Estimation of Turbomachinery Losses Through Cascade Testing

“ESTIMATION OF TURBOMACHINERY FLOWESTIMATION OF TURBOMACHINERY FLOW LOSSES THROUGH CASCADE TESTING”

A lecture by

KMM SWAMY & R SENTHIL KUMARANS i ti t P l i Di i iScientists Propulsion Division,

National Aerospace Laboratories

for two day seminar ony

Loss Mechanisms in Steam and Gas Turbines

held at M.S.Ramaiah School of Advanced Studies

Date: 18 07 2009Date: 18-07-2009

Types of losses in turbomachinery

Losses associated with boundary layers / viscous phenomena- Friction, wakes, separation, secondary flows, mixing

Losses associated with compressibility effectsLosses associated with compressibility effects- Shock losses

Miscellaneous losses- Tip clearance flows, disk-friction, partial admission, incidence

18/07/2009 LOSS MECHANISMS IN STEAM AND GAS TURBINES KMM SWAMY

Representation of loss and efficiency

Turbine CompressorPressure loss coefficient ω = (P01 - P02) / (P02 - p2)

E l ffi i t

Compressor

Pressure loss coefficient ω = (P01 - P02) / (P01 - p1)

Energy loss coefficientζ = (h2 - h2s) / ½ C2

2

Efficiency(h h ) / (h h )

Energy loss coefficientζ = (h2 - h2s) / ½ C1

2

Efficiencyηt = (h01 ‐ h02 ) / (h01 ‐ h02s) ηc = (h02s ‐ h01 ) / (h02 ‐ h01)


Linear Cascade Annular cascade tunnel

Stages of tests to understand turbo machinery flows

Quick and easy techniqueexcellent for parametric studySimulation of 3D flow not possible

A closer approximation to actual conditionModel design and experimentation complexDoes not include the rotation effect

Low speed large scale test rigCl t th i diti

High speed rigM lCloser to the engine condition

Enables detailed measurementsSimulates engine Reynolds number

More complexDetailed measurements difficultCloser to engine condition

rem

ent

High speed rigs

Engine

e of

mea

su

Low speed large scale Rigs

g p g

Ease

Linear Cascade

Annular Cascade

Flow field complexity


Linear cascade model & cascade testing

sγ

S - Pitchγ - Stagger

T bi bl d Li C d d l

A linear cascade model is an array of aerofoils stacked at uniform pitch and stagger representing a section of a turbo machinery blade row.

Turbine blade row Linear Cascade model

Linear cascade testing is a simplified experimental method for evaluating aerodynamic performance of turbo machinery aerofoils where Coriolis effectsand curvilinear effects are ignored.

The three-dimensional flows can be simplified to two-dimensional flowsby using linear cascades.


Cascade tests for Axial machines and radial machines

Axial machines

The blade row is unrolled from a cylinder by a simple transformationby a simple transformation

x = z, y = r θ

Radial machines

Data obtained from conventionali l d h ll b li d baxial cascades shall be applied by

conformal transformation from radial (z = reiθ) to axial plane (ζ = ξ + iη)

Where,

ζ = ln z, ξ = ln r η = θ


SIGNIFICANCE OF CASCADE TESTSFlow parameters such as inlet flow angle true relative Mach numberFlow parameters such as inlet flow angle, true relative Mach number,

true Reynolds number etc., can be simulated with ease

Can provide aerodynamic performance data like blade loading / lift

coefficient, profile loss / drag coefficient and flow deflection

Easy to map pressure and velocity distributions over the aerofoils and in the

passagep g

Detailed studies on laminar, transition & turbulent boundary layers over

turbo machinery aerofoils can be carried out

Separation and vortex formation studiesSeparation and vortex formation studies

Local boundary layer profile and shear stress measurements over the aerofoils can

also be made

It is simple to generate data at off design conditions

Ideal method for comparison of different profiles for the same design or in other words

optimization of aerofoilsp

Can provide data bank for validating CFD codes


Limitations of Cascade testingCurvilinear and Coriolis effects are ignored

Predominantly a cold flow test method

Offers no information on three dimensional flow structure

Lack of information on unsteady flow fields

A very difficult process while applied to radial flow machinesA very difficult process while applied to radial flow machines

