The Pennsylvania State University
The Graduate School
Department of Aerospace Engineering
DESENSITIZATION OF OVER TIP LEAKAGE IN AN AXIAL TURBINE
ROTOR BY TIP SURFACE COOLANT INJECTION
A Thesis in
Aerospace Engineering
by
Nikhil Molahally Rao
© 2005 Nikhil Molahally Rao
Submitted in Partial Fulfillment of the Requirements
for the Degree of
Doctor of Philosophy
August 2005
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The thesis of Nikhil Molahally Rao was reviewed and approved* by the following:
Cengiz Camci Professor of Aerospace Engineering Thesis Advisor Chair of Committee
Dennis K. McLaughlin Professor of Aerospace Engineering
Lyle N. Long Professor of Aerospace Engineering
Savas Yavuzkurt Professor of Mechanical Engineering
Timothy F. Miller Senior Research Associate
George S. Lesieture Professor of Aerospace Engineering Head of the Department of Aerospace Engineering
*Signatures are on file in the Graduate School
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ABSTRACT
Mechanical energy extraction in axial flow turbine rotors occurs through a change
in angular momentum of the working fluid. The gap between the turbine rotor and the
stationary casing is referred to as the tip gap. High pressure turbine blades are typically
un-shrouded and pressure driven flow through the tip gap is termed as over tip leakage.
Over tip leakage reduces efficiency of the turbine stage and also causes thermal distress
to blade tip surfaces. The gap height typically increases over the operational life of a
turbine, leading to increased efficiency drop. The thermal load on the tip surface also
increases with increasing gap height and is exacerbated by the radial transport of high
temperature fluid found in the core of the combustor exit flow. Thus over tip leakage not
only decreases stage efficiency, but also constrains it by limiting the maximum cycle
temperature.
Reducing the sensitivity of turbine performance to the effects of the tip gap is
termed Tip Desensitization. An experimental investigation of tip desensitization through
coolant injection from a tip surface trench was conducted in a large scale, low speed,
rotating research turbine facility. Five out of twenty nine rotor blades, referred to as
cooled blades, are provided with coolant injection at four locations, at 61%, 71%, 81%,
and 91% blade tip axial chord length. At each of the first three locations the coolant jets
are directed towards the blade pressure-side, while coolant is exhausted radially at the last
location.
Qualitative information of tip surface flow was obtained through the
implementation of surface flow visualization using an oil and pigment mixture.
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Quantitative measurements of absolute total pressure were conducted in a plane located at
30% axial chord downstream of the rotor exit plane using a fast response total pressure
probe aligned to the absolute tip velocity vector. Time accurate, phase locked total
pressure was measured and averaged over 200 rotations. Total pressure defects due to
over tip leakage vortex, passage secondary flow, and blade wakes are clearly observed in
total pressure maps over the entire circumference and a range of radial locations. The
effect of coolant mass flow rate and injection location was investigated by coolant
injection from a single cooled blade with a gap height of 1.40% blade height. Coolant
mass flow rates in the range of 0.4% - 0.7% of turbine mass flow rate were investigated.
Coolant injection from all five cooled blades was also investigated. The effect of casing
endwall surface roughness was also studied, without coolant injection.
The sensitivity of total pressure defect, due to over tip leakage, to tip gap height is
reduced by both coolant injection and roughening of the casing surface. The total
pressure defect due to the large gap height of 1.40% blade height is reduced to levels
comparable to the defect due to a gap height of 0.72% blade height. The strong total
pressure gradient that characterizes the leakage vortex due to the gap height of 1.40%
blade height is considerably diminished by both coolant injection and roughening of the
casing surface. Coolant injection from 81% chord location is most effective in reducing
both the total pressure defect and the total pressure gradient. Casing surface roughness
significantly shifts the leakage vortex towards blade suction surface reducing its
interaction with the upper passage vortex. The benefit of casing surface roughness is
greater at larger gap heights.
v
Surface flows on the rotor and casing endwalls, blade tip surface for gap heights
ranging from 0.72% - 1.4% blade height, and blade tip surface with coolant injection at a
gap height of 1.4% blade height were visualized. Highly overturned flows are observed
on both endwalls, caused by secondary flow exiting the stator. Important rotor endwall
features such as the horseshoe vortex system, cross-passage secondary flow, and the path
of the pressure-side leg of the horseshoe vortex are identified. A distinct reattachment
line occurs on the tip surface of un-cooled blades, at approximately twice the gap height
from the pressure-side corner. Surface patterns clearly indicate the presence of a gap
vortex and chord-wise flow on the tip surface of un-cooled blades.
Coolant jets are turned by the gap flow, towards suction-side of the tip gap.
Cooling films form on the tip surface due to coolant jets from the first two locations
while intense mixing of coolant and main gap flow is observed around the 81% chord
location. The coolant jets prevent reattachment on the tip surface and also affect pressure-
side corner separation of the gap flow. Suction surface traces with coolant injection
indicate that roll up of gap flow in the neighboring passage is not continuous.
Coolant injection directed towards the blade pressure-side corner and artificially
introduced surface roughness could, individually, be effective techniques to operate high
pressure axial flow turbines at large tip gap heights without the associated penalty on
efficiency. The combination of coolant injection and artificially introduced surface
roughness could lead to greater loss reduction. The specific coolant scheme used has been
shown, qualitatively, to have thermal benefits in addition to reducing the total pressure
defect associated with the roll up of gap flow into a vortex.
vi
TABLE OF CONTENTS
LIST OF FIGURES ..................................................................................................... ix
LIST OF TABLES.......................................................................................................xvi
NOMENCLATURE ....................................................................................................xvii
ACKNOWLEDGEMENTS.........................................................................................xxi
Chapter 1 Introduction ................................................................................................1
1.1 Axial Turbine Blade Terminology..................................................................4 1.2 Turbine Rotor Passage Flow...........................................................................6 1.3 Over Tip Leakage (OTL)................................................................................7
1.3.1 Aerodynamic Losses ............................................................................11 1.3.2 Thermal Effects ....................................................................................13
1.4 Tip Desensitization .........................................................................................14 1.5 Film Cooling of Turbine Blade Tips ..............................................................17 1.6 Objectives of Current Research ......................................................................18
Chapter 2 Facility Description ....................................................................................20
2.1 The Axial Flow Turbine Research Facility (AFTRF) ....................................20 2.2 The Turbine Stage...........................................................................................23
2.2.1 Stage Aerodynamic Design ..................................................................26 2.2.2 AFTRF Tip Clearance Distribution......................................................29
2.3 Tip Cooled Blades ..........................................................................................31 2.4 Air Transfer System (ATS) ............................................................................36 2.5 Flow Visualization..........................................................................................36 2.6 Instrumentation ...............................................................................................38
2.6.1 Monitoring Instrumentation..................................................................39 2.6.2 Performance Measurement ...................................................................40
2.6.2.1 High-Frequency Total Pressure Probe .......................................41 2.6.3 Coolant Mass Flow Meter ....................................................................42 2.6.4 Rotating Frame Pressure Transducer....................................................43
2.7 Data Acquisition .............................................................................................44 2.8 Data Processing ..............................................................................................47 2.8 Operation ........................................................................................................49
Chapter 3 Rotor Flow Visualization ...........................................................................51
3.1 Rotor Hub Endwall Flow Visualization .........................................................53 3.1.1 Visualization with the Oil-Dot Technique ...........................................53 3.1.2 Visualization with the Oil-Film Technique..........................................57
vii
3.2 Blade Tip Surface Flow Visualization............................................................65 3.3 Rotor Casing Surface Flow.............................................................................72
3.3.1 Oil-Film Visualization of Rotor Casing Surface Flow.........................73
Chapter 4 Effect of Tip Gap Height on Over Tip Leakage.........................................79
4.1 Oil Film Based Tip Surface Flow Visualization.............................................79 4.1.1 Operation ..............................................................................................80 4.1.2 Tip Surface Flow Visualization; Large Gap Height (t/h = 1.4%) ........81
4.1.2.1 Blade Pressure Surface...............................................................81 4.1.2.2 Blade Tip Surface.......................................................................82 4.1.2.3 Heat Transfer Implications.........................................................87
4.1.3 Tip Surface Flow Visualization; Small Gap Height (t/h = 0.71%) .....88 4.1.4 Tip Surface Flow Visualization of Other Gap Heights ........................90 4.1.5 Suction Surface Traces from Oil Film Based Tip Surface Flow
Visualization...........................................................................................93 4.2 Total Pressure Measurements .........................................................................96
4.2.1 Baseline, No Injection ..........................................................................97 4.2.1.1 Region above 85% Blade Height ...............................................99 4.2.1.2 Region 75% - 85% Blade Height ...............................................100 4.2.1.3 Passage Core ..............................................................................100
4.2.2 Repeatability.........................................................................................102 4.2.3 Effect of the Tip Gap Height ................................................................104
Chapter 5 Effect of Coolant Mass Flow Rate on Over Tip Leakage ..........................112
5.1 Visualizing the Effect of Coolant Injection ....................................................113 5.1.1 Effect of Tip Trench .............................................................................114 5.1.2 Injection at Minj = 0.4% at Gap Height of t/h = 1.40% ........................116
5.1.2.1 Injection Prior to Start-up...........................................................116 5.1.2.2 Injection at Operating Speed ......................................................119
5.1.3 Visualization at Other Injection Rates..................................................122 5.1.4 Suction Surface Traces .........................................................................125 5.1.5 Heat Transfer Implication.....................................................................127
5.2 Total Pressure Measurement...........................................................................128 5.2.1 Comparison of Averaged Values..........................................................135
Chapter 6 Effect of Injection Location on Over Tip Leakage ....................................138
6.1 Injection from Individual Holes......................................................................139 6.2 Injection from Combination of Holes.............................................................146
Chapter 7 Multiple Cooled Blades and the Effect of Casing Surface Roughness ......154
7.1 Baseline...........................................................................................................155 7.2 Variation of Coolant Mass Flow Rate ............................................................157
viii
7.3 Effect of Casing Surface Roughness ..............................................................163 7.3.1 Smooth Casing Surface ........................................................................163 7.3.2 Fine Surface Roughness (220 Grit) ......................................................166 7.3.3 Coarse Surface Roughness (100 Grit) ..................................................168
7.4 Comparison of the Averaged Total Pressure Coefficient ...............................169
Chapter 8 Summary and Conclusions.........................................................................173
8.1 Summary.........................................................................................................174 8.1.1 Surface Flow Visualization ..................................................................174 8.1.2 Total Pressure Measurement ................................................................177
8.2 Conclusions.....................................................................................................181 8.3 Recommendations for Future Work ...............................................................183
Bibliography ................................................................................................................185
Appendix A Total Pressure Probe Characteristics......................................................190
A.1 Angular Sensitivity ........................................................................................190 A.2 Frequency Spectrum of Flow Field ...............................................................191 A.3 Uncertainty Analysis......................................................................................193
Appendix B AFTRF Tip Clearance Distribution........................................................196
Appendix C Total Pressure Coefficient Contour Map................................................200
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LIST OF FIGURES
Figure 1.1: Axial Turbine Rotor Passage.....................................................................5
Figure 1.2: Design Pressure Distribution on Turbine Blade........................................5
Figure 1.3: Turbine Rotor Passage Flow, From Yamamoto [4]. .................................9
Figure 1.4: Tip Gap Flow Conceptual Model From Bindon [7]..................................9
Figure 2.1: Schematic of the Axial Flow Turbine Research Facility. .........................21
Figure 2.2: Turbine Casing Frame For Measurement Windows. ................................21
Figure 2.3: AFTRF Instrumentation Window. ............................................................24
Figure 2.4: Blade Velocity Triangles...........................................................................29
Figure 2.5: AFTRF Tip Clearance Distribution...........................................................31
Figure 2.6: Cooled Blade Tip Surface. ........................................................................33
Figure 2.7: Cooled Blade Tip Coolant Injection Arrangement....................................34
Figure 2.8: Cross-Sectional View of Cooled Blade.....................................................35
Figure 2.9: Air Transfer System (ATS). ......................................................................37
Figure 2.10: Schematic of Monitoring Instrumentation. .............................................38
Figure 2.11: Schematic of Performance Instrumentation. ...........................................39
Figure 2.12: Total Pressure Probe Static Calibration...................................................42
Figure 2.13: Coolant Discharge Measurement Orifice Calibration. ............................44
Figure 2.14: Coolant Mass Flow (Orifice) Meter Calibration Minj vs. VDC...............45
Figure 3.1: Rotor Endwall Surface Flow Visualization Using the Oil-Dot Technique. ............................................................................................................55
Figure 3.2: Rotor Endwall Static Pressure Distribution, From Xiao [3]. ....................56
Figure 3.3: Oil Film On Rotor Endwall Before Test. ..................................................57
Figure 3.4: Near Leading Edge Surface Flow Features on the Rotor Endwall Visualized Using the Oil-Film Technique............................................................59
x
Figure 3.5: Blade Passage Surface Flow Patterns on the Rotor Endwall Visualized Using the Oil-Film Technique. .............................................................................61
Figure 3.6: Near Trailing Edge Surface Flow Patterns on the Rotor Endwall Visualized Using the Oil-Film Technique............................................................63
Figure 3.7: Blade Suction Surface Trace Formed During Rotor Endwall Surface Flow Visualization................................................................................................64
Figure 3.8: Oil Dots on Blade (B21) Pressure Surface Before and After Test Run. ...66
Figure 3.9: Tip Surface Flow Visualization (t/h = 1.40%) by Oil Dots Applied Near Blade Tip......................................................................................................67
Figure 3.10: Near Leading Edge Detail of Tip Surface Flow Visualization Using the Oil-Dot Technique (t/h = 1.40%)....................................................................69
Figure 3.11: Tip Surface Flow Visualization (t/h = 1.40%) Using the Oil-Dot Technique; Oil Applied Near Blade Root. ...........................................................70
Figure 3.12: Tip Surface Flow Visualization (t/h = 0.71%) Using the Oil-Dot Technique. ............................................................................................................71
Figure 3.13: Casing Surface Flow Visualization During Turbine Operation Using the Oil-Dots Technique.........................................................................................75
Figure 3.14: Rotor Casing Surface Flow Visualization Using the Oil-Dot Technique. ............................................................................................................76
Figure 3.15: Rotor Casing Surface Flow Visualization Using the Oil-Film Technique; Before and After Test. .......................................................................77
Figure 3.16: Rotor Casing Endwall Surface Flow Visualization Using the Oil-Film Technique.....................................................................................................78
Figure 4.1: Oil Film on Blade Pressure Surface Before and After Test. .....................82
Figure 4.2: Surface Flow Patterns on Tip Surface of Blade (B21) With a Gap Height of t/h = 1.40%, Visualized Using the Oil-Film Technique.......................86
Figure 4.3: Surface Flow Patterns on Front Half of Tip Surface of Blade B21, Using the Oil-Film Technique. .............................................................................86
Figure 4.4: Surface Flow Patterns on Rear Half of Tip Surface of Blade B21............87
Figure 4.5: Surface Flow Patterns on Tip Surface of Blade (B7) With a Gap Height of t/h = 0.71%, Visualized Using the Oil-Film Technique.......................90
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Figure 4.6: Surface Flow Patterns on the Tip Surface of Blade (B2) With a Gap Height of t/h = 0.81%, Visualized Using the Oil-Film Technique.......................92
Figure 4.7: Surface Flow Patterns on Tip Surface of Blade (B21 at reduced gap height) With a Gap Height of t/h = 1.2%, Visualized Using the Oil-Film Technique. ............................................................................................................92
Figure 4.8: Influence of Gap Height on the Location of the Visualized Reattachment Line on the Blade Tip Surface. ......................................................93
Figure 4.9: Suction Surface Traces Formed During Oil-Film Based Tip Surface Flow Visualization; Gap Heights, t/h = 1.4% and t/h = 1.2%. .............................95
Figure 4.10: Suction Surface Traces Formed During Oil-Film Based Tip Surface Flow Visualization; Gap Heights, t/h = 0.81% and t/h = 0.72%. .........................95
Figure 4.11: Effect of the Tip Gap Height on Measurements From Suction Surface Trace Formed During Oil-Film Based Visualization of Tip Surface Flow. .....................................................................................................................96
Figure 4.12: Total Pressure Coefficient Contours With No Coolant Injection (Base1)..................................................................................................................98
Figure 4.13: Secondary Flow Vectors At Rotor Exit From LDA Measurements by Ristic et al [54]......................................................................................................101
Figure 4.14: Total Pressure Coefficient Contours With No Coolant Injection (Base3)..................................................................................................................103
Figure 4.15: Radial Distribution of the Rotor Averaged Total Pressure Coefficient; Baseline Repeatability. .....................................................................104
Figure 4.16: Repeatability of Wake Profiles at r = 0.96h With No Coolant Injection. ...............................................................................................................107
Figure 4.17: Repeatability of Wake Profiles at r = 0.57h With No Coolant Injection. ...............................................................................................................107
Figure 4.18: Repeatability of the Passage Averaged Coefficient For Cooled Blade B21........................................................................................................................108
Figure 4.19: Total Pressure Coefficient Contours With Tip Gap Height of Cooled Blade B21 Reduced to t/h = 0.72%. .....................................................................108
Figure 4.20: Effect of Reducing the Tip Gap Height of Blade B21 On the Wake Profile at r = 0.96h. ...............................................................................................109
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Figure 4.21: Effect of Reducing the Tip Gap Height of Blade B21 On the Wake Profile at r = 0.57h. ...............................................................................................109
Figure 4.22: Effect of the Tip Gap Height On the Passage Averaged Coefficient of Cooled Blade B21.................................................................................................110
Figure 4.23: A Comparison of the Passage Averaged Coefficient Distribution For Blade B7 and Blade B21.......................................................................................110
Figure 4.24: Variation in the Area Averaged Total Pressure Coefficient with Tip Gap Height. (Area = 20% span*1 passage). .........................................................111
Figure 5.1: Surface Flow Visualization of the Effect of Tip Trench on Cooled Blade B21. ............................................................................................................115
Figure 5.2: Surface Flow Visualization of the Effect of Coolant Injection at Minj = 0.4% From Cooled Blade B21; Injection While Turbine at Rest. ........................117
Figure 5.3: Surface Flow Visualization of the Effect of Coolant Injection at Minj = 0.4% From Cooled Blade B21; Injection While Turbine at Operating Speed.....................................................................................................................120
Figure 5.4: Surface Flow Visualization of the Effect of Coolant Injection at Minj = 0.5% From Cooled Blade B21; Injection While Turbine at Rest......................123
Figure 5.5: Surface Flow Visualization of the Effect of Coolant Injection at Minj = 0.6% From Cooled Blade B21; Injection While Turbine at Rest......................124
Figure 5.6: Surface Flow Visualization of the Effect of Coolant Injection at Minj = 0.7% From Cooled Blade B21; Injection While Turbine at Rest......................124
Figure 5.7: Suction Surface Traces for Minj = 0.4% and Minj = 0.5%. ........................126
Figure 5.8: Suction Surface Traces for Minj = 0.6% and Minj = 0.7%. ........................127
Figure 5.9: Total Pressure Coefficient Contours with Coolant Injection at Minj = 0.41%. ...................................................................................................................129
Figure 5.10: Wake Profile at r = 0.96h, Without and With Coolant Injection at Minj = 0.41% and Minj = 0.52%.............................................................................130
Figure 5.11: Total Pressure Coefficient Contours With Coolant Injection at Minj = 0.52%. ...................................................................................................................131
Figure 5.12: Total Pressure Coefficient Contours With Coolant Injection at Minj = 0.63%. ...................................................................................................................132
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Figure 5.13: Wake Profile at r = 0.96h, Without and With Coolant Injection at Minj = 0.63% and Minj = 0.72%.............................................................................133
Figure 5.14: Total Pressure Coefficient Contours With Coolant Injection at Minj = 0.72%. ...................................................................................................................134
Figure 5.15: Wake Profile at r = 0.57h, Without and With Coolant Injection.............134
Figure 5.16: Effect of Coolant Injection On the Passage Averaged Coefficient of Cooled Blade B21.................................................................................................136
Figure 5.17: Area Averaged Coefficient For Blade B21 With Coolant Injection at a Tip Gap Height of t/h = 1.40%. .........................................................................137
Figure 6.1: Total Pressure Coefficient Contours for Coolant Injection From Blade B21 and Location H1 at 61% Cax. ........................................................................140
Figure 6.2: Total Pressure Coefficient Contours for Coolant Injection From Blade B21 and Location H2 at 71% Cax. ........................................................................141
Figure 6.3: Total Pressure Coefficient Contours for Coolant Injection From Blade B21 and Location H3 at 81% Cax. ........................................................................142
Figure 6.4: Total Pressure Coefficient Contours for Coolant Injection From Blade B21 and Location H4 at 91% Cax. ........................................................................142
Figure 6.5: Effect of Injection Hole Location on the Passage Averaged Coefficient of Cooled Blade B21. ...........................................................................................145
Figure 6.6: Effect of Injection Location on the Wake Profile at r = 0.96h..................145
Figure 6.7: Total Pressure Coefficient Contours for Coolant Injection From Blade B21 and Locations H1+H2. ..................................................................................147
Figure 6.8: Total Pressure Coefficient Contours for Coolant Injection From Blade B21 and Locations H1+H3. ..................................................................................148
Figure 6.9: Total Pressure Coefficient Contours for Coolant Injection From Blade B21 and Locations H2+H3. ..................................................................................148
Figure 6.10: Total Pressure Coefficient Contours for Coolant Injection From Blade B21 and Locations H1+H2+H3..................................................................149
Figure 6.11: Total Pressure Coefficient Contours for Coolant Injection From Blade B21; Full Injection......................................................................................150
xiv
Figure 6.12: Effect of Injection Location Combinations on the Passage Averaged Coefficient of Cooled Blade B21. ........................................................................151
Figure 6.13: Rotor Averaged Coefficient With Combined Injection. .........................152
Figure 6.14: Effect of Injection Location Combinations on the Wake Profile at r = 0.96h. ....................................................................................................................153
Figure 7.1: Total Pressure Coefficient With Multiple Cooled Blades; No Coolant Injection (Baseline), Minj = 0.0%..........................................................................156
Figure 7.2: Wake Profile at r = 0.96h Comparing Baseline Distributions of Multiple Cooled Blade and Single Cooled Blade (B21). .....................................157
Figure 7.3: Total Pressure Coefficient With Multiple Cooled Blades; Minj = 0.43%..158
Figure 7.4: Total Pressure Coefficient With Multiple Cooled Blades; Minj = 0.62%. ...................................................................................................................160
Figure 7.5: Total Pressure Coefficient With Multiple Cooled Blades; Minj = 0.72%. ...................................................................................................................160
Figure 7.6: Wake Profiles at r = 0.96h for Multiple Blade Coolant Injection. ............162
Figure 7.7: Passage Averaged Coefficient Comparison for Multiple Blade Coolant Injection. ...............................................................................................................162
Figure 7.8: Total Pressure Coefficient With a Smooth Plastic Layer On the Casing Inner Surface.........................................................................................................164
Figure 7.9: Wake Profiles at r = 0.96h Comparing the Influence of Casing Surface Roughness.............................................................................................................165
Figure 7.10: Wake Profiles at r = 0.57h Comparing the Influence of Casing Surface Roughness................................................................................................166
Figure 7.11: Total Pressure Coefficient With Fine Sandpaper (220 Grit) On the Casing Inner Surface.............................................................................................168
Figure 7.12: Total Pressure Coefficient With Coarse Sandpaper (100 Grit) On the Casing Inner Surface.............................................................................................170
Figure 7.13: Passage Averaged Coefficient For Blade B21 With Different Casing Roughness Treatments..........................................................................................171
Figure 7.14: Rotor Averaged Coefficient With Different Casing Roughness Treatments. ...........................................................................................................172
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Figure A.1: Probe Response to Incidence. (Squares denote rotor averaged Cpt and circles denote passage averaged Cpt.). .................................................................192
Figure A.2: Frequency Spectrum of Rotor Exit Flow Near Rotor Tip. .......................193
Figure B.1: Clearance Gap Variation Along Blade Axial Chord Length For TCL1...199
Figure C.1: Total Pressure Contour Map Of the Entire Rotor Exit Flow Field In the Measurement Plane.........................................................................................201
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LIST OF TABLES
Table 2-1: AFTRF Overall Stage Design Characteristics............................................25
Table 2-2: Design Thermodynamic Parameters. .........................................................26
Table 2-3: AFTRF Design Coefficients.......................................................................26
Table 2-4: AFTRF Nozzle Design...............................................................................27
Table 2-5: AFTRF Rotor Design. ................................................................................28
Table 2-6: Radial Variation of Turbine Stage Design Parameters From 1-D Mean-line Analysis. ........................................................................................................28
Table 2-7: Estimate of Coolant Discharge Mass Flow Rate, From Pudupatty [43]. ..34
Table 2-8: Some Calculated Parameters From Low Speed DAS. ...............................46
Table 5-1: Test Matrix of Flow Visualization with Coolant Injection. .......................114
Table 6-1: Test Matrix for Effect of Injection Location. .............................................139
Table A-1: Uncertainty and Nominal Values in Measured Parameters.......................195
Table A-2: Uncertainty in Derived Parameters ...........................................................195
Table B-1: Gap Height Variation Along Blade Axial Chord For Clearance Distribution TCL1 ................................................................................................197
Table B-2: Gap Height Variation Along Blade Axial Chord For Clearance Distribution TCL3 ................................................................................................198
xvii
NOMENCLATURE
Cax Blade tip axial chord length, m.
Cp Specific heat at constant pressure, kJ/kg. K
Cps Static pressure coefficient.
Cpt Total pressure coefficient.
Cpt, R Rotor averaged total pressure coefficient.
Cpt,, A Area averaged total pressure coefficient.
Cpt,P Passage averaged total pressure coefficient.
d Differential operator.
h Blade height, m.
h0 Total or stagnation enthalpy, kJ/kg.
h1 Distance of upper boundary of suction surface oil trace from tip surface,
m.
h2 Width of suction surface oil trace, m.
i Circumferential index.
j Radial index.
lc Distance of reattachment line measured from pressure-side corner,
perpendicular to camber-line.
lx Distance of reattachment line measured from pressure-side corner,
perpendicular to blade axial chord.
