1

Numerical Studies of Two Approaches for Surge Control in Axial

Compressors

Vishwas Iyengar, Jimmy Tai, Saeid Niazi and Lakshmi N. Sankar

School of Aerospace Engineering

Georgia Institute of Technology, Atlanta, GA 30332-0150

Summary

Axial compression systems have a limited operation range at low-mass flow rates,

sometimes due to occurrence of rotating stall or modified surge. Active bleed control methods

have been developed to eliminate the stall/surge phenomena. While being quite effective, the

bleed control methods have some limitations. ‘Self-Recirculation’ treatment is one such concept

that has been proposed by researchers. The aim of this study is to numerically model these two

concepts on the Rotor 67 configuration and assess the relative advantages and drawbacks.

Preliminary results are presented for the performance map of the Rotor 67 configuration.

Results are also presented for surge control with the aid of active bleed control. Calculations are

in progress for the passive self-recirculation treatment concept.

Introduction

High-speed, high pressure ratio axial compression systems are widely used in many

aerodynamic applications. However they have a limited range of stable operations at low-mass

flow rates because of fluid dynamic instabilities. These instabilities cause deteriorations in

system operations due to the compressor reaching a state of rotating stall or surge. Rotating stall

is essentially a 2-D unsteady local phenomenon where the flow is no longer uniform, where as

surge is global 1-D instability that can affect the whole system. Modified surge, a combination of

rotating stall and surge can also occur in some instances. Throughout the aero-propulsion history,

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much work has been done to broaden the compressor operating range by developing control

strategies to alleviate the rotating stall and surge.

In the last two decades considerable progress has made in understanding and modeling of

axial compressor stall and surge. With the growth in Computational Fluid Dynamics (CFD) as an

effective analysis technique, in CFD we have a resourceful tool which provides us a way to

simulate and understand the complex flow phenomena in turbomachinery. 3-D codes that are

capable of analyzing unsteady turbomachinery flow with single and multiple blade passages have

been developed by several researchers1-5. But these have been limited to modeling steady flow

phenomena in axial and centrifugal compressors, near or at design conditions. Active control

methodologies, which can alleviate the stall/surge phenomena, such as bleeding and air injection

can be studied extensively by CFD as long as an appropriate numerical model is available.

It is well known that the adverse flow conditions in tip clearance can cause the surge/stall

phenomena and reduce the pressure rise, flow range and efficiency of the turbomachinery.

Blockage development in a transonic axial compressor rotor was studied by Suder6 (1997) and

the impact of shock on blockage was investigated. The flow through the tip clearance region for

Rotor 37 was computed and compared to aerodynamic probe and laser anemometer data by

Chima7 (1996). The interaction of the tip vortex; the passage shock and the casing boundary

layer were described using the computation. End wall and casing treatments for Rotor 67 may be

found in Ref 8.

To model unsteady flow in air breathing vehicles the present authors have developed a 3-

D unsteady flow solver, GT-TURBO3D. Ref. 9 and 10 report results from the application of this

solver to a NASA low-speed centrifugal compressor. In Ref 11 (previously published by current

author) numerical simulation of rotating stall and surge alleviation in carried out for axial

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compressor (Rotor 67). It was found that the use of bleed valves located on the diffuser walls can

eliminate the adverse phenomena of stall/surge and stable operation can be restored. Open-loop

and closed-loop case were examined, calculation show that both types of bleeding eliminate both

rotating stall and modified surge.

Scope of Present Work

Active Control approaches, while effective, have two drawbacks. Firstly, since the air is

being bleed, we have a loss of high pressure air which hinders the total pressure rise from the

compressor. And secondly there is still the need for an active controller and actuator. Passive

concepts which involve no (or very few) moving parts will be preferable. One such concept

called the “self-recirculation” was recently proposed by Hathaway13 (2002) and has already been

applied by him to the Rotor 67 configuration. The purpose of the present work is to

independently apply this concept to the Rotor 67 and Rotor 37 configurations, and to compare

this concept to the active bleeding that has already been studied for Rotor 67.

Configurations Being Studied

Figure 1: NASA Rotor 37 Configuration Figure 2: NASA Rotor 67 Configuration13

51.4 cm

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Two standard configurations (Figure 1 and 2), named NASA Rotor 37 and NASA Rotor 67 are

being used in this study. These configurations were chosen because there has been extensive

work (computational and experimental) done on them. A limited amount of unsteady data is also

available. An algebraically generated H-H-O grid is being used in this work.

The Rotor 67 grid has 125 cells in the streamwise direction, 63 and 41 cells in the radial

and circumferential directions, respectively. A hyperbolic tangent node point distribution near

the surfaces was used to provide adequate resolution of the boundary layers. The clearance gap

was spanned by six cells in the radial direction.

