School of Aerospace Engineering Alex Stein, Saeid Niazi, and Lakshmi N. Sankar School of Aerospace Engineering Georgia Institute of Technology lsankar/MURI

School of Aerospace Engineering

Alex Stein, Saeid Niazi, and Lakshmi N. Sankar

School of Aerospace EngineeringGeorgia Institute of Technology

http://www.ae.gatech.edu/~lsankar/MURI

Supported by the U.S. Army Research Office Under the Multidisciplinary University Research Initiative (MURI) on Intelligent Turbine Engines

Numerical Studies of Stall and Surge Numerical Studies of Stall and Surge Alleviation in CompressorsAlleviation in Compressors

School of Aerospace EngineeringOverviewOverview

Objectives and Motivation Rotating Stall and Surge Flow Solver and Boundary Conditions DLR High-Speed Centrifugal Compressor

• Unsteady Surge Simulations• Surge Control Using Air-Injection

NASA Axial Rotor 67 Results• Peak Efficiency Conditions• Off-design Conditions• Bleed Valve Control

Conclusions


Objectives and MotivationObjectives and Motivation

• Develop a numerical scheme to model and understand compressor stall and surge.

• Explore active and passive control strategies (Bleed Valve, Air-Injection) to extend useful operating range of compressors.

Lines of ConstantRotational Speed

Lines of ConstantEfficiency

Ch

oke

L

imit

Su

rge

Lim

it

Flow Rate

To

tal P

ress

ure

Ris

e

Desired Extension of Operating Range


Motivation and ObjectivesMotivation and Objectives

Compressor instabilities can cause fatigue and damage to entire engine

School of Aerospace EngineeringRotating StallRotating Stall

• Rotating stall is a 2-D unsteady local phenomenon

• Types of rotating stall:

•Part-span•Full-span

1

2

1

2

1

2

Blade 1 sees a high

Blade 1 stalls. Blade 1 recovers.Blase 2 stalls.

t=0 t= 0+ t=0++

School of Aerospace EngineeringSurgeSurge

Mild Surge Deep Surge

Time

Flow Rate

Period of Deep Surge Cycle

Flow Reversal

Limit CycleOscillations

Pressure Rise

Flow Rate

MeanOperating Point Peak

PerformancePressure Rise

Flow Rate

Time

Flow Rate

Period ofMild Surge Cycle


• Diffuser bleed valves•Pinsley, Greitzer, Epstein (MIT)•Prasad, Neumeier, Haddad (GT)

• Movable plenum walls•Gysling, Greitzer, Epstein (MIT)

• Guide vanes•Dussourd (Ingersoll-Rand Research Inc.)

• Air-injection•Murray (CalTech)•Fleeter, Lawless (Purdue)•Weigl, Paduano, Bright (MIT & NASA Lewis)

How to Control InstabilitiesHow to Control Instabilities

Bleed Valves

Movable Plenum Walls

Guide Vanes

Air-Injection


GTTURBO3D Flow SolverGTTURBO3D Flow Solver• Reynolds averaged Navier-Stokes equations in finite volume

representation.

• A Four Point Stencil is used to compute the inviscid flux terms at the cell faces according to Roe’s formulation (Third-order accurate in space, first- or second-order accurate in time)

• The viscous fluxes are computed to second order spatial accuracy.

• Turbulence is modeled by one-equation Spalart-Allmaras model

• Code can handle multiple computational blocks and inlet distortions

School of Aerospace EngineeringBoundary Conditions (GTTURBO3D)Boundary Conditions (GTTURBO3D)

Outflow boundary(coupling with plenum)

Periodic Boundaryat compressor inlet

Solid Wall Boundaryat compressor casing

Periodic Boundaryat diffuser

Solid Wall Boundaryat impeller blades

Periodic Boundaryat clearance gap

Solid Wall Boundaryat compressor hub

Inflow Boundary

Zonal Boundary


Outflow BC (GTTURBO3D)Outflow BC (GTTURBO3D)

Plenum Chamber•u(x,y,z) = 0 •pp(x,y,z) = const.•isentropic

ap, Vp

mc

.

mt

.