Minj Coolant injection mass flow rate, [-]
N Rotor speed, rpm.
xviii
p local static pressure, Pa
P0, p0 Total or stagnation pressure, Pa.
qm Mean wheel speed based dynamic pressure, Pa.
r Radius, m.
r,θ,x Cylindrical coordinates.
T Temperature, K.
t Tip gap height, m.
T0 Total or stagnation temperature, K.
U Blade speed, m/sec.
V Absolute velocity vector, m/sec
v Specific volume, m3/kg.
Vax Axial component of absolute velocity vector, m/sec.
Vt Tangential component of absolute velocity vector, m/sec.
W Relative velocity vector, m/sec
x,y,z Cartesian coordinates.
GREEK SYMBOLS
∅ Diameter, m.
ψ Loading coefficient.
α Absolute flow angle.
β Relative flow angle.
φ Flow coefficient.
xix
γ Ratio of specific heats.
ηtt Total-Total efficiency.
ρ Density, kg/m3.
SUBSCRIPTS
1 Stage inlet.
2 Nozzle exit / rotor inlet
3 Rotor exit.
amb Ambient (pressure or temperature).
hub Rotor hub.
m mid-span
max Maximum.
tip Rotor tip.
ACRONYMS
B(#) Blade number.
C.L. Centerline
EES Engine equivalent speed.
EGV Exit guide vane.
H(#) Injection hole number.
HP High pressure.
Hp Pressure-side leg of horseshoe vortex.
Hs Suction-side leg of horseshoe vortex.
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L.E. Blade leading edge.
LP Low pressure.
N Nozzle.
NGV Nozzle guide vane.
OTL Over tip leakage.
PS Blade pressure surface or pressure-side
R Rotor.
SS Blade suction surface or suction-side.
T.E. Blade trailing edge.
TCL Tip clearance.
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ACKNOWLEDGEMENTS
I would like to thank Dr. Cengiz Camci for his guidance and encouragement in
helping me achieve my academic goals. I would also like to thank Dr. McLaughlin, Dr.
Long, Dr. Yavuzkurt, and Dr. Miller for serving on my doctoral committee and for
providing me with valuable suggestions.
This achievement was also made possible by the financial support of the
Department of Aerospace Engineering, through teaching assistantships. I would also like
to thank the staff of the Department of Aerospace Engineering for their cheerful
disposition, patience in answering my many questions, and encouragement.
I would particularly like to thank Mark Catalano and Rick Auhl for their
assistance in overcoming the technical difficulties. The maintenance and modifications of
the research facility were promptly addressed by Harry Houtz. In addition to this, I
appreciate his interest in my research and his willingness to share his experience with me.
I would also like to acknowledge J. D. Miller for providing me with the necessary
resources and for assisting me in overcoming technical difficulties.
My family has been a constant source of inspiration and determination, both of
which were needed in attaining this goal. Thank you for your faith in my abilities. I am
also extremely grateful to be able to share this with numerous friends.
1
Chapter 1
Introduction
Transformation of the energy potential of fuels into useful forms of energy has
been an important aspect of civilization ever since the Industrial Revolution. The
invention of the steam engine provided great impetus to the Industrial Revolution as it
enabled the harnessing of thermal energy for mechanical drive applications. Similarly,
the gas turbine engine greatly accelerated developments in aviation. In addition to
propulsion of large and medium aircrafts, gas turbine engines are also used extensively in
electrical power generation and in marine applications. Thus improving the performance
of the gas turbine engine has great economic and environmental value.
Gas turbine engines belong to a class of rotating, energy conversion systems
called turbomachines. The energy transfer in turbomachines is achieved by a change in
the angular momentum of the working fluid, in a rotating blade row termed the rotor.
Energy is extracted from the working fluid in turbines, while energy is transferred to the
fluid in compressors and pumps. Turbine engine flows are contained within a duct or
engine casing. A gap is therefore necessary between the rotor and the stationary casing.
This gap is responsible for the phenomenon of Over Tip Leakage (OTL), which in turn
deteriorates the aerodynamic and thermal performance of the axial flow turbine
The necessary gap between the turbine rotor and stationary casing is referred to as
the tip gap or the tip clearance. In multi-stage turbines it is typical to have the blade tips
of low pressure (LP) stages covered by a shroud and the tip gap is formed between the
2
shroud outer surface and the casing. The shroud constrains the blades against large
displacements due to vibration, as the LP turbine blades tend to be longer than those in
high pressure (HP) stages. It is possible to shroud the LP turbine due to the relatively
lower gas temperature and rotational speed. On the other hand, HP turbines form the front
end of multi-stage turbines and extract energy from hot gas exiting the combustor, where
gas temperatures can be higher than metal melting point temperatures. Effective cooling
is necessary for proper functioning of the HP turbine stages and a shrouded turbine rotor
would necessitate cooling of the shroud. Cooling air, which is extracted from the engine
compressor section, is accounted for as a loss of cycle work. The addition of metal at the
maximum radius might also constrain the design performance due to increased blade root
stress and differential thermal growth during engine transients. Hence, shrouded HP
turbine designs are uncommon. The space between the blade tip profile and stationary
casing in HP turbine rotors is referred to as the tip gap and flow through this gap is
termed over tip leakage (OTL). In multi-stage turbines the annulus area also changes
significantly along the length of the turbine section. The annulus height of HP turbines is
small and a gap of the same physical dimension accounts for a larger part of the annulus
area. Thus, the influence of over tip leakage on the passage core flow is relatively
stronger in HP stages.
The total-total stage efficiency (ηtt) is defined in Equation 1-1. The stage
temperature and pressure are denoted by T and p respectively, while the subscripts “01”
and “03” refer to the stage inlet and exit conditions respectively. Efficiency increases
with either an increase in inlet temperature or a decrease in exit temperature. Efficiency
drop due to over tip leakage results from the exit total temperature (T03) of the fluid that
3
participates in over tip leakage remaining elevated, since energy is not fully extracted
from it. Another source of energy loss due to over tip leakage in transonic axial flow
turbines has been recently reported by Thorpe, et al [1]. They show, both analytically and
through measurement of casing temperature, that work is done on the fluid within the tip
gap of a transonic rotor, leading to increase in total temperature of the working fluid.
The stage inlet temperature (T01) may be limited by a number of factors including
thermal considerations and the net efficiency gain due to cooling. Over tip leakage causes
thermal deterioration of the blade tip surface and according to Bunker [2] might
necessitate a decrease in operating temperature (T01) over time, causing a further
reduction in efficiency. The effect of lowering the inlet temperature may also be
considered from the perspective of entropy generation. Gibbs’ equation as given by
Equation 1-2 shows that entropy generation is inversely proportional to the initial
temperature. Losses may be minimized by increasing the cycle maximum temperature.
Thus, improvements in efficiency are limited by the maximum permissible operating
temperature.
⎟⎟⎟
⎠
⎞
⎜⎜⎜
⎝
⎛ −
⎟⎟⎠
⎞⎜⎜⎝
⎛−
⎟⎟⎠
⎞⎜⎜⎝
⎛−
=
γ
γη
1
01
03
01
03
pp1
TT1
tt (1-1)
pTCsT vddd p −= (1-2)
4
The research reported in this thesis is specific to un-shrouded, high pressure
turbine rotors. This chapter summarizes the current understanding of the over tip leakage
process, the aerodynamic and thermal effects of over tip leakage, previous studies aimed
at minimizing the effects of over tip leakage, and the objective and organization of this
thesis.
1.1 Axial Turbine Blade Terminology
A typical turbine blade passage is shown in Figure 1.1. The blade profiles shown
form the tip section of the turbine blades in the rotating turbine research rig at The
Pennsylvania State University. The direction of rotation is from bottom to top, as shown
by the blade speed vector U and W denotes inlet velocity in the relative frame of
reference. The blade passage, highlighted by a hatched pattern, is bounded by pressure
surface (PS) on the top and suction surface (SS) at the bottom. A pressure distribution is
generated on each blade surface, as shown in Figure 1.2, as the flow is turned in the blade
passage. The design blade surface pressure distribution shown was obtained from Xiao
[3]. The pressure is shown in terms of a pressure coefficient, as defined by Equation 1-3,
where 1p is the passage averaged local inlet pressure and Um is the wheel speed at mean
radius. Pressure driven flow, from blade pressure-side to blade suction-side due to the
pressure difference across the gap is referred to as over tip leakage or tip leakage flow.
21
5.0 mps U
ppCρ−
= (1-3)
5
Figure 1.1: Axial Turbine Rotor Passage.
-2.00
-1.50
-1.00
-0.50
0.00
0.50
0.00 0.20 0.40 0.60 0.80 1.00 1.20
x/Cax
Cps
SSPS
LE TE
Figure 1.2: Design Pressure Distribution on Turbine Blade.
6
1.2 Turbine Rotor Passage Flow
Turbine rotor passage flows are highly three dimensional and a brief description is
given with reference to Figure 1.3. The schematic of passage flow, from Yamamoto [4],
summarizes the observations of several researchers. Since the function of the passage is
to guide and turn the fluid, the primary flow follows streamlines defined by the blade
surfaces bounding the passage. However, viscous effects, streamline curvature, and
presence of strong pressure gradients generate significant recirculatory flows that are
generally termed as secondary flows.
The inlet boundary layer (1) stagnates as it approaches the leading edge and rolls
up into a horseshoe vortex (2). The two legs of the horseshoe vortex, pressure-side leg (3)
and suction-side leg (4) eventually merge near the blade suction surface and may become
part of a larger secondary flow structure, the rotor endwall or hub passage vortex (9). The
passage vortex is a result of pressure-driven cross-passage flow in the rotor endwall
boundary layer (8). Similarly, there exists a casing endwall or tip passage vortex. Over tip
leakage is also shown in Figure 1.3. The tip leakage vortex (12) and the tip passage
vortex rotate in opposite directions. The interaction between these vortices generates
limiting streamlines on the blade suction surface. In general, the vortices transport fluid
with high momentum and temperature into the boundary layers and deposit the low
momentum boundary layer fluid into the mainstream. The rotational kinetic energy is
dissipated without the extraction of useful work. The interaction between vortices also
leads to increased turbulence and mixing.
7
1.3 Over Tip Leakage (OTL)
The process of over tip leakage in actual engines is very complex and is affected
by compressibility, shocks, rotational effects, and the influence of the stationary casing.
The physics and effects of OTL flow have been studied primarily in cascades and for
incompressible flows. Results from rotating rig experiments are limited and as such no
data exists for flow within the tip gap or near tip surface flow.
Tip leakage flow methodology was investigated by Booth, et al [5] in different
water flow rigs. This study experimentally validated Rains’ [6] hypothesis that leakage
flow rate may be predicted based on an inertial balance. Booth, et al [5] confirmed that
pressure driven flow occurred normal to the gap while transverse velocity was preserved
across the gap. Gap discharge coefficient was found to be independent of gap height, inlet
boundary layer, and cross-flow velocity when the gap discharge coefficient was defined
using the static pressure across the gap. The gap mass flow rate was quantified by a
velocity calculated from Bernoulli’s equation used in conjunction with a discharge
coefficient.
The first detailed measurement of OTL flow within the tip gap, reported by
Bindon [7], conducted in a linear cascade of turbine blades showed that the gap flow
constitutes of fluid originating from two sources, as shown in Figure 1.4. The inlet
boundary layer passes through the tip gap in the front part of the blade, while pressure
driven flow commences at approximately 30% axial chord length. The pressure driven
flow separates at the pressure-side corner of the blade, reattaches on the tip surface and
flows towards suction-side of the tip gap. A gap vortex, caused by the separation, exists
8
in the tip gap. A favorable pressure gradient within the separated region generates a
chord-wise flow on the tip surface, towards the blade trailing edge. Low momentum fluid
within the gap vortex is entrained by the main gap flow around the mid-chord region of
the blade tip surface, where the driving pressure difference reaches a maximum. In
addition to the separation vortex the wall static pressure on the tip surface, near the
pressure-side corner was shown to be significantly lower than both the suction-side
pressure and the cascade exit pressure. This low pressure was attributed to the contraction
effect of the separation bubble in the tip gap. The gap flow diffuses in the tip gap,
generating a wake near the tip surface. Thus, the gap flow is characterized by a wake
region (d) near the blade tip surface and a low loss, jet (c) in the remainder of the gap.
Velocity measurements at different heights within the tip gap were conducted by
Yaras et al [8] in a linear turbine cascade. Bulk of the gap flow was found to occur above
the gap separation vortex. The change in velocity from gap inlet to gap exit was measured
to be small, suggesting that flow was fully accelerated at gap inlet. Close to the tip
surface the velocity vectors were observed to change rapidly and a chord-wise flow was
detected within the separation vortex. The centerline of the gap separation vortex was
located at about 20% gap height. The presence of a chord-wise flow and gap vortex are
also supported by surface flow visualization by Sjolander and Cao [9]. The visualization,
conducted in an idealized tip gap, showed a highly organized flow in the near pressure-
side corner of the tip surface, where flow separates due to the sharp corner of the tip gap.
9
Figure 1.3: Turbine Rotor Passage Flow, From Yamamoto [4].
Figure 1.4: Tip Gap Flow Conceptual Model From Bindon [7].
10
The linear cascade investigations while providing detailed understanding of the
gap flow do not capture the effect of casing relative motion. This motion is in a direction
opposite to the leakage flow and its effect would depend on the gap height. The effect of
casing motion was investigated in a linear cascade by Yaras and Sjolander [10]. Casing
motion at 100% engine equivalent speed (EES) was found to achieve a global reduction
in gap velocity. This was attributed to a reduced pressure effect, caused by an enhanced
passage vortex. Increasing casing relative speed also confined the gap vortex closer to the
tip, primarily due to the reduction in gap velocity. Correspondingly, blade loading near
the tip was shown to be reduced by Yaras, et al [11]. Flow downstream of the blades
displayed greater vorticity due to the passage vortex and lower vorticity due to the tip
leakage vortex. The increased vorticity of the passage vortex was attributed to the relative
casing motion, while decrease in leakage flow vorticity was attributed to the reduction in
gap mass flow rate. Water model cascade tests with a moving belt by Graham [12]
confirmed that the gap mass flow rate decreased with both a decrease in gap height and
an increase in casing relative speed. A cut-off speed was also determined, at which speed
the gap flow was completely eliminated at gap height of 1%. Morphis and Bindon [13]
investigated the effect of relative casing motion in an annular cascade and showed that
the low pressure at the PS corner was not affected by the influence of the casing.
However, higher pressure was measured in the reattachment region behind the bubble
and was attributed to a reduction in leakage flow.
While the effect of the relative casing motion may be simulated in cascades by
incorporating moving belts, the effect of centrifugal and coriolis forces on the passage
flow is not accounted for in such studies. Yamamoto [14] compared the leakage vortex
11
location in a stationary cascade with that in a rotor and found that the leakage vortex was
confined to the tip region of the passage, possibly due to centrifugal effects in rotors. In
stationary cascades the vortex responded to the span-wise pressure gradient and moved
radially inwards towards the hub. Kaiser and Bindon [15] noted that the energy
associated with the over tip leakage fluid in rotating rigs is higher than that found in
cascades, since energy extraction from over tip leakage flow is incomplete. Velocity
measurements in the rotating frame by McCarter [16], in the large scale, low speed axial
turbine at Penn State showed the path of the leakage vortex along the blade passage was
not as steep as that measured in cascades. The losses due to the tip leakage vortex also
increased further downstream of the rotor. Thus, it is important for tip desensitization
investigations to be conducted in a rotating environment.
1.3.1 Aerodynamic Losses
Losses in a turbine stage result from different mechanisms and those occurring in
vane and blade passages may be broadly split into profile losses, secondary flow and tip
leakage losses, and losses due to shocks. Over tip leakage flow is a significant contributor
to losses in turbine stages. Waterman [17] and Booth [18] suggest that a third of the stage
losses may be attributed to tip leakage flows. Considering losses in the rotor alone,
Schaub, et al [19] indicate that OTL may be responsible for as much as 45% of loss
within the rotor passage. Most HP rotors are designed to operate with gap heights of 1% -
2% blade height, which makes the losses associated with over tip leakage
disproportionately large.
12
Loss measurements by Bindon [7] in a linear cascade indicate that total pressure
loss due to the clearance gap vary linearly with gap size, with the lowest gap size
investigated being 0.3%. The total loss due to the tip gap was also separated into internal
gap loss (39%), suction corner mixing loss (48%), and endwall/secondary loss (13%).
The internal gap loss is attributed to the low momentum fluid within the separation
bubble. Mixing losses were seen to arise only over the last 20% of axial chord. McCarter
et al [20] measured the total pressure loss in a large scale, low-speed rotating rig
(AFTRF) at Penn State. They showed that the leakage vortex strengthens dramatically at
around 80% axial chord, causing mixing losses due to interaction between leakage and
passage flows, to rise. Losses due to leakage vortex were found to be about 25% higher
than those caused by the passage vortex.
The loss generation also depends on the profile thickness. Booth et al [5] found
that performance dropped as blade tip profile thickness was reduced while maintaining
loading distribution. A similar result was obtained by Sjolander & Cao [9] who showed
that the mass averaged losses increased as the blade tip profile thickness decreased in
relation to the gap height. Graham [12] found that flow in the gap over thin tipped blades
showed little effect of the moving endwall, thereby concluding that the pressure-driven
flow was essentially unaffected by surface friction.
The interaction between the passage vortex and the leakage vortex in linear
cascades was found to increase, with increasing blade turning and increasing gap height
by Yamamoto [4]. The vorticity from the passage vortex was observed to decrease by
McCarter, in the region that the leakage vortex vorticity increased. The interaction
between the two vortices was found to occur at and beyond the rotor exit plane. An
13
important conclusion was that secondary flows in the rotating rig dissipated slower than
the secondary flows in cascades. Additionally, the tip leakage vortex dissipated slower
than the upper passage vortex.
1.3.2 Thermal Effects
A discussion on high-pressure turbine blade over tip leakage flow is incomplete
without the inclusion of the thermal effects of OTL. Even though this thesis does not
report any quantitative results from heat transfer measurements on turbine blade tips a
brief description of some of the public domain knowledge is provided to complete the
picture. As noted by Bunker [2] blade tips are exposed to hot gases on all sides and the
tip surface experiences some of the highest and lowest heat transfer coefficients leading
to severe thermal gradients. Tip clearance increases with operational hours due to casing
rub and oxidation of the tip surface. This leads to decreased efficiency. One of the most
common effects of OTL is the burn-out of the pressure-side corner that has been linked to
the chord-wise flow on the tip surface by Bindon [21], and Sjolander and Cao [9]. This is
one of the reasons limiting cycle maximum temperature and in turn the efficiency of the
work extraction process in turbine blade passages. Heat transfer coefficients measured by
Bunker et al [22] on the tip surface of a blade with sharp edges show the extreme
variation in heat transfer coefficient over the tip surface. Bunker [23] also notes that the
leakage vortex that forms near the suction surface and its interaction with the tip passage
vortex is responsible for increased oxidation of the suction surface.
14
1.4 Tip Desensitization
The clearance between the blade tips of a high pressure, un-shrouded, axial gas
turbine rotor and the stationary casing has been shown, in the previous sections, to be a
source of efficiency loss and a limiting factor to improving efficiency through increased
cycle maximum temperature. The gap height is also subject to growth due to both tip
surface material oxidation and tip surface rubbing into the casing during engine thermal
transients. The efficiency drop as a function of tip gap height is linear, as shown in
Bunker. More importantly the slope of the line is greater for smaller engines and high
pressure turbine stages due to the small annulus areas. Thus, tip desensitization may be
defined as reducing the sensitivity of turbine performance to the effects of the tip gap
height.
Most of the tip desensitization techniques reported involve modification of the tip
surface geometry in an effort to decrease the leakage mass flow rate, by decreasing the
discharge coefficient. A comprehensive study of various tip geometries by Booth et al [5]
included, among other geometries, tip surface extensions (winglets), squealer tips, and
double squealer or grooved tips. The single and double squealer configurations were
found to give the best reduction in discharge coefficient. However, their performance was
found to be very sensitive to geometry and gap Reynolds number. Winglets were found
to generate greater improvement due to both a reduction in discharge coefficient and a
reduced pressure drop across the tip gap surface. Bindon [24] visualized flow over
squealer tips using smoke in a linear cascade and found that at large gap heights the
leakage flow essentially passed over the tip surface as if the surface were flat. Linear
15
cascade results of Heyes et al [25] show that suction-side squealers are better than
pressure-side squealers.
The heat transfer on tip surfaces with squealer cavities has also been studied both
experimentally, in cascades, and through numerical simulation. Bunker and Bailey [26]
concluded that squealer tips decrease the heat load on the tip surface. Cavity surface heat
transfer coefficients decreased with increase in cavity depth and the distribution of heat
transfer coefficients was also more uniform. The overall heat load on the cavity wall also
decreased with increase in cavity depth. Numerical simulations by Acharya et al [27] also
indicate that the aerodynamic benefits of double squealer tips decrease with increase in
gap height. The reductions in leakage mass flow rate and tip heat transfer coefficients
obtained over flat tip geometry grew smaller as the gap height of double squealer tips was
increased. The presence of a squealer cavity decreased the suction surface heat transfer
coefficients due to modification in the leakage vortex strength. Papa et al [28] found that
for a blade tip surface with squealers the starting point of the leakage flow moved
rearwards with increasing gap height.
Morphis and Bindon [13] experimented with pressure-side corner rounding in an
effort to reduce separation related losses in the tip gap. In the 1% - 2% gap height range
the total-total stage efficiency of a single stage was found to improve with rounding-off.
The efficiency of the second stator row was found to improve with rounding-off of the
pressure-side corner of the first stage blades. Numerical simulation by Ameri [29]
showed that leakage mass flow rate increased by about 25% when the pressure-side
corner was rounded-off, due to an increase in the gap discharge coefficient. The addition
of a camber-line sealing strip decreased the leakage mass flow rate, however the mass
16
flow rate was still greater than that through the gap over a sharp edged, flat tip blade. The
reduction in total pressure loss caused by a camber-line strip on the tip surface with
pressure-side corner rounding-off was not proportional to the reduction observed in gap
mass flow rate. The tip surface heat transfer was found to increase due to rounding-off of
the pressure-side edge, by Bunker et al [22].
Research at Penn State has consisted of both experimental and computational
investigations of tip desensitization. The effect of various tip geometries was studied in
the low-speed rotating rig at Penn State by Dey [30]. The tip geometries studied included
pressure-side and suction-side extensions (winglets), and single and double squealers.
The axial coverage of these tip modifications was also investigated. The height of the tip
gap was found to be an important parameter and the effectiveness of squealers degraded
quickly as gap height was increased. The first investigation of coolant injection from a
rotating rig was done by Dey and Camci [31]. No appreciable change in rotor exit total
pressure was measured and this was attributed to the coolant injection holes being too
small. Computational investigation of tip desensitization through tip surface chamfering
by Tallman and Lakshminarayana [32] indicate that OTL losses may be reduced due to
turning of the leakage flow towards the camber-line.
Of the desensitization techniques reviewed only the double squealer tip finds
actual application in service. Double squealer rims offer the advantages of minimizing
contact surface area and partly shielding cavity surface from thermal degradation. The
other techniques, except tip coolant injection from flat tips, present added cooling
requirements that may not be easily addressed. In a comprehensive review of various
desensitization methods Harvey [33] notes that attempts to reduce the discharge
17
coefficient have proven contradictory. In fact, the use of shrouds is pointed out as
advantageous from the point of view of efficiency retention during service. However
turbine shrouds, in addition to requiring cooling, increase tip mass, which can introduce
mechanical complications. Furthermore, the desire to increase firing temperatures might
at some stage make shrouds untenable in turbine designs, particularly HP turbines.
1.5 Film Cooling of Turbine Blade Tips
It is typical for modern high pressure turbine blades to eject coolant into the tip
gap from radial bores on the tip surface. This may be done either for cooling of the blade
tip surface, or for purging dirt to maintain the effectiveness of the internal cooling
system. More recently the effect of radial coolant injection on blade tip surface heat
transfer and tip gap aerodynamics has been studied.
Experimental investigation of a modeled blade tip is reported in Kim et al [34].
Kwak and Han [35] investigated film cooling of gas turbine blade tips in a linear cascade.
Blade tip heat transfer coefficients were found to increase with gap height. Film cooling
from radial holes along the camber line was shown to be more effective at larger
clearances and at higher blowing ratios. Near tip coolant injection from the pressure
surface was found to increase film effectiveness over the tip surface. Hohlfeld et al [36]
numerically simulated the film cooling effect of dirt purge holes on turbine blade tips.
Coolant ejected from the purge holes served to block the leakage flow at a tip clearance
of 0.54% blade height. However, as gap height was increased to 1.64% span the blocking
effect decreased. At the large clearance, shroud cooling effectiveness increased with
18
blowing ratio, while tip surface cooling effectiveness first decreased and then increased.
Acharya et al [27] also computationally simulated film cooling of turbine blade tips. Film
cooling at three different gap heights was simulated. Coolant injection was found to alter
the leakage vortex and also decrease the heat transfer coefficient along the coolant
trajectory. Film cooling effectiveness was found to increase slightly with gap size. Two-
dimensional computations of over tip leakage with coolant injection by Koschel et al [37]
show that the coolant discharge has a minor effect on the gap discharge mass flow rate.
The influence of wall motion appeared to be more pronounced with coolant injection.
The effect of coolant injection into the tip gap was simulated by Chen et al [38] using
two-dimensional Navier-Stokes equations. A single radial jet from a tip surface slot was
found to decrease the gap inlet mass flow rate. The discharge mass flow rate remained
more or less the same and was independent of the coolant mass flow rate. The effect of
the secondary jet was found to be sensitive to the position and width of the tip surface
slot.
