Download - School of Aerospace Engineering Computational Analysis of Stall and Separation Control in Centrifugal Compressors Presented By Alexander Stein School of.

School of Aerospace EngineeringComputational Analysis of Stall and Computational Analysis of Stall and

Separation Control in Separation Control in Centrifugal CompressorsCentrifugal Compressors

Presented By

Alexander SteinSchool of Aerospace EngineeringGeorgia Institute of Technology

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

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School of Aerospace EngineeringOutline of PresentationOutline of Presentation

•Research objectives and motivation

•Background of compressor control

•Introduction of numerical tools

•Configurations and validation results•DLR high-speed centrifugal compressor (DLRCC)

•NASA Glenn low-speed centrifugal compressor (LSCC)

•Off-design results without control•Surge analysis

•Off-design results with air injection control•Steady jets

•Pulsed jets

•Conclusions and recommendations

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School of Aerospace Engineering

Motivation and ObjectivesMotivation and Objectives• Use CFD to explore and understand compressor stall and surge

• Develop and test control strategies (air injection) for centrifugal compressors

• Apply CFD to compare low-speed and high-speed configurations

Lines of ConstantRotational Speed

Lines of ConstantEfficiency

Cho

ke

Lim

it

Sur

ge L

imit

Flow Rate

Tot

al P

ress

ure

Ris

e

Desired Extension of Operating Range

4


Motivation and ObjectivesMotivation and Objectives

Compressor instabilities can cause fatigue and damage to entire engine

5


Greitzer’s Phenomenological ModelGreitzer’s Phenomenological Model

PlenumVp

Throttle

Ac

Length Lc

Compressor

Helmholtz-Resonator Model

cc

p

ch AL

V

a2

U

L2

UB

Non-dimensional B-Parameter (Greitzer):

Assumptions:•Compressor modeled as actuator disk

•Fluid inertia contained in pipes•Spring-like properties confined to plenum

B < Bcritical => Rotating StallB > Bcritical => Surge

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School of Aerospace EngineeringWhat is Surge?What is Surge?

Mild Surge Deep Surge

Time

Flow Rate

Period of Deep Surge Cycle

Flow Reversal

Pressure Rise

Flow Rate

MeanOperating Point

Limit CycleOscillations

Pressure Rise

Flow Rate

PeakPerformance

Time

Flow Rate

Period ofMild Surge Cycle

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• Diffuser Bleed Valves•Pinsley, Greitzer, Epstein (MIT)•Prasad, Neumeier, Haddad (GT)

• Movable Plenum Wall•Gysling, Greitzer, Epstein (MIT)

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

• Air Injection•Murray (CalTech)•Weigl, Paduano, Bright (NASA Glenn)•Fleeter, Lawless (Purdue)

How to Alleviate SurgeHow to Alleviate Surge

Bleed Valves

Movable Plenum Walls

Guide Vanes

Air-Injection

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Numerical Formulation (Flow Solver)Numerical Formulation (Flow Solver)

t

qdV E ˆ i F ˆ j G ˆ k n dS R ˆ i S ˆ j T ˆ k

n dS

Reynolds-averaged Navier-Stokes equations in finite volumerepresentation:

• q is the state vector.• E, F, and G are the inviscid fluxes (3rd order accurate). R, S,

and T are the viscous fluxes (2nd order accurate).• A one-equation Spalart-Allmaras model is used.• Code can handle multiple computational blocks and rotor-

stator-interaction.

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Boundary Conditions (Flow Solver)Boundary Conditions (Flow Solver)

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

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Outflow Boundary Condition (Flow Solver)Outflow Boundary Condition (Flow Solver)

Plenum chamber•up(x,y,z) = 0 •pp(x,y,z) = const.•isentropy

ap, Vp

mc

.

mt

.

Outflow boundary

)mm(V

a

dt

dptc

p

2pp

Conservation of mass and isentropic expression for speed of sound:

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NASA Low-Speed Centrifugal CompressorNASA Low-Speed Centrifugal Compressor

•Designed and tested at NASA Glenn

•Mild pressure ratio•Ideal CFD test case

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NASA Low-Speed Centrifugal CompressorNASA Low-Speed Centrifugal Compressor•20 Main blades•55 Backsweep•Grid 129 x 61 x 49 (400,000 nodes)

•A grid sensitivity study was done with up to 3.2 Million nodes.