CFD Outflow Boundary

)mm(V

a

dt

dptc

p

2pp

Conservation of mass:


DLR High-Speed Centrifugal CompressorDLR High-Speed Centrifugal Compressor

40cm

•Designed and tested by DLR (Germany)

•High pressure ratio•AGARD test case


DLR High-Speed Centrifugal CompressorDLR High-Speed Centrifugal Compressor

•24 main blades•30 backsweep•CFD-grid 141 x 49 x 33 (230,000 grid-points)

Design Conditions:•22,360 RPM•Mass flow = 4.0 kg/s•Total pressure ratio = 4.7•Adiab. efficiency = 83%•Exit tip speed = 468 m/s•Inlet Mrel = 0.92


DLR-High-Speed-Results (Design Conditions)DLR-High-Speed-Results (Design Conditions)

Static Pressure Along ShroudStatic Pressure Along Shroud

Excellent agreement between CFD and experiment

0

0.5

1

1.5

2

2.5

3

0 0.2 0.4 0.6 0.8 1

Meridional Chord, S/Smax

Lo

ca

l S

tati

c P

res

su

re,

p/p st

d Experiment

CFD


DLR-High-Speed-Results (Off-Design Conditions)DLR-High-Speed-Results (Off-Design Conditions) Performance Characteristic MapPerformance Characteristic Map

Unsteady fluctuations are denoted by size of circles

Fluctuations at 3.1 kg/sec are 30 times larger than at 4.6 kg/sec

3

3.5

4

4.5

5

5.5

2 2.5 3 3.5 4 4.5 5

Corrected Mass Flow (kg/s)

To

tal P

ress

ure

Rat

io

Experiment

CFD

Design

Surge Choke


DLR-High-Speed-Results (Surge Conditions)DLR-High-Speed-Results (Surge Conditions)

Mild surge cycles develop

Surge amplitude grows to 60% of mean flow rate

Surge frequency = 94 Hz (1/100 of blade passing frequency)

t/2


DLR-High-Speed-Results (Air-Injection-Setup)DLR-High-Speed-Results (Air-Injection-Setup)

Injection angle, = 5º3 to 6% injected mass flow rate

0.04RInlet

Casing

5°

Rotation Axis

Impeller

RInlet

School of Aerospace EngineeringDLR-High-Speed-Results (Air-Injection)DLR-High-Speed-Results (Air-Injection)

Different yaw angles, 3% injected mass flow rateDifferent yaw angles, 3% injected mass flow rate

Yaw angle directly affects the unsteady leading edge vortex shedding


DLR-High-Speed-Results (Air-Injection)DLR-High-Speed-Results (Air-Injection) Different yaw angles, 3% injected mass flow rateDifferent yaw angles, 3% injected mass flow rate

Optimum:Surge amplitude/main flow = 8 %Injected flow/main flow = 3.2 %Yaw angle = 7.5 degrees


DLR-High-Speed-Results (Air-Injection)DLR-High-Speed-Results (Air-Injection) Yaw angle vs. angle of attack, 3% injected mass flow rateYaw angle vs. angle of attack, 3% injected mass flow rate

-40

-30

-20

-10

0

10

20

-20 0 20 40 60

Yaw Angle, Degrees

Lo

ca

l A

ng

le o

f A

tta

ck

, D

eg

ree

s

Injection yaw angle directly affects leading edge angle of attack

=> maximum control for designer


DLR-High-Speed-Results (Air-Injection)DLR-High-Speed-Results (Air-Injection) Angle of attack is directly altered by injectionAngle of attack is directly altered by injection

-20

0

20

40

60

80

100

120

0% 20% 40% 60% 80% 100%Time (in percentage of Tsurge)

An

gle

of

Att

ack,

Deg

rees No injection

3.2% Injection,7.5 degr. Yaw

Optimum injection yaw angle of 7.5 degrees yields best result

School of Aerospace Engineering Axial Compressor (NASA Rotor 67)Axial Compressor (NASA Rotor 67)

• 22 Full Blades

• Inlet Tip Diameter 0.514 m

• Exit Tip Diameter 0.485 m

• Tip Clearance 0.61 mm• Design Conditions:

– Mass Flow Rate 33.25 kg/sec

– Rotational Speed 16043 RPM (267.4 Hz)

– Rotor Tip Speed 429 m/sec

– Inlet Tip Relative Mach Number 1.38

– Total Pressure Ratio 1.63

– Adiabatic Efficiency 0.93

Hub

4 Blocks73X32X21Total of 196,224 cells

School of Aerospace Engineering Literature Survey of NASA Rotor 67Literature Survey of NASA Rotor 67

• Computation of the stable part of the design speed operating line:

• NASA Glenn Research Center (Chima, Wood, Adamczyk, Reid, and Hah)• MIT (Greitzer, and Tan)• U.S. Army Propulsion Laboratory (Pierzga) • Alison Gas Turbine Division (Crook)• University of Florence, Italy (Arnone )• Honda R&D Co., Japan (Arima)

• Effects of tip clearance gap: • NASA Glenn Research Center (Chima and Adamczyk)

• MIT (Greitzer)

• Shock boundary layer interaction and wake development: • NASA Glenn Research Center (Hah and Reid).