1.6 Objectives of Current Research
The design of modern high pressure turbine blades is complex and as noted by
Bunker [2] is governed by various, diverse parameters, including aerodynamic and
thermal variables. The most common blade tip designs include un-shrouded blade tips,
with and without squealer rims, and a few shrouded designs. The effectiveness of
squealer rims with cavity cooling is to some extent controlled by the necessary hading of
turbine casing. Furthermore, cooling of the blade tip region is almost always necessary.
19
The primary objective of this experimental investigation is to study the effect of
tip coolant injection on OTL in a rotating environment. The configuration used in Dey
and Camci [31] is modified to allow for greater coolant mass flow rates. The study
consisted of both quantitative and qualitative measurements. Time accurate total pressure
measurements were conducted downstream of rotor exit using a high-frequency
transducer aligned with the absolute flow. The measurements without and with coolant
injection are analyzed for performance benefits. Surface flow visualization was also used
to qualitatively understand the interaction between injected coolant and gap flow, not
only from an aerodynamic view point but also from the thermal benefits that may be
present. Some of the results presented and discussed in this thesis may also be found in
Rao and Camci [39], [40], [41], and [42].
The organization of this thesis is as follows. Chapter 2 describes the facility,
aerodynamic design of the turbine stage, instrumentation, and operation of the facility.
Flow visualization on the rotor endwall, the casing endwall, and the blade tip surface is
reported in Chapter 3. The effect of tip gap height is the subject of Chapter 4, where the
results presented include total pressure measurements and surface flow patterns. Chapter
5 reports the effect of coolant injection from a single rotor blade, where the effect of
coolant mass flow rate on OTL is discussed. The effect of coolant injection location on
the measured total pressure defect due to over tip leakage is discussed in Chapter 6.
Chapter 7 discusses coolant injection from multiple blades and the effect of casing
surface roughness. A summary of the results, conclusions and future work
recommendations are presented in Chapter 8.
20
Chapter 2
Facility Description
The facility used in this investigation is the Axial Flow Turbine Research Facility
(AFTRF) at the Pennsylvania State University. A description of the facility, turbine stage
design, and instruments used is included in this chapter.
2.1 The Axial Flow Turbine Research Facility (AFTRF)
The AFTRF is a large scale, low speed, single stage, axial flow turbine facility. A
schematic of the facility is shown in Figure 2.1. The shaded regions denote rotating
components. The main components in the flow path of this facility are listed below in
order,
a) Inlet section,
b) Turbine stage,
c) Transition ducting with acoustic muffler,
d) Two-stage axial fan, not shown in the figure,
e) Exhaust diffuser, not shown in figure
The facility inlet is housed in a large stagnation chamber covered with foam cloth.
Air flows through a smooth bell-mouth section with an axi-symmetric center-body that
terminates at the stator. The centre-body houses slip-ring units for data transfer from
rotating frame to stationary frame.
21
Figure 2.1: Schematic of the Axial Flow Turbine Research Facility.
Figure 2.2: Turbine Casing Frame For Measurement Windows.
Rotating Hub
Stub Shaft
Bell-mouth Inlet
Centre-body
0.5715 m
0.1524 m
0.2985 m0.2794 m
0.1334 m
Total pressure probe
Split line between rotating & stationary hub Nozzle ring
C.L.
Rotating hub
22
The turbine stage consists of a stator row (N) that accelerates the flow into the
rotor (R). The turbine outer casing is provided with a rectangular access frame that
measures (0.5715 m x 0.2794 m), as shown in Figure 2.2. The frame is symmetric about
the horizontal plane and starts at 0.1524 m upstream of stator row. Windows for probe
based measurements, shown in Figure 2.3, or optical access are mounted in this frame.
The windows are fabricated so that the inner surface of the windows are curved to the
turbine casing outer diameter (∅max = 0.9166 m) and fits in flush. A row of exit guide
vanes (EGV’s), three chord lengths downstream of the rotor remove the swirl induced in
the flow by the rotor blades. Air then passes through the transition ducting, which is lined
with acoustic damping liner to reduce noise generation. Airflow through the facility is
induced by the suction provided by a four stage axial fan. The combined flow capacity of
the fans is 10 m3/sec and they generate a combined pressure rise of 74.7 mm Hg
(approximately 40” water column). The fan duct is covered externally by thermal lagging
material to ensure adiabatic boundary conditions in the fan duct. Air energized by the
fans is then exhausted to atmosphere through the diffuser, which has an external moving
endplate that provides a means to control mass flow rate.
A rotating instrumentation drum is mounted on to the turbine disk. Instruments,
such as pressure transducers, hot-wire anemometers, probe traverse, mass flow measuring
devices, etc, are mounted in the instrumentation drum. All instruments are wired to a slip-
ring unit on the main shaft. An outer drum that forms the rotating hub, as marked in
Figure 2.1, extends 133 mm downstream of the rotor exit plane and forms the inner
diameter (∅hub = 0.6706 m) of the rotor annulus. The rotor endwall (hub) surface extends
23
beyond this in the stationary frame, with a small gap provided at the split line with the
rotating hub as marked in Figure 2.2.
A stub shaft, with a pulley, is flange mounted to the downstream end of the
rotating instrumentation drum. Torque is transmitted from stub-shaft to a brake shaft by
the belt & pulley system. A water-cooled, eddy current brake absorbs power generated by
the turbine. The brake is capable of maintaining the turbine speed constant to within ±1
rpm. The brake shaft is equipped with an inline torque meter for torque measurement. A
BEI systems optical shaft encoder is mounted in the stationary frame aft of the pulley.
The encoder shaft is coupled to a stub shaft mounted to the turbine pulley.
2.2 The Turbine Stage
Overall characteristics of the turbine stage are shown in Table 2-1. The rotor tip
radius is 0.4582 m and the inner radius is 0.3353 m, giving a hub-tip ratio of 0.7317. This
ratio is representative of high-pressure turbine stages, where the stage inlet density is high
and hence the annulus height tends to be small. The turbine generates about 60 kW of
power while operating at a nominal speed of 1300 rpm and a nominal through flow rate
of 11.05 kg/s. The design total-total isentropic efficiency of the turbine stage is 0.893.
Design performance data of the turbine stage is shown in Table 2-2. The design
inlet conditions to the turbine stage are that of standard atmosphere at mean sea level.
Low inlet to exit pressure and temperature ratios implies that compressibility effects are
insignificant. Additionally, there is minimal density change through the stage allowing
for a constant mean radius throughout the stage.
24
Figure 2.3: AFTRF Instrumentation Window.
Nozzle ring area Casing inner face
∅ = 0.9166 m Over tip region
Total pressure probe
Window View From OUTSIDE
Stationary frame probe traverse
Probe stem
25
For a specific turbine design, where the geometry and gas are fixed, the pressure
ratio, temperature ratio, and efficiency depend on the flow function and the speed
function as given by Equation 2-1. The definition and value of the flow function and the
speed function are listed in Table 2-3, along with other design coefficients. Typical inlet
conditions experienced were P01 = 98.0 kPa (980 mbar) and T01 = 301.15 K (28° C). The
departure from design inlet operating conditions can be corrected for by changing the
turbine mass flow rate and rotor speed such that the ratios on the RHS of Equation 2-1
are kept constant. Design performance is achieved by operating the turbine stage at a
mass flow rate of 10.466 kg/sec and a rotor speed of 1327 rpm. These two parameters
may also be represented in non-dimensional form by φ, the flow coefficient.
⎟⎟⎠
⎞⎜⎜⎝
⎛=
0101
01
01
03
01
03
TN,
PTm
,TT,
PP &
fη (2-1)
Table 2-1: AFTRF Overall Stage Design Characteristics.
Power; P (kW) 60.6
Rotor Tip Radius; rtip (m) 0.4582
Rotor Hub Radius; rhub (m) 0.3353
Rotor hub-tip ratio 0.7317
Mass Flow Rate; m& (kg/sec) 11.05
Rotational Speed; N (rpm) 1300
Total-Total Isentropic Efficiency; ηtt 0.893
26
2.2.1 Stage Aerodynamic Design
Turbine stage nozzle guide vane (NGV) design parameters are listed in Table 2-4.
The stator ring has 23 nozzle vanes that turn the inlet axial air stream by about 70°. The
specific vane exit angle is chosen in order to maximize vane efficiency. Design Reynolds
numbers at nozzle inlet and exit are representative of modern high pressure stages. The
spacing between the stator row and rotor is variable between 20% and 50% blade tip
Table 2-2: Design Thermodynamic Parameters.
Total Inlet Pressure;Po1 (kPa) 101.36
Total Inlet Temperature;To1 (K) 289
Total Pressure Ratio; Po1/Po3 1.0778
Total Temperature Ratio; To3/To1 0.981
Pressure Drop; Po1-Po3 (mm Hg) 56.04
Table 2-3: AFTRF Design Coefficients.
Flow Function; m& √T / P (kg√oK m2 / kN sec) 1.85
Speed Function; N/√T (rpm/√oK) 76.47
Flow Coefficient; φm = (V2ax/U)m 0.568
Work Coefficient; ψm = ∆ho/Um2 1.85
Specific Work Output; ∆ho/ m& (kJ/kg) 5.49
27
axial chord length. The spacing used in the current investigation is 20% blade tip axial
chord length.
Turbine rotor design data is listed in Table 2-5. The rotor blade count is 29. The
maximum relative Mach number in the rotor is 0.24. The blades were designed to have a
nominal tip clearance of 0.9 mm or 0.76% blade height (h = 0.1229 m), on a rotor
average basis. Blade velocity triangles at three sections are shown in Figure 2.4. Blade
inlet angles change considerably from hub to tip, giving the blade a twisted, three-
dimensional shape at the leading edge. The rotor inlet geometric parameters indicate a
small departure from free-vortex conditions (r * Vθ = Constant). The design absolute
velocity vector at rotor exit is aligned at 25.16° to axial at the tip and at 35.13° to axial at
the hub. The AFTRF turbine stage is a reaction stage, which means that part of the static
pressure drop occurs across the turbine rotor, as shown in Table 2-6. Thus, more than half
the static pressure drop at the tip radius takes place in the turbine rotor, while very little
static pressure drop occurs within the rotor at the hub.
Table 2-4: AFTRF Nozzle Design.
Number of Vanes 23
Stator Zweifel Coefficient 0.7247
Chord; (m) 0.1768
Spacing; (m) 0.1308
Maximum Thickness; (mm) 38.81
Turning Angle; 70o
Reynolds Number ( ÷105) inlet / exit (3~4) / (9~10)
Stator Efficiency; ηs 0.994
28
Table 2-5: AFTRF Rotor Design.
Number of Blades 29
Rotor Zweifel Coefficient 0.9759
Relative Mach Number 0.24
Blade Height; hb (m) 0.1229
Tip Clearance; (mm) 0.9
Turning Angle; Tip / Hub 95.42o / 125.69o
Chord; (m) 0.1287
Axial Tip Chord; (m) 0.084
Spacing; (m) 0.1028
Maximum Thickness; (mm) 22
Reynolds Number ( ÷105) inlet / exit (2.5~4.5) / (5~7)
Rotor Efficieny; ηR 0.8815
Table 2-6: Radial Variation of Turbine Stage Design Parameters From 1-D Mean-line Analysis.
Parameter HUB MID TIP
Loading, ψ 2.156 1.854 1.635
Reaction, R 0.184 0.382 0.507
Flow Coefficient, φ 0.754 0.568 0.429
Turning, ∆β 125.09 110.54 86.20
29
2.2.2 AFTRF Tip Clearance Distribution
The distribution of tip gap height in the turbine rotor is shown in Figure 2.5. The
gap height was measured at the split line of the measurement window using feeler gages.
The values reported here are within ±0.0254 mm (0.001”) or t/h = 0.02% of the actual
value. Measurements were made in three equal regions along blade axial chord (near
leading edge, mid-chord, near trailing edge) and an average was computed for each
Figure 2.4: Blade Velocity Triangles.
U
W2
V2
V3
W3
30
region. The values reported in Figure 2.5 are average values for the tip surface. The
variation of gap height of each blade, with axial distance from the leading edge, is shown
in Appendix B. Four distributions were active over the course of the project. The design
tip clearance for the AFTRF is 0.9 mm or t/h = 0.76%. Five blades in the rotor, referred
to as, “cooled blades,” were modified to have a slightly larger tip gap. TCL1, the dashed
blue line, shows a single blade (B21) with a large gap height of t/h = 1.40%, whereas the
other blades have a gap height closer to the design gap height. The gap height of the other
cooled blades was reduced to the levels shown by applying precision plastic layers to the
tip surface with the help of double-sided tape of thickness 0.102 mm (0.004”). The
distribution TCL2, solid green line, was obtained by reducing the gap height of blade B21
to t/h = 0.72%, using precision plastic layers and double-sided tape. These two
distributions are relevant to results in Chapter 4, and results from “isolated injection,”
presented in Chapter 5 and Chapter 6. TCL3 represents gap height as measured from bare
metal tip surface of each blade. Thus, this represents the maximum tip clearance levels
achievable in the rig in this configuration and was used for coolant injection from all
cooled blades, referred to as “multiple injection,” results of which are presented in
Chapter 7. The last distribution, TCL4, was obtained by applying precision plastic layer
to the casing surface in the measurement window region and is relevant to casing surface
roughness results presented in Chapter 7.
31
2.3 Tip Cooled Blades
As mentioned in the previous section five blades, referred to as cooled blades,
were modified to a slightly larger tip clearance gap of 1.65 mm or t/h = 1.34%. These
blades were also modified to include features for injecting coolant air from the tip surface
of the blades. The tip surface of one of the cooled blades is shown in Figure 2.6. The tip
trench has four injection holes, three of which connect to the radial plenum chambers
shown in the inset of Figure 2.6. The fourth injection hole, marked H4, is located at the
end of the trench, near the trailing edge and serves as its own plenum. The radial plenums
0.30%
0.50%
0.70%
0.90%
1.10%
1.30%
1.50%
1.70%
0 5 10 15 20 25 30
Blade Number
Gap
Hei
ght (
t/h) %
TCL1
TCL2
TCL3
TCL4
COOLED BLADES
Figure 2.5: AFTRF Tip Clearance Distribution.
32
are blocked by a plug, inserted and permanently fixed after all modifications were
completed. All radial plenum chambers open into a common blade root plenum. A
standard ¾” Aluminum tube is fitted into the base of the blade root plenum to allow for
air lines to be connected to the chamber.
Geometric details of the injection scheme are shown in Figure 2.7 & Figure 2.8.
The tip trench is 60 mm long and 2 mm wide, as shown in Figure 2.7. The trench extends
from blade mid-chord to 0.91 Cax and is at a 60° angle with respect to the axial direction.
The three injection locations (H1 – H3) within the trench are made up of two holes, each
of ∅ = 0.762 mm, while the last (H4) is a single hole of ∅ = 1.8 mm. The number of
holes at each location and their diameters are specified in the table in Figure 2.7.Based on
a network analysis by Pudupatty [43], the discharge mass flow rates from each injection
location are as shown in Table 2-7. The trailing edge location H4 accounts for about half
the total discharge area and a third of the total coolant mass flow rate. While H3 has the
same discharge area as H1 and H2, the radial plenum diameter to this location is smaller,
as shown in Figure 2.7, thereby resulting in a small reduction in the mass flow rate.
The four injection sets are located 0.10 Cax apart, starting at 0.61 Cax. The distance
is measured from blade leading edge to the mid point of the line joining the centers of the
two injection holes. As shown in the cross-sectional details in Figure 2.8, the two
injection holes at H1 are inclined at 10° to each other and at 45° to the tip surface,
directed towards the blade pressure-side corner. It must be noted that the jets are not
blocked by the edge of the tip trench. The two holes open into the corresponding radial
plenum. This detail is repeated at H2 and H3.
33
Figure 2.6: Cooled Blade Tip Surface.
34
Figure 2.7: Cooled Blade Tip Coolant Injection Arrangement.
Table 2-7: Estimate of Coolant Discharge Mass Flow Rate, From Pudupatty [43].
Injection location Area (mm2) / Fraction of total area
Radial plenum area
(mm2)
Fraction of total mass flow rate
H1 0.912073 / 17.24% 12.4898 22.56%
H2 0.912073 / 17.24% 12.4898 22.56%
H3 0.912073 / 17.24% 7.05539 22.32%
H4 2.55431 / 48.28% 2.55431 32.55%
35
Figure 2.8: Cross-Sectional View of Cooled Blade.
36
2.4 Air Transfer System (ATS)
Figure 2.9 describes the air transfer system used for bringing coolant air from the
stationary frame to the plenum chamber of the cooled blades, in the rotor frame of
reference. The shaded regions in Figure 2.9 denote rotating components, while the un-
shaded regions denote stationary components. A custom designed air chamber is located
near the downstream bearing of the turbine rotor assembly. The rotating face of the
chamber is connected to the rotating instrumentation drum, while the stationary part of
the chamber is fixed to the bearing housing. Two precision seal systems minimize
leakage between the rotating and stationary surfaces of the air transfer system. The two
stationary seals work against plasma coated surfaces on the turbine shaft and rotating
drum of the AFTRF. Five barbed pipe fittings are threaded into the rotating part of the
ATS. High-pressure nylon tubes are connected between the fittings and metallic straight
connectors mounted in the rotating drum. One of these connectors is modified by
inserting an orifice for mass flow rate measurement. High-pressure tubes then supply the
air to the blade plenum chambers.
2.5 Flow Visualization
Surface flow visualization experiments were conducted using a mixture of oil and
pigment. The oil used was ash less dispersant SAE 40 Aviation oil (Aeroshell oil W80).
Initial experiments conducted resulted in the conclusion that oil by itself was not a viable
medium for surface flow visualization due to the high wall shear stresses encountered in
37
the rotor blade passages. Two types of pigments, Titanium White artists’ oil color
containing titanium oxide and zinc oxide, and Zinc Yellow Hue artists’ oil color
containing Arylide yellow and Zinc Oxide were used for surface flow visualization. The
visualization material used contained about 500 mm3 of pigment mixed in with oil and
weighed approximately 2 gm in a volume of 10 ml. The mixture with Zinc Yellow Hue
was observably more viscous than the mixture with Titanium White Artists’ oil color.
Figure 2.9: Air Transfer System (ATS).
Air-Flow
38
2.6 Instrumentation
A comprehensive instrumentation and measurement system, schematics of which
are shown in Figure 2.10 and Figure 2.11, was set-up as part of the current research
effort. The monitoring system shown in Figure 2.10 is used to control operation of the
facility, while the performance measurement system shown in Figure 2.11 is used to
measure the turbine rotor exit flow field to assess the effect of desensitization.
Encoder Signal Unit
Thermocouple Signal Conditioning Unit
ANALOG OUTPUT
DIGITAL OUTPUT
LOW SPEED DASTIMING I/O CONNECTIONS
ANALOG INPUTS
P03
P01
Inlet Pitot-Static Probe
V2ax / Um
6-Channel Pressure
Transducer
ρV12 / 2
1 / rev. Trigger Pulse
6000 / rev. Clock P
ulse
To High Speed DAS
Exit Pitot-Static Probe
Inlet Thermocouple
Exit Thermocouple
Room Thermocouple
Bearing Thermocouple
OMEGAThermocouple Display
Units
ρV32 / 2
T01
Troom Tbrg
T03
Himmelstein Torque Meter + Display Unit
Power + Torque
Figure 2.10: Schematic of Monitoring Instrumentation.
39
2.6.1 Monitoring Instrumentation
The measurements for monitoring test conditions consist of pressures and
temperatures at stage inlet and exit, bearing temperature, ambient conditions, and shaft
output, as shown in Figure 2.10. A low speed data acquisition system described in
Section 2.7 acquires and processes the data.
Turbine inlet flow conditions are measured using a single Pitot-static probe and a
single K-type thermocouple, both located at turbine mean radius and about 1.5 vane tip
KRONHITE Filter
Kulite Power & Signal Unit
Total Pressure Probe with Kulite Sensor
VALIDYNE Pressure
Transducer
Encoder Signal Unit
Analog Signal From Low-Speed
DAQ
DIGITAL OUTPUT
HIGH SPEED DAS
TIMING I/O CONNECTIONS
ANALOG INPUTSP03
P01
Inlet Pitot-Static Probe
V2ax / U
m
HONEYWELL Pressure
Transducer
ρV1 2 / 2
1 / rev. Trigger Pulse
6000 / rev. Clock Pulse
Stepper Motor Controller
Figure 2.11: Schematic of Performance Instrumentation.
40
axial chord lengths upstream of the nozzle. The Pitot-static probe is connected to pressure
transducers to measure inlet total pressure and inlet kinetic energy per unit volume, as
shown schematically in Figure 2.10. Inlet total temperature, measured using the
thermocouple, is used to control turbine operating speed. Similarly, the rotor exit total
pressure, kinetic energy per unit volume, and total temperature are measured using a
single Pitot-static probe and K-type thermocouple located at mean radius and 0.6 rotor tip
axial chord length downstream of the rotor. The pressure transducers used in conjunction
with the Pitot-static probes are a Validyne transducer and Honeywell transducers.
Bearing temperatures are monitored using K-type thermocouples on the bearing
housings. Shaft power and torque are measured by an inline torque meter connected to a
Himmelstein display. The power and torque readings are not connected to the data
acquisition system and are noted down.
2.6.2 Performance Measurement
The effect of the implemented desensitization methods on over tip leakage flow is
assessed by measuring the total pressure downstream of the rotor exit plane. The
instruments used consist of a high-frequency total pressure probe and inlet pressure
probe. The signals from these instruments are connected to a high speed data acquisition
system, which is described in Section 2.7. An analog voltage signal generated by the
monitoring system is also connected to this system.
41
2.6.2.1 High-Frequency Total Pressure Probe
A high frequency response pressure sensor, XCS-062-5D, manufactured by Kulite
Semiconductors is used to measure the exit total pressure field. The sensor is a sealed
cylindrical tube of diameter 1.6 mm, rated at 5 psid, and operates in a differential mode.
A protective B-screen mounted in front of the sensor diaphragm reduces the diaphragm
frequency response from 150 kHz to approximately 20 kHz, Kulite [44]. The sensor is
powered by a regulated 10 VDC supply and the transducer signal is amplified through an
instrumentation amplifier. Static calibration of the probe using a manometer yields a
linear calibration curve, as shown in Figure 2.12. The zero offset of the transducer is
subtracted from measured voltage when the transducer is pressurized.
The transducer is housed in a 3.5 mm outside diameter tube with a square cut
face, such that it is flush with the square face of the tube and all gaps between the sensor
and tube are sealed to prevent cavities. The differential pressure tube is left open to the
room, which means that one end of the diaphragm is at Pamb. The probe is mounted in a
single axis, stepper motor driven, linear traverse system on the casing window. The
traverse with probe is shown in Figure 2.3. This allows the probe to be traversed with
increments of 1/16th inch and greater. The probe face is positioned 30% chord
downstream of the rotor exit, at mid-pitch with respect to the upstream nozzle passage,
and aligned to the absolute tip flow vector at the rotor exit, as shown in Figure 2.2.
Characteristics of the total pressure probe, such as angular sensitivity and frequency
response are included in Appendix A. Through measurements at rotor exit the probe is
shown to be insensitive to incidence angles in a range ±15.
42
2.6.3 Coolant Mass Flow Meter
Coolant mass flow rate to cooled blade B21 is measured by an ∅ 8.0 mm orifice
installed on the supply line in the rotating frame. The orifice is installed in a ∅ 9.525 mm
(3/8”) pipe with two static pressure taps located 2D upstream and 5D downstream of the
orifice plate. The orifice meter is calibrated to relate the pressure difference (∆P) between
the static pressure taps to the standard volume flow rate through the orifice. The
calibration chart shown in Figure 2.13 relates the pressure difference to the non-
dimensional mass flow rate Minj. Minj as defined in Equation 2-2 is the ratio of orifice
R2 = 1.00
-3.000
-2.500
-2.000
-1.500
-1.000
-0.500
0.000
0.00 0.50 1.00 1.50 2.00 2.50
Applied Gage Pressure, dP (kPa)
Mea
sure
d V
olta
ge C
hang
e Fr
om V
0, dV
(VD
C)
dP (Pa) = 925.352 dV (VDC)V0 = Voltage at (dP = 0).
Figure 2.12: Total Pressure Probe Static Calibration.
43
mass flow rate to turbine design mass flow rate, assuming coolant is being injected from
all 29 blades in the blade row. The orifice meter pressure taps are connected to a PSI
systems pressure transducer in the rotating frame, as detailed in the Section 2.6.4. The
calibration equation used to set the coolant mass flow rate during the test is obtained from
a modified curve-fit relating dependence of Minj on the transducer signal due to applied
pressure difference, as shown in Figure 2.14. The actual turbine mass flow rate varies for
each test run and may also be different from the design mass flow rate. Thus, the value of
Minj reported with the results has been corrected for the actual operating conditions.
2.6.4 Rotating Frame Pressure Transducer
A PSI Systems ESP-48 transducer mounted in the rotating drum is used for
measuring the static pressures from the orifice meter. The transducer has 48 sensing ports
and one reference port and can measure a pressure difference of ±1 psi between any of
the sensing ports and the reference port. The static pressure taps from the orifice meter
are connected to the PSI transducer in the differential mode, to directly measure the
pressure difference. The upstream static pressure tap is connected to the reference port,
while the downstream static pressure tap is connected to a sensing port. The transducer is
powered by a 10VDC regulated supply. Channel indexing is achieved by an 8-bit digital
signal.
T
c inj m
29*m M&
&=
(2-2)
44
2.7 Data Acquisition
Data acquisition is accomplished by using two National Instruments DAQ boards
and associated LabView programs. A low speed, 8-channel, board PCI-6024E in a PC-
AT computer is used for acquiring data for test monitoring. Analog inputs to this board
consist of inlet and exit total pressure, inlet and exit kinetic energy per unit volume, total
temperatures at stage inlet and exit, inlet chamber temperature, and bearing temperature.
Digital input consists of a 6000 per rev pulse generated by the shaft encoder. The output
functions consist of an analog output voltage that is proportional to the flow coefficient
0.00
0.50
1.00
1.50
2.00
2.50
3.00
0.00 0.50 1.00 1.50 2.00 2.50
Wall Static Pressure Difference Across Orifice Plate, (psid)
Min
j, (%
)
Figure 2.13: Coolant Discharge Measurement Orifice Calibration.
