study of high speed flow behavior through axial …€¦ · increasing the aoa, the amount of...

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International Journal of Mechanical Engineering and Technology (IJMET)

Volume 9, Issue 5, May 2018, pp. 371–382, Article ID: IJMET_09_05_041

Available online at http://www.iaeme.com/ijmet/issues.asp?JType=IJMET&VType=9&IType=5

ISSN Print: 0976-6340 and ISSN Online: 0976-6359

© IAEME Publication Scopus Indexed

STUDY OF HIGH SPEED FLOW BEHAVIOR

THROUGH AXIAL COMPRESSOR CASCADE

BLADES

Vinayaka N, Akshaya C

Department of Aeronautical Engineering, Nitte Meenakshi Institute of Technology,

Yelahanka, Bengaluru, Karnataka, India

Shiv Pratap Singh Yadav

Department of Mechanical Engineering, Nitte Meenakshi Institute of Technology,

Yelahanka, Bengaluru, Karnataka, India

H N Reddappa

Department of Mechanical Engineering, Bangalore Institute of Technology,

V V Puram, Bengaluru, Karnataka, India

ABSTRACT

The present work focuses on understanding the behavior of high speed flow

(supersonic) through the compressor cascade at different Angle of Attacks (AOA).

Flow behavior plays a vital role in compressor overall performance. For the present

work, two main parameters of fluid flow, Velocity and Static Pressure are considered

for performing flow analysis. To increase the pressure raise per stage, blade speed

has to be increased. But in our study, blade speed is kept constant and the effect of

angle of attack was analyzed to determine the raise in static pressure. The

Computational Fluid Dynamics (CFD) approach was employed to perform flow

analysis. When analyzed, it was found that a Normal Shock Wave is formed. Upon

increasing the AOA, the amount of reduction in velocity and raise in static pressure

occurred. Further increase in AOA resulted in normal shock wave to move towards

the leading edge of the compressor blade. It was found at 20 degree AOA the pressure

raise was reduced because of flow separation leading to blade stall. The raise in static

pressure was mainly due to the shock diffusion.

Keywords: Supersonic Flow, Angle of Attack, Flow Analysis, Normal Shock Wave,

Flow Separation.

Cite this Article: Vinayaka N, Akshaya C, Shiv Pratap Singh Yadav and H N

Reddappa, Study of High Speed Flow Behavior through Axial Compressor Cascade

Blades, International Journal of Mechanical Engineering and Technology, 9(5), 2018,

pp. 371–382.

http://www.iaeme.com/IJMET/issues.asp?JType=IJMET&VType=9&IType=5

Study of High Speed Flow Behavior through Axial Compressor Cascade Blades


1. INTRODUCTION

For an aircraft to fly, the most important component are its engines. Engines basically work

on the principle of Newton’s third law. Jet engine work on Brayton cycle. There are two types

of compressors, axial compressor and centrifugal compressor. Here in this study, Axial

Compressor is chosen for the flow analysis. The first stage of the compressor experiences

supersonic flow velocities across their blades due to the higher thrust requirements. In this

study, the behavior of Axial Compressor Cascade blade at higher velocities was analyzed. The

scope was to find out the flow behavior at high speed through the compressor cascade and to

understand how the flow behaves when it passes through compressor cascade at near

supersonic velocities at various AOA. The behavior of the Normal shock wave at different

angle of attacks is of prime importance for this analysis. Hence, we performed this analysis of

near supersonic flow at 295.75 m/s through the designed compressor cascade at different

angle of attacks.

CFD flow analysis was carried out on compressor cascade at low subsonic speed at

various incidence angles by K.M.Panday[1]. The behavior of the fluid parameters at various

angles of attacks at subsonic speeds was analyzed. Accordingly, the maximum raise in static

pressure happens between 4 to 6 degrees. It was observed that the static pressure rises with

increase in angle of attack, while the velocity and dynamic pressure decreases. It was found

out that due to the turbulence, a greater intermixing of the fluid molecules occurs, leading to

increased resistance to the flow.

A research paper by John F, Klapproth[2], throws light on investigation of Mach number

effects on compressor blades. But to maintain the constant thrust and exhaust velocity, the

mass flow rate has to be increased due to reduction in diameter. Due to this, it results a high

intake velocity which may lead to supersonic flows on the first stage of compressor blades.