• End-wall and casing treatment: • NASA Glenn Research Center (Adamczyk)

• MIT (Greitzer)

School of Aerospace EngineeringRelative Mach Contours at Mid-SpanRelative Mach Contours at Mid-Span

(Peak Efficiency)(Peak Efficiency)

Spatially uniform flow at design conditions

IV

III

II

I

LETE


0.8

1

1.2

1.4

1.6

-125 -50 25 100 175 250% C h o r d

M

CFD

Experiment

30% Pitch

Relative Mach Number at 90% Radius (Peak Efficiency)

TELE

0.8

1

1.2

1.4

1.6

-125 -50 25 100 175 250% C h o r d

M

CFD

Experiment

50% Pitch

TELE

School of Aerospace EngineeringShock-Boundary Layer InteractionShock-Boundary Layer Interaction

(Peak Efficiency) (Peak Efficiency)

LE

TE

Shock

Near Suction Side


LE

TE

Shock

Velocity Profile at Mid-PassageVelocity Profile at Mid-Passage (Peak efficiency) (Peak efficiency)

•Flow is well aligned.•Very small regions of separation observed in the tip clearance gap(Enlarged view)

-50

-30

-10

10

30

50

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

% Mass Flow rate Fluctuations

% P

ress

ure

Flu

ctua

tion

s

Fluctuations are very small (2%)


LE

TE

Clearance Gap

Enlarged View of Velocity Profile in Enlarged View of Velocity Profile in the Clearance Gap (Peak efficiency)the Clearance Gap (Peak efficiency)

•The reversed flow in the gap and the leading edge vorticity grow in size and magnitude as the compressor operates at off-design conditions


Peak Efficiency

Controlled

A

Performance Map (NASA Rotor 67)Performance Map (NASA Rotor 67)

measured mass flow rate at choke: 34.96 kg/s

CFD choke mass flow rate: 34.76 kg/s1.3

1.4

1.5

1.6

1.7

1.8

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

3% Bleed Air

Near Stall

Unstable Conditions

BC

Choke m

m

D


IIIIIIIVLE

TE

I

II

III

IV

Location of the Probes for Observing Location of the Probes for Observing the Pressure and Velocity Fluctuationsthe Pressure and Velocity Fluctuations

The probes are located at 30% chord upstream of the rotor and 90% span. They are fixed in space.

School of Aerospace EngineeringOnset of the Stall (Clean Inlet)Onset of the Stall (Clean Inlet)

•Probes show identical fluctuations.

•Flow while unsteady, is still symmetric from blade to blade.

IIIIII

IV

0.5

0.8

1.1

1.4

1.7

0.00 0.36 0.73 1.09 1.45 1.82

Pre

ssur

e

t/2

I

II

III

IV


IIIIIIIV

Onset of the Stall (Disturbed Inlet)Onset of the Stall (Disturbed Inlet)

•Inlet distortion simulated by dropping the stagnation pressure in one block by 20%.

•Flow is no longer symmetric from blade to blade.

•Frequency of rotating stall is N, where : blade passage frequency.

0.4

0.7

1

1.3

1.6

0.00 0.36 0.73 1.09 1.45 1.82 2.18

Pre

ssur

e

t/2


Bleed Valve ControlBleed Valve Control

• Pressure, density and tangential velocities are extrapolated from interior. .• Un = mb/(Ab)

One Tip Chord

Hub

Shroud


Bleed Valve ControlBleed Valve Control

-50

-30

-10

10

30

50

-40 -20 0 20 40

-50

-30

-10

10

30

50

-40 -20 0 20 40

% Mass Flow Rate Fluctuations

% Total Pressure

Fluctuations

Without Bleed Valve

With Bleed Valve

3% bleed air reduces the total pressure fluctuations by 75%


Bleed Valve ControlBleed Valve ControlAxial Velocity Near LEAxial Velocity Near LE

% F

rom

Hub

After 1.5 Rev.

After 0.5 Rev.

Bleed Valve.


•A 3-D numerical flow solver has been developed to investigate compressor instabilities.

•The flow solver has been applied to obtain a detailed understanding of surge and rotating stall phenomena in axial and centrifugal compressors.

•Air-injection and bleeding have been numerically analyzed as compressor control schemes. •Surge margin extension was achieved for both compression systems.

•The proper application of air-injection is sensitive to the injection-parameters (e.g. yaw angle).

ConclusionsConclusions

Documents

School of Aerospace Engineering Alex Stein, Saeid Niazi, and Lakshmi N. Sankar School of Aerospace Engineering Georgia Institute of Technology lsankar/MURI