Can be an expensive exercise

Cascade test data require appropriate treatment if used for through flow analysis like stream line curvature method

Streamlines across aStreamlines across a multistage turbomachine


Cascade wind tunnel - Classification Cascade d tu e C ass cat o

1. Subsonic, transonic & supersonic

2 Bl d & k d2. Blowdown & suck down

3. Open circuit & closed circuit (Variable density)

4. Medium of operation: Air, steam, combusted gas products etc.,

NAL Cascade Wind Tunnels

a) Subsonic cascade Tunnel (SCT)b) Transonic Cascade Tunnel (TCT)


NAL TRANSONIC CASCADE TUNNEL (TCT)18/07/2009 LOSS MECHANISMS IN STEAM AND GAS TURBINES KMM SWAMY

NAL - TRANSONIC CASCADE TUNNEL SPECIFICATIONS

T t S ti 153 500 *Test Section - 153 x 500 mm*Blade chord - 40 to 80 mmProbe traverse - 220 mm in 150 secondsSpan wise traverse - 75 mmAir storage volume - 2800 cubic metersStorage pressure - 11 atmTotal temperature - 300 KMass flow (Typical) - 5 to 15 Kg/s

* Maximum


FOR TURBINE CASCADES:

Inlet Mach number Up to chokingInlet Mach number - Up to chokingOutlet Mach number - Up to 1.5Reynolds number - 0.3 to 2.5 millions (outlet)

FOR COMPRESSOR CASCADES:

Inlet Mach number - Up to 0.85Reynolds number - 0.7 to 1.3 millions (inlet)Reynolds number - 0.6 to 1.1 millions(outlet)


Instrumentation for cascade tunnelsPressure probesessu e p obes

Keil probePitot probe

Three hole probe

Total pressure / temperature rake

Boundary layer probetemperature rake

Courtesy: M/S United sensor corporation


Five hole probes


ESP pressure scanner 16 channel intelligent pressure scannerESP pressure scanner 16 channel intelligent pressure scanner

Kulite pressure transducer Three sensor hot wire probe

Courtesy: M/S Scanivalve corporation, Kulite & Dantec


Flow visualization techniques for cascade tunnels

Smoke flow visualization

Tuft flow visualization

Oil flow visualization

Schlieren techniqueSchlieren technique

Background Oriented Schlieren technique

Interferograms

Particle image velocimetry

LASER Doppler Velocimetry SMOKE FLOW VISUALIZATION OVER A TURBINE CASCADE


INTERFEROGRAM OF A TURBINE CASCADE

Vortices

SCHLIEREN PHOTOGRAPH OF A TURBINE ROTOR CASCADE


OIL FLOW VISUALIZATION ON A TURBINE CASCADE

CALIBRATION OF PRESSURE PROBES

Combined pressure probes are used for loss (fom total pressure) and flow deflection measurements during cascade tests

These probes have to be calibrated as they are employed in non-nulling mode

FACILITIES AT NAL FOR CALIBRATING PRESSURE PROBESFACILITIES AT NAL FOR CALIBRATING PRESSURE PROBES

Induction tunnel

A straight 5 hole 3D probe calibrated in the new facilityA straight 5-hole 3D probe calibrated in the new facility


SAMPLE CALIBRATION CURVES OF A FIVE HOLE 3D PROBE


SCHEMATIC OF A TURBINE NOZZLE CASCADEIN NAL TRANSONIC CASCADE TUNNEL


Typical Wake Traverse of a Transonic Gas Turbine Stator Cascade

0.97

1.00

Rat

io

yp

0.94

tal P

ress

ure

0.88

0.91

02Y/

P01,

Tot

BLADE 1 BLADE 2 BLADE 3

0.850 0.5 1 1.5 2 2.5 3

P0

PitchP01: 1440 5 mm Hg M1: 0 533 Beta1: 63 1 Deg Pitch

VARIATION OF TOTAL PRESSURE RATIO WITH PROBE TRAVERSE

P01: 1440.5 mm Hg M1: 0.533 Beta1: 63.1 Deg P02: 1389.8 mm Hg M2: 1.052 Beta2: 67.4 Deg



-63

-60

(Deg

)