45
and digital lines for indexing channels on the rotating frame pressure transducer. The
input data is acquired and averaged over 60 seconds. A virtual instrument written using
LabView processes the inputs and displays various parameters, some of which are shown
in Table 2-8, along with the use of each parameter.
y = 1.094x0.4257
R2 = 0.9949
0.00
0.50
1.00
1.50
2.00
2.50
3.00
0.00 1.00 2.00 3.00 4.00 5.00 6.00
Transducer Signal, (VDC)
Min
j, (%
)
0.332
Figure 2.14: Coolant Mass Flow (Orifice) Meter Calibration Minj vs. VDC.
46
A high speed, 4-channel, simultaneous sampling DAQ board, the PCI-6110E is
used for performance improvement measurements. The analog inputs include signal from
the high-frequency total pressure probe, inlet kinetic energy per unit volume, inlet total
pressure, and a voltage signal from the low-speed data acquisition system. Timing and
I/O operations are controlled by a one per rev., trigger pulse and a 6000 per rev. clock
pulse. Digital output lines are used to control the stepper motor.
The associated Labview instrument initiates data acquisition at each radial
position of the total pressure probe upon being triggered and acquires 6000 points per
revolution of the rotor. This is an important feature implemented into the current data
acquisition process. Typically, during a single run the turbine speed is varied with
temperature, to hold the speed function constant. Additionally, the turbine speed for
different runs tends to be different. The use of the encoder pulse to control the scan clock
ensures that each data point falls in the same circumferential bin, independent of the rotor
speed. The unsteady total pressure downstream of the rotor is acquired 6000 times per
Table 2-8: Some Calculated Parameters From Low Speed DAS.
Parameter Operational Importance
Pressure Ratio, Po3/Po1 -
Temperature Ratio, To3/To1 -
Flow Coefficient at mid-span, φ Controls acquisition of data on high speed DAS
Rotor Speed, N Rotor speed control
Actual Mass Flow Rate, m& Controls acquisition of data on high speed DAS
47
revolution, which corresponds to a nominal sampling rate of 132 kHz. The data is
ensemble averaged over 200 ensembles. Stepper motor control also rests with this virtual
instrument.
2.8 Data Processing
The unsteady total pressure downstream of rotor exit is ensemble averaged and
converted to a non-dimensional total pressure coefficient as given by Equation 2-3. The
pressure drop across the stage is a measure of both the energy extracted from the air as
well as the energy loss in the stage. The advantage of the specific non-dimensional form
is that this value is invariant with operating conditions. Um may be expressed in terms of
the speed function as shown in Equation 2-4. The air density term in the numerator of
Equation 2-3 may be canceled out by taking inlet total pressure out in the numerator and
multiplying and dividing the RHS of Equation 2-3 by R. Then the total pressure
coefficient is given by Equation 2-5. The pressure ratio term in the numerator is a
function of the flow function and the speed function, which if maintained constant should
result in the same pressure ratio, irrespective of operating conditions. Since all terms in
the denominator are constant, maintaining the flow coefficient and hence the stage total
pressure ratio is important to ensure repeatability in the measured total pressure
coefficient. The uncertainty in total pressure coefficient, calculated by the uncertainty
propagation method of Kline and McClintock [45], is δCpt = ±0.024 or ±0.58%.
48
The measured total pressure coefficient is circumferentially averaged over each
passage to obtain a passage averaged coefficient as defined in Equation 2-6, at each radial
position. The average is essentially the mean of 207 total pressure coefficient values. The
number of circumferential points per passage is obtained by dividing the total number of
points per rotation (6000) by the number of blades (29) and truncating the fraction. The
radial distribution of the passage averaged coefficient isolates the effect of OTL in each
passage and allows for comparison of tip gap behavior with and without coolant
injection.
2
0103
21
),(),(m
pt
U
PjiPjiCρ
−= (2-3)
des
operdes
mm T
TN
60r2U ∗=π (2-4)
60r2A where
TN
A21
1P
),(P
),( m2
des
des
01
03
π=
⎟⎟⎠
⎞⎜⎜⎝
⎛
⎟⎟⎠
⎞⎜⎜⎝
⎛−
=
jiR
jiC pt (2-5)
∑+
=207
, ),()(i
iptPpt jiCjC (2-6)
49
For each test the total pressure coefficient is circumferentially averaged over the
entire rotor (6000 points) to obtain the rotor averaged coefficient as defined in
Equation 2-7. The radial distribution of the rotor averaged coefficient is assumed to be
invariant if the flow within the rotor passages is locally modified. Thus the effect of
changing the gap height of cooled blade B21 and the effect of coolant injection from B21
should produce no effect on the rotor averaged coefficient. This allows each data set to be
compared for repeatability and is used as a criterion for acceptable data.
An area averaged coefficient is also calculated for individual passages as defined
in Equation 2-8. The average is computed over an area of one passage (207 points) and a
height of 20% span starting at the tip. This coefficient gives a single value than captures
the effect over the entire passage.
2.8 Operation
Operation for total pressure measurements consists of starting up one of the axial
flow fans and allowing the turbine rotor to spin at part speed and load for 5 minutes. The
second fan is then started and the rotor is allowed to spin for an additional 5 minutes
∑=
=6000
1, ),()(
iptRpt jiCjC (2-7)
∑ ∑=passage heightblade
ptApt jiCC%20
, ),( (2-8)
50
before increasing speed to value indicated by inlet total temperature. The facility is then
allowed to run at constant speed for about 30 minutes in order to let the ambient air, the
facility, and the total pressure probe to achieve thermal equilibrium. Turbine speed is
corrected if necessary and data acquisition is commenced when the measured flow
coefficient is close to the required value. Data acquisition is triggered by a one per rev
trigger pulse from the shaft encoder and 6000 data points are acquired per revolution.
Upon completion of data set at a given radius the probe is automatically traversed radially
to the next radial position. After each movement a 5 second delay occurs before data
acquisition. Two radial positions are measured in a minute of run time. During the test,
speed changes are effected depending on the change in temperature. For tests involving
coolant injection the coolant flow is started at least 5 minutes before data acquisition.
This allows the coolant flow to stabilize.
Operation procedure for flow visualization experiments is detailed in subsequent
chapters. The start-up process is much quicker and the turbine rotor is brought to
operating speed within 2 minutes. Visual observations were conducted during the
operation to help terminate the test.
51
Chapter 3
Rotor Flow Visualization
Surface flow visualization is a qualitative flow measurement technique that is
used frequently in the study of internal and external flows. Although surface flow
visualization in the study of turbine passage flows is common in cascades, results from
rotating turbomachinery flows are rare. Langston [46] notes that surface flow
visualization has been extensively used in cascades to explain complex cascade passage
flows and also to guide the alignment of probes for measurements near surfaces.
Merzkirch [47] defines mechanical interaction techniques for surface flow visualization
as those that consist of application of dots or film of oil and pigment mixture to the
surface being studied. The slope of the surface shear stress lines indicates the ratio of
two-dimensional shear stresses, as given by Equation 3-1, from Langston [46]. The flow
is assumed to be in the x-z plane, with y-axis normal to the plane. The velocities are u in
the x-direction and w in the z-direction.
0
0)(
=
=
∂∂
∂∂
=
y
y
yu
yw
dxxdz (3-1)
52
An ink-dot-film technique of surface flow visualization was developed by Eckerle
and Langston [48] and used to study the formation of horseshoe vortex due to the
interaction of endwall boundary layer and a vertical cylinder. Aunapu et al [49], [50] used
a similar technique in their investigation of cascade endwall flow and the effect of an
endwall fence on cascade secondary flows. Allen and Kofskey [51] employed smoke
flow visualization to study the development and interaction of secondary and tip leakage
flows in a rotating rig. Dring and Joslyn [52] studied blade surface and tip surface flows
by injecting ammonia from surface pressure taps. The ammonia interacted with diazo
paper applied to the endwall surfaces and indicated the direction of surface streamlines.
Results from the application of mechanical interaction techniques to visualize flows on
rotating surfaces are rare in the public domain.
In this chapter the visualization of rotor endwall flow, blade tip surface flow, and
turbine casing endwall flow is presented. It is believed that the surface flow visualization
results discussed in this and subsequent chapters are the first such surface flow
visualization results from a rotating rig. The visualization material used consists of a
mixture of oil and pigment as detailed in Section 2.5 and following Merzkirch’s [47]
terminology the techniques used will be referred to as oil-dot technique or oil-film
technique. The surface shear stress lines formed by the visualization material are referred
typically referred to as surface or limiting streamlines. They will be referred to as streaks
in this document.
53
3.1 Rotor Hub Endwall Flow Visualization
Rotor endwall (hub) surface flow visualization was conducted using oil dots and
oil film, applied to the rotating endwall surface. Oil dots provide discrete surface markers
that enable unambiguous representation of surface flow. The origin of the visualization
material and surface streamline is easily identified when using dots. The information
gained from this is useful in analyzing the patterns left in the oil-film, which provides a
better global view of the surface flows. The origin of the visualization material is not
clear when using the oil-film technique and as suggested by Langston [46] the results
must be interpreted carefully since the accumulation of visualization material might be
due to absence of wall shear, as in separated flow, or due to application of transported
visualization material. In all tests the timing was determined by visual observations and
ranged from 10 minutes of steady operation when using oil dots to 20 minutes when
using oil film.
3.1.1 Visualization with the Oil-Dot Technique
The rotor endwall surface was cleaned and the visualization mixture was applied
to the endwall surface in the form of a grid of dots. The visualized streaks on the rotor
endwall are shown in Figure 3.1. The negative of the original image is presented for
improved clarity and hence the blade surfaces appear to be in shadow.
Figure 3.2, from Xiao [3], shows the static pressure distribution on the rotor
endwall. Points A and B denotes local maxima in static pressure, while point C denotes
54
the minimum static pressure measured on the rotor endwall. Streaks near the rotor inlet
plane, shown in Figure 3.1, illustrate the complex boundary layer development in this
region. The oil streaks at the leading edge of blade B23 are tilted in the z-direction, up to
about mid-pitch (Point 1), as the inlet flow moves around the suction surface. The flow
from mid-pitch to about ¾ pitch (Point 2) is directed axially and towards the suction
surface of B23 in the last quarter of the passage width. This behavior is consistent with
the static pressure distribution in Figure 3.2 where strong pitch-wise gradients occur near
the blade leading edge and a large, mostly uniform pressure field occurs centered around
mid-pitch. The lowest static pressure in the front half of the passage is measured in an
area approximately symmetric about the point of maximum blade curvature. Figure 3.1
shows that all visualization material in the passage up to the dotted white line is swept
into this area, marked by the dashed white curve.
Visualization material along the pressure surface – endwall corner is swept
towards the suction surface of blade B23. The orientation of the streaks changes rapidly
from the leading edge to a maximum angle of 75° near the point of maximum curvature
on the pressure surface. This may be explained by the strong, cross-passage pressure
gradients seen in Figure 3.2. The highest static pressure is measured at B and the isobars
up to mid-pitch have a strong axial orientation, indicating that the boundary layer flow is
mostly in the negative z-direction. The pressure field changes character around mid-chord
with stronger chord-wise gradients appearing. This is reflected in the oil streaks in
Figure 3.1. Streaks near the suction surface – endwall corner are mostly parallel to the
suction surface. Oil streaks in the passage are directed towards rotor exit, rather than the
suction surface of blade B23.
55
Figure 3.1: Rotor Endwall Surface Flow Visualization Using the Oil-Dot Technique.
1
2
B23
B22
x
z
56
Figure 3.2: Rotor Endwall Static Pressure Distribution, From Xiao [3].
57
3.1.2 Visualization with the Oil-Film Technique
Figure 3.3 shows the rotor endwall with an initial oil film layer applied over the
surfaces of two blade passages. Oil upstream of the leading edge is applied with a
tangential orientation, while oil within the passage follows approximately the contour of
the passage. This reduces the ambiguity that the observed patterns are due to paint brush
strokes during application. While the oil film is thicker in Passage 1, the thickness of oil
application is reasonably uniform in each passage.
Figure 3.3: Oil Film On Rotor Endwall Before Test.
Tangential brush strokes
Brush strokes aligned with passage curvature
58
Figure 3.4 shows the surface streaks formed in the oil film, in the near leading
edge region. Streaks upstream of rotor inlet and on the rotating hub are oriented mostly in
the axial direction, indicating that inlet boundary layer development on the rotating hub is
predominantly in the axial direction. As with visualization using oil dots, the orientation
of the oil streaks varies in the pitch-wise direction. In region 1, the streaks are inclined
towards the pressure surface, while in region 2 the streaks are pointing away from the
blade pressure surface. In between these regions and immediately upstream of the blade
leading edge the inflow is axial. A detailed description of inlet boundary layer behavior
as it approaches a bluff body is given by Eckerle and Langston [48]. The inlet boundary
layer stagnates as it approaches the blade leading edge (LE) and eventually separates
from the rotor endwall. This leads to the accumulation seen within the oval region. The
separated flow then impinges on the leading edge and is turned back towards the
stagnating flow. As these opposing flows meet they create a saddle point and fluid is
turned to flow around the blade leading edge. This leads to the formation of the two legs
of the horseshoe vortex.
The pressure-side leg (Hp) is driven across the passage by transverse pressure
gradient and is clearly identified in the patterns formed on the rotor endwall surface. The
vortex impinges closer to the pressure surface of the blade. Two dividing lines are
formed, one between Hp and the inlet flow, and the other between Hp and cross-passage
flow. The vortex moves in a straight line except for a small shift to the right, probably
caused by the strong cross-passage flow. The suction-side leg (Hs) of the horseshoe
vortex is not as distinct. However, close to the leading edge the dividing line separating
inlet boundary layer and Hs is identified by noting that the inlet boundary layer and the
59
re-circulating flow approach the line from opposing directions, as shown by the dotted
lines. The vortex impinges on the blade suction surface and is subject to stretching due to
acceleration. A large amount of visualization material is washed away from the rotor
endwall, indicating increased wall shear. This is consistent with the presence of thin
boundary layers due to acceleration around the suction surface curvature.
Figure 3.4: Near Leading Edge Surface Flow Features on the Rotor Endwall Visualized
Using the Oil-Film Technique.
2
1
Saddle Point
Hp
Hs
L.E.
x
z
60
Figure 3.5 shows an overall picture of the rotor endwall surface flow. The
pressure-side leg of the horseshoe vortex is marked by continuous curves. The curves are
drawn in regions where the visualization material is completely washed out, as in the
region close to the leading edge marked by a dotted circle. The pressure-side leg of the
horseshoe vortex divides the surface flow in to two distinct regions. On the impingement
side of the vortex the passage boundary layer is highly skewed by cross-flow from
pressure surface to suction surface. Similar boundary layer behavior is also observed in
cascade tests as reported by, among others, Gregory-Smith [53].
Flow on the other side of the pressure-side leg is blocked from following the
passage flow and instead flows towards the suction surface. The lower dividing line
becomes a little indistinct as the vortex approaches the blade suction surface, while the
upper dividing line is seen to continue downstream, as observed by the lack of
visualization material parallel to the blade split line. The rotor endwall surface static
pressure distribution in Figure 3.2 shows that strong chord-wise gradients occur in this
region, attributed to the passage vortex. Thus it is expected that increased wall shear due
to the entrainment of cross-flow by the pressure leg of the horseshoe vortex causes
greater oil to be removed in this region. Three dashed, white curves also indicate the
streak curvature, as the cross-flow approaches the split line. The last, dotted curve drawn
turns towards rotor exit after the split line. Indeed, downstream of this curve it can be
observed that visualization material crosses and also accumulates in the split line.
Therefore, the dotted curve probably represents the location where the passage vortex
lifts off the rotor endwall. The dividing line of the suction leg of the horseshoe vortex is
indistinguishable beyond point A. The continuous curve drawn leading up to this point
61
indicates a bifurcation pattern observed close to the leading edge. Thus it is suggested
that the suction leg of the horseshoe vortex lifts off the endwall in this region.
Accumulation of visualization material in this region of the rotor endwall also indicates a
reduction in wall shear.
Figure 3.6 displays surface flow features in the near trailing edge region of the
blade passage. The secondary flow is seen to encompass the entire passage. Some oil
Figure 3.5: Blade Passage Surface Flow Patterns on the Rotor Endwall Visualized Using
the Oil-Film Technique.
A
x
z
62
accumulation is evident on the endwall-blade corner, indicating reduced wall shear. The
green, dashed curve drawn separates the surface flow into two regions. Wall shear stress
above this curve appears to be lower than that below the curve. This could be a result of
higher velocity away from the pressure surface due to the significantly lower endwall
static pressures shown in Figure 3.2. At rotor exit, flow from pressure surface to mid-
pitch appears to exit the rotor at an angle of 64°, close to the design relative flow angle of
65°, as marked out. The blade wake is identified by the region of low removal rate of
visualization material. Flow closer to the suction surface is overturned, due to the passage
cross flow, and not much visualization material reaches beyond the rotor exit plane.
Visualization material is also swept up the blade suction surface, as shown in
Figure 3.7. The tip gap height for this blade is (t/h = 0.72%). The most significant
observation is the large white streak, enclosed between the two continuous curves drawn,
just below the blade tip, starting at about mid-chord and extending the length of the blade
surface. The mostly radial lines are drawn out by the visualization material as it is forced
up the blade suction surface by centrifugal action on the oil reaching the rotor endwall –
suction surface corner. It is clear that in the enclosed region the visualization material is
subject to intense, probably turbulent, shear that prevents the formation of the radial
streamlines. Due to its proximity to the blade tip the streak is believed to be caused by the
interaction between the tip leakage vortex and the tip passage vortex. The streak starts off
from the tip, in a radially inward trajectory before turning to move in the stream-wise
direction. The size of the streak also increases with chord-wise distance and at the trailing
it extends from about 13% blade height to about 25% blade height.
63
Figure 3.6: Near Trailing Edge Surface Flow Patterns on the Rotor Endwall Visualized
Using the Oil-Film Technique.
Blade Wake
64°
x
z
64
Figure 3.7: Blade Suction Surface Trace Formed During Rotor Endwall Surface Flow
Visualization.
Lines caused by visualization material from rotor endwall moving up the blade surface due to centrifugal action
Streak of visualization material due to interaction between tip leakage and tip passage vortices
65
3.2 Blade Tip Surface Flow Visualization
Blade tip surface flow visualization was conducted by applying oil dots on the
pressure surface of certain rotor blades. The visualization material moves up the blade
pressure surface under the action of centrifugal forces and is eventually thrown off the
blade surface at the pressure-side corner. The effect of centrifugal force on the
visualization material is greatest at the start of the experiment due to its finite volume.
During operation the thickness of the visualization material decreases and the effect of
aerodynamic forces is greater than that at start-up. Some of the material is carried on to
the tip surface by leakage flow entering the gap. The visualization material deposited is
then subject to the action of aerodynamic shear forces on the tip surface. The use of oil
dots gives discrete markers on the surface and information gained from this is then
applied towards understanding patterns that are caused by the application of oil film on
the pressure surface. These results will be discussed in subsequent chapters.
Figure 3.8 shows the pressure surface of blade B21 (t/h = 1.40%), before and after
the test run. The dots were applied about 1” below the tip surface, in two staggered rows.
The motion of the dots up the blade surface is governed by the ratio of the centrifugal
force to aerodynamic shear. It is clear from the second picture that the centrifugal force is
the dominant force on the visualization material over most of the blade surface, as the
dots move in discrete lines, directed radially outward. Towards the trailing edge however
the aerodynamic forces increase due to acceleration of flow over the rear half of the
blade. The lines are seen to tilt towards the trailing edge and visualization material in the
last 3-4 columns does not reach the tip surface.
66
The tip surface patterns for this test are shown in Figure 3.9. All dot traces, except
those close to the trailing edge, deposit material on to the tip surface. A continuous
limiting streamline is visible on the tip surface along the blade length indicating chord-
wise flow in the near pressure-side corner of the tip surface. Oil streaks directed towards
the suction-side corner of the tip gap are seen to originate from this limiting streamline.
As will be shown in the next chapter a strong re-circulatory pattern is observed on the tip
surface and the limiting streamline is likely formed at the edge of the re-circulating flow
and the corner separated flow. This chord-wise flow was also measured by Bindon [7] in
a linear cascade setup and observed in surface flow visualization of an idealized tip gap
by Sjolander and Cao [9].
Figure 3.8: Oil Dots on Blade (B21) Pressure Surface Before and After Test Run.
BEFORE RUN
AFTER RUN
67
Figure 3.9: Tip Surface Flow Visualization (t/h = 1.40%) by Oil Dots Applied Near
Blade Tip.
0.2 Cax.
0.5 Cax.
Limiting streamline moves towards suction-side of tip surface
H2 at 0.71 Cax.
H1 at 0.61 Cax.
Low momentum region
Minimal transfer of oil on to tip
surface
68
The limiting streamline is first visible, as shown in the enlarged detail in
Figure 3.10, close to 0.2 Cax. It stays very close to the pressure-side corner until about
mid-chord and then moves observably towards the blade suction surface. The location of
the limiting streamline in relation to the trench shows that it passes closest to the injection
hole at 71% chord. The limiting streamline is not observed over the last 5% of the blade
chord length. In this region, visualization material from the pressure-surface does not
reach the tip surface as most of it is turned around the trailing edge. Additionally, as will
be shown in the next chapter, gap flow fails to reattach on to the tip surface in this region.
A second test was conducted by applying the oil dots near the blade root. The
blade tip surface was also coated with smooth, flat black paint. The tip surface pattern
from this is shown in Figure 3.11. The repeatability between the two tests is very good.
The limiting streamline is seen to follow the same path as shown in Figure 3.9. There is a
noticeable shift towards the blade suction-side corner between 50% and 60% chord and
then the limiting streamline runs fairly parallel to the pressure-side corner. The lack of oil
up to mid-chord is believed to be due to insufficient visualization material reaching the
tip surface.
The tip surface flow patterns were also investigated on a blade with half the gap
height, t/h = 0.71%. The surface flow pattern is shown in Figure 3.12. A curved streak is
observed on the tip surface at about 0.3% Cax. The limiting streamline beginning from
this point terminates at about 0.4 Cax in a long straight streak across the tip surface. The
lack of curvature of this streak suggests that the near surface gap flow is mainly
responding to the gap pressure differential and is not affected by the chord-wise pressure
differential across the gap. The break in the limiting streamline might be due to a sudden
69
increase in gap velocity in this region. Beyond this point the limiting streamline begins
again at 0.45 Cax and is continuous up to the trailing edge. It is closer to the blade
pressure-side corner, in comparison to the large gap height previously discussed. The
streaks that cross the tip surface after mid-chord show greater inclination towards the
trailing edge, when compared to the single streak at 0.4 Cax.
Figure 3.10: Near Leading Edge Detail of Tip Surface Flow Visualization Using the Oil-
Dot Technique (t/h = 1.40%).
≅ 0.2 Cax.
70
Figure 3.11: Tip Surface Flow Visualization (t/h = 1.40%) Using the Oil-Dot Technique;
Oil Applied Near Blade Root.
0.5 Cax.
≅ 0.55 Cax.
Limiting streamline moves towards suction-side of tip surface
71
Figure 3.12: Tip Surface Flow Visualization (t/h = 0.71%) Using the Oil-Dot Technique.
0.3 Cax.
0.4 Cax.
0.45 Cax.
72
3.3 Rotor Casing Surface Flow
Rotor casing surface flows were visualized using both the oil-dot technique and
the oil-film technique. The oil dots were placed in three staggered rows along the rotor
axis and covered the rotor footprint. Figure 3.13 is an image taken during while the
turbine was running. Position of a few oil dots is marked with circles. The slope of the
surface streaks depends on the ratio of the shear stress in the x-direction and in the z-
direction. The characteristic velocity in the x-direction is the axial component of velocity
and in the z-direction is the tangential component added vectorially to the blade speed.
Visualization material displaces almost linearly in the first 30% blade chord and
then turns smoothly towards the axial direction. Most of the oil dots coalesce in this
region to generate a single streak. The orientation, with respect to the axial direction, of
the linear portion of the streaks is marked at various positions. The first two streaks, S1
and S2, are inclined at 87° degrees and 84° degrees respectively, which is greater than the
70° degree design angle. The proximity of these dots to the vane suction surface suggests
that this orientation is a result of overturning caused by vane passage secondary flow.
Similar pattern of overturned flow was observed on the rotor endwall with oil film
visualization. Indeed, these streaks merge immediately downstream of the vane trailing
edge and follow a path that is parallel to the vane trailing edge camber-line.
The third and fourth oil streaks, S3 and S4, which are immediately upstream of the
rotor inlet plane, are measured at 75° and 66°, respectively. Near blade mid-chord the
streak (S5) is measured at 46°, indicating almost equal shear stresses in the xy and zy
planes. The tangent to S4, measured at mid-chord is also inclined at 47° and as observed
73
the shear stress in the xy plane becomes more dominant with chord-wise distance. Dots in
the last 1/3 of the blade passage follow a very shallow path and are quickly turned
towards the axial, and then in the –z-direction beyond the rotor exit plane. The angle
measured for the streak S6 is -26°, which is close to the design exit angle at blade tip. It is
observed that not all the streaks are turned away from the direction of rotation. This
behavior is emphasized in Figure 3.14 acquired after the rotor was stopped. The two
streaks SA and SB indicate that the surface flow is as per design. The streaks close to SA
show similar orientation, while the streaks just below SB are directed axially. This axial
orientation changes as we approach streak SA. It is believed that the axial orientation of
the streaks is caused by the highly overturned flow from the nozzle guide vane exit flow
field being dominated by strong endwall cross flows.
3.3.1 Oil-Film Visualization of Rotor Casing Surface Flow
Rotor casing flow visualization was also conducted by applying an oil film to the
transparent casing window, as shown in the image on the left in Figure 3.15. The initial
oil film coverage is a little over one nozzle (N) pitch circumferentially, and about 1.5
rotor tip axial chord length, in the axial direction. Application of the visualization
material (brush strokes) is axially oriented. The image on the right in Figure 3.15 was
acquired after the test was completed.