The use of guide vanes is eliminated by operating at higher Mach number range. Excess

amount of tip loading can also be minimized when operating at higher pressure ratios.

William B. Briggs[3] made an experimental analysis on NACA 65(12)10 blade cascade at

inlet Mach number from 0.12 to 0.89. This analysis was performed to show the behavior of

turning angle, blade wake, pressure distribution and static pressure at different Mach numbers.

It was found that until the strong shock was developed in the cascade, it had an insignificant

effect on turning angle and blade wake. And these two parameters were directly proportional

to the Mach number. Then the turning angle had a sudden decrease in its value as the Mach

number was suddenly increased. As the inlet Mach number increased, the blade pressure

coefficient increased and shifted rearwards. Further increase in Mach number resulted in

stagnation point moving towards the convex side of the blade. Nivin Francis, J. Bruce Ralphin

Rose [4] carried out a computational analysis on swept and leaned transonic compressor

cascade blades. They analyzed the shock wave interaction with the modified cascade blades

and with Rotor 35. They concluded by saying that the improvement in Mach number for

swept and lean rotor than rotor 35, and the boundary layer growth was also minimized. An

experimental analysis was carried out on 2-D low speed compressor cascade by John R,

Erwin, Melvyn savage, and James Emery J [5]. The results showed that, at low inlet angles

and high solidities the A614b blade appeared to offer better high-speed capabilities than either

the A2IB or the Al0 blades. Peng Lin, Cong Wang, Yong Wang [6], worked on rotating stall in

axial compressors induced due to the rotating stall inception processes and proposed a

distortion model to analyze the phenomenon. They concluded by saying that, distortion screen

covered less angles than distorted sectors. H A Schreiber and B. Kusters [7] did a

computational analysis to understand the behavior of interaction of supersonic flow through

the cascade and the boundary layer generated on the blades. A solver which is of two-

dimensional multi block Navier–Stokes equations has been successfully applied.

Vinayaka N, Akshaya C, Shiv Pratap Singh Yadav and H N Reddappa


Investigations showed that, an accurate resolution and results was obtained by the solver,

which reproduced the shock wave across the blade passage. Hengtao Shi, Baojie Liu and

Xianjun Yu [8] carried out a computational analysis on a supersonic cascade and proposed a

shock loss model which is of 1D, to optimize the structure of a shock wave generated in and

across the cascade. According to the conclusion of their paper, the total pressure recovery co-

efficient can be obtained at an optimal shock angle at a fixed transonic Mach number.

Roberto Biollo and Ernesto Benini, [9] has done a study on recent advancement in

transonic compressor blades. The conclusion states that, a significant aerodynamics

improvement has been done due to the understanding of loss mechanisms at supersonic flows

using computational methods and optical measurement techniques. John A, Slrs Jisky,

William Roberts B. [10] has done an experimental analysis on smoke flow visualization for

determining the shock losses across the compressor cascade. Using this technique the flow

behavior through compressor cascade can be visualized easily. Results also states that this

technique can also be applied to determine the wave drag of aircraft or the cascade blades.

A numerical analysis was carried out on the effects of low-aspect ratio and aerodynamic

sweep on axial transonic compressor design based on RANS Equations by S. L. Puterbaugh,

C Hah and Wadia A R[11]. From analysis it was found that, the upswept rotor had the higher

stall margin than the aft-swept rotor. The results also indicated that the rotors efficiency was

not improved by the design of aft-swept blades. There was a study carried out at subsonic

axial flow on unsteady supersonic cascade by M. Verdon, Joseph and James E [12]. From

results, the flutter boundary of supersonic fans was found out. Copenhaver W W and Wadia A

R [13] here, in this paper the author has analyzed the effect of cascade area ratio on the

transonic compressor performance. The results states that, reducing the trailing edge effective

camber, expressed as throat-to-exit area ratio, results in an improvement in peak efficiency

level without significantly lowering the stall line. Aldo Rona, Renato Paciorri and Marco

Geron [14], in this experiment the author have designed and tested the axial compressor

cascade blades in wind tunnel. Pressure measurements at the cascade outlet and synchronous

spark schlieren visualization of the test section, with and without the optimized slotted tail-

board, have confirmed the gain in pitch-wise periodicity predicted by the numerical model.

