor blade are same . First the analysis has been carried out for the single stage axial flow compressor having an AR2 through the stage (IGV, Rotor &stator) and then rotor blade AR2 kept constant and change the AR3 for stator blade and IGV. The study has predicted for estimating the performance characteristics, such as power, pressure rise and diffusion factor of the compressor stage is presented. This study shows that best operating conditions is getting in case of AR 323 for the referred single stage axial flow compressor. Keywords— Axial flow compressor, Computational Fluid Dynamic, AxStream

Aspect

ratio,

I. INTRODUCTION An axial compressor is an important part of any efficient gas turbine. Axial flow compressors are the fluid pumping machinery where the fluid enters and exits axially to the rotor axis. The unique features like high mass flow rate for a small frontal area and high efficiency ratio with higher mass flow rate makes multistage axial flow compressors a perfect choice for gas turbines used in jet engines. The performance and reliability of a gas turbine heavily relies on its axial compressor module. Higher stability margin is paramount for successful operation of an axial compressor and also for its reliability. The efficiency and stability margin of an axial compressor is mainly defined by aero thermal design. Every axial flow compressor has a limit on its mass flow, at both higher and lower side for a specific speed called its operating range, within which it can operate successfully. The maximum mass flow rate is called choke limit. The peak efficiency point is normally referred as the design point. The minimum mass flow rate at which the compressor can run stably is called the stall point. 259

International Journal of Emerging Technology and Advanced Engineering Website: www.ijetae.com (ISSN 2250-2459, ISO 9001:2008 Certified Journal, Volume 3, Issue 8, August 2013) The objective of this paper is CFD analysis of single stage subsonic axial flow compressor at various aspect ratio (AR) configurations such as single stage axial flow compressor having an AR2 through the stage (IGV, Rotor &stator) and then rotor blade AR2 kept constant and change the AR 3 for rotor blade and IGV. Description Of The Compressor : The configuration studied in this work is a single stage axial Flow subsonic compressor (fig. 1). The nominal rotational speed is 14443 rpm for a mass flow rate of 15.47 kg/s. The stage total pressure ratio is 1.2 having stage efficiency 0.90 % and Rotor tip diameter is 485.15 mm

Fig.2 Design space view with filtered solutions

Preliminary design will generate number of designs basing on given data and constraints. The generate design shows the applied design having best efficiency among all these design points as shown in fig.2. The characteristic curve check performance of machine from the generate solutions. It is also indicate that the curve passing through the design point.

Fig. 1 Axial flow compressor

II. NUMERICAL ANALYSIS A numerical analysis has been carried out using AxStream code for varying aspect ratio for single stage axial flow compressor with fixed AR 2 through the stage and by considering the AR 3 for IGV and stator blade while AR 2 for rotor blade to obtained the performance characteristics of the compressor stage.

Fig. 3 Performances curves for compressor design

B. Post design In this mode we are able to change aspect ratio of given compressor and run the streamline calculations. The paper will illustrate the aspect ratio variation for the stage. In this module vary aspect ratio for single stage axial flow compressor with fixed AR 2 through the stage and performed further procedure and get the results then again come to post design mode and by considering the AR 3 for IGV and stator blade while AR 2 for rotor blade to obtained the results. The result with aspect ratio variation will show the various flow parameters i.e. static pressure distribution in compressor, total pressure and absolute pressure distribution, meridonal velocity and Mach no. is shown in fig. 4. The meridonal velocity is maximum in the position of IGV and rotor interaction gap and minimum at the inlet of the IGV of the given compressor.