74
An image of the outer casing flow patterns taken after the test was completed and
rotor was stopped is shown in Figure 3.16. The orientation of the streaks observed is very
similar to that observed with the oil-dot visualization. A region of decreased shear stress
is identified by a pair of white dotted lines. The fluid in this region appears to originate
from near the suction surface of nozzle guide vane N1 and the visualization material ends
up at a location exactly downstream of nozzle guide vane N2. This confirms the previous
hypothesis that the axial orientation of the streaks downstream of the rotor is caused by
the highly overturned flow exiting the nozzle passage. The reason for the greater
visualization material within this feature suggests that it is the path of the vane tip
passage vortex. Downstream of rotor exit there are regions of differing oil accumulation.
Oil streaks leading into region A appear to cover the greatest distance, carrying more
visualization material. It is clear from this picture that the wall shear imposed on the
visualization material is significantly larger in the footprint of the rotor than downstream
of the rotor.
75
Figure 3.13: Casing Surface Flow Visualization During Turbine Operation Using the Oil-
Dots Technique.
S1 (87°)
S2 (84°)
S3 (75°)
S4 (66°)
Streamlines converge
S5 (48°)
S6 (-26°)
Vane T.E.
Vane T.E.x
z
76
Figure 3.14: Rotor Casing Surface Flow Visualization Using the Oil-Dot Technique.
SA (-26°)
Vane (N2) T.E.
Vane (N1) T.E.
SB
Axially directed streamline(s)
x
z
77
Figure 3.15: Rotor Casing Surface Flow Visualization Using the Oil-Film Technique;
Before and After Test.
Blade Trailing
Edge
Vane Trailing
Edge
78
Figure 3.16: Rotor Casing Endwall Surface Flow Visualization Using the Oil-Film
Technique.
N1
N2
A
79
Chapter 4
Effect of Tip Gap Height on Over Tip Leakage
The results presented in this chapter discuss the effect of tip gap height on over tip
leakage (OTL) flow and the observed total pressure downstream of the rotor. Changes in
the tip surface flow due to variation in gap height were investigated qualitatively, by
using the oil-film technique of surface flow visualization. As introduced previously, OTL
flow is ejected into the passage from the blade suction-side and forms a vortex in the
blade passage thereby affecting the passage flow. The variations in leakage vortex
footprint are studied in the stationary frame by measuring total pressure downstream of
the rotor. A high resolution total pressure map at stage exit in the cold research turbine
could be used as a measurement of aerodynamic efficiency. One of the objectives of this
chapter is to set-up baseline data to compare to the effect of coolant injection.
4.1 Oil Film Based Tip Surface Flow Visualization
The oil film technique of surface flow visualization is implemented to study over-
tip leakage. The oil dot technique discussed in the previous chapter was able to
distinguish the chord-wise flow on the tip surface. However identification of micro-flow
features, as observed in the case of rotor endwall flow visualization, is difficult when
using discrete flow markers.
80
4.1.1 Operation
The blade surfaces are first coated with smooth flat black paint to allow for better
contrast with the white pigment used and to avoid reflections during image acquisition.
Visualization material is then applied on the pressure surface, with a soft brush as evenly
as possible. Images of the blade pressure surface with the un-disturbed oil film were
recorded before each test. The optical window was then fastened and the first fan stage
was started, which brought the turbine rotor to a stable speed of about 1240 rpm in
approximately 30 seconds. The second fan stage was then started and both the turbine
rotor speed and flow rate stabilized sufficiently in 30 seconds. Thereafter, the rotor speed
was increased to the corrected operating speed. Thus, the time required for the rotor to
reach operating speed was less than 90 seconds in a total run time of about 20 minutes. A
strobe light controlled by the AFTRF encoder one per rev., pulse was used to conduct
visual observations of the blades during the tests.
During the startup, the thickness of the visualization material on the blade
pressure surface is reduced, as the centrifugal action forces oil to move up the blade
surface and eventually the oil is ejected off the surface. It was observed that some of the
oil started splattering on the casing when the speed reached about 700 rpm. Some of the
oil thrown off the pressure surface is carried into the tip gap by the leakage flow. Thus,
by the time the rotor reached operating speed some amount of visualization material is
carried on to the tip surface. The thickness of the visualization material deposited on the
tip surface is extremely small and was observed to move in well defined dots across the
tip surface, under the influence of the aerodynamic shear generated by the gap flow. The
81
test was terminated when no motion was observed for over a minute, which on average
was 20 minutes after startup. The extremely thin streak patterns observed are thus due to
steady or time averaged flow in the tip gap. It must be noted that the mixture never dried
hard and could be smeared even after the test. Tests were also conducted to ensure that
the oil deposited on the tip surface did not dry by the time the rotor reached operating
conditions.
4.1.2 Tip Surface Flow Visualization; Large Gap Height (t/h = 1.4%)
Flow visualization was done on blades with different gap heights from a
maximum of 1.4% to a minimum of 0.72%. The surface patterns are most distinct at the
largest gap height and are discussed first in detail. At the smaller gap heights many of the
same features are present and the discussion focuses more on drawing out the differences.
4.1.2.1 Blade Pressure Surface
Figure 4.1 shows images of the blade pressure surface with oil film before (on the
left) and after (on the right) the visualization experiment. The initial oil film is brush
painted so that it has no preferred direction. Care is taken to ensure reasonably uniform
film thickness. A definite radial orientation of the oil film is visible in the image acquired
after the test. The oil layer on the surface of the rotating concave surface is forced
radially outwards due to centrifugal forces on the oil film. During startup the
aerodynamic shear, imposed by the passage flow in the relative frame, on the oil layer is
82
less significant in comparison to the centrifugal forces due to the initial thickness of the
oil film. The patterns on the pressure surface are not indicative of the passage flow.
However, the curvature of the patterns is indicative of the relative strengths of
aerodynamic and centrifugal forces imposed on the oil film. The blade surface is visible
in some areas where all the paint has washed out.
4.1.2.2 Blade Tip Surface
Figure 4.2 shows the tip surface patterns on the test blade (B21) with a sharp
corner, flat tip, and a tip gap height of t/h = 1.40%. The trench is blocked by applying a
smooth and thin layer of tape on the tip platform and cutting it to the tip shape. An
Figure 4.1: Oil Film on Blade Pressure Surface Before and After Test.
BEFORE TEST
AFTER TEST
83
extremely thin layer of smooth flat black paint is then sprayed on the tape. A grid of lines
is drawn at intervals of 0.05 Cax to allow positioning of various features.
One of the general observations that can be made is the oil accumulation on the
tip surface all along the PS edge of the blade. As compared to the rest of the tip surface,
only a few streaks are seen to originate from the accumulated oil up to 0.30 Cax along the
pressure-side corner. Additionally, these streaks do not extend to the suction-side corner
of the tip surface. As discussed in the previous chapter, the oil-dot technique also showed
a limiting streamline originating at about 0.20 Cax. Gap flow in this region comprises
mainly of the inlet flow, thereby generating lower wall shear on the tip surface. The
streaks that do form indicate turning of gap flow towards the camber-line as it progresses
to the suction-side of the tip gap. Blade tip and casing surface pressure measurements by
Xiao et al [54] in the AFTRF indicate minimal gap pressure differential in this region.
Similar behavior is also observed in linear cascade, shroud surface pressure
measurements by Bunker et al [22] and numerical simulation of flow in the Bunker
cascade by Ameri [29]. The numerical simulation also shows that gap flow streamlines in
the front quarter of the tip surface originate primarily from the rotor inlet flow. A definite
reattachment and recirculation pattern is observed on the tip surface from about 0.3 Cax,
extending the length of the tip surface.
Enlarged images of the surface patterns are shown in Figure 4.3 and Figure 4.4 for
clarity. A dashed curve near the leading edge of the tip surface, in Figure 4.3, indicates
the oil streak due to inlet flow passing through the tip gap. Close to the suction surface of
the gap the streak curves towards the exit and this is likely due to the gap flow
responding to the chord-wise pressure gradient set up on the suction surface. The
84
continuous curve traced on the tip surface, in Figure 4.3, is a limiting streamline between
flow towards the blade suction-side corner and re-circulating flow towards the pressure-
side corner. Oil carried by the leakage flow entering the tip gap, around the pressure-side
corner is deposited on to the tip surface. Subsequently, wall shear forces the oil towards
either the suction-side corner or the pressure-side corner. Wall shear near reattachment is
low and hence, according to Yang [55], is expected to yield a wide zone of oil
accumulation. In the present study however, the position of the reattachment is believed
to be fairly accurate due to the bifurcation pattern observed. Chord-wise orientation of
streaks at reattachment may indicate a three-dimensional reattachment. The direction of
streaks in the recirculation zone indicates that flow within the recirculation bubble is
towards the trailing edge. The recirculation patterns between 0.4 Cax and 0.5 Cax are not
as well defined. The smearing could be a result of the separation bubble lifting off the tip
surface and being ejected into the passage, as conceptualized by Bindon [7].
Measurements by Xiao [3] in the AFTRF show tightly spaced pressure contours in this
region of the tip gap. The rapid acceleration of fluid entering the gap is accompanied by
increasing wall shear, causing more of the visualization material to be washed away.
Immediately behind this smeared region and in the direction of the leakage flow is an
area that contains very little visualization material. Since visual observations indicated no
wetness, due to oil, it is possible that most of the oil carried into the gap does not reach
the tip surface in this region.
The features from blade mid-chord to the trailing edge are very similar, oil
accumulation at the pressure-side corner, a reattachment line on the tip surface, and over
tip leakage flow towards the suction-side corner are observed in Figure 4.4. The near
85
pressure-side corner oil accumulation is caused by both the re-circulating flow and the
separated flow in this region. The location of the reattachment line normal to the
pressure-side corner changes very little along the blade length. Similar behavior was
observed with the line denoting chord-wise flow in the previous chapter. It is believed
that the limiting streamline obtained through the oil-dot visualization occurs between the
recirculation bubble and flow within the separation bubble near the pressure-side corner.
Reattachment occurs at about 4% Cax (2*gap height) from the pressure-side corner
between 0.3 Cax and 0.65 Cax. The distance decreases at about 0.75 Cax along PS edge,
where the reattachment occurs at about 3% Cax. The reattachment occupies half of the
gap at about 0.8 Cax and the gap flow is fully separated in the last 5% of the blade chord.
The oil streaks leading towards the suction-side corner are more or less normal to
the camber-line up to about 0.8 Cax along camber-line, after which the streaks are turned
more towards the camber-line. An interesting observation is the path followed by the oil
after impinging on the tip platform. The turning of the streaks is seen to be more gradual
for flow towards the suction-side corner than in the direction of the bubble. Therefore, the
re-circulating flow experiences greater acceleration towards the pressure-side corner.
This behavior is consistent with the significantly low wall static pressure that occurs in
the near pressure-side corner of the tip surface, as measured by Xiao [3] in the AFTRF.
After about 0.6 Cax the streaks turn more sharply towards the suction-side corner, due to a
drop in chord-wise pressure gradient in the region of separated flow on the tip surface.
The surface streamlines observed are similar to the streak patterns, on a plane parallel to
and very close to the tip surface, obtained from a steady computational simulation by
Prasad and Wagner [56].
86
Figure 4.2: Surface Flow Patterns on Tip Surface of Blade (B21) With a Gap Height of t/h = 1.40%, Visualized Using the Oil-Film Technique.
Figure 4.3: Surface Flow Patterns on Front Half of Tip Surface of Blade B21, Using the Oil-Film Technique.
Recirculation
Normal Leakage Direction
Reattachment Inlet Boundary Layer Fluid
Accumulation Due to Separation
Normalized distance markers
Chord-wise Flow
87
4.1.2.3 Heat Transfer Implications
The tip surface flow patterns observed also have important heat transfer
implications. Low heat transfer areas may be expected in about the first 1/3rd of the tip
surface due to inlet boundary layer flow through the gap. Impingement of over tip
leakage flow causes higher heat transfer to the region near the pressure-side corner of the
blade. Acceleration of fluid into the recirculation bubble results in higher wall shear,
enhancing heat transfer to the tip surface around the reattachment line. Tip heat transfer
coefficients measured by Kwak et al [57] in a linear cascade show a region of maximum
heat transfer coefficient beginning at the pressure-side corner and extending towards the
blade suction surface. This region also spans most of the profile length. Additionally,
chord-wise flow in the recirculation bubble could trap high temperature fluid entering the
blade row, thereby exposing the near pressure-side corner of the tip surface to high total
temperatures. Previous hot streak studies, for example Prasad and Hendricks [58],
indicate that the maximum temperature in the core of the passage flow entering the blade
Figure 4.4: Surface Flow Patterns on Rear Half of Tip Surface of Blade B21.
Reattachment lineFully separated flow region0.95Cax
Reattachment lineFully separated flow region0.95Cax
Gap flow turning towards T.E.
Normal gap flow direction.
88
row usually migrates to the PS corner of the blade tip. Fully separated gap flow, as
observed in the last 5% blade chord, would lead to decreased heat transfer coefficients
due to low momentum activity in this region.
4.1.3 Tip Surface Flow Visualization; Small Gap Height (t/h = 0.71%)
The tip surface flow patterns on blade B7 (t/h = 0.71%), with half the clearance
gap as that of the test blade (B21) are shown in Figure 4.5. The patterns on blades B21
(t/h = 1.4%), B7 (t/h = 0.71%), and B2 (t/h = 0.81%) were obtained in a single test.
Minimal streaks are formed in the region up to about 0.25 Cax along the pressure-side
edge. The existing streaks are curved, indicating acceleration towards the suction-side of
the gap. Normal to 0.35 Cax along the pressure-side, a circle identifies a region of
smeared surface lines, similar to that shown in Figure 4.3 and attributed to the possible
ejection of the separation bubble. Thus, there seems to be a shift towards the leading edge
of some of the consistent flow features when the gap height is reduced.
The region near the pressure-side corner indicating reattachment and recirculation
is reduced at the smaller clearance gap, which causes greater oil accumulation along the
pressure-side edge. Results shown in Chapter 3, where the chord-wise, limiting
streamline formed closer to the pressure-side edge, support this observation.
Reattachment occurs at about 2% Cax (2.1*gap height) from the pressure-side edge. In a
direction normal to the pressure-side corner and in the region between 0.65 Cax and 0.7
Cax along the pressure-side there is a region of higher shear, as evidenced from the
observation that more of the tip surface is visible. This is not observed for large tip
89
clearance (t/h = 1.4%) and hence would seem to be characteristic of smaller clearance
gaps. In the oil-dot visualization shown previously it was observed that the only streaks
seen on the tip surface were in this region. As will be shown later in the paper, suction
surface traces show that significant leakage flow appears to enter the passage just
upstream of this region. Hence, the higher velocity and shear, in this region could be due
to proximity of the leakage vortex to the suction surface. Measured blade pressure
coefficient in Xiao, et al [54] shows a significant dip in the suction surface pressure
around 0.6 Cax. The return of dense streak patterns to the left of this region would
indicate the movement of the leakage vortex away from blade suction surface. The
leakage vortex forms farther away from the suction surface when the gap height is large
(t/h = 1.4%) and hence has a reduced influence on the gap flow.
There is no indication of a fully separated flow over the tip surface at the small
gap height, as even at the trailing edge there is a distinct direction to the streaks
indicating flow post reattachment towards the suction-side edge. Oil Streaks closer to the
trailing edge are however not oriented normal to the camber-line. This is attributed to the
increased tangential momentum of the leakage flow as it enters the tip gap. It was shown
earlier that pressure surface patterns qualitatively indicate increased wall shear due to
passage flow. The gap flow starts to turn away from normal to camber-line at about 0.8
Cax along the camber line. Bindon [7] indicates that in the near TE region of the tip, flow
in the separation bubble may be moving away from the trailing edge. The orientation of
streaks in the recirculation zone indicates however that the flow is indeed towards the
trailing edge. While it is not clear from the oil patterns, it is expected that the bubble is
ejected into the blade wake.
90
4.1.4 Tip Surface Flow Visualization of Other Gap Heights
Visualization was also carried out for two other gap heights, of t/h = 0.81% and
t/h = 1.2%. The tip surface patterns are shown in Figure 4.6 and Figure 4.7 respectively.
A contoured, precision plastic layer was applied to the tip surface of B21 with double-
sided adhesive tape, to obtain the gap height of t/h = 1.2%. Blade B2 was used without
modification for t/h = 0.81%.
At both gap heights the oil accumulation at pressure-side corner is observed,
along with the re-circulation. The differences between gap height 0.71% and 0.81% are
minimal. It does appear that the reattachment line is a little farther away from the
pressure-side corner, as the recirculation patterns are better defined at this gap height.
There is a region of greater oil removal from 0.65 Cax to 0.8 Cax, similar to that observed
at the smallest gap height. Gap flow is still fully attached near the trailing edge.
Figure 4.5: Surface Flow Patterns on Tip Surface of Blade (B7) With a Gap Height of t/h = 0.71%, Visualized Using the Oil-Film Technique.
0.1
0.2 0.3
0.4 0.5 0.6
0.7
0.8
0.9
1.0
Region of higher wall shear stress
91
The tip surface patterns in Figure 4.7 describe a flow field similar to that
discussed up to now. With increase in gap height the recirculation patterns tend to be
more distinct, as the reattachment line moves farther away from the pressure-side corner
leading to lesser accumulation on the tip surface. The high shear area identified at the
smaller gap heights is not evident on the tip surface. There are indications of separated
flow in the last 5% blade chord.
The distance of the reattachment line from the pressure-side corner was measured
on the images at 0.6 Cax and is presented in Figure 4.8. The distance is measured normal
to the axial chord (lx) and also normal to the camber-line (lc). The measurement is
rendered non-dimensional with respect to the blade tip axial chord length and the gap
height. The variation of the measurement with gap height is plotted. The variation with
gap height of the distance normalized by blade tip axial chord is linear. The exception to
this behavior is the measurement at t/h = 0.81%, which is believed to be due to
measurement accuracy. The linear fit is not unexpected as previous research has indicated
that characteristic measures of OTL, such as gap mass flow rate, total pressure losses and
efficiency drop, vary linearly with gap height. The location of the reattachment line
measured normal to camber-line and normalized with gap height does not change much
with gap height. Indeed the outlier in this case is again the measurement at t/h = 0.81%.
Thus it can be concluded that the reattachment occurs about 2*t from the pressure-side
corner. Of course it must be noted that this might only be valid for a range of gap
pressure difference.
92
Figure 4.6: Surface Flow Patterns on the Tip Surface of Blade (B2) With a Gap Height of t/h = 0.81%, Visualized Using the Oil-Film Technique.
Figure 4.7: Surface Flow Patterns on Tip Surface of Blade (B21 at reduced gap height) With a Gap Height of t/h = 1.2%, Visualized Using the Oil-Film Technique.
0.1
0.2 0.3
0.4 0.50.6
0.7
0.8
0.9
1.0
Region of higher wall shear stress
Re-attachment line
Accumulation due to corner separation
0.1
0.2 0.3
0.40.5 0.6
0.7
0.8
0.9
1.0
More distinct re-attachment line
Accumulation due to corner separation
93
4.1.5 Suction Surface Traces from Oil Film Based Tip Surface Flow Visualization
Visualization material carried across the tip surface is deposited on the blade
suction surface in the form of a trace, similar to that observed in the case of rotor endwall
flow visualization. Images of the suction surfaces of the blades at all gap heights were
also acquired and are shown in Figure 4.9 and Figure 4.10. The oil that is deposited on
the suction surface is carried over by the leakage flow as it exits the gap suction-side and
enters the passage. It was shown in Chapter 3 that visualization material from oil film on
0.0%
2.0%
4.0%
6.0%
8.0%
10.0%
12.0%
0.5% 0.7% 0.9% 1.1% 1.3% 1.5%
Tip Gap Height, t/h [-]
Loc
atio
n of
Rea
ttac
hmen
t, [-
]
0.00
0.50
1.00
1.50
2.00
2.50
3.00
Loc
atio
n of
Rea
ttac
hmen
t, l c
/ t [-
]Measured normal to axial chord, lx/t
Measured normal to camber-line, lc/t
lc / t
Figure 4.8: Influence of Gap Height on the Location of the Visualized Reattachment Line on the Blade Tip Surface.
94
the rotor endwall was forced up the blade suction surface and resulted in a similar trace.
This was attributed to visualization material being transported in the interaction zone
between the tip leakage vortex and the tip passage vortex. It must be reiterated that the
oil-pigment mixture used for gap heights of 0.71%, 0.81%, and 1.40% was from the same
batch. It appears that as the gap height is decreased the start point of the trace is better
defined. The traces appear to originate on the suction surface in the region 0.65 Cax to
0.75 Cax. McCarter [16] showed through velocity measurements in the AFTRF that the
leakage flow at 0.7 Cax is very close to the suction surface and develops rapidly at 0.8
Cax.
The distance from the tip surface to the upper boundary of the trace (h1) and width
of the trace (h2) were measured off the blade surfaces and are presented in Figure 4.11 as
a function of the gap height. The measurements are normalized by the blade height (h)
and both measurements vary linearly with gap height. As the gap height is reduced the
lower boundary of the trace moves up towards the blade tip and the width of the trace
grows smaller. Thus, it might be concluded that as the gap height is decreased the leakage
vortex is closer to the blade surface, as well as closer to the blade tip. The measurements
for blade B7 (t/h = 0.71%) are smaller than that observed in Figure 3.7 and this is
attributed to the origin of the visualization material. In the results presented here the oil is
carried by the gap flow, while in Figure 3.7, visualization material is present on the
suction surface, in addition to OTL capturing oil thrown off the blade suction surface.
95
Figure 4.9: Suction Surface Traces Formed During Oil-Film Based Tip Surface Flow
Visualization; Gap Heights, t/h = 1.4% and t/h = 1.2%.
Figure 4.10: Suction Surface Traces Formed During Oil-Film Based Tip Surface Flow
Visualization; Gap Heights, t/h = 0.81% and t/h = 0.72%.
h2
h1
96
4.2 Total Pressure Measurements
Quantitative results from total pressure measurement are presented in this section.
Total pressure measurements downstream of the rotor are recorded for all 29 blade
passages using a phase-locked measurement technique. The results are presented in form
of contour plots that show the total pressure distribution in the r-θ plane, radial
distributions of passage averaged and rotor averaged coefficients, circumferential
distribution of total pressure coefficient at individual radial locations, and area averaged
total pressure coefficient for individual passages. The relevant equations are in Section 2..
The term “Base#” is used to refer to tests conducted without coolant injection and with
0.02.04.06.08.0
10.012.014.016.0
0 0.5 1 1.5Gap Height, (t/h %)
SS T
race
Mea
sure
men
t, (h
1 /h %
; h2
/h %
)
h1/h
h2/h
Figure 4.11: Effect of the Tip Gap Height on Measurements From Suction Surface Trace
Formed During Oil-Film Based Visualization of Tip Surface Flow.
97
the cooled blade B21 gap height of t/h = 1.40%. Initial data sets cover a span from 0.051h
to 0.981h. This gives an overall picture of passage flows, including near rotor endwall
flow. Since variations below 0.438h were found to be unaffected by gap height, within
uncertainty limits, most of the data sets cover a span from 0.438h to 0.981h.
The radial distributions of the passage averaged and rotor averaged total pressure
coefficients are plotted as continuous curves, rather than points. The curve itself is
generated by a three point moving average of the data with the end points unchanged.
The passage averaged coefficient is computed for the passage that is bounded on the
suction surface by the blade number referenced. So, a passage averaged coefficient for
B21 is the measured total pressure coefficient averaged across the passage that is
bounded by the suction surface of B21 and pressure surface of B20. Thus, the passage
defined contains the leakage vortex of B21. The passage is defined to start at the edge of
blade wake and core passage flow and contain 207 points. The circumferential
distributions obtained at a fixed radial position are termed as “wake plots” throughout the
manuscript.
4.2.1 Baseline, No Injection
Baseline tests were conducted with a test blade (B21) tip gap height of t/h =
1.40%. The rotor tip clearance distribution is TCL1, as shown in Figure 2.5. Thus, the
baseline is relevant to only the isolated coolant injection study. The measurements were
conducted using the high response total pressure probe described in Section 2.6.2.1. The
probe is in the stationary frame of reference and samples the total pressure field 6000
98
times for each revolution of the rotor. Thus a complete map of the rotor exit total pressure
field is obtained, an example of which is briefly discussed in Appendix C. The contour
plots presented show only a sector containing the cooled blade.
Figure 4.12 is a contour plot of the total pressure coefficient, as defined in
Equation 2-3. Radial locations corresponding to the hub, the casing, and 50%, 75%, and
85% blade span are shown in solid curves. The number (#) and gap height (t/h) of each
blade is shown above the casing boundary. The direction of rotation is from right to left,
as represented by the blade speed vector below the hub boundary. The total pressure
distribution over five passages, with the cooled blade B21 at the centre, shown is
designated Base1. Solid boundaries in the tip region and dashed curves in the mid-span
region are drawn to visually track changes in the total pressure field, which is possible
due to the phase-locked measurement technique employed. These details are consistently
reproduced in all subsequent contour plots.
Figure 4.12: Total Pressure Coefficient Contours With No Coolant Injection (Base1).
99
4.2.1.1 Region above 85% Blade Height
This region is dominated by two distinct flow regimes. The tip leakage vortex,
seen in the wake region of each blade, is characterized by low total pressure and a near
elliptical shape. The shape results from the total pressure data being projected onto a
plane normal to the rotor axis. The leakage vortices in each blade passage are seen as the
lowest total pressure regions in the passage. The size of the leakage vortices and the total
pressure defect increase with increasing gap height. The entrainment of low momentum
fluid by the leakage flow and diffusion beyond the rotor exit plane has caused the blade
wake to be completely overshadowed by the leakage flow.