2. OBJECTIVES

1. To analyze the static pressure variation for high speed flow at various angle of attacks.

2. To study the velocity variation at high speed flow at various angle of attacks.

3. To obtain the value of angle of attack at which flow separation starts.

4. To study the shock wave behavior at various angle of attacks.

3. METHODOLOGY

The best suited airfoil series for the axial compressor blades are NACA 65 series [15], in

which NACA 65210 airfoil is chosen for the flow analysis. This specific blade was chosen

because the airfoil has its maximum thickness at 10% of chord and the lift co-efficient is 0.2.

The obtained co-ordinates from airfoil tools website were imported to the ANSYS design

modeler and the cascade was modeled. Then the control volume is generated across the

cascade through which the fluid flows. The control volume is a rectangular 2-D duct inclined

at an angle to match the 0 angle of attack for the cascade blades.



Figure 1 Geometric modeling of cascade blade Figure 2 Mesh across the model

Before analyzing the obtained geometry in ANSYS FLUENT, the model has to be

meshed. An Unstructured mesh was generated in the geometry using ICEM CFD. Triangular

elements were used to mesh the geometry. Resulting in 627495 elements (cells) and 313737

nodes. Boundary conditions chosen are:-

1. Inlet Velocity Inlet 295.75m/s.

2. Outlet Pressure outlet 0 bars.

3. Walls Wall.

4. Rotor 1,2,3 Walls.

5. Operating Pressure 101.3 KPa.

Analysis of the meshed geometry was done using ANSYS FLUENT commercial Package.

4. RESULTS AND DISCUSSIONS

Computational Fluid Dynamics (CFD) analysis was carried out at an incremental value of

AOA by 5 degree, continuing from 0 degree to 25 degree. The contours of static pressure and

velocity for the above range are plotted from figure 3 to figure 26. The obtained results

showed that, as the high speed flow reaches the cascade leading edge, the flow accelerates

resulting in increase in velocity. Due to this increase in velocity which is more than the local

sonic Mach number, there is a formation of shock wave. But the diffusion of the flow takes

place. This way of diffusing the airflow is known as shock diffusion. Once the flow passes

through the shock wave, the velocity get decreases and the static pressure increases. As angle

of attack was increased, the increase in static pressure was enhanced and the shock wave

started moving towards the leading edge of the cascade blade. At an angle of 20 degrees, the

flow got separated resulting in pressure loss.



4.1. Test Case 1: AOA, 𝛼= 0°

Figure 3 Contour of Static pressure Figure 4 Chord Length vs. Static Pressure

Figure 3 shows the variation of static pressure at 0 degree AOA. The above figure 4 is the

scatter plot of variation of static pressure along the chord length. The sudden raise in static

pressure as shown in graph is the indication of presence of shock wave. The position of shock

wave is at 25mm from trailing edge. The graph shows some negative values which indicate

that the pressure is below atmospheric pressure and positive values represent above

atmospheric pressure.

Figure 5 Contour of Velocity Vectors Figure 6 Chord Length vs. Velocity

The above figure 5 shows the velocity vectors variation across the compressor cascade at

0 degree AOA. Figure 6, the scatter plot shows the variation of velocity through the

compressor cascade blades graphically. The exit velocity as indicated by graph is 232.705m/s.

Hence, there is a decrement of 63.05m/s velocity. The decrement of velocity is basically due

to the flow passing through the normal shock wave. The velocity increases as it reaches the

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leading edge of the blade from inlet. Then as it passes through the shock wave, the velocity

has decreased.



Figure 7 shows the variation of static pressure at 5 degree AOA. The above figure 8

shows the scatter plot of variation of static pressure at 5° angle of attack graphically. The

amount of pressure rise, above atmospheric, has risen from 6.6532 KPa (at 0° AOA) to

10.4362 KPa. The position of shock wave is at 31 mm from trailing edge. The shock wave has

moved 6mm towards the leading edge from trailing edge. Even at this AOA, the pressure raise

is due to shock diffusion.