The grid for NASA Rotor 37 was generated with the same methodology that was used for

NASA Rotor 67. It has 60x36x21 grid points in each flow passage in the streamwise, spanwise

and pitchwise directions, respectively. Five points were placed in the clearance gap. As in the

case of NASA Rotor 67, the tip clearance gap was modeled and periodic boundary conditions

were applied.

Sample body fitted grid for these configurations are shown in figures 3 and 4.

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Figure 3: Rotor 37 Grid

Figure 4: Rotor 67 Grid

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Results In Hand

At this writing, the following results are in hand:

• Performance maps for Rotor 37 and Rotor 67, with comparisons with experiments

• Active Flow Control (Bleeding) for Rotor 67

These results are briefly discussed below. The full paper will contain additional details of the

simulation.

Rotor 37 Performance Map

Bright et al.15 at NASA Glenn Research Center have obtained surge and stall data for this

compressor. The experimental data are for the Stage 37, which includes Stator 37 and Rotor 37.

The flow solver was applied to the NASA Rotor 37 at 70% design speed, which is about 12000

RPM and at 100% design speed which is 17188.7 RPM16. From the results, the compressor

performance map was extracted and is shown in Figure 5 and 6. Clearly from the plot we see

good agreement between the computed and measured data for both operating conditions.

Computed and Measured characterstic performance map at 70% design speed (Rotor 37)

1

1.2

1.4

1.6

10 12 14 16 18

Mass Flow Rate (kg/sec)

Tota

l Pre

ssur

e Ra

tio

Computed Experimental

Figure 5: Computed and measured characteristic performance map at 70% design speed (Rotor 37).

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Measured characteristic performance map at Design Speed (Rotor 37)

1

1.2

1.4

1.6

1.8

2

2.2

2.4

19 19.5 20 20.5 21

Mass Flow Rate (kg/sec)

Tota

l Pre

ssur

e Ra

tio

Experimental Computed

Figure 6: Computed and measured characteristic performance map at Design Speed (Rotor 37).

Rotor 67 Performance Map

The design rotational speed is 16043 RPM, and the tip leading edge speed is 429 m/sec

with a tip relative Mach number 1.38. The flow solver was applied to NASA Rotor 67, the

calculated total pressure ratio at full speed are plotted verses normalized mass flow and

compared to experimental data in Figure 7. These are previously published results and can be

found in Ref 12, and provide a starting point for the present study.

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Figure 7: Comparison of measured and computed characteristic performance map (Rotor 67)12.

Three operating points A, B and C indicated on Figure 7 were studied in detail. Point ‘A’

corresponds to a stable operating point for the compressor and is related to the peak efficiency

condition. It was found that decreasing the mass flow through the compressor from point ‘A’ to

point ‘B’ triggers the onset of the rotating stall, further decreasing the mass flow from point ‘B’

to point ‘C’ causes the compressor to reach a instable state of modified surge. The figure below

shows the variation of mass flow rate with time at the operating point C. It is clear that the

compressor is experiencing extreme variations in mass flow rate, indicative of surge. Figure 8

shows the variation of mass flow rate with the fluctuation in pressure at point C.

.

.

Chokedm

m

A

B C

1.3 1.35 1.4

1.45 1.5

1.55 1.6

1.65 1.7

1.75 1.8

0.82 0.84 0.86 0.88 0.9 0.92 0.94 0.96 0.98 1

Tot

al P

ress

ure

Rat

io

CFD Experiment

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Figure 8 Variation of mass flow rate and Pressure fluctuations at operating point C (Rotor 67).

Active Flow Control of Rotor 67 Configuration: A closed loop control for enhancing the

performance of the rotor 67 configuration at the mass flow rate corresponding to operation point

C was subsequently implemented. Computational probes were used to measure the pressure

upstream of the compressor face at each time step. If the pressure values at the probes were out

of the permitted fluctuation limits, the bleed valve removed some air from the compressor, as

specified by

pAKm bbb ∆=.

Where, Kb is the bleed valve constant and is related to the valve geometry, Ab is the bleed valve

and ∞−=∆ ppp b , bp and ∞p are the pressure at the bleed location and ambient pressure,

respectively. Figure 9 shows the upper and lower permitted limit of the inlet pressure

fluctuations.

-50

-30

-10

10

30

50

-40 -30 -20 -10 0 10 20 30 40

% Pressure Fluctuations

% Mass Flow Rate Fluctuations

Without control, operating point C

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Figure 9: Upper and lower limit of pressures that trigger closed-loop control (Rotor 67).

Figure 10: Variation of mass flow rate and Pressure fluctuations under closed-loop control (Rotor 67).