-66

63

low

Ang

le

-72

-69

y, O

utle

t F


-75

72

0 0.5 1 1.5 2 2.5 3

Bet

a2

Pit hP01 1440 H M1 0 33 B 1 63 1 D Pitch

VARIATION OF OUTLET FLOW ANGLE WITH PROBE TRAVERSE

P01: 1440.5 mm Hg M1: 0.533 Beta1: 63.1 Deg P02: 1389.8 mm Hg M2: 1.052 Beta2: 67.4 Deg



1 10

1.15

1.20m

ber

yp

1.00

1.05

1.10

et M

ach

Num

0.90

0.95

M2y

, Out

le


0.80

0.85

0 0.5 1 1.5 2 2.5 3

PitchP01: 1440.5 mm Hg M1: 0.533 Beta1: 63.1 Deg

VARIATION OF OUTLET MACH NUMBER WITH PROBE TRAVERSE

P02: 1389.8 mm Hg M2: 1.052 Beta2: 67.4 Deg


0.10

CIE

NT

EFFECT OF OUTLET MACH NUMBER ON PRESSURE LOSS COEFFICIENTOF A TURBINE ROTOR CASCADE

0 06

0.08

SS C

OEF

FIC

)]

0.04

0.06

RA

GED

LO

SP0

/ (P

02-p

2)

66.0 Deg63.1 Deg60.0 Deg

BETA1 set at

0.02

NTU

M A

VER

[DP

0.000.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0 1.1 1.2

MO

MEN

OUTLET MACH NUMBER 'M2'


70

eg)

EFFECT OF OUTLET MACH NUMBER ON OUTLET FLOW ANGLEOF A TURBINE ROTOR CASCADE

68

RA

GED

B

ETA

2' (D

e

BETA1 set at

64

66

NTU

M A

VER

W A

NG

LE ' 66.0 Deg

63.1 Deg60.0 Deg

BETA1 set at

62MO

MEN

UTL

ET F

LOW

600.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0 1.1 1.2

OU



0.7

EFFECT OF OUTLET MACH NUMBER ON INLET MACH NUMBEROF A TURBINE ROTOR CASCADE

0.5

0.6

VER

AG

EDM

BER

'M1'

BETA1 set at

0.3

0.4

MEN

TUM

AV

T M

AC

H N

U

66.0 Deg

63.1 Deg

60.0 Deg

BETA1 set at

0.1

0.2MO

MIN

LE

0.00.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0 1.1 1.2



EFFECT OF OUTLET MACH NUMBER ON SURFACE MACH NUMBER DISTRIBUTION

OF A GAS TURBINE PROFILE

1 2

1.4

OF A GAS TURBINE PROFILE

058, 0.276 & 0.383057, 0.333 & 0.476056, 0.387 & 0.570055 0 429 & 0 669

Beta1 set at 60 Deg.

Run No, M1 & M2

1.0

1.2r

055, 0.429 & 0.669054, 0.462 & 0.765053, 0.479 & 0.858052, 0.487 & 0.967051, 0.494 & 1.055

0.6

0.8

ach

Num

ber

0.4

0.6

M

0.0

0.2

0.0 0.2 0.4 0.6 0.8 1.0

X / Cax


EFFECT OF INCIDENCE ON PRESSURE LOSS COEFFICIENT OF A

5 0

6.0

(%)

GAS TURBINE NOZZLE VANE PROFILE M2: 0.94

4.0

5.0

OEF

FIC

IEN

T

3.0

RE

LOSS

CO

1.0

2.0

PRES

SUR

0.0-20 -15 -10 -5 0 5 10 15 20

INCIDENCE ANGLE (Deg.)

Pr. loss coeff

( g )


EFFECT OF INCIDENCE ON OUTLET FLOW ANGLE

70OF A NOZZLE VANE PROFILE

Beta2

M2: 0.94

68

69

GLE

(Deg

)

67

68

T FL

OW

AN

66

OU

TLE

65-20 -15 -10 -5 0 5 10 15 20

INCIDENCE ANGLE (Deg)INCIDENCE ANGLE (Deg)


Effect of Incidence on Surface Mach Number Distribution of a Gas Turbine Nozzle Vane Profile