The leakage vortex of B21 occupies about 15% of the blade span and extends well
into the blade passage. The vortex has a well defined core and strong gradients in both
radial and circumferential directions. The total pressure at the core is approximately 1*qm
lower than the maximum total pressure coefficient measured. The minimum test blade
total pressure, measured in the core of the tip vortex, is also lower than that of the
neighboring blades shown because of the much larger gap height of this blade.
The second distinct flow feature in the last 15% blade height is the intermediate
total pressure zone between the two adjacent tip vortices. This zone (Cpt=-3.85, green) is
bounded by the outer casing that is in relative motion with respect to the rotating passage,
the core flow in the middle of the passage and the two subsequent tip vortices. The
secondary flow pattern near the outer casing is more distorted in comparison to its
counterpart near the hub surface because of the leakage flow. Velocity measurements by
McCarter et al. [20] show higher axial and reduced tangential velocities in this region.
100
This is consistent with under-turning of the passage flow due to the blockage presented
by the tip leakage vortex. Such under-turning would be the greatest in the passage
containing the test blade leakage vortex, and hence the measured total pressure would be
lower than that of neighboring passages.
4.2.1.2 Region 75% - 85% Blade Height
The main flow feature present in this region is the tip-side passage vortex, as
marked in the figure. The identity of this structure is also confirmed by results in
McCarter et al [20]. It exists just below the tip leakage vortex and has a smaller total
pressure defect than that of the leakage vortex. The interaction of the passage vortex with
the leakage vortex is greatest for the test blade, as seen by the poor definition of the
passage vortex in the wake. Another interesting characteristic of this region is a band of
lower total pressure that runs across the passage, culminating in the passage vortex. This
suggests that the secondary flow effect is greatest in this region, with the flow probably
being over-turned as it migrates towards the suction side of the passage.
4.2.1.3 Passage Core
The passage core occupies about 50% of the passage. It has a well defined core of
high total pressure at around 50% blade height. Just below 75% span, close to the blade
suction-side, a dip in the passage core is observed. This is possibly due to a vortical
structure. The extent of the dip suggests that this structure is formed early in the blade
101
passage and is convected downstream. Below 50% span the passage core is strongly
affected by the hub-side passage vortex, which rotates in the same direction as the
leakage vortex. Although the hub-side passage vortex is not as clearly discernible as the
tip-side passage vortex, its effect on the passage core is clear. A better definition of the
hub-side passage vortex is clearly observed in the past, phase-locked LDA measurements
shown in Figure 4.13, obtained by Ristic et al. [59] in the same facility. It must be noted
that the nominal absolute flow angle at the hub is 35°, which is a 10° incidence on the
probe. Additionally, the passage flow tends to be over-turned closer to the hub, due to the
rotor endwall passage vortex. In the passage containing the test blade, the passage core is
shifted radially downwards, probably by the radially inward migration of fluid due to the
blockage of the tip leakage vortex. The core in the passage formed by B21 & B22 is at a
higher radius, suggesting a strong radial flow towards the tip due to the enlarged
clearance of the test blade.
Figure 4.13: Secondary Flow Vectors At Rotor Exit From LDA Measurements by Ristic
et al [54].
Hub-side passage vortex
102
4.2.2 Repeatability
Repeatability of baseline data is presented through contour plot in Figure 4.14,
radial distribution of rotor averaged pressure coefficient in Figure 4.15, wake plots at two
radial locations in Figure 4.16 and Figure 4.17, and radial distribution of passage
averaged coefficient in Figure 4.18. A total of three baseline data sets were acquired over
a six month period. Figure 4.14 is a contour plot from the last data set, designated Base3,
shows good comparison of the total pressure coefficient distribution with Base1. While
the radial extent of the data set is shortened, the magnitude of total pressure coefficient
and the geometric definition of flow structures match well with that shown in
Figure 4.12. The radial distribution of circumferentially averaged total pressure
coefficient, along all 29 passages in the rotor, also shows good repeatability. In the tip
vortex dominated zone and core of the passages the repeatability is sufficiently
maintained for three baseline experiments performed at different dates. While general
repeatability is good, greater deviation occurs for Base1 in the region between 70% and
80% blade height, which in some instances is at the limit of the uncertainty band. Overall
repeatability for the measurements is deemed to be good within a 95% confidence level.
The circumferential distribution of total pressure coefficient at selected radial
locations of r/h = 0.96 and r/h = 0.57 is shown in Figure 4.16 and Figure 4.17,
respectively. These locations correspond respectively to the location of the leakage vortex
dominated zone and the core flow, where minimal influence of the leakage vortex is
expected. Data shown extends from blade B19 to blade B23. Overall, the repeatability is
very good at both radial locations. The influence of the test blade, with a relatively large
103
tip clearance of t/h = 1.40%, is noticeable in Figure 4.16 as the largest total pressure
defect induced by the large tip leakage vortex. However, when the same distribution at a
lower radius (r/h=0.57) is examined, the tip vortex influence diminishes.
Radial distributions of circumferentially averaged total pressure coefficient for the
passage defined by blade B20 and blade B21 (test blade) are shown in Figure 4.18 . The
distribution is very similar to the results from all 29 passages (rotor averaged) except that
the tip region results are influenced by the large tip clearance of blade B21 (t/h=1.40%).
The effect of the increased clearance of the test blade is smoothed out in the rotor
averaged coefficient, and hence the peak in the passage averaged coefficient distribution
for r/h>0.85 may be interpreted as a total pressure defect zone. As in the case with rotor
averaged coefficient, the repeatability between the three baseline sets is very good, except
in the region between 70% and 80% blade height where Base1 displays a greater
difference, which in some instances is near the limit of the uncertainty band.
Figure 4.14: Total Pressure Coefficient Contours With No Coolant Injection (Base3).
104
4.2.3 Effect of the Tip Gap Height
Figure 4.19 shows a contour plot of total pressure coefficient obtained with the
test blade tip clearance is reduced to t/h=0.72%, from t/h=1.40%. The clearance was
reduced by applying precision plastic layer on to the tip surface with double-sided
adhesive tape. The effect of reducing the tip clearance is evident in the distribution and
level of measured total pressure at stage exit. The solid lines marked around the tip vortex
zones are from the baseline case and used for comparison purposes. The tip leakage
vortex of cooled blade B21 is much smaller in size and closer to the suction-side of the
Figure 4.15: Radial Distribution of the Rotor Averaged Total Pressure Coefficient;
Baseline Repeatability.
Uncertainty band
105
passage. The total pressure measured in the vortex is higher than that of the baseline by
about 0.2 qm. The passage flow above 85% span should experience lesser under-turning
and hence register as a higher absolute total pressure, in comparison to the baseline. The
tip-side passage vortex of blade B21 is better defined due to reduced interaction with the
leakage vortex. The passage core itself has rotated in a counterclockwise direction due to
the reduced blockage of the leakage flow. Importantly, the tip leakage vortices in the
neighboring passages are unaffected and hence the influence of changing the gap height
appears to be localized to the passage bounded by blades B20-B21.
The wake plot at r/h = 0.96, presented in Figure 4.20 clearly shows the movement
of the wake of blade B21 towards the suction surface of the blade. The blue triangles
indicate a significant total pressure defect in the tip leakage vortex when the clearance is
large, at t/h=1.40%. The total pressure defect is reduced (orange squares) and the
magnitude is similar to that of the neighboring blades. The wake plot at r/h=0.57,
Figure 4.21, shows a slight drop in the total pressure at the mid-span. This is probably
due to the aforementioned rotation in the passage flow due to a reduction in the influence
of the leakage flow on the passage flow. Tip vortex impact at this lower radial position is
not measurable in the passage of the test blade.
The radial distribution of the passage averaged pressure coefficient in Figure 4.22
shows that there is a definite shift in the flow field towards the blade tip, in addition to
the increase in total pressure. This confirms that the lower total pressure seen in the wake
plots is due to a change in the passage flow structure. The defect due to the tip-side
passage vortex is also better defined than in the baseline sets because the tip
vortex/passage vortex interaction is reduced when the tip gap is reduced to t/h=0.72% .
106
Figure 4.23 compares the radial distribution of the passage averaged coefficient
from two tests. The curves in RED were obtained with tip gap height of blade B21 at t/h
= 1.40%, while the curves in GREEN were obtained with tip gap height of blade B21 at
t/h = 0.72%. The passage averaged coefficient of passage containing the tip leakage
vortex of blade B7 (t/h = 0.71% in dash-dot curves) is compared with that of passage
containing the tip leakage vortex of blade B21 (t/h = 1.40%, continuous and dashed
curves). Blade B7 is diametrically opposite to blade B21 in the 29 blade rotor. The
repeatability of the radial distributions for blade B7 is good, as expected, since the effect
of changing the gap height of blade B21 is expected to create changes locally. The radial
distribution of blade B7 is almost identical to that of blade B21 with a tip gap height of
t/h = 0.72%. The tip gap heights of neighboring blades B6 is t/h = 0.73%, B8 is t/h =
0.71% , B20 is t/h = 0.77%, and B22 is t/h = 0.83%. Thus, it appears that the gap height
of neighboring blades has little influence on over tip leakage flow, as long as the
variation in gap heights is within 15%.
A comparison of the area averaged coefficient as a function of tip gap height is
shown in Figure 4.24. The area average is computed using Equation 2-8, over one
passage containing the tip leakage vortex of blades with varying tip gap heights and a
height of 20% span, from 0.8h – 1h. The variation of the area averaged coefficient is
quite linear. The maximum deviation of the area averaged coefficient based on the curve-
fit, from the actual measured data is 0.2% of the measurement at the tip gap height of t/h
= 0.92%. This trend agrees well with that observed in the measurements from flow
visualization.
107
Blade Number
Tota
lPre
ssur
eC
oeff
icie
nt,C
pt
19 20 21 22 23-4.6
-4.5
-4.4
-4.3
-4.2
-4.1
-4
-3.9
-3.8
-3.7
Base1: t/h = 1.40%, Minj = 0Base2: t/h = 1.40%, Minj = 0Base3: t/h = 1.40%, Minj = 0
Figure 4.16: Repeatability of Wake Profiles at r = 0.96h With No Coolant Injection.
Blade Number
Tota
lPre
ssur
eC
oeff
icie
nt,C
pt
19 20 21 22 23-4.3
-4.2
-4.1
-4
-3.9
-3.8
-3.7
-3.6
-3.5
-3.4
-3.3
Base1: t/h = 1.40%, Minj = 0Base2: t/h = 1.40%, Minj = 0Base3: t/h = 1.40%, Minj = 0
Figure 4.17: Repeatability of Wake Profiles at r = 0.57h With No Coolant Injection.
108
Figure 4.18: Repeatability of the Passage Averaged Coefficient For Cooled Blade B21.
Figure 4.19: Total Pressure Coefficient Contours With Tip Gap Height of Cooled Blade B21 Reduced to t/h = 0.72%.
Uncertainty band
109
Blade Number
Tota
lPre
ssur
eC
oeff
icie
nt,C
pt
19 20 21 22 23-4.6
-4.5
-4.4
-4.3
-4.2
-4.1
-4
-3.9
-3.8
-3.7
Base3: t/h = 1.40%, Minj = 0t/h = 0.72%, Minj = 0
Figure 4.20: Effect of Reducing the Tip Gap Height of Blade B21 On the Wake Profile at
r = 0.96h.
Blade Number
Tota
lPre
ssur
eC
oeff
icie
nt,C
pt
19 20 21 22 23-4.3
-4.2
-4.1
-4
-3.9
-3.8
-3.7
-3.6
-3.5
-3.4
-3.3
Base3: t/h = 1.40%, Minj = 0t/h = 0.72%, Minj = 0
Figure 4.21: Effect of Reducing the Tip Gap Height of Blade B21 On the Wake Profile at
r = 0.57h.
110
Figure 4.22: Effect of the Tip Gap Height On the Passage Averaged Coefficient of
Cooled Blade B21.
Figure 4.23: A Comparison of the Passage Averaged Coefficient Distribution For Blade
B7 and Blade B21.
-4.16 -4.04
111
-4.11-4.10-4.09-4.08-4.07-4.06-4.05-4.04-4.03
0 0.5 1 1.5 2Tip Gap Height, (t/h %)
Are
a A
vera
ged
Coe
ffic
ient
, Cpt
,A
Figure 4.24: Variation in the Area Averaged Total Pressure Coefficient with Tip Gap Height. (Area = 20% span*1 passage).
112
Chapter 5
Effect of Coolant Mass Flow Rate on Over Tip Leakage
The effect of coolant injection on over tip leakage is studied in two parts. This
chapter presents the effect of coolant mass flow rate on over tip leakage flow. Coolant is
injected from a single blade, referred to as the test blade (B21), and the series of tests are
referred to as isolated injection. Precision plastic layers were attached with double-sided
adhesive tape to the tip surfaces of the four other blades with injection capability. This
makes the four blades inactive from the coolant injection point of view and also changes
the rotor clearance distribution. The clearance distribution with these modifications is
TCL1, as specified in Figure 2.5. It was shown in Chapter 4 that the total pressure
signature of the tip leakage vortex of cooled blade B21 was not affected by the gap size
of blades preceding it. A single large disturbance was also preferable to maintain flow
repeatability of the rotor and allows for the measurement of the effect of an isolated blade
injecting coolant into the tip gap.
The effect of coolant mass flow rate is investigated both qualitatively, using flow
visualization, as well as quantitatively by measuring total pressure downstream of the
rotor exit. The mass flow rate of coolant injection into the tip gap is an important
parameter. Air used for turbine hot-section cooling is high-pressure air bled from the
engine compressor section. Thus, a part of total cycle work is trapped in the coolant air.
Since coolant does not participate in the work generation, cooling air represents a loss in
113
efficiency or in available work. This loss must however be compared to the benefits that
are accrued from coolant injection, for example, air used in internal blade cooling and
film cooling of turbine blades allows turbine inlet temperatures to be higher than metal
melting point, which in turn increases the thermal efficiency of the engine. In the case of
tip injection the efficiency lost due to injecting coolant into the tip gap must be balanced
against loss reduction in the high-pressure turbine stages due to its effect on the OTL
flow, as well as the heat transfer benefits. After the leading edge, the tip surface is the
hottest part of a turbine blade and heat transfer benefits might indicate that higher inlet
temperatures are possible, thereby increasing the overall thermal efficiency of the gas
turbine engine.
5.1 Visualizing the Effect of Coolant Injection
The flow visualization technique employed to study OTL flow was described in
the previous chapter. The advantage of using flow visualization with coolant injection is
that in addition to obtaining information on the fluid dynamic interaction between the
coolant jets and gap flow, potential heat transfer benefits may be identified, in the
rotating frame. As before the blade surfaces were coated with smooth flat black paint
prior to application of the visualization material. The optical window was then fastened
and the first fan was started, followed by the second fan in 30 seconds. The rotor speed
was increased to operating speed after flow through the facility had stabilized.
Coolant mass flow rate injected is stated as a percent of turbine mass flow rate,
assuming coolant is injected from all blades as shown in Equation 2-2. Four injection
114
mass flow rates, 0.4%, 0.5%, 0.6%, and 0.7% were tested. In all cases, except one,
coolant injection was initiated and stabilized prior to start-up. For the case of Minj = 0.4%
however, a test was conducted where injection was commenced after the rotor had
reached operating speed. The various test cases are summarized in Table 5-1 .
5.1.1 Effect of Tip Trench
Tip surface flow over a flat tip at a gap height of t/h = 1.40% was discussed in
Chapter 4. The effect of the tip trench was investigated using flow visualization, as
shown in Figure 5.1, at the gap height of t/h = 1.40%. As compared to the flat tip surface
flow patterns Figure 4.2, there is, as expected, little change in surface flow in the front
half of the blade. A distinct reattachment line starts to form near 0.35 Cax and the re-
circulation pattern is clearly visible. The reattachment line intersects with the near
pressure-side edge of the tip trench at about 0.6 Cax. In the region from 0.5Cax, where the
Table 5-1: Test Matrix of Flow Visualization with Coolant Injection.
Test Case Blade # t/h% Minj% Description
T1 21 1.40 0 Tip with trench
T2 21 1.40 0.4 Injection at N = 0
T3 21 1.40 0.4 Injection at operating speed
T4 21 1.40 0.5 Injection at N = 0
T5 21 1.40 0.6 Injection at N = 0
T6 21 1.40 0.7 Injection at N = 0
115
trench starts, to 0.55Cax the streaks in the leakage flow direction deposit visualization
material in the trench. It appears that gap flow does not reattach on the tip surface from
0.55 Cax to the trailing edge.
The streak pattern between the trench and the pressure-side corner indicates
recirculation up to H1, suggesting that the gap vortex is still present and is separated from
the tip surface. Recirculation is also observed beyond H2 at 71% chord. Between H1 and
H2 the region between the trench and the pressure-side corner shows more accumulation
than on the flat tip. It is likely that the trench weakens the recirculation and hence reduces
wall shear stress in this region. There is no accumulation in the trench, except close to 0.5
Cax, another indication that the gap vortex is lifted off the tip surface. One of the
interesting changes from the flat tip is the region of low oil concentration to the left of
H1. The reason for this is unknown and it is not observed at the other injection locations.
A limiting streamline, similar to that seen with oil dot visualization, is observed on the tip
Figure 5.1: Surface Flow Visualization of the Effect of Tip Trench on Cooled Blade B21.
H1H2H3H4
Oil in trench RecirculationChord-wise flow
0.1
0.2 0.3
0.40.50.6
0.9
1.0
116
surface injection hole H2. In visualization without the trench, recirculation due to the gap
vortex accumulates visualization material near the pressure-side corner of the blade and
likely prevents the formation of a clear chord-wise limiting streamline. Hence, this
observation confirms that the gap vortex influence on the tip surface is reduced.
5.1.2 Injection at Minj = 0.4% at Gap Height of t/h = 1.40%
This particular injection case was tested in two different ways, as described in
Table 5-1. Figure 5.2 shows the surface flow patterns when injection was commenced
before startup, while Figure 5.3 was obtained for injection after the turbine rotor reached
operating speed.
5.1.2.1 Injection Prior to Start-up
Figure 5.2 shows the surface patterns on the blade tip with injection before
startup. Surface flow patterns up to about 0.6 Cax along the PS edge are similar to those
observed for tip surface with trench and no injection, indicating no effect of coolant
injection in this region. A re-circulatory pattern appears around 0.35 Cax and extends up
to 0.6 Cax. Beyond this point the observed patterns are distinctly different due to the
effect of the coolant jets. The pressure-side corner accumulation is intermittent. Jet
penetration into the leakage flow, due to coolant injection from H1, leaves an arc-like
imprint in the accumulated oil, on the tip surface, between the trench and the pressure-
side corner. Separation is not completely eliminated due to the limited coverage
117
available. Additionally, divergence of the jets from each other allows oil to accumulate in
between the jets.
A large area with little to no streaks is seen between the trench and blade suction-
side corner, in the region surrounding H1. Thus the coolant jets spread out as they turn
and flow towards the suction-side corner of the blade. The leading jet does not lean
towards the trailing edge and appears to be turned back sharply. The trailing jet on the
other hand does have a 10° angle towards the trailing edge and is turned more gradually.
The lack of accumulation at 0.65 Cax along the pressure-side corner is due to the trailing
jet in H1 blocking the passage flow as it enters the gap. Once this jet is turned back it
covers a larger area on the tip platform than the leading jet. Immediately behind the
injection holes there are very few streaks, when compared to the no injection cases. The
streaks formed are located in between the injection holes and indicate low momentum
activity. The leakage flow that passes between the jets diffuses out on the tip surface and
Figure 5.2: Surface Flow Visualization of the Effect of Coolant Injection at Minj = 0.4% From Cooled Blade B21; Injection While Turbine at Rest.
Gap flow turning around coolant jet
Oil accumulates between jets
Little to no streaks
Trench accumulation
0.1
0.3 0.4
0.6
0.70.8
0.9
1.0
118
also gives up some energy in overcoming the shear imposed by the coolant jets. The
chord-wise flow on the tip surface would also be blocked by the coolant jets. No oil-dot
based visualization was conducted and hence it is difficult to say with any certainty the
path taken by this fluid in the separated flow region near the pressure-side corner of the
tip surface.
Accumulation of visualization material, due to pressure-side corner separation, is
observed between H1 and H2. The streaks in this region however are not as sharp as
those observed in the forward half of the blade and might be a result of visualization
material accumulated in the trench. It was shown earlier that with no injection there was
no accumulation in the trench and hence path of leakage flow is altered by the coolant
jets. Visualization material accumulates on the tip surface up to the trench and is
probably driven by gap flow forced in to this region by jet blockage. In the previous
chapter it was discussed that visualization material accumulates near the pressure-side
corner of the tip surface partly due to the separation effect and partly due to paint being
deposited by the re-circulating flow. The extent of deposition indicates that recirculation
is augmenting the material deposited by the separated flow.
The path of leakage flow around the leading jet from H2 is discernible by the
curvature of the streaks, particularly around the leading jet. The curvature suggests flow
of leakage fluid around the blockage presented by the coolant jets. The accumulation at
the pressure-side corner is lesser than that observed near H1. The jets are again turned
back, leaving a large clear region between the trench and the suction-side corner. There is
greater accumulation in the trench between H2 and H3 and consequently, the streak
patterns between trench and suction-side corner are indistinct. Tip separation is reduced
119
along the pressure-side corner from about 0.72 Cax to 0.85 Cax, the region of influence of
H2 and H3.
Injection from H3 is not much different in the effect it has on the gap flow. The
coolant jets start out towards the pressure-side and are turned to flow towards the suction-
side. The streak patterns between the trench and the suction-side corner, in the region
around H3, are considerably smeared and could indicate considerable mixing between
gap flow and coolant fluid. The location of this injection hole is close to the impingement
location of normal gap flow. The last injection hole (H4) is radial and hence allows for
separation to occur between 0.85 Cax and 0.9 Cax. It appears however to block the leakage
flow beyond 0.9 Cax, forcing the leakage to turn towards the blade trailing edge.
5.1.2.2 Injection at Operating Speed
Figure 5.3 shows tip surface flow patterns for the case when the injection is
initiated after the rotor reached operating speed. The patterns are qualitatively different.
In this case, by the time the injection is initiated, there is some deposition and shearing of
visualization material on the tip platform. Injection is fully stabilized 30 seconds after
rotor speed is set to operating speed. Subsequently, with injection it is expected that areas
of low flow momentum activity will display greater concentration of paint, while a lower
concentration of paint will indicate flow with greater momentum.
Interrogating the images with this perspective it is clear that both Figure 5.2 and
Figure 5.3 essentially represent the same flow patterns. Observations are confined to the
area beyond mid-chord. The first set of injection jets are turned towards the suction-side
120
by the pressure driven gap flow. Accumulation behind H1 is dense with no clear streak
pattern indicating low flow velocities in this region. It might be recalled that in the
previous discussion the patterns in this region were interpreted to show low wall shear
due to wake like flow after leakage fluid passes between the coolant jets. The
accumulation is particularly high behind the trailing jet of H1, which is expected since it
is along the path the leakage flow is expected to take. Penetration of the jets in to the
separation zone is indicated by the removal of visualization material close to the
pressure-side corner. To the left of H1 and between the trench and suction-side corner the
patterns are lighter than those observed in the forward part of the blade, which can only
happen if due to the greater wall shear generated by the coolant jets being turned around
to form a film over the tip surface. In Figure 5.2 this appears as a region clear of streaks.
Figure 5.3: Surface Flow Visualization of the Effect of Coolant Injection at Minj = 0.4% From Cooled Blade B21; Injection While Turbine at Operating Speed.
Accumulation due to low momentum activity
Coverage of coolant jets
121
There is very little accumulation in the trench, to the left of H1 this is expected as
injection was started after rotor reached operating speed. The accumulation that does
exist indicates that some visualization material is still being carried over from the
pressure surface even two minutes into the test. The effect of lower momentum flow is
also seen in the pattern between H2 and the SS edge. The accumulation of paint due to
lower momentum flow occurs to the left of the trailing jet, instead of behind it. The
patterns between H2 and H3, behind the trench are more distinct streak patterns than
those observed in the previous injection test. There is also some accumulation within the
trench. The mixing effect of the coolant jets is in clear evidence near the location H3. The
streaks that would have formed prior to injection, between the trench and the SS edge, are
completely smeared indicating much greater shear than without injection and hence
attributable to coolant injection. Accumulation along the pressure-side corner is
intermittent, indicating jet penetration into the leakage flow. To the left of the last
injection hole, at 0.95 Cax, a dense line of accumulation extends across the tip surface. In
Figure 5.2 this same feature is seen as a dark line due to lack of paint accumulation. This
line is probably caused by a change in the dominant shear direction due to the coolant jet
at H4. The dominant shear direction is expected to be normal to the pressure-side corner
and blockage due to the coolant jet at H4 causes the leakage flow to turn towards the
trailing edge.
One of the important pieces of information that this test yields is that the
visualization material is not dry and is able to trace out the path of the tip surface flow.
Since injection was commenced after the rotor reached the operating speed, if the
visualization material had dried out then the pattern should resemble that shown in
122
Figure 5.1 . However, the pattern is distinctly different and the physical reasoning behind
the appearance of the pattern is supported by Figure 5.2. Hence, it is concluded that the
visualization material is not dry, two minutes into the test and is influenced by wall shear.
5.1.3 Visualization at Other Injection Rates
The other injection rates tested were, Minj = 0.5%, 0.6%, and 0.7%. These results
are shown in Figure 5.4, Figure 5.5, and Figure 5.6, respectively. In each case, leakage
flow up to about 0.6 Cax is unaffected by the injection. While there is some change in the
density of streaks and their definition, the important features like reattachment and
recirculation are identifiable. The variations are believed to be caused partly by variations
in the mixture and partly by the operating temperature. These effects were not studied, as
it is believed that the available information is sufficient to make consistent observations.
Focusing on the flow patterns seen near injection location H1, higher jet
momentum has almost eliminated accumulation due to pressure-side corner separation.