A. Preliminary Design Preliminary design starts from specification of technical requirement and setting up design task and compressor conceptual layout that includes: • Boundary conditions (inlet pressure, temperature, pressure raise ratio etc). • Conceptual design and sizing layout i.e. quantity of modules (group of stages) inside compressor, number of stages in each group, work coefficient • The designer has to decide what kind of geometrical parameters should be used as design constraints, i.e. specific diameter and its ranges or exact value, and blade heights or angles based on requirements or assumptions. 260

International Journal of Emerging Technology and Advanced Engineering Website: www.ijetae.com (ISSN 2250-2459, ISO 9001:2008 Certified Journal, Volume 3, Issue 8, August 2013)

Fig.4 pressure and velocity distribution

Fig. 6 Camber line curvature

The fig. 4 are indicate that the static pressure is maximum at the stator outlet of the compressor and minimum at the inlet guide vane inlet of compressor. The maximum relative pressure is at the inlet of the rotor and minimum at the inlet of inlet guide vane of compressor, similarly total absolute pressure is at the stator outlet of the compressor and minimum at the inlet guide vane inlet of compressor. The relative Mach number is maximum at the inlet of rotor which is indicate the high velocity at the inlet of rotor and minimum at the inlet of IGV and the stator outlet which is indicate the low velocity at the entrance and exit of the axial flow compressor.

The profile thickness distribution is shown in fig.7 that mode is used to smooth the flow.

Fig. 7 Profile thickness distribution

In the stream wise chart appearance the Mach number distribution and relative pressure distribution is shown in fig. 8 & 9.

Fig.5 Mach number

C. Profiling On Plain Sections The main objective of this mode is to perform profiling on plain profiles section to obtain optimal flow characteristics and pressure distributions. On the next step 3D blade design, stacking and shaping are performed, and complete geometry ready for further analysis. Camber profiling - Camberline profiling mode is appears as default for compressors cascades profiling as equal thickness distribution on both sides of camberline. Camberline curvature distribution can be adjusted as well.

Fig. 8 Mach number distribution

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Fig. 11 stress at LE & TE before smoothening Fig. 9 Pressure distribution

The stress at the LE & TE before smoothening is shown in fig. 11 that is show the stress distribution along the relative blade height. This high stress is reduce by editing profile of the blade and reduce the stress of the blade profile as shown in fig.12.

D. 3D Blade Design In the mode of 3D design procedure, smoothing the geometrical parameters along height and stacking axis redistribution are available. When profiling on individual sections is finished, the user can check smoothness and adjust parameters distribution along the blade on Blade 3D pane. In this mode provide smooth distribution of angles and other geometrical parameters and Provide smooth stress distribution along blade height within margins of admissible stress.

Fig. 12 stress at LE & TE after smoothening

E. Structural and Modal Analysis 3D structural and modal analysis will carried out for performing analysis. Modal analysis will give the frequency at various mode shown in fig. 5.This mode also shows the assembly of the IGV, rotor and stator for detail study of at various modes of frequency. The next mode of study will give the final result of the compressor with various flow parameters. Initially meshing is done by fine meshing with mesh type H to improve the accuracy of results.

Fig. 10 3D blade profile

The stress distribution is shown in fig. 10 that is indicate the area of blade with smooth distribution. The fig. indicate that the stress distribution is minimum at the side of LE and maximum at the side of TE.

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International Journal of Emerging Technology and Advanced Engineering Website: www.ijetae.com (ISSN 2250-2459, ISO 9001:2008 Certified Journal, Volume 3, Issue 8, August 2013) Modal analysis will give the frequency at various mode shown in fig. 13.This mode also shows the assembly of the IGV, rotor and stator for detail study of at various modes of frequency. The next mode of study will give the final result of the compressor with various flow parameters.

The flow diagrams for different parameters are shown in figure. The maximum static pressure and total absolute pressure is developed at the rotor i.e. near to entrance of rotor blade as shown in fig. 15 & 16.

Fig. 16 Total Relative Pressure

The flow diagram also indicate that the maximum pressure is developed in the rotor blade as shown in fig, 16.

Fig. 13 Modal analysis

F. Flow Analysis The CFD flow analysis will give a flow of fluid in compressor. This analysis will show designed flow streamlines. The results are shown in the fig below in flow diagrams Fig. 17 Absolute Mach number

The absolute Mach number is maximum at near the exit of rotor blade in fig. 17. The relative Mach number is maximum in rotor near and middle of rotor blade.