1.2

1.4

-15 Deg-10 Deg0 Deg10 Deg

INCIDENCE

M2ref: 0.94

1.0

10 Deg15 Deg

0.6

0.8

ach

Num

ber

0.4

Ma

0.0

0.2

0.0 0.2 0.4 0.6 0.8 1.0

X/Cax


Effect of free stream turbulence

An experimental study was conducted in a two-dimensional linear cascade, focusingon the suction surface of a low pressure turbine blade. Flow Reynoldsnumbers, based on exit velocity and suction length, have been varied from 50,000 to300 000 The freestream turbulence intensity was varied from 1 1 to 8 1 percent300,000. The freestream turbulence intensity was varied from 1.1 to 8.1 percent.Separation was observed at all test Reynolds numbers. Increasing the flow Reynoldsnumber, without changing freestream turbulence, resulted in a rearward movement ofthe onset of separation and shrinkage of the separation zone. Increasing thefreestream turbulence intensity without changing Reynolds number resulted infreestream turbulence intensity, without changing Reynolds number, resulted inshrinkage of the separation region on the suction surface. The influences on theblade's wake from altering freestream turbulence and Reynolds number are alsodocumented. It is shown that width of the wake and velocity defect rise with adecrease in either turbulence level or chord Reynolds number

“An Experimental Investigation of the Effect of Freestream Turbulence on the Wake of a Separated Low-Pressure Turbine Blade at Low Reynolds Numbers”

decrease in either turbulence level or chord Reynolds number.

Murawski CG, Vafai K J. Fluids Eng. -- June 2000 -- Volume 122, Issue 2, 431


Effect of free stream turbulence

Tip clearance losses represent a major efficiency penalty of turbine blades. Thispaper describes the effect of tip clearance on the aerodynamic characteristics ofan unshrouded axial-flow turbine cascade under very low Reynolds numbery yconditions. The Reynolds number based on the true chord length and exit velocityof the turbine cascade was varied from 4.4×104 to 26.6×104 by changing thevelocity of fluid flow. The freestream turbulence intensity was varied between 0.5%and 4.1% by modifying turbulence generation sheet settings. Three-dimensionaly y g g gflow fields at the exit of the turbine cascade were measured both with and withouttip clearance using a five-hole pressure probe. Tip leakage flow generated a largehigh total pressure loss region. Variations in the Reynolds number and freestreamturbulence intensity changed the distributions of three-dimensional flow, but had no

“Effects of Reynolds Number and Freestream Turbulence on Turbine Tip Clearance Flow”T k ki M t J T b h J 2006 V l 128 I 1 166

y g ,effect on the mass-averaged tip clearance loss of the turbine cascade.

Takayuki Matsunuma J. Turbomach. -- January 2006 -- Volume 128, Issue 1, 166


An experimental and analytical study has been performed on the effect of ReynoldsEffect of free stream turbulence

number and free-stream turbulence on boundary layer transition location on thesuction surface of a controlled diffusion airfoil (CDA). The experiments wereconducted in a rectilinear cascade facility at Reynolds numbers between 0.7 and3.0×106 and turbulence intensities from about 0.7 to 4 percent. An oil streaktechnique and liquid crystal coatings were used to visualize the boundary layer state.For small turbulence levels and all Reynolds numbers tested, the accelerated frontportion of the blade is laminar and transition occurs within a laminar separationbubble shortly after the maximum velocity near 35–40 percent of chord. For highturbulence levels (Tu>3 percent) and high Reynolds numbers, the transition regionmoves upstream into the accelerated front portion of the CDA blade. For thoseconditions, the sensitivity to surface roughness increases considerably; at Tu=4percent, bypass transition is observed near 7–10 percent of chord. Experimentalresults are compared to theoretical predictions using the transition model, which isimplemented in the MISES code of Youngren and Drela. Overall, the results indicatethat early bypass transition at high turbulence levels must alter the profile velocitydistribution for compressor blades that are designed and optimized for high Reynolds

“Effects of Reynolds Number and Free-Stream Turbulence on Boundary Layer Transition in a Compressor Cascade”Schreiber HA etal J Turbomach -- January 2002 -- Volume 124 Issue 1 1

numbers.