The above figure 9 shows the velocity vectors variation across the compressor cascade at

5 degree AOA. Figure 10 shows the scatter plot of variation of velocity. The velocity has

reduced from 232.705m/s (at 0° AOA) to 213.190m/s. The decrement of velocity is basically

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due to the flow passing through the normal shock wave. Hence, the total velocity decrement

from leading edge to the trailing edge is 82.56 m/s.

4.3. Test Case 3: AOA, = 10°


The contours of variation of static pressure is shown in figure 11 at 100 AOA. The above

figure 12 shows the scatter plot of variation of static pressure at 10° angle of attack. As the

above graph indicates that the flows static pressure decreases first due to the increment in

velocity at the leading edge and then it increases as it approaches the shock wave. The amount

of pressure raise above atmospheric pressure at this angle of attack is 13.647 KPa.


The above figure 13 shows the velocity vector variation across the compressor cascade at

10 degree AOA. Figure 14 is the scatter plot of variation of velocity. The velocity has reduced

from 213.190 m/s (at 10° AOA) to 193.478m/s. The decrement of velocity is basically due to

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the flow passing through the normal shock wave. Hence, the total velocity decrement from

leading edge to the trailing edge is 102.272 m/s.

4.4. Test Case 4: AOA, α= 15°


Figure 15 shows the variation of static pressure at 15 degree AOA. The above graph in

figure 16 is indicating that, as the angle of attack was further increased, the shock wave has

moved further from trailing edge. Figure 16 shows the variation of static pressure graphically.

The position of shock wave was 40mm from trailing edge in the previous test case. But now

the shock wave is at 53mm far from trailing edge. The shock wave has moved 12mm from its

previous position. The static pressure at the exit of the cascade is 19.740.2 KPa above

atmospheric pressure.



15 degree AOA. The above graph shows that the velocity has reduced compared to the

previous AOA. The velocity at exit of the cascade has decreased from 193.478 m/s to 173.757

m/s. At supersonic flows, the reduction of velocity is due to the normal shock wave.

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Figure 19 shows the variation of static pressure at 20 degree AOA. The above graph in

figure 20 shows that there is an increment in static pressure. Figure 20 shows the variations of

static pressure at different positions of cascade blades. The position of the shock wave is at 69

mm from trailing edge. At this AOA, the amount of increment in static pressure raise has

decreased when compared to previous simulations. There is a raise in static pressure, but

lesser than pervious angle of attacks. Hence, the cascade’s exit static pressure is 19.740 KPa.

The reason behind this is that there is a small amount of flow separation which has started on

the suction side top surface of blades.



20 degree AOA. In figure 22, the graph is indicating the velocity at the exit of the cascade is

154.754m/s. At 15° AOA, the velocity was 173.757 m/s, but now it has reduced. In the above

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vector plot the graph is clearly indicating a very small amount of flow separation on the upper

surface of the blade 2 and 3.



Figure 23 shows the variation of static pressure at 25 degree AOA. In the above figure 24,

the graph shows that the static pressure has increased after passing through the shock wave,

but it is also indicating that there is a lot of variation of static pressure downstream of the

shock wave. This fluctuation is basically caused due to the flow separation. We can clearly

see from the contour and vector plots that the flow has separated at the leading edge. Because

of this, there is a circulation induced at the separated region resulting in pressure loss.



25 degree AOA. In figure 26, we can see that there is a variation of velocity after the passing

of air through the shock wave. The velocity at the exit of the cascade is 139.433m/s. This is

lesser when compared to the previous test cases. But the difference of amount of decrement

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when compared to the past two test cases, has reduced. This reduction is basically due to the

flow separation. In the above vector plot, we can clearly see the flow separation.

Table 1 Position of shock wave at various angles of attacks

Test Cases Angle of Attack Position of shock wave

from trailing edge (in mm)

Test Case 1 0° 25

Test Case 2 50

31

Test Case 3 100 40

Test Case 4 150

53

Test Case 5 200

69

Test Case 6 250

81

Table 1 shows the position of shock wave from trailing edge in mm, obtained from flow

simulation results at various AOA in range from 00 to 25

0.

5. CONCLUSION

Analysis of compressor cascade blades has been done at higher velocities which were near

supersonic. It was found that at 295.75m/s velocity, the raise in static pressure is due to

normal shock wave or shock diffusion, which was produced inside the compressor cascade.