Design Conditions:•1,862 RPM•Mass flow = 30 kg/s•Total pressure ratio = 1.19•Adiab. efficiency = 92.2%•Tip speed = 492 m/s•Inlet Mrel = 0.31

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0.9

0.92

0.94

0.96

0.98

1

1.02

1.04

0 0.2 0.4 0.6 0.8 1

0.9

0.92

0.94

0.96

0.98

1

1.02

1.04

0 0.2 0.4 0.6 0.8 1

Validation Results (Low-Speed)Validation Results (Low-Speed)Blade Pressure Computations vs. MeasurementsBlade Pressure Computations vs. Measurements

p/p

Meridional Chord

5% Span 49% Span

0.9

0.92

0.94

0.96

0.98

1

1.02

1.04

0 0.2 0.4 0.6 0.8 1

79% Span Mass flow = 30 kg/sec (design) CFD pressure side CFD suction side Exp pressure sideExp suction side

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Validation Results (Low-Speed)Validation Results (Low-Speed)Blade Pressure Computations vs. MeasurementsBlade Pressure Computations vs. Measurements

p/p

Meridional Chord

5% Span 49% Span

79% Span Mass flow = 23.5 kg/sec(75% of design mass flow) CFD pressure side CFD suction side Exp pressure sideExp suction side

0.9

0.92

0.94

0.96

0.98

1

1.02

1.04

0 0.2 0.4 0.6 0.8 1

0.9

0.92

0.94

0.96

0.98

1

1.02

1.04

0 0.2 0.4 0.6 0.8 1

0.9

0.92

0.94

0.96

0.98

1

1.02

1.04

0 0.2 0.4 0.6 0.8 1

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DLR High-Speed Centrifugal CompressorDLR High-Speed Centrifugal Compressor

•Designed and tested by DLR

•High pressure ratio•AGARD test case

40cm

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DLR High-Speed Centrifugal CompressorDLR High-Speed Centrifugal Compressor•24 Main blades•30 Backsweep•Grid 141 x 49 x 33 (230,000 nodes)

•A grid sensitivity study was done with up to 1.8 Million nodes.

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

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School of Aerospace EngineeringValidation Results (High-Speed) Validation Results (High-Speed)

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

Experiment

CFD

Loca

l Sta

tic P

ress

ure,

p/p

std

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School of Aerospace EngineeringValidation Results (High-Speed)Validation Results (High-Speed)

Same momentum deficit was observed experimentally in other configurations.

Near suction side

Mid-passage

Near pressure side

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

Computational and experimental data are within 5%

Fluctuations at 3.2 kg/sec are 23 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

Mass Flow (kg/sec)

Tot

al P

ress

ure

Rat

io

Experiment

CFD

A

BCD

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

Large limit cycle oscillations develop

Oscillations remain bound=> mild surge

-30 -20 -10 0 10 20 30

-20

-10

0

10

20

Mass Flow Fluctuations (%)

A: 4.6 kg/sec

Pre

ssur

e R

ise

Flu

ctua

tion

s (%

)

-30 -20 -10 0 10 20 30

-20

-10

0

10

20


B: 3.8 kg/sec

Pre

ssur

e R

ise

Flu

ctua

tion

s (%

)

-30 -20 -10 0 10 20 30

-20

-10

0

10

20


D: 3.2 kg/sec

Pre

ssur

e R

ise

Flu

ctua

tion

s (%

)

-30 -20 -10 0 10 20 30

-20

-10

0

10

20


C: 3.4 kg/sec

Pre

ssur

e R

ise

Flu

ctua

tion

s (%

)

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Mild surge cycles develop

Surge amplitude grows to 60% of mean flow rate

Surge frequency = 90 Hz (1/100 of blade passing frequency)

Off-Design Results (High-Speed) Off-Design Results (High-Speed) Mass Flow FluctuationsMass Flow Fluctuations

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School of Aerospace EngineeringOff-Design Results (High-Speed)Off-Design Results (High-Speed)

Flowfield vectors show a large separation zone near the leading edge

Velocity vectors colored by Mrel

at mid-passage

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School of Aerospace EngineeringOff-Design Results (High-Speed) Off-Design Results (High-Speed)

Stagnation Pressure ContoursStagnation Pressure Contours

•Vortex shedding causes reversed flow•Origin of separation occurs at leading edge pressure side

View

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Flowfield stalls but no surge occurs

This is in accordance with Greitzer’s B-Criterion:

sec/kg19m

sec/kg15m

sec/kg13m

criticalcc

p BAL

V

a2

UB

Off-Design Results (Low-Speed)Off-Design Results (Low-Speed) Velocity Vectors at Design SpeedVelocity Vectors at Design Speed