Schreiber HA etal. J. Turbomach. January 2002 Volume 124, Issue 1, 1


Measurements of pressure distributions profile losses and flow deviation were

Effect of surface roughnessMeasurements of pressure distributions, profile losses, and flow deviation werecarried out on a planar turbine cascade in incompressible flow to assess the effects ofpartial roughness coverage of the blade surfaces. Spanwise-oriented bands ofroughness were placed at various locations on the suction and pressure surfaces ofthe blades Roughness height spacing between roughness elements and band widththe blades. Roughness height, spacing between roughness elements, and band widthwere varied. A computational method based on the inviscid/viscous interactionapproach was also developed; its predictions were in good agreement with theexperimental results. This indicates that good predictions can be expected for avariety of cascade and roughness configurations from any two-dimensional analysisvariety of cascade and roughness configurations from any two dimensional analysisthat couples an inviscid method with a suitable rough surface boundary-layer analysis.The work also suggests that incorporation of the rough wall skin-friction law into athree-dimensional Navier-Stokes code would enable good predictions of roughnesseffects in three-dimensional situations. Roughness was found to have little effect oneffects in three dimensional situations. Roughness was found to have little effect onstatic pressure distribution around the blades and on deviation angle, provided that itdoes not precipitate substantial flow separation. Roughness on the suction surfacecan cause large increases in profile losses; roughness height and location of theleading edge of the roughness band are particularly important. Loss increments dueleading edge of the roughness band are particularly important. Loss increments dueto pressure-surface roughness are much smaller than those due to similar roughnesson the suction surface.

1. “Measurements and prediction of the effects of surface roughness on fil l d d i ti i t bi d ”profile losses and deviation in a turbine cascade”

KIind RJ etal. J. Turbomach 1998, vol. 120, pp. 20-27


Effect of surface roughnessThe aerodynamic performance of a turbine blade was evaluated via total pressure lossmeasurements on a linear cascade. The Reynolds number was varied from 600 000 to1 200 000 to capture the operating regime for heavy-duty gas turbines. Four differenttypes of surface roughness on the same profile were tested in the High Speed CascadeWind Tunnel of the University of the German Armed Forces Munich and evaluatedagainst a hydraulically smooth reference blade. The ratios of surface roughness tochord length for the test blade surfaces are in the range of Ra/c=7.610−06–7.910−05.The total pressure losses were evaluated from wake traverse measurements. The lossincrease due to surface roughness was found to increase with increasing Reynoldsnumber. For the maximum tested Reynolds number of Re=1 200 000 the increase intotal pressure loss for the highest analysed surface roughness value of Ra=11.8 m wasfound to be 40% compared to a hydraulically smooth surface. The results of themeasurements were compared to a correlation from literature as well as towell-documented measurements in literature. Good agreement was found for highReynolds numbers between the correlation and the test results presented in this paperand the data available from literature.

“Surface Roughness Effects on Turbine Blade Aerodynamics”Frank Hummel etal. J. Turbomach JULY 2005, Vol. Copyright © 2005 by ASME 127 / 453


Surface isentropic Mach number distributionpfor β1=133.3 deg, Ma2,th=0.85 in dependenceon Reynolds number

Courtesy:Frank Hummel etal. J. Turbomach JULY 2005, Vol. Copyright © 2005 by ASME 127 / 453


Re2 th=900 000Re2 th=600 000 Re2,th=1200 000.

Total pressure loss from wake traverse measurements of a double Pitot probefor test blade rough part compared to smooth part Ma2 th=0 75 β1=133 3°

Re2,th=900 000.Re2,th=600 000. Re2,th 1200 000.

for test blade, rough part compared to smooth part. Ma2,th 0.75, β1 133.3

Courtesy:Frank Hummel etal. J. Turbomach JULY 2005, Vol. Copyright © 2005 by ASME 127 / 453


EFFECT OF INLET BOUNDARY LAYERSC O OU SMotivation: To study the performance of compressor aerofoil sections near the walls with the influence of boundary layers and secondary flows.