Fig. 14 Static Pressure

Fig. 18 Relative Mach number

The analysis of single stage axial flow compressor for static pressure, total absolute and relative pressure and absolute and relative Mach number are shown in the chart below. Fig. 15 Total Absolute Pressure

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Fig. 23 Relative Mach number

Fig. 19 Static Pressure

III. RESULTS AND DISCUSSION The result shows that the comparison between two single stage axial flow having different same and different AR thought the stage Sr no. Property

Fig. 20 Total Absolute Pressure

Unit

323

222

01 Stat. Pressure at Outlet

kPa

113.4292

113.3725

02 Total Pressure at Outlet

kPa

121.6002

121.5510

03 Power

kW

277.9688

279.4439

1.1194

1.1188

05 Diffusion Factor (by NASA)

0.2881

0.2874

06 Diffusion Factor by De Haller (W2/W1) 07 Total Pressure Rise

0.8229

0.8227

20.2743

20.1260

04 Total-Static Pressure Ratio

-

kPa

The results are based comparison between two single stage axial flow compressor having AR 232 and AR 222 are analyze. The results are shows static pressure at outlet, total pressure at outlet and total pressure rise is maximum in case of AR 232 as compare to AR 222. Power required to run the compressor is minimum in case of AR 323.Further more diffusion factor by NASA and De Haller is more in AR 232 compressor. Through this study, it has been clearly brought out that, the AR 323 has optimum results as compare to AR 222. This paper give a direction to research work. Hence future work will be to do the analysis taking aspect ratios AR 333 and compare with AR 323 and so it may be give different direction to research work.

Fig. 21 Total Relative Pressure

Fig. 22 Absolute Mach number

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International Journal of Emerging Technology and Advanced Engineering Website: www.ijetae.com (ISSN 2250-2459, ISO 9001:2008 Certified Journal, Volume 3, Issue 8, August 2013) J. H. Horlock and G J. Fahmi, 1967 “A Theoretical Investigation of the Effect of Aspect Ratio on Axial Flow Compressor Performance”, London: her majesty’s stationary office. [7] Herrig, L.J., Emery, J.C., and Erwin, J.R., 1955. “Systematic Two Dimensional Cascade Tests of NACA 65 Series Compressor Blades at Low Speed,” NACA R.M. E 55H11. [8] G.T.S. Fahmi, 1971. “ The Performance Of Axial-Flow Compressors Of Differing Blade Aspect Ratio” London : Her Majesty’s Stationery Office. [9] Dieter Bestle, 2011. “Curvature driven two-dimensional multiobjective optimization of compressor blade sections” Aerospace Science and Technology. [10] Lingen Chen, Jun Luo , Fengrui Sun & Chih Wu, 2005. “Optimized efficiency axial-flow compressor” Applied Energy 81. [11] Axial flow compressor design, Turbomachinery Flow Path Conceptual Design Suite, AxSTREAM. [12] HIH Saravanamuttoo, GFC Roger and Hcohen, “Gas Turbine Theory” 5 th Edition, PEARSON Education. [6]

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Boyce, M.P. 2005. Gas Turbine Engineering Handbook, Second Edition, Butterworth- Hienemann 2005. Werner R. Britsch, Walter M. Osborn and Mark R. Laessig, Sep 1979. “Effects of Diffusion Factor, Aspect Ratio, and Solidity on Overall Performance of 14 Compressor Middle Stages” NASA Technical Paper 1523. Hanoca P*, Shobhavathy M T, 2011. “CFD analysis to investigate the effect of axial spacing in a single stage transonic axial flow compressor”, Symposium on Applied Aerodynamics and Design of Aerospace Vehicle (SAROD 2011) November 16-18, Bangalore, India. Ronuld J. Steinke und James E. Crozcse, May 1967. “Analytical studies of aspect ratio and curvature variation for axial flow Compressor inlet Stages under High Loading”. G.T.S. Fahmi, 1967 “The performance of axial flow compressor of different blade aspect ratio” London: her majesty’s stationary office.

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