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School of Aerospace EngineeringOff-Design Results (Low-Speed)Off-Design Results (Low-Speed)

Velocity vectors at 200% design speed at mid-passage

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Performance Characteristic MapPerformance Characteristic Map

1

1.2

1.4

1.6

1.8

2

2.2

10 20 30 40 50Mass Flow Rate (kg/s)

To

tal P

ress

ure

Ra

tio

Exp

CFD

200% Design Speed

Design Speed

Unsteady fluctuations are denoted by size of circles

Surge fluctuations at 200% design speed are 7 times larger than at 100% design speed

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Comparison of Different Shaft SpeedComparison of Different Shaft Speed

Relative Mach NumberNear Leading Edge Hub

Relative Mach NumberNear Leading Edge Tip

B-Parameter

100%DesignSpeed

0.13 0.23 1.95

200%DesignSpeed

0.26 0.475 3.9

Conclusions:• Compressibility effects are fundamental for surge• For surge to occur B > Bcritical

cc

p

ch AL

V

a2

U

L2

UB

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Air Injection SetupAir Injection Setup

Systematic study: injection rate and yaw angle were identified as the most sensitive parameters.

Related work: Rolls Royce,Cal Tech, NASA Glenn /MIT,

0.04RInlet

Casing

5°

Rotation Axis

Impeller

RInlet

Yaw Angle

Main Flow

Injected Fluid Sheet

Compressor Face

Compressor Casing

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School of Aerospace EngineeringAir Injection Results (High-Speed)Air Injection Results (High-Speed)

Different Yaw Angles, 3% Injected Mass Flow RateDifferent Yaw Angles, 3% Injected Mass Flow Rate

-20

0

20

40

60

80

100

120

0% 20% 40% 60% 80% 100%

Time (in Percentage of Tsurge)

Inci

denc

e A

ngle

(D

egre

e)No Injection

3.2% Injection

Yaw angle directly affects incidence angle=> Maximum control for designer

-40

-30

-20

-10

0

10

20

-20 -10 0 10 20 30 40 50

Yaw Angle (Degree)

Inci

denc

e A

ngle

(D

egre

e)

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Different Yaw Angles, 3% Injected Mass Flow RateDifferent Yaw Angles, 3% Injected Mass Flow Rate

Positive yaw angle is measured in opposite direction of impeller rotation

-25

0

25

50

75

100

-20 0 20 40 60

Yaw Angle (Degree)

Red

ucti

on in

Sur

ge

Am

plit

ude

(%)

2

2.5

3

3.5

4

4.5

0 10 20 30 40

yaw = 45 deg

yaw = 7.5 deg

yaw = -15 deg

Rotor Revolutions, t/

Mas

s F

low

(kg

/sec

)

Optimum yaw angle of 7.5deg. yields best result

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Leading edge separation is suppressed by injection

Velocity vectors colored by Mrel

at mid-passage

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

Leading edge reversed flow regions has vanished

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School of Aerospace EngineeringAir Injection Results (Parametric Studies)Air Injection Results (Parametric Studies)

High-Speed Compressor Low-Speed Compressor

•An optimum yaw angle exists for both compressors.•A reasonable amount (3% to 5%) of injected air is sufficient in both configurations to suppress surge.

-200

2040

12

34

50

50

100

150

200

Injection Rate (%)

Non

dim

. Sur

ge A

mpl

itud

e (%

)

20

40

60

80 3 4 5 6 7

0

100

200

300

Injection Rate (%)N

ondi

m. S

urge

Am

plit

ude

(%)

Yaw Angle(Deg.)Yaw Angle

(Deg.)

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School of Aerospace EngineeringAir Injection Results (Neural Network Model)Air Injection Results (Neural Network Model)

A Neural Network can be trained to model the injection maps:•Include more parameters (shaft speed, throttle settings, etc.)•Use NN-model as a controller in a real engine•Training of such a controller by CFD is much cheaper than by experiments

Input Hidden Layer Hidden Layer Output Layer

Yaw Angle

Injection Rate

W

b

W

b

Surge Amplitude

W

b+++

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School of Aerospace EngineeringAir Injection Results (Neural Network Model)Air Injection Results (Neural Network Model)

High-Speed Compressor Low-Speed Compressor

Reasonable agreement between CFD injection performance maps and NN models is observed.

Injection Rate (%)

Non

dim

. Sur

ge A

mpl

itud

e (%

)

Yaw Angle(Deg.)