Use of flat plates (extension plate) and trip wires to generate boundary layers with displacement thickness of 1% & 3% of spanwith displacement thickness of 1% & 3% of span

FLOWTunnel wall

Boundary layer probe

Partition plate

Extension plateTrip wire

Partition plate

Boundary layer With trip wire

Boundary layer without trip wire

Cascade side plate Cascade blade

Schematic of a compressor cascade with a controlled inlet boundary layer

18/07/2009 LOSS MECHANISMS IN STEAM AND GAS TURBINES KMM SWAMY18/07/2009 LOSS MECHANISMS IN STEAM AND GAS TURBINES KMM SWAMY

32

NALCD [2004] ‐ CONFIGURATION : 3BOUNDARY LAYER

[With Partition Plates AR=1.5,& Inlet Extension Plates Of 140 mm]14-12-2004 BETA1 Set At 40 Deg.

EFFECT OF INLET BOUNDARY LAYERS

20

24

28

n mm) No Trip

2mm Trip3 mm trip

M = 0.6

4

8

12

16

SPAN (i 5mm Trip6 mm trip

0

0.4 0.5 0.6 0.7 0.8 0.9 1.0V/Vmax

3.5

4.0

PAN) M = 0.6

2.0

2.5

3.0HICKN

ESS (%

OF SP

0.0

0.5

1.0

1.5

DISPLACE

MEN

T T

0 1 2 3 4 5 6 7DIAMETER OF TRIP WIRE (in mm)

A CDA compressor cascade with flat extension plate and trip wire


COMPARISON OF RESULTS (RM2-PROFILE)EFFECT OF TRAILING EDGE GEOMETRY

COMPARISON OF RESULTS (RM2 PROFILE) Round Trailing Edge & Cut Trailing edge

At Design Incidence62

BETA 1 : 45.5 deg

60

GLE

, deg

SCTERTE

56

58

OU

TLET

AN

G

54

56

BET

AT2

, O M1 (Design) = 0. 406M2 (Design) = 1.284Beta 1 (M) = 43.00 degBeta 2 (M) = 62.00 deg

520.2 0.4 0.6 0.8 1 1.2 1.4 1.6

M2 OUTLET MACH NUMBERM2, OUTLET MACH NUMBER

FIG. VARIATION OF OUTLET FLOW ANGLE WITH OUTLET MACH NUMBER


COMPARISON OF RESULTS (RM2-PROFILE)EFFECT OF TRAILING EDGE GEOMETRY

COMPARISON OF RESULTS (RM2 PROFILE) Round Trailing Edge & Cut Trailing edge

At Design Incidence 0.3

SCTERTE

BETA 1 : 45.5 deg

0.2

0.25

OSS

CO

EFF. RTE

Pr. Loss = (P01-P02)/(P02-p2)

0.1

0.15

PRES

SUR

E L

M1 (Design) = 0. 406M2 (Design) = 1.284Beta 1 (M) = 43.00 degBeta 2 (M) = 62.00 deg

0.05

OM

EGA

, P

00 0.2 0.4 0.6 0.8 1 1.2 1.4 1.6

M2, OUTLET MACH NUMBER FIG. VARIATION OF PRESSURE LOSS COEFF. WITH OUTLET MACH NUMBER


Unlike the conventional method of heating the main flow or usingEFFECT OF COOLANT FLOWS ON TURBINE CASCADE

Unlike the conventional method of heating the main flow or usingCarbon-di- oxide as the coolant to simulate the density ratios, aningenious method of having the main flow at room temperature andcooling the coolant to a lower temperature has been adopted tocooling the coolant to a lower temperature has been adopted tosimulate the density ratios.

Coolant to mainstream temperature ratios of 0.5 and 0.9 were i l t dsimulated.

The actual aspect ratio of trailing edge slots of the NGV was maintained

using two partition plates in the cascade assembly.

Configurations:

I - Base profile, without coolant flowII - LE & TE coolant flows at Tc/Tg = 0.9, Pc/Pg = 1.02III - LE & TE coolant flows at Tc/Tg = 0.5, Pc/Pg = 1.04IV - TE coolant flow at Tc/Tg = 0.9, Pc/Pg = 1.02V - TE coolant flow at Tc/Tg = 0.5, Pc/Pg = 1.04


Simulation of actual coolant to gas density ratios in cascade tests

Motivation: To study the effect of coolant flows on the loss characteristics of gas turbine profiles

An ingenious method of having the main flow at room temperature and cooling the coolant to a lowertemperature was used to simulate the temperature ratios. The coolant air was passed through ah t h i d i b th f li id it t tt i l t theat exchanger immersed in a bath of liquid nitrogen to attain low temperatures.