As seen in figure 10, the close loop control discussed earlier did a good job of reducing

the fluctuations in mass and pressure. The amplitude of oscillations was reduced by nearly a

factor of 2. Thus, the simple closed loop control is indeed an effective way of alleviating surge. It

does have the disadvantage of removing useful compressed air from the engine, and requires

sensors and bleed valve actuators. The second concept under study is aimed at removing these

shortcomings.

0 5 10 15 20 250.45

0.5

0.55

0.6

0.65

0.7

0.75

0.8

0.85

0.9

0.95

Rotor Revolution, Ωt/2π

refPP

Pressure lower limit

Pressure upper limit

-50

-30

-10

10

30

50

-30 -20 -10 0 10 20

% Mass Flow Rate Fluctuations

% Pressure Fluctuations

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Results in Progress

Calculations are currently in progress for an independent evaluation of the self-

recirculation technique for the Rotor 67 configuration. This concept is schematically shown in

figure 11.

Figure 11: Self-Recirculation Casing Treatment Model

Although it is a passive concept, it may be modified by changing the recirculation path

(or by opening/closing valves located along the passageway). The full paper will contain these

results, along with a comparison of this concept to the active control concept referred to earlier.

The following additional information will be provided:

1. Complete mathematical and numerical formulation.

2. A detailed discussion of the self-recirculating casing model.

3. Detailed studies of the phenomena that lead to stall and surge in Rotor 67.

4. Results of the application of the casing model on Rotor 67.

Should time permit, results for the Rotor 37 configuration will also be presented.

Hub

Case

Injection Bleed

Rotor

In Flow

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References

1. Chima, R. V., and Yokota, J. W., “Numerical Analysis of Three-Dimensional Viscous

Internal Flow,” AIAA Journal, Vol. 28, No. 5, 1990, pp. 798-806.

2. Hall, E. J., “Aerodynamic Modeling of Multistage Compressor Flow Fields - Part 1:

Analysis of Rotor/Stator/Rotor Aerodynamic Interaction,” ASME paper 97-GT-344, 1997.

3. Dawes, W. N., “A Numerical Study of the 3D Flowfield in a Transonic Compressor Rotor

With a Modeling of the Tip Clearance Flow,” AGARD Conference Proceedings n 401,

Neuilly sur Seine, Fr.

4. Hah, C., and Wennerstrom, A. J., “Three-Dimensional Flowfields Inside a Transonic

Compressor With Swept Blades,” Journal of Turbomachinery, Vol. 113, 1991.

5. Hathaway, M. D., and Wood, J. R., “Application of a Multi-Block CFD Code to Investigate

the Impact of Geometry Modeling on Centrifugal Compressor Flow Field Predictions,”

Transactions of the ASME, Vol. 19, Oct 1997, pp. 820-830.

6. Suder, K.L., “Blockage Development in a Transonic, Axial Compressor Rotor,” Turbo-Expo

’97, Orlando, Florida.

7. Chima, R.V., “Calculation of Tip clearance effects in a Transonic Compressor Rotor,”

NASA TM 107216, May 1996

8. Crook, A. J., Greitzer, E. M., Tan, C. S., and Adamczyk J. J., “Numerical Simulation of

Compressor Endwall and Casing Treatment Flow Phenomena,” Journal of Turbomachinery,

Vol. 115, July 1993, pp. 501-511.

9. Niazi, S., Stein, A., and Sankar, L. N., “Development and Application of a CFD Solver to the

Simulation of Centrifugal Compressors,” AIAA paper 98-0934, 1998.

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10. Stein, A., Niazi, S., and Sankar, L.N., “Computational Analysis of Stall and Separation

Control in Centrifugal Compressors,” AIAA Paper 98-3296, July 1998.

11. Niazi, S., “Numerical Simulation of Rotating Stall and Surge alleviation in axial

compressors” PhD thesis, July 2000.

12. Niazi, S., Stein, A., and Sankar, L.N., “Computational Analysis of Stall Control Using Bleed

Valve in a High-Speed Compressor” AIAA Paper 2000-3507.

13. Hathaway, M.D., “Self-Recirculating Casing Treatment Concept for Enhanced Compressor

Performance” Turbo-Expo 2002, Amsterdam, Holland.

14. Strazisar, A. J., Wood, J. R., Hathaway, M. D., Suder, K. L., "Laser Anemometer

Measurements in a Transonic Axial Flow Fan Rotor," NASA Tp-2879, Nov. 1989.

15. Bright, B. M., Qammar, H. K., and Hartley, T. T., “Dimension Determination of Precursive

Stall Events in a Single Stage High Speed Compressor,” NASA TM 107268, 1996.

16. Reid, L. and Moore, R. D., “Design and Overall Performance of Four Highly Loaded, High-

Speed Inlet Stages for an Advanced High-Pressure-Ratio Core Compressor,” NASA Paper

TM-1337, 1978.