Injection Rate (%)N

ondi

m. S

urge

Am

plit

ude

(%)

Yaw Angle(Deg.)

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0 10 20 30 40 50-100

-50

0

50

100

Air Injection Results (Pulsed Jets)Air Injection Results (Pulsed Jets)

Surge fluctuations decrease as long as the injection phase was lagged 180 deg. relative to the flow=> suggests feedback control

)tsin(007.0023.0m

)t(msurge

flow

inj

With Phase Angle Adjustments

Without Phase Angle Adjustments

Non

dim

. Su

rge

Flu

ctu

atio

ns

(%)

Rotor Revolutions, t

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Amplitude of pulsed jets has a stronger impact than mean injection rate=> reduction in

external air requirements by 50%

)tsin(015.0015.0m

)t(msurge

flow

inj

0 5 10 15 20 25-100

-50

0

50

100


With Phase Angle Adjustments

Without Phase Angle Adjustments

Non

dim

. Su

rge

Flu

ctu

atio

ns

(%)


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School of Aerospace EngineeringAir Injection Results (Pulsed Jets)Air Injection Results (Pulsed Jets)

A short boost from the injected air is sufficient to suppress surge onset

)tsin(015.0015.0m

)t(msurge

flow

inj

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School of Aerospace EngineeringAir Injection Results (Pulsed Jets)Air Injection Results (Pulsed Jets)

No separation occurs

)tsin(015.0015.0m

)t(msurge

flow

inj

40


• Jets pulsed at higher frequencies are more effective than low-frequency jets (increased mixing, higher turbulent intensity)

• There is a practical limitation on the highest possible frequency

)t4sin(007.0023.0m

)t(msurge

flow

inj

0 5 10 15 20 25 30-15

-10

-5

0

5

10

15

Air Injection Results (Pulsed Jets)Air Injection Results (Pulsed Jets)N

ond

im. S

urg

e F

luct

uat

ion

s (%

)


41


1.5% injected mass is sufficient to suppress surge

)t4sin(015.0015.0m

)t(msurge

flow

inj

0 5 10 15 20-15

-10

-5

0

5

10

15

Non

dim

. Su

rge

Flu

ctu

atio

ns

(%)



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School of Aerospace EngineeringConclusionsConclusions

•A Viscous flow solver has been developed to •obtain a detailed understanding of surge in centrifugal compressors.•determine fluid dynamic factors that lead to stall onset.

•The non-dimensional B-Parameter is a useful way to determine a priori which configuration will surge.

•Steady jets are effective means of controlling surge:•Alter local incidence angles and suppress boundary layer separation.•Yawed jets are more effective than parallel jets.•An optimum yaw angle exists for each configuration.•Air injection can be modeled by a multi-parameter neural network.

•Pulsed jets yield additional performance enhancements:•Lead to a reduction in external air requirements.•Jets pulsed at higher frequencies perform better than low-frequency jets.

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School of Aerospace EngineeringRecommendationsRecommendations

•Perform studies that link air injection rates to surge amplitude via a feedback control law.

•Use flow solver to analyze and optimize other control strategies, e.g. inlet guide vanes, synthetic jets, casing treatments.

•Employ multi-passage flow simulations to study rotating stall and appropriate control strategies.

•Study inflow distortion and its effects on stall inception.

•Improve turbulence modeling of current generation turbomachinery solvers. Analyze the feasibility of LES methods.

44


How to Control Surge (Active Control)How to Control Surge (Active Control)

Controller Unit

Bleed Air

PressureSensorsAir

Injection

45

School of Aerospace EngineeringLiterature Survey on Air InjectionLiterature Survey on Air Injection

Rolls Royce (Day et al., 1997):Injection into Tip Region is More Effective than Injection into the Core Flow

Cal Tech (Murray et al., 1997):Steady Air Injection Reduces Bandwidth Requirements for Bleed Valves

NASA/MIT (Bright et al., 2000):Effectiveness of Air Injectors is Independent of 1.) Azimuthal Jet Arrangement2.) Number of Jets

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School of Aerospace EngineeringNumerical Formulation (Flow Solver)Numerical Formulation (Flow Solver)

* * * *i-1 i i+1 i+2

Cell Face i+1/2

Stencil for qleftStencil for qright

Left Right

A Four Point Stencil is Used to Compute the Inviscid Flux Termsat the Cell Faces According to Roe’s Flux Splitting Scheme::