The actual aspect ratio of trailing edge slots of the NGV was maintained using two partition plates in the cascade assembly.


Insulated coolant feed lines Thermocouple connections Heat exchanger

Coolant flow control valves


EFFECT OF COOLANT FLOWS ON TOTAL PRESSURE RATIO

1

EFFECT OF COOLANT FLOWS ON TOTAL PRESSURE RATIOOF A TURBINE NOZZLE CASCADE

Beta1: -1.5 Deg., M2:1.1

0.95

sure

Rat

io

Config. I

Config II

0.9

1, T

otal

Pre

ss

Config. II

Config. III

Config. IV

Config. V

0.85

P02Y

/P01

0.81.0 1.2 1.4 1.6 1.8 2.0 2.2 2.4 2.6 2.8 3.0

Pitch

P02/P01 V/S PROBE TRAVERSEP02/P01 V/S PROBE TRAVERSE


EFFECT OF COOLANT FLOWS ON OUTLET FLOW ANGLEOF A TURBINE NOZZLE CASCADE Beta1: -1 5 Deg M2:1 1

-64

-62

-60

e

OF A TURBINE NOZZLE CASCADE Beta1: -1.5 Deg, M2:1.1

-68

-66

-64

t Flo

w A

ngle

Config. I

Config II

-72

-70

eta2

Y, O

utle

t Config. II

Config. III

Config. IV

Config. V

-78

-76

-74Be

-801.0 1.2 1.4 1.6 1.8 2.0 2.2 2.4 2.6 2.8 3.0

PitchOUTLET FLOW ANGLE V/S PROBE TRAVERSE


EFFECT OF COOLANT FLOWS INTEGRATED LOSS COEFFICIENT

0 16

0.20

EFFECT OF COOLANT FLOWS – INTEGRATED LOSS COEFFICIENTBETA1 set at -1.5 Deg.

0.12

0.16

P0/(P

02-p

2)]

Config. IConfig. II

0 04

0.08

MEG

A =

[dP Config. III

Config. IVConfig. V

0.00

0.04

0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0 1.1 1.2

O


VARIATION OF PRESSURE LOSS COEFFICIENT WITH OUTLET MACH NUMBER


EFFECT OF COOLANT FLOWS INTEGRATED

75

EFFECT OF COOLANT FLOWS – INTEGRATED OUTLET FLOW ANGLE

BETA1 set at -1.5 Deg.

73

74

GLE

, Bet

a2

72

73

FLO

W A

NG Config. I

Config. II

Config. III

Config. IV

Config. V

71

OU

TLET

700.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0 1.1 1.2


VARIATION OF OUTLET FLOW ANGLE WITH OUTLET MACH NUMBER


EFFECT OF AVDR


EFFECT OF AVDR


USE OF PARTITION PLATES


CDNAL CASCADE PROFILE AT DESIGN INCIDENCE`AT DESIGN INCIDENCE

DIFFERENT AVDR1.3

Y

BETA 1 = 43.7 DEG

1.2

TY D

ENSI

TY

AVDR

1.1

AL

VELO

CI

RA

TIO 1.177

1.2481.285

AVDR

1

AVD

R, A

XIA

0.90.4 0.5 0.6 0.7 0.8 0.9 1

M1, INLET MACH NO.

FIG. 60 VARIATION OF AXIAL VELOCITY DENSITY RATIO WITH INLET MAC


CDNAL CASCADE PROFILE AT DESIGN INCIDENCE

EFFECT OF AVDR

0 25

0.3

FF

1.1771.248

AVDRBETA 1 = 43.7 DEG

0.2

0.25

LOSS

CO

E 1.2481.285OMEGA = (P01-P02)/(P01-P1)

0.1

0.15

PRES

SUR

E

0.05

0

OM

EGA

, P

00.4 0.5 0.6 0.7 0.8 0.9 1

M1, INLET MACH NO.

G 61 O O SS OSS CO C OFIG. 61 VARIATION OF PRESSURE LOSS COEFF. WITH INLET MACH NO.


CDNAL CASCADE PROFILE EFFECT OF AVDR

1.6

1

1.2

1.4

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Date post:	01-Dec-2015
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Estimation of Turbomachinery Losses Through Cascade Testing

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