Immediately behind the injection holes there is greater oil concentration than that seen for
the lowest injection rate, and this increases with increase in the injection rate. At the
highest injection rate of 0.7%, the streaks behind the injection holes show a converging
pattern. On either side of the jets, there is evidence of film formation, with the leading jet
displaying better coverage than the trailing jet. This is opposite to that observed for Minj =
0.4%. Additionally, it appears that the influence of the trailing jet begins farther
downstream along the blade chord than for Minj = 0.4%. This is again consistent in that
123
greater jet momentum will better resist turning due to leakage flow. There is also
significant accumulation in the trench.
Streaks formed due to injection from H2 and H3 are similar in nature to that near
H1. There is again very little accumulation near the pressure-side corner. The wake like
behavior of the streamlines between the trench and the suction-side corner is more
pronounced, especially near H2, as observed by the greater accumulation very near the
suction-side corner of the tip surface. In this region the leakage flow is expected to be
lower in momentum than that near H1 due to the natural reduction in driving pressure
differential across the tip surface. The most interesting change is exhibited by injection
from H3. The jet trajectories are defined better at the higher injection rates of Minj = 0.6%
and Minj = 0.7%. The leading jet appears to be more effective in covering the tip surface
than the trailing jet. The influence of the trailing jet is farther to the left of the jet.
Injection from H4 appears to stagnate around the hole and flow over the SS edge into the
passage.
Figure 5.4: Surface Flow Visualization of the Effect of Coolant Injection at Minj = 0.5% From Cooled Blade B21; Injection While Turbine at Rest.
PS corner accumulation eliminated
Trench accumulation
124
Figure 5.5: Surface Flow Visualization of the Effect of Coolant Injection at Minj = 0.6% From Cooled Blade B21; Injection While Turbine at Rest.
Figure 5.6: Surface Flow Visualization of the Effect of Coolant Injection at Minj = 0.7% From Cooled Blade B21; Injection While Turbine at Rest.
0.1
0.2 0.3
0.40.5 0.6
0.7
0.8
0.9
1.0
0.1
0.3 0.40.5
0.6 0.7
0.8
0.9
1.0
PS corner accumulation and streamlines near T.E.
125
5.1.4 Suction Surface Traces
The suction surface traces of the tip leakage vortex are shown in Figure 5.7 for
Minj = 0.4%, 0.5%, and Figure 5.8 for Minj = 0.6%, 0.7%. For Minj = 0.4% the streak
pattern is similar to that obtained without tip injection and is seen from approximately 0.6
Cax. This is actually closer to the leading edge than in the case of t/h = 0.71%. It appears
that blockage due to coolant injection from H1 is causing the leakage flow to turn away
from the trailing edge. Visual observations showed that there was more oil deposition.
This could result from the leakage vortex moving closer to the suction surface. The width
of the trace, at the trailing edge, is about 10% blade height and the clear region measures
5% blade height from the tip. In comparison to the flat tip case there is some movement
of the vortex closer to the blade tip. The trace appears to start closer to the trailing edge
for Minj = 0.5%. Indeed, the trace is not continuous, rather shows multiple leakage entry
paths, at 0.8 Cax and 0.88 Cax along the SS edge, into the passage. There is a oil free
region between the two entry points and the streak extends 8% blade height from the tip.
The greatest oil deposition occurs after 0.95 Cax.
Injection at Minj = 0.6% also displays multiple entry points along the suction
surface, the first one located at 0.75 Cax along SS edge. Maximum oil deposition occurs
beyond 0.95 Cax. The trace measures 10% blade height from the tip. At Minj = 0.7 % the
first clear entry point is at 0.8 Cax. There is a very light trace that appears to start at 0.73
Cax. It was noticed that even when using the same mixture, the sharpness of the patterns
varied, probably due to initial room temperature. This might be a reason for the traces to
show up much later along the suction surface for two higher injection rates. The trace
126
extends 12% blade height from the tip for the highest injection rate. The maximum paint
deposition occurs after 0.95 Cax along SS edge. It is clear that other than at the lowest
injection rate, the injection jet near the trailing edge does cause leakage flow near the
trailing edge to be channeled into the passage. From a heat transfer perspective these
results show that the heat transfer coefficient on the suction surface may be greater than
that for t/h=0.71%. It also identifies a potential problem area, close to the trailing edge
region where the heat transfer may actually be increased beyond that experienced by a
blade with t/h=0.71%.
Figure 5.7: Suction Surface Traces for Minj = 0.4% and Minj = 0.5%.
Minj = 0.4% Minj = 0.5%
127
5.1.5 Heat Transfer Implication
The tip trench is located ideally, to prevent strong reattachment on the tip surface.
Additionally, coolant injection reduces the effects of flow separation, albeit
intermittently, and reattachment on the tip surface. The recirculation of high temperature
fluid in the separation bubble is also affected. In the previous chapter, the reattachment
and associated recirculation were identified as the possible reasons for the enhanced heat
transfer observed on the near pressure-side corner of the tip surface. Thus, the location of
the trench and coolant injection should decrease the heat transfer coefficient on the tip
surface influenced by the coolant jets. The surface flow patterns also indicate that in the
region between the tip trench and the suction-side corner of the tip surface a reduction in
tip heat transfer is possible. Blockage of the leakage flow leads to low momentum
Figure 5.8: Suction Surface Traces for Minj = 0.6% and Minj = 0.7%.
Minj = 0.6% Minj = 0.7%
128
activity between the injection holes and the suction-side corner. Additionally, the coolant
jets turn towards the suction-side corner and form a film over the tip surface. Both these
effects would also reduce heat transfer to the tip surface.
The picture is not all positive however. Clearly the presence of reattachment near
the trailing edge is of concern as this will lead to greater heat transfer coefficients in this
region of the tip surface. However, with the temperature of the fluid considerably lower
than inlet temperatures, it is possible that this modification will not cause undue heating
of the tip surface. Radial holes are probably not the most effective way to cool the tip
surface, while jets angled towards the pressure-side and also towards or away from the
trailing edge appear to allow better area coverage.
5.2 Total Pressure Measurement
As in the previous chapter the results are presented in form of contour plots, radial
distributions of passage averaged coefficient, wake plots, and area averaged total pressure
coefficient. The comparisons made in this section are with respect to the baseline data
shown in Figure 4.14. Injection was initiated from the test blade only.
Contour plot for injection at Minj = 0.41% is shown in Figure 5.9. There is
considerable reduction in the tip leakage vortex of the test blade, as measured by the total
pressure downstream of the rotor exit. The minimum total pressure in the wake increases
by about 0.2 qm. This effect is similar in magnitude to that obtained with reduced tip
clearance of t/h = 0.72% and no injection. However, the flow structure is less like a
vortex with a smaller pitch-wise footprint. Additionally, the severe gradient observed in
129
the tip leakage vortex due to the large gap height of t/h = 1.40% is also eliminated. The
vortex appears to hold its position, as there is no observable movement towards the blade
suction surface, as seen when the gap height was reduced to t/h = 0.72%. Injection has no
effect on either the passage core flow or the interaction between the leakage vortex and
the tip side passage vortex. The wake plot in Figure 5.10 shows the reduction in wake
depth with tip injection as compared to the baseline. While there is some thinning of the
wake, the wake does not move to the right.
Figure 5.9: Total Pressure Coefficient Contours with Coolant Injection at Minj = 0.41%.
130
When the coolant mass flow rate is increased to Minj = 0.52%, Figure 5.11, the
leakage vortex core shows a slightly increased total pressure drop, as compared to Minj =
0.41%. The measured total pressure is however greater than that measured for the large
tip clearance. The shape of the vortex is also better defined than for Minj = 0.41%. The
size of the leakage vortex is smaller than that observed at the large gap height however
the vortex shows no appreciable movement towards the blade suction surface. The
interaction between the tip side passage vortex and the leakage vortex is still strong and
no effect is observed on the passage core. The wake plot, Figure 5.10, only serves to
confirm that the energy defect, while not as great as that observed for the large tip gap is
slightly greater than that obtained at the lower injection rate.
Blade Number
Tota
lPre
ssur
eC
oeff
icie
nt,C
pt
19 20 21 22 23-4.6
-4.5
-4.4
-4.3
-4.2
-4.1
-4
-3.9
-3.8
-3.7
Base3: t/h = 1.40%, Minj = 0t/h = 1.40%, Minj = 0.41%t/h = 1.40%, Minj = 0.52%
Figure 5.10: Wake Profile at r = 0.96h, Without and With Coolant Injection at Minj =
0.41% and Minj = 0.52%.
131
Injection at higher coolant mass flow rates begins to change the nature of the
leakage vortex from the test blade. As shown in Figure 5.12, at a coolant mass flow rate
of Minj = 0.63%, the tip leakage vortex from the test blade is not only reduced, but also
shifted more towards the casing. This observation may be expected when the gap height
is reduced. The tip-side passage vortex of the test blade is better defined due to reduced
interaction between the two flow structures. Injection also affects the tip leakage vortex
due to blade B22. The tip vortex of blade B22 is shifted towards the blade suction surface
and has a higher value of total pressure associated with it. Figure 5.13 compares the wake
profile with injection Minj = 0.63% () and the baseline and it is clear that the total pressure
defect of blade B21 is reduced. The reduction is however not as great as compared to
Figure 5.11: Total Pressure Coefficient Contours With Coolant Injection at Minj = 0.52%.
132
injection at 0.41%. The wake shift and defect reduction in the wake of blade B22 is also
seen in the wake profile.
At the highest coolant mass flow rate of Minj = 0.72%, shown in Figure 5.14, the
tip leakage vortex from blade B21 is shifted towards the casing and also appears to have
moved out into the passage. The increase in total pressure coefficient is similar to that
observed for coolant mass flow rate of Minj = 0.63%. The higher jet momentum however
affects the tip leakage vortex of blade B22 even more. This leakage vortex is shifted even
more, this is more apparent when the position of the leakage vortex is related to that of
the tip-side passage vortex. The structure also shows signs of complete mixing and a
greater total pressure coefficient. It appears that the jet momentum is also affecting the tip
Figure 5.12: Total Pressure Coefficient Contours With Coolant Injection at Minj = 0.63%.
133
leakage vortex of blade B23, as a slight shift is observed. In the baseline it was shown to
be inside the tip-side passage vortex, while injection at Minj = 0.72% has actually moved
it to the outside, probably due to jet penetration.
In all cases, the passage core flow is not affected much, as seen by the boundaries
representing the baseline in the contour plots. Indeed when comparing the wake plots
with and without injection at a radial location of 0.57h, shown in Figure 5.15, it is seen
that there is practically no change in the wake structure close to mid-span due to injection
from the blade tip.
Blade Number
Tota
lPre
ssur
eC
oeff
icie
nt,C
pt
19 20 21 22 23-4.6
-4.5
-4.4
-4.3
-4.2
-4.1
-4
-3.9
-3.8
-3.7
Base3: t/h = 1.40%, Minj = 0t/h = 1.40%, Minj = 0.63%t/h = 1.40%, Minj = 0.72%
Figure 5.13: Wake Profile at r = 0.96h, Without and With Coolant Injection at Minj =
0.63% and Minj = 0.72%.
134
Figure 5.14: Total Pressure Coefficient Contours With Coolant Injection at Minj = 0.72%.
Blade Number
Tota
lPre
ssur
eC
oeff
icie
nt,C
pt
19 20 21 22 23-4.3
-4.2
-4.1
-4
-3.9
-3.8
-3.7
-3.6
-3.5
-3.4
-3.3
Base3: t/h = 1.40%, Minj = 0t/h = 1.40%, Minj = 0.41%t/h = 1.40%, Minj = 0.52%t/h = 1.40%, Minj = 0.63%t/h = 1.40%, Minj = 0.72%
Figure 5.15: Wake Profile at r = 0.57h, Without and With Coolant Injection.
135
5.2.1 Comparison of Averaged Values
The radial distribution of passage averaged values of total pressure coefficient for
passage containing tip leakage vortex of blade B21, shown in Figure 5.16 compares the
distributions for gap heights of t/h = 1.40%, t/h = 0.72%, and coolant injection at a gap
height of t/h = 1.40%. One of the general observations that may be made is that coolant
injection reduces the distinct peak due to the large leakage vortex of blade B21. The
radial distributions with coolant injection are very close to that obtained at gap height of
t/h = 0.72%. In addition, the averaged total pressure coefficient in the leakage vortex
affected zone is similar to that in the tip-side passage vortex zone. The radial distributions
of passage averaged coefficient with coolant injection at gap height of t/h = 1.40% and
that due to gap height of t/h = 0.72% lie within the uncertainty band in the region from
0.75h – 1h. Thus, it appears that in the range of coolant mass flow rates studied there is
about equal effect and this effect is similar to that of reducing the gap height of the
cooled blade B21 to t/h = 0.72%.
Figure 5.17 compares the area averaged total pressure coefficient in the passage
containing the tip leakage vortex of cooled blade B21. The area averaged coefficient is
shown for tests without and with injection at the large gap height of t/h = 1.40%, and also
at the small gap height of t/h = 0.72%. The extent of the area average covers one passage
and 20% blade height. The small gap height (in blue) is used as a reference value against
which the effectiveness of coolant injection may be compared qualitatively. All area
averaged values due to coolant injection at gap height of t/h = 1.40% lie on the line
representing the averaged value of gap height t/h = 0.72%. Thus, it is possible to
136
conclude that the aerodynamic effect of the large gap height of t/h = 1.40% has been
modified by coolant injection to that due to a gap height of t/h = 0.72%.
Figure 5.16: Effect of Coolant Injection On the Passage Averaged Coefficient of Cooled Blade B21.
Uncertainty band
137
-4.15
-4.10
-4.05
-4.00
-3.95
-3.90
0 1 2 3 4 5
[-]
Are
a A
vera
ged
Tot
al P
ress
ure
Coe
ffic
ient
Minj = 0.41%Minj = 0.52%
Minj = 0.63%
Minj = 0.72%
t/h= 1.40%
t/h= 0.72%
Figure 5.17: Area Averaged Coefficient For Blade B21 With Coolant Injection at a Tip Gap Height of t/h = 1.40%.
138
Chapter 6
Effect of Injection Location on Over Tip Leakage
The chord-wise location of coolant jets is also an important parameter in
optimizing coolant injection. The focus of this chapter is the effect of injection from
individual locations and combinations of locations on OTL flow. As before, coolant is
injected from a single cooled blade, referred to as the test blade (B21) and the rotor
clearance distribution is TCL1, as shown in Figure 2.5. Test cases are referred to by the
number of the injection locations active in each case. Thus, H1 refers to injection from
the location 61% chord (H1) and H1+H3 refers to combined injection from the locations
61% chord (H1) and 81% chord (H3). Table 6-1 summarizes the various tests conducted.
Injections holes were rendered inactive by blocking the holes with silicone sealant. The
sealant was allowed at least 24 hours to set and the seal was pressure tested at 40 psig,
both before and after the test. Thus it was ensured that coolant was injected from only the
desired locations. Only total pressure measurements were done to study the effect of
injection location. In all cases the ATS pressure was maintained constant at 10 psig. This
was done since in actual engines the coolant stream total pressure is maintained constant.
The choice of pressure was made to conform to the lowest injection mass flow rate tested
and also to maintain the total pressure ratio to below critical pressure ratio. The injection
rate was stabilized for at least 5 minutes before data acquisition was commenced.
139
6.1 Injection from Individual Holes
Figure 6.1 shows the total pressure contours when injecting coolant from H1 only,
located at 61% chord. Comparing the pitch-wise extent of the vortex to the boundary
marking the vortex with no injection, it is apparent that coolant injection has reduced the
size of the vortex. The minimum total pressure measured in the leakage vortex has not
changed appreciably, however the core of the vortex appears to have shifted closer to the
test blade suction surface. The strong gradient characterizing the tip leakage vortex due to
a large tip gap height is also observed.
Table 6-1: Test Matrix for Effect of Injection Location.
Test Case Active Injection Hole
Minj%
H1 H1 0.2
H2 H2 0.2
H3 H3 0.2
H4 H4 0.2
H1+H2 H1 and H2 0.3
H1+H3 H1 and H3 0.3
H2+H3 H2 and H3 0.3
H1+H2+H3 H1, H2, and H3 0.35
Full Injection All locations 0.42
140
The pitch-wise extent of the leakage vortex is also reduced when injecting from
H2 (at 71% chord) only as shown in Figure 6.2. The reduction in total pressure defect is
greater than when injecting from H1 only and the gradient across the tip leakage vortex is
smoother. Figure 6.3 shows the effect of injecting from H3 (at 81% chord). The reduction
in pitch-wise extent is similar to that observed for H2 injection. The total pressure defect
however has been significantly reduced, by about 20% qm. This difference is similar to
that observed, as shown in the previous chapter, both when the gap height was reduced to
t/h = 0.72% and also with coolant injection at Minj = 0.42% into a gap height of t/h =
1.40%. The vortex core is more uniform with smoother gradients across the vortex. The
passage flow above 85% blade height and between blades B20 and B21 appears to be less
energetic. Coolant injection from H4 has very little effect on the leakage vortex, as shown
Figure 6.1: Total Pressure Coefficient Contours for Coolant Injection From Blade B21 and Location H1 at 61% Cax.
141
in Figure 6.4. The size of the vortex is about the same as that in the baseline and the low
total pressure core has moved back to a location comparably similar to that observed in
the baseline. There does appear to be some reduction in the total pressure defect. The
more significant effect of coolant injected from H4 is the reduction in total pressure
measured in the passage above 85% blade height. It is not possible to confirm that this is
due to additional total pressure loss.
Figure 6.2: Total Pressure Coefficient Contours for Coolant Injection From Blade B21 and Location H2 at 71% Cax.
142
Figure 6.3: Total Pressure Coefficient Contours for Coolant Injection From Blade B21 and Location H3 at 81% Cax.
Figure 6.4: Total Pressure Coefficient Contours for Coolant Injection From Blade B21 and Location H4 at 91% Cax.
143
The reduction in the passage area occupied by the leakage vortex and the
movement the tip leakage vortex towards the blade suction surface can occur if the gap
normal momentum is reduced, similar to when the gap height is reduced. Thus coolant
injection from locations H1, H2, and H3 does cause reduction in gap normal velocity,
either by forcing the gap flow to turn towards axial or by reducing gap mass flow rate and
producing a wake like behavior. The gap thickness to gap height ratio and the gap
pressure differential is greatest at H1. The gap flow may be expected to contain a wake
region due to diffusion beyond the vena-contracta, as reviewed earlier. Coolant jets at H1
originate very close to the location of reattachment and may cause a more intense
diffusion of the gap flow around the blockage presented by the coolant jets. This
reduction in normal gap momentum manifests itself as a reduction in the pitch-wise
extent of the tip leakage vortex.
The greatest defect reduction is observed to occur due to coolant injection at the
81% blade axial chord location. Flow visualization results indicate that injection hole H3
is located close to the reattachment line. Additionally, the tip gap surface flow is seen to
turn towards the trailing edge, due to a drop in pressure difference across the gap. Flow
visualization also indicates intense mixing of gap flow and injected coolant. It would
appear then that the total pressure defect observed in the measurement plane is caused
due to the gap flow diffusing close to the suction-side of the gap, where coolant injection
prevents a strong recirculation and subsequent diffusion. As reviewed earlier the leakage
vortex was seen to increase dramatically in size between 80% and 90% chord and the
area mass averaged shed vorticity also more than doubled in this region. The trailing edge
jet injects coolant radially into the gap where the gap flow is fully separated and the
144
negligible effect may be attributable to either entrainment of coolant by the leakage flow
or the fact that there isn’t much diffusion within the gap occurring in the first place.
The radial distribution of the passage averaged total pressure coefficient for the
passage bounded by the suction-side of the test blade is shown in Figure 6.5. The effect
of coolant injection from individual locations is large enough to affect the passage
averaged total pressure for individual injection from the locations H1, H2, and H3. Also
shown in Figure 6.5 is the radial distribution of passage averaged total pressure
coefficient at a tip gap height of t/h = 0.72%. The reduction in total pressure defect due to
coolant injection from individual locations is not quite as large as that obtained by
reducing the gap height of blade B21 from t/h = 1.40% to t/h = 0.72%. Below 90% span
the values follow that of the baseline pretty closely. The total pressure defect reductions
are also localized to about the last 10% of blade height. In the case of injection from H4
there is little change from the baseline, with the near casing total pressure actually lower
by about 2%. This is caused partly by the reduced total pressure observed in between the
two tip leakage vortices of blades B20 and B21, as shown in Figure 6.4.
The wake structure downstream of the blades is shown in Figure 6.6. The change
in wake depth follows closely what has been observed in the contour plots. The greatest
reduction in wake depth is observed for coolant injection from H3, while the other three
cases show similar reduction in depth. The similarity in the wake depth of injection cases
H1, H2, and H4 is more likely due to movement of the vortex, as clearly the effect of H4
injection is negligible.
145
Figure 6.5: Effect of Injection Hole Location on the Passage Averaged Coefficient of
Cooled Blade B21.
Blade Number
Tota
lPre
ssur
eC
oeff
icie
nt,C
pt
19 20 21 22 23-4.6
-4.5
-4.4
-4.3
-4.2
-4.1
-4
-3.9
-3.8
-3.7
-3.6
Base3: t/h = 1.40%, Minj = 0t/h = 1.40%, H1t/h = 1.40%, H2t/h = 1.40%, H3t/h = 1.40%, H4
Figure 6.6: Effect of Injection Location on the Wake Profile at r = 0.96h.
Uncertainty band
146
6.2 Injection from Combination of Holes
Isolated injection studies indicated that coolant injection from the near trailing
edge, radial hole did not affect the tip leakage vortex significantly. Hence in studying the
combinations, H4 was not considered.
Figure 6.7 shows the effect of combined injection from H1 & H2. The pitch-wise
extent of the tip leakage vortex is reduced more than that when individually injecting
from H1. In addition, the vortex has shrunk in its span-wise extent. The decrease in
measured total pressure drop, in the leakage vortex region, is also greater than the
individual injection cases. Higher total pressure from the passage flow is seen within the
continuous curve enclosing the baseline tip leakage vortex. Coolant injection from the
combination of H1+H3 also decreases the area occupied by the tip leakage vortex of
cooled blade B21, as shown in Figure 6.8. The reduction in the total pressure defect
associated with the leakage vortex of blade B21 is greater than that achieved by the
combination of H1+H2. The tip leakage vortices of blades B21 (t/h = 1.40%) and blade
B19 (t/h = 0.92%) resemble each other closely, indicating an effective reduction in gap
height is achievable through coolant injection from the combination of H1+H3.
The total pressure map with coolant injection from the combination H2+H3 is
shown in Figure 6.9. The area occupied by the tip leakage vortex of cooled blade B21 is
smaller in comparison to that with no injection. The area of low total pressure is however
greater than that observed in the two previous combinations, where H1 was active.
Additionally, higher total pressure (GREEN Zone) is not observed within the boundary
denoting the tip leakage vortex of blade B21 without injection. This compares well with
147
the observations made for isolated injection, where coolant injection from H1 caused the
greatest reduction in pitch-wise extent of the tip leakage vortex of blade B21. It is not
unexpected that the tip leakage vortex of cooled blade B21 is significantly reduced in size
when coolant is injected from the combination of H1+H2+H3, as shown in Figure 6.10.
The associated total pressure defect is also reduced, along with a drop in the total
pressure gradient across the tip leakage vortex. Higher total pressure (GREEN Zone) is
also observed within the boundary of the tip leakage vortex with no coolant injection.
Figure 6.7: Total Pressure Coefficient Contours for Coolant Injection From Blade B21 and Locations H1+H2.
148
Figure 6.8: Total Pressure Coefficient Contours for Coolant Injection From Blade B21 and Locations H1+H3.
Figure 6.9: Total Pressure Coefficient Contours for Coolant Injection From Blade B21 and Locations H2+H3.
149
The effect of coolant injection from all locations is shown in Figure 6.11. The
measured coolant mass flow rate is Minj = 0.42%. The effect of coolant injection at a
mass flow rate of Minj = 0.41% was previously discussed with reference to Figure 5.9.
The total pressure distribution in the tip leakage vortex of cooled blade B21 in
Figure 6.11 is remarkably similar to that observed for the tip leakage vortex of blade B21
in Figure 5.9. This further supports the good repeatability between tests as discussed in
Section 4.2.2. The area of influence of the tip leakage vortex of blade B21 is significantly
reduced. The minimum total pressure measured in the tip leakage vortex is about 0.2 qm
greater than that measured with no coolant injection.
Figure 6.10: Total Pressure Coefficient Contours for Coolant Injection From Blade B21 and Locations H1+H2+H3.
150
The variation in passage averaged total pressure distribution, due to coolant
injection from various combinations of injection locations is shown in Figure 6.12. The
passage average is computed for the passage containing the tip leakage vortex of cooled
blade B21. All combinations tested increase the averaged total pressure in the span-wise
region from 0.8h – 1h. The total pressure in this region is affected by the tip leakage
vortex, the tip-side passage vortex, and the interaction between these secondary flow
structures. Thus the combined effect of the injection locations tested affects the passage
flow more than individual injection. The improvement due to all combinations tested is
about the same, since all the curves are contained within the uncertainty band. The effect
of the tested combinations is also similar to the effect of reducing the gap height of the
test blade from t/h = 1.40% to t/h = 0.72%.
Figure 6.11: Total Pressure Coefficient Contours for Coolant Injection From Blade B21; Full Injection.
151
In the case of H1+H3 it is seen that below 84% blade height there appears to be
significant total pressure recovery. This however is not due to injection alone. Figure 6.13
shows the rotor averaged total pressure coefficient for the various combinations. In
general there is good repeatability, except in the case of H1+H3, where the curve below
88% blade height is shifted to the right. The shift is just within the uncertainty band and
hence the data is treated as having greater uncertainty. The same behavior was observed
Figure 6.12: Effect of Injection Location Combinations on the Passage Averaged
Coefficient of Cooled Blade B21.
Uncertainty band
152
in the passage averaged results for this case and hence it is not possible to separate the
effect of coolant injection, if any, on the flow in this region.