Third-Order Accurate in Space

• Turbulence is Modeled by One-Equation Spalart-Allmaras Model• Code Can Handle Multiple Computational Blocks and Rotor-

Stator-Interaction

47


Overview of ConfigurationsOverview of ConfigurationsCentrifugal Compressor C FD Configurations

U n ders tan d in g of S tall O n set Im p ro ved In jectio n Schem es

Im p ro ved B lad e Des ign D evelo p In jec tio n Crite r ia

Code Validationvs. Experimental Data

- D e sig n O p era tion- O ff-De s ign O p era tion

Id e n tif ica tio n o f P a ram ete rs th a tL e ad to In sta b ilit ies

- a t O ff-D es ig n -S p e ed

S ta b ility-A n a lys is a t P o in t o fM a x im um P re ssu re R ise

A ir-In je c tio n C o ntro l S ch e m e- D e riva tion o f In je ctio n P e rfo rm a nce M a p Us in g S te ad y B lo w ing

NASA Low SpeedCentrifugal Compressor (LSCC)

Code Validationvs. Experimental Data

- D e sig n O p era tion

Id e n tif ica tio n o f P a ram ete rs th a tL e ad to In sta b ilit ies

- a t O ff-D es ig n -S p e ed

S ta b ility-A n a lys is a t P o in t o fM a x im um P re ssu re R ise

A ir-In je c tio n C o ntro l S ch e m e- D e riva tion o f In je ctio n P e rfo rm a nce M a p Us in g S te ad y B lo w ing

- O p tim iza tio n o f C o n tro l S ch e m e U sin g P u lsed Je ts

DLR 4.7:1Centrifugal Com pressor (DLRCC)

48

School of Aerospace EngineeringThe Present ApproachThe Present Approach

Air-injection contro l

G TTURBO3 DV isco u s f lo w so lve r

- S im u la te & a n a lyse f low f ie ld- Id en tify se p ara tion reg io ns

G TSYS3DE ig en m od e so lve r

- e va lu a te sys te m e ige n va lu es- a n a lyse sys te m s ta b ility

Experim entsW o rk a t M IT , C a lTe ch , N A S A L e w is- va lid a te f lo w so lve r- co rre la te in je c tio n re su lts

Understanding of stall onset Im proved injection schem es

Im proved blade design Develop injection criteria

The Tools

The Results

49

School of Aerospace EngineeringValidation Results (Low-Speed)Validation Results (Low-Speed)

Velocity Vectors in Meridional Planes Velocity Vectors in Meridional Planes

Wake-likemomentum deficit

4% away

from

Pressure

Side

50% away

from

Pressu

re Side97% aw

ay fr

om

Pressu

re Side

Leading Edge

Trailing EdgeClearance GapFlow ProducesVelocity Deficit

Same Phenomenonwas Observed Experimentally

50


51


52

School of Aerospace EngineeringEigenmode Analysis (GTSYS3D)Eigenmode Analysis (GTSYS3D)

• Calculates eigenvalues/-vectors of the compression system matrix

• Based on small perturbation Euler model:q = q0 + q

• The resulting form is:d/dt(q) = Aq

where: - q is the state vector of small perturbations- A is the system matrix of size

5N1N2N3 x 5N1N2N3

53


Off-Design Results (High-Speed)Off-Design Results (High-Speed)System eigenvalues at stable condition (4.6 kg/sec) System eigenvalues at stable condition (4.6 kg/sec)

-0.6

-0.4

-0.2

0

0.2

0.4

0.6

-2.5 -2 -1.5 -1 -0.5 0

Re

Im

9x4x3

-0.6

-0.4

-0.2

0

0.2

0.4

0.6

-2.5 -2 -1.5 -1 -0.5 0

ReIm

18x7x5

•Mostly acoustic modes with Re < 0 (damping, stable)•Complex conjugate pairs are oscillatory•Simple poles (Im = 0) near origin are unstable

54

School of Aerospace EngineeringOff-Design Results (High-Speed) Off-Design Results (High-Speed)

System eigenvalues during surge cycle System eigenvalues during surge cycle

-0.08

-0.04

0

0.04

0.08

-1 -0.5 0

ReIm

-0.08

-0.04

0

0.04

0.08

-1 -0.5 0

Re

Im

-0.08

-0.04

0

0.04

0.08

-1 -0.5 0

Re

Im

-0.08

-0.04

0

0.04

0.08

-1 -0.5 0

Re

Im

At beginning of surge cycle

After 25% of surge cycle



•Most acoustic (damping) modes have vanished

•Simple pole at origin destabilizes system