As the angle of attack was increased, the raise in static pressure happened due to the

strengthening of normal shock wave. The velocity was reduced when the flow passed through

the normal shock wave. It was found that, at 20 degree angle of attack, the flow started

separating, but it had minor losses. At 25 degree angle of attack the flow had separated

drastically leading to the loss of static pressure raise. As the angle of attack was increased, the

shock wave started moving from trailing edge to the leading edge. This phenomenon had

happened due to the increased acceleration of flow on the leading edge of cascade blade as

angle of attack is varied, resulting in less pressure at upstream compared to the pressure at the

trailing edge. Hence, from the present research work, it has been concluded that to avoid flow

separation, the best suitable AOA range is 15 degree to 19.5 degree.

REFERENCES

[1] Pandey, K. M., Chakraborty, S., and Deb, K. CFD Analysis of Flow through Compressor

Cascade. International Journal of Soft Computing and Engineering, 2(1), 2012, pp. 362-

371.

[2] John Klapproth, F,. General considerations of Mach number effects on compressor-blade

design. National Advisory Committee for Aeronautics, Research Memorandum, NACA

RM E53L23A, 1954, pp. 1-24.

[3] William Briggs, B,. Effect of Mach number on the flow and application of compressibility

corrections in a two dimensional subsonic-transonic compressor cascade having varied

porous-wall suction at the blade tips. National Advisory Committee for Aeronautics,

2648, 1952, pp. 1-24.

[4] Nivin Francis, Bruce Ralphin Rose, J. CFD Analysis of Swept and Leaned Transonic

Compressor Rotor. International Journal of Engineering Trends and Technology, 4(6),

2013, pp. 2524-2528.

[5] John, R, Erwin, Melvyn savage, and James Emery, J. Two-dimensional compressor low-

speed cascade investigation of NACA blade sections having a systematic variation in

mean-line loading. National Aeronautics and Space Administration, NASA Technical

Note 3817, 1956, pp. 1-129.



[6] Peng LIN, Cong WANG, Yong WANG. A high-order model of rotating stall in axial

compressors with inlet distortion.Chinese Journal of Aeronautics,30(3),2017, pp.898-906.

[7] Kusters, B. and Schreiber, H. A. Compressor Cascade Flow with Strong Shock-

Wave/Boundary-Layer Interaction. American Institute of Aeronautics and Astronautics

Journal, 36(11), 1998, pp. 2072-2078.

[8] Baojie Liu, Hengtao Shi and Xianjun Yu. A new method for rapid shock loss evaluation

and reduction for the optimization design of a supersonic compressor cascade.

Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace

Engineering, 0(0), 2017, pp. 1-19.

[9] Roberto Biollo, Ernesto Benini. Recent advances in transonic axial compressor

aerodynamics. Progress in Aerospace Sciences, 56, 2013, pp. 1-8.

[10] James Crouse, E., William Roberts, B., and John Slrs Jisky, A., High Speed Smoke Flow

Visualization for the determination of Cascade Shock Losses. American Institute of

Aeronautics and Astronautics Journal, 79-0042, 1979, pp. 1-10.

[11] Hah, C., Puterbaugh. S. L., and Wadia, A. R. Control of shock structure and secondary

flow field inside transonic compressor rotors through aerodynamic sweep. The American

Society of Mechanical Engineers, 98-GT-561 345, 1998, pp. 1-15.

[12] Joseph, M. Verdon., and James, E. McCune. The study of unsteady supersonic cascade in

subsonic axial flow. Aeronautics and Astronautics Journal,75-22, 1975,pp.1-13.

[13] Wadia, A. R., and Copenhaver, W. W. An Investigation of the Effect of Cascade Area

Ratios on Transonic Compressor Performance. The American Society of Mechanical

Engineers, 118. 1996, pp. 760-770.

[14] Aldo Rona, Renato Paciorri and Marco Geron. Design and Testing of a Transonic Linear

cascade Tunnel with Optimized Slotted Walls. The American Society of Mechanical

Engineers, 128, pp. 23-34.

[15] Wright, L. C. Blade Selection for a Modern Axial-Flow Compressor. AiResearch

Manufacturing Company, pp. 603-626.

study of high speed flow behavior through axial …€¦ · increasing the aoa, the amount of...

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