Wake plots for the injection combinations are shown in Figure 6.14. At the near-
tip radial location of 0.96h the wake depth is reduced for all combinations tested. The
greatest reduction occurs for the full blowing case and the least for the combination of
H1+H2. However, there is not much difference between the cases. The most interesting
feature, in comparing Figure 6.6 and Figure 6.14, is that the wake profile of B21 is
observably shifted towards the blade suction-side. This behavior is not observed for the
individual injection cases. This means that combined injection is successful in moving
the tip leakage vortex closer to the blade suction surface.
Figure 6.13: Rotor Averaged Coefficient With Combined Injection.
Uncertainty band
153
Blade Number
Tota
lPre
ssur
eC
oeff
icie
nt,C
pt
19 20 21 22 23-4.6
-4.5
-4.4
-4.3
-4.2
-4.1
-4
-3.9
-3.8
-3.7
-3.6
Base3: t/h = 1.40%, Minj = 0t/h = 1.40%, H1+H2t/h = 1.40%, H1+H3t/h = 1.40%, H2+H3t/h = 1.40%, H1+H2+H3t/h = 1.40%, H1+H2+H3+H4
Figure 6.14: Effect of Injection Location Combinations on the Wake Profile at r = 0.96h.
154
Chapter 7
Multiple Cooled Blades and the Effect of Casing Surface Roughness
The results presented so far dealt with coolant injection from a single cooled
blade (B21) and the tests were referred to as isolated injection tests. It was shown that
higher injection rates had some effect on the leakage flow in neighboring passages, due to
increased momentum of the coolant jets. As noted in Chapter 2, five blades (B17-B21)
were modified for coolant injection. The tip gap on four of the five cooled blades was
reduced by applying precision plastic shims on the blade tip surfaces. In order to study
the effect of coolant injection from multiple blades these shims were removed and the
effect of multiple cooled blades on the tip leakage flow over cooled blade B21 is
presented in this chapter. Coolant mass flow rate measurement was done only on the
supply line to blade B21, as described in Chapter 2. The removal of the plastic shims on
four of the five cooled blades resulted in the clearance distribution TCL3, shown in
Figure 2.5.
A preliminary investigation of the effect of casing surface roughness was also
conducted by artificially roughening the casing inner surface. The roughness was
introduced by applying various grades of sandpaper using double-sided adhesive tape.
The nominal clearance distribution for the surface roughness study is TCL4 in Figure 2.5.
155
7.1 Baseline
The total pressure coefficient contour map with five cooled blades and no coolant
injection is shown in Figure 7.1. The discussion in this section is centered on the
differences between Figure 7.1 and Figure 4.14, where the large tip gap was present on
cooled blade B21 only. Note that the gap height for blades B19 and B20 is greater than
that in the results presented in previous chapters. The effect of enlarging the tip gap
height of these blades increases the area occupied by their respective tip leakage vortices.
The tip leakage vortices of blade B19 and B20 also shift to the left and away from the
blade suction surface, as shown by the boundary of the baseline tip leakage vortices from
Figure 4.14. It was shown in Chapter 4 that the tip leakage vortex of cooled blade B21
moved significantly towards the blade suction surface when the tip gap height was
reduced from t/h = 1.40% to t/h = 0.72%. Thus, the behavior observed here is consistent
with the change in the tip gap height.
The position of the tip leakage vortex of cooled blade B21 has not changed much
and the slight shift towards the left may be attributed to the increased tip gap height of
blade B20. Increasing the gap height of blade B20 leads to lower momentum in the tip-
side passage vortex, allowing the leakage vortex of blade B21 to affect more of the
passage. The total pressure associated with the leakage vortex of blade B21 is unchanged
in magnitude, while the total pressure coefficient of blades B19 and B20 has dropped by
0.2qm, in comparison to Figure 4.14. For the blade sector shown, the largest total pressure
defect is seen for blade B19, which has the largest tip gap. The GREEN zone between
leakage flow structures of blades B20 and B21 has also increased considerably.
156
The wake structure of blades B19 and B20 is also different, in comparison to the
results presented in Figure 4.14. The distinct tip-side passage vortex has vanished and is
replaced by a zone of uniformly low total pressure. This occurs due to the increased
interaction of the tip leakage vortex with the tip-side passage vortex. The passage core
between blades B20 and B21 indicates a radial outward shift, due to a reduction in the
influence of the tip-side passage vortex on the core flow.
The change in the wake profile at the span-wise location of 0.96h is shown in
Figure 7.2. The comparison of the isolated injection baseline and multiple injection
baseline shows that blade B21 is mostly unaffected by the increased tip gap height of
blades B17-B20. However the wake profiles for blade B17-B20 show considerable
Figure 7.1: Total Pressure Coefficient With Multiple Cooled Blades; No Coolant Injection (Baseline), Minj = 0.0%.
157
increase in depth, particularly for blade B19. The tip gap height of blade B19 was
increased from t/h = 0.92% to t/h = 1.54%. The shift in the wake profile to the left is also
observed and is on average about 20% pitch.
7.2 Variation of Coolant Mass Flow Rate
Coolant mass flow rates of 0.43%, 0.62%, and 0.72% were investigated. From
previous coolant injection results it was seen that coolant mass flow rate of Minj = 0.41%
had the most beneficial effect on the leakage vortex. Injection at a coolant mass flow rate
Figure 7.2: Wake Profile at r = 0.96h Comparing Baseline Distributions of Multiple
Cooled Blade and Single Cooled Blade (B21).
158
of Minj = 0.63% appeared to be the most effective overall, while the coolant mass flow
rate of Minj = 0.72% had the greatest effect on the tip leakage flow structure of blade B22.
The total pressure contour map with multiple blade coolant injection at Minj =
0.43% is shown in Figure 7.3. The tip leakage vortices of the cooled blades (B17 – B21),
with enlarged tip gap heights, are seen to have moved to the right and closer to the blade
suction surfaces. The reduction in total pressure defect is 0.15qm, on average. Multiple
blade coolant injection appears to have an effect on the tip leakage flow structures of
blades B22 and B23, this is not seen when coolant was injected from cooled blade B21
only at Minj = 0.41%. Thus, multiple blade coolant injection appears to be generating a
strong flow in the near casing region, opposite in direction to the leakage flow, thereby
affecting the tip leakage flow over blades with no coolant injection. A comparison with
Figure 5.9, where coolant was injected from blade B21 only at Minj = 0.41%, shows that
coolant injection at the lowest coolant mass flow rate has an almost identical effect.
Figure 7.3: Total Pressure Coefficient With Multiple Cooled Blades; Minj = 0.43%
159
An increase in the coolant mass flow rate to Minj = 0.62% produces a greater
reduction in the area occupied by the tip leakage vortex of cooled blade B21, as shown in
Figure 7.4. This was not observed when injecting from a single cooled blade (B21), as
discussed in Chapter 4. The tip leakage vortex of blade B21 displays a greater movement
to the right and is almost a part of the blade wake. It may be recalled that with isolated
injection the tip leakage vortex retained a compact structure and was displaced closer to
the casing. This would also suggest that multiple blade injection is setting up a near
casing flow that is opposing the leakage flow. Thus, when injecting at Minj = 0.62%, the
tip leakage vortex of blade B21 shows greater reduction in the total pressure defect. A
visual comparison shows that the total pressure defect associated with the tip leakage
vortex of blade B21 (t/h = 1.40%) with coolant injection at Minj = 0.62%, in Figure 7.4, is
much smaller than that associated with the total pressure defect of blade B23 (t/h =
0.77%) in Figure 7.1. The higher kinetic energy (GREEN) zone in between adjacent tip
leakage vortices of blades B20 and B21 occupies a larger area, due to a reduction in the
influence of the tip leakage vortex of blade B21 on the passage flow.
The characteristics of the tip leakage vortex of blade B21 with coolant injection at
a mass flow rate of Minj = 0.72% are very similar to that observed at the coolant mass
flow rate of Minj = 0.62%, as shown in Figure 7.5. The area occupied by the vortex, total
pressure defect, and location of the vortex are almost identical in both cases. The
increased momentum of the coolant streams appears to affect the tip-side passage vortex.
The tip-side passage vortex of blade B22 occurs at a lower radial location, closer to 75%
span, thereby reducing the interaction between the tip leakage and tip-side passage
vortices.
160
Figure 7.4: Total Pressure Coefficient With Multiple Cooled Blades; Minj = 0.62%.
Figure 7.5: Total Pressure Coefficient With Multiple Cooled Blades; Minj = 0.72%.
161
The wake profiles of coolant injection from five cooled blades and the
corresponding baseline is shown in Figure 7.6. The five cooled blades, with large tip gap
heights, are clearly distinguished in the baseline by the much deeper wakes observed for
blades B17-B21. The wake depth is reduced at all coolant mass flow rates and the
movement of the wake profile to the right is also observed. The change in wake depth of
blade B21 indicates a total pressure defect reduction of approximately 0.3 qm, greater
than the 0.2 qm reduction observed with isolated injection. The circumferential shift in the
wake profile also increases with increase in the coolant mass flow rate. The effect of
coolant injection of the tip leakage flow over blade B22 is also observed at the selected
radius. The shift in the wake profile of blade B22 is more pronounced at the highest
coolant mass flow rate. The reduction in total pressure defect, while greater than the
uncertainty band, is small.
The radial distribution of the passage averaged coefficients shown in Figure 7.7
compares the effect of multiple blade coolant injection to that of no coolant injection. The
effect of multiple cooled blades on blade B21 is presented. Total pressure gains are
observed in the region from 0.85h to 1h at all coolant mass flow rates. As with coolant
injection from blade B21 only, variation in coolant mass flow rate appears to have
negligible effect on the averaged total pressure. There is little effect of coolant injection
over the rest of the passage height over which the data is available. Consistently higher
total pressure is measured in the near casing region for all coolant mass flow rates due to
energizing of the passage flow by the higher momentum present in the coolant jets.
162
Figure 7.6: Wake Profiles at r = 0.96h for Multiple Blade Coolant Injection.
Figure 7.7: Passage Averaged Coefficient Comparison for Multiple Blade Coolant
Injection.
Uncertainty band
163
7.3 Effect of Casing Surface Roughness
The effect of casing surface roughness on over tip leakage flow was investigated
by applying coarse and fine sand paper to the inner surface of the casing window, with
the aid of double-sided adhesive tape. This method of artificially roughening the casing
inner surface changes the tip clearance distribution. A smooth surface of equivalent tip
clearance was also tested by applying a precision plastic shim of the same thickness. This
was done to isolate the effect of artificially roughening the casing surface from the effect
of tip gap height reduction due to treatment thickness. The nominal thickness of the
applied casing treatment is 0.3175 mm or 0.26% blade height and the resulting tip
clearance distribution is TCL4, as shown in Figure 2.5. Only total pressure measurements
with no coolant injection were conducted.
7.3.1 Smooth Casing Surface
The total pressure coefficient distribution in the measurement plane with a smooth
casing surface at a uniformly reduced tip gap height is shown in Figure 7.8. The
treatment thickness is 0.26% of blade height. The tip gap height of each blade is noted
above the casing boundary and the clearance of cooled blade B21 is t/h = 1.14%. The tip
gap height reduction leads to the expected drop in total pressure defect for all blades, as
compared to the total pressure coefficient observed in Figure 7.1. The smallest tip gap
height obtained is t/h = 0.51% for blade B23 and consequently the corresponding tip
leakage vortex is very weak and contained within the wake. It must be noted that the total
164
pressure defect due to leakage vortex of blade B21 (t/h = 1.14%) is greater than that of
blade B21 with a tip gap height of t/h = 0.72%, which was discussed in reference to
Figure 4.14. The expected movement of the tip leakage vortices to the right is also
observed. The higher total pressure zones between subsequent tip leakage vortices are
more energized. The tip-side passage vortices in the wakes of blades B22 and B23 are
quite distinct and occur below 0.85h.
The wake profiles shown in Figure 7.9 compare the baseline for multiple cooled
blades with the wake obtained by uniformly reducing the tip gap height through
application of the smooth surface treatment. Also shown are the wake profiles due to
Figure 7.8: Total Pressure Coefficient With a Smooth Plastic Layer On the Casing Inner Surface.
165
artificially introduced surface roughness, which will be discussed later. The reduction in
tip gap height causes the expected reduction in wake depth and circumferential
movement of the wake profile to the right. The reduction in wake depth for the cooled
blade B21 is 0.14 qm. The wake profile near mid-span is shown in Figure 7.10 and
indicates no effect at mid-span due to the uniform reduction in tip gap height.
Figure 7.9: Wake Profiles at r = 0.96h Comparing the Influence of Casing Surface
Roughness.
166
7.3.2 Fine Surface Roughness (220 Grit)
The total pressure coefficient distribution shown Figure 7.11 was obtained after
applying 220 Grit sandpaper to the casing inner surface. The thickness of the applied
treatment is 0.26% blade height. It is clear that the artificially introduced surface
roughness has considerable effect on both the leakage flow and the tip-side passage
vortex. The tip leakage vortices of the cooled blades shown are greatly reduced in size.
The total pressure defect is also substantially reduced, along with the gradients. The tip
leakage vortex is contained within the wake and is located to the right of the tip-side
passage vortex. The tip leakage vortices of blades B22 and B23 are completely mixed in
Figure 7.10: Wake Profiles at r = 0.57h Comparing the Influence of Casing Surface
Roughness.
167
with the wake fluid. It is of course unclear whether the leakage flow is eliminated or if it
is completely mixed in with the wake. The increased surface roughness is expected to
locally increase turbulent kinetic energy within the tip gap, reducing the normal
momentum exiting the tip gap suction-side corner. Thus the secondary kinetic energy
associated with the tip leakage vortex in the blade passage is reduced and the over tip
leakage fluid is found within the blade wake.
The tip-side passage vortices appear to be stronger and especially so at the larger
tip gap heights. This is partly due to the reduced interaction between the tip leakage flow
and the tip-side passage secondary flow. The passage cores are shifted considerably to the
left and this is probably due to the effect of the tip-side passage vortex on the passage
core.
The wake profiles in Figure 7.9 show that the artificially introduced surface
roughness leads to considerable reduction in wake depth, in the range of 0.3 qm to 0.4 qm.
The wake depth of cooled blade B21 is reduced by 0.38 qm, which is greater than the 0.14
qm reduction obtained by the effect of reducing the gap height only, as discussed earlier.
Thus, the effect of artificially roughening the casing inner surface is about 1.7 times the
effect of reducing the tip gap height. The total pressure defect reduction is also greater at
the larger tip gap heights, indeed for the wake profiles shown there is small difference
between the smooth wall and fine roughness wake depths at the smaller tip gap heights.
There is however a global shift in the wake profiles to the right. The effect on the passage
core is also observed in the wake profiles at 0.57h, shown in Figure 7.10. The wake
profile may be observed to have shifted to the left, a behavior also seen in the contour
plot. The wake depth however remains almost the same.
168
7.3.3 Coarse Surface Roughness (100 Grit)
Figure 7.12 is a contour plot of the total pressure field downstream of the rotor
with coarse sand paper applied on to the casing. While the treatment thickness (0.3% h) is
slightly greater than the previous two treatments, the difference (0.04% h) is believed to
be too small to have an effect on the tip gap flow. Hence, any differences seen may be
attributed to the increased surface roughness.
The effect of coarse roughness on the tip leakage flow appears to be similar to
that of fine sandpaper. The tip leakage vortices are greatly reduced in area and the total
pressure field within the vortices is uniform. The tip leakage vortices at the small gap
Figure 7.11: Total Pressure Coefficient With Fine Sandpaper (220 Grit) On the Casing Inner Surface.
169
heights are again mixed in with the blade wakes. The tip-side passage flow also appears
to be more uniform across the passage. The tip-side passage vortices are stronger due to
the reduced interaction with the tip leakage flow. The tip-side passage vortices have also
moved to a lower radial location. The reduced interaction between the secondary flows
near the blade tip is a result of reduction in leakage flow momentum exiting the tip gap,
caused by greater turbulent kinetic energy within the tip gap. There is considerable
spacing between the two tip secondary flow structures, indicating that the increase in
surface roughness is responsible for this shift. The movement of the passage core to the
left is also observed and hence it can be concluded that this is due to the increased surface
roughness.
The wake profiles at r = 0.96 h, due to the increased surface roughness are quite
similar to those obtained with the fine surface roughness as shown in Figure 7.9. This
implies that the effect of surface roughness is not enhanced significantly by increasing
the roughness quality of the flow path. The wake profiles close to mid-span, shown in
Figure 7.10, also indicate that the effect of surface roughness is unchanged by the
increase in the roughness quality of the casing inner surface.
7.4 Comparison of the Averaged Total Pressure Coefficient
The passage averaged total pressure coefficient, computed for the passage
containing the tip leakage vortex of cooled blade B21 (t/h = 1.40%, without casing
treatment) with different casing roughness treatments is shown in Figure 7.13. The
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baseline with five cooled blades (solid RED) and coolant injection at Minj = 0.62%
(dotted GREEN) are also shown for comparison. The decrease in tip gap height by
0.26%h without artificial roughness (smooth casing) reduces the total pressure defect
observed in the region from 0.85h to 1h. The reduction is however at the limit of the
uncertainty band. The introduction of surface roughness doubles the total pressure defect
reduction over the entire passage, from 0.05qm to 0.1qm at the location 0.93h. The
benefits of artificial roughness are comparable to that obtained through multiple blade
coolant injection in the region influenced by the tip leakage vortex.
Figure 7.12: Total Pressure Coefficient With Coarse Sandpaper (100 Grit) On the Casing Inner Surface.
171
The rotor averaged total pressure coefficient for the multiple cooled blade
baseline and the three casing treatments is shown in Figure 7.14. The total pressure defect
due to the five large tip gap heights is visible in the rotor averaged coefficient. This is in
contrast to the rotor averaged coefficient in Figure 4.15. Thus, the performance of the
rotor may be expected to decrease when the rotor tip clearance distribution is changed
from TCL1 to TCL3. The artificial introduction of surface roughness appears to have no
observable effect in the region of the tip leakage vortex, over that of reducing the tip gap
height by 0.26%h. This supports the earlier conclusion that artificial roughening of the
casing surface is most beneficial at the larger gap heights.
Figure 7.13: Passage Averaged Coefficient For Blade B21 With Different Casing
Roughness Treatments.
Uncertainty band
172
Figure 7.14: Rotor Averaged Coefficient With Different Casing Roughness Treatments.
Uncertainty band
173
Chapter 8
Summary and Conclusions
Turbomachines are widely used in the transfer of thermal energy to mechanical
energy in power generation and aircraft engines. Axial flow turbines extract energy from
the working fluid by effecting a change in the angular momentum of the flow. The
efficiency of this ideally isentropic process depends on how closely the blade passage
flow occurs to the design streamlines. Secondary flow in blade passages, such as passage
vortices, horseshoe vortices and over tip leakage cause the fluid to deviate considerably
from design streamlines thereby reducing the efficiency. The gap between rotating
turbine blades and stationary casing, in high pressure, un-shrouded, axial flow turbines is
called the tip gap or tip clearance. Flow in this gap from blade pressure-side to blade
suction-side, termed over tip leakage, is pressure driven, generates considerable losses,
and increases the thermal load in regions of the tip surface. The tip gap height also
increases with service, thereby causing further decrease in engine efficiency and also the
exhaust gas temperature. Reducing the effects of this necessary gap on turbine
performance is termed tip desensitization and an experimental investigation of tip surface
coolant injection as a method for desensitizing turbine blade tips was reported in this
thesis. Many of the extensively researched desensitization techniques serve to block over
tip leakage mass flow rate by reducing gap discharge coefficient. These methods
typically present increased cooling requirements. Even the most commonly used tip
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surface geometry that of double squealer tips is subject to burn-out and thermal
degradation. Tip surface coolant injection directed towards blade pressure-side corner,
reported in this thesis, aims at modifying existing, radial coolant injection schemes and
thereby does not incur the penalty of additional cooling requirements.
8.1 Summary
Two different measurement approaches were used to study tip desensitization in
this study. Qualitative investigation of the effects of tip gap height and tip surface coolant
injection was studied using surface flow visualization. Surface flows on the rotor endwall
and turbine casing were also visualized. The visualization techniques used consisted of
oil-dot technique and the oil-film technique. Quantitative measurements consisted of
obtaining high resolution, total pressure distributions at 30% chord downstream of the
rotor exit using a fast response total pressure probe aligned with the absolute tip velocity
vector. The influence of tip gap height, tip surface coolant injection from rotor blades,
and the effect of casing surface roughness were investigated in detail.
8.1.1 Surface Flow Visualization
Surface flows on the rotor endwall, turbine casing, and rotor blade tip surfaces
were visualized. Rotor endwall surface flow visualization showed the formation of the
horseshoe vortex upstream of the blade leading edge. Variations in boundary layer
growth on the rotating hub, upstream of the rotor inlet plane were also captured. The path
175
of the pressure-side leg of the horseshoe vortex is well defined. The cross-passage
boundary layer flow that leads to the passage vortex was discussed with reference to
previously obtained rotor endwall static pressure distributions. At rotor exit the flow in
the rotating frame was shown to exit the blade passage near design angle from pressure
surface up to mid-pitch and then the flow was highly overturned due to the passage
secondary flow. Visualization of the interaction zone between tip leakage vortex and tip
passage vortex was also possible. A streak of visualization material was observed to
extend from about 60% blade tip axial chord length all the way to the blade trailing edge.
The streak diverged along its length and its lower boundary was measured at 75% span at
the trailing edge.
Rotor tip surface visualization was performed by applying the oil and pigment
mixture to the blade pressure surface and allowing rotational effects and leakage flow
entering the gap to transport oil on to the tip surface. Visualization material carried on to
the tip surface when using the oil dot technique is discrete. Chord-wise flow within the
separation zone on the tip surface transported the oil in a well defined streamline. The
streamline moves further away from the pressure-side corner, close to 0.5 blade tip axial
chord and subsequently runs parallel to the blade profile, maintaining its distance from
the pressure-side corner. The distance of the streamline from the pressure-side corner also
varies with gap height, being closer to the blade pressure-side corner at smaller gap
heights. Two gap heights were investigated. The distance of the chord-wise streamline
from the pressure-side corner was measured to be 1.4*t for a gap height of t/h = 1.40%
and 1.1* t for a gap height of t/h = 0.72%. In terms of blade tip axial chord length the
distances are 3% Cax for t/h = 1.40% and 1.2% Cax for t/h = 0.72%.
176
Oil film applied to blade pressure surface clearly identifies a separation region
near the pressure-side corner, followed by a well defined reattachment line on the tip
surface. Re-circulating flow is also clearly identified by this technique. The regions
where reattachment and re-circulation are observed are closely coupled with areas of high
heat transfer rates on the blade tip surface.
The position of the reattachment line from blade pressure-side corner was
measured at 60% Cax for gap heights (t/h) of 0.72%, 0.81%, 1.2%, and 1.40%. The
distance measured normal to blade tip axial chord length, when normalized by the blade
tip axial chord length, was found to increase linearly with gap height. The distance almost
doubled when the gap height was increased by almost twice. The distance measured
normal to the camber-line and normalized by gap height was found to stay reasonably
constant at about 1.95*t – 2.0*t.
Turbine casing surface flow visualization was also conducted using the two
techniques. Oil dots indicate the highly overturned nozzle exit flow in the near suction
surface region at vane exit. Flow angles at rotor exit are near design angles for flow
originating away from the nozzle suction surface. The highly over-turned nozzle exit
flow is seen at rotor exit as being under-turned, with a predominantly axial direction. This
underscores the importance of conducting probe based measurements with the probe
positioned at nozzle mid-pitch. A region of low momentum fluid originating from the
suction-side of nozzle vane was tracked up to blade mid-chord.
Surface flow visualization with coolant injection showed that pressure-side corner
separation was substantially reduced in the region influenced by the coolant jets, at all
coolant mass flow rates tested. Recirculation was completely eliminated in the last 40%
177
of blade tip axial chord length. This should reduce not only the gap mixing losses, but
also decrease the heat transfer rate to the tip surface in this region. Flow in the last 1/3rd
of the blade is substantially changed, including the elimination of fully separated tip
region.
Coolant jets form localized films on the tip platform, at all injection rates. The
leading jet of each injection set forms a film to the right of the hole, while the trailing jet
forms a film to the left, due to its inclination towards the trailing edge.
At 0.4% coolant mass flow rate the trailing jets cover the largest area on the tip
surface. At the other three coolant mass flow rates the best surface area coverage is
obtained from the leading jets. Thus, orientation of cooling jets appears to depend upon
the amount of coolant injected. Increasing the coolant mass flow rate also appears to
cause more mixing between the coolant and the gap flow. Leakage flow is blocked by the
coolant jets, leading to low momentum activity between the injection holes and the SS
corner.
8.1.2 Total Pressure Measurement
Total pressure measurements of the flow field 30% blade tip axial chord lengths
downstream of the rotor exit plane were conducted using a high-frequency, total pressure
probe aligned with the angle of the tip velocity vector in the stationary frame of
reference. Time accurate, phase-locked, total pressure measurements are averaged over
200 rotor revolutions and non-dimensionalized with a mean wheel speed based dynamic
pressure (qm). The total pressure probe attached to the outer casing of the turbine stage is
178
traversed in the radial direction. Thus, a complete two-dimensional mapping of the rotor
exit total pressure field for all 29 passages is possible. Distinct flow structures such as tip
leakage vortices, passage vortices, blade wake, and core flow were detected by
employing a phase-locked ensemble averaging technique. Circumferential averaging of
the total pressure coefficient was done for individual passages and for the entire rotor.
The rotor averaged coefficient was used to ascertain repeatability of data, while the
passage averaged coefficient enabled distinguishing the effect on individual passages.
At a gap height of t/h = 1.40% the tip leakage vortex is relatively large, occupying
about 15% blade span. The tip leakage vortex for this case presents a significant blockage
to the passage flow. The total pressure defect due to the large vortex is about 20% qm
greater than that of blades with nominal blade clearance.
Reduction of tip clearance effectively reduced the blockage presented by the tip
clearance flow to the passage flow. The total pressure drop measured in the leakage
vortex was reduced by about 20% qm. The size of the leakage vortex was greatly reduced,
by almost 50%, at the smaller gap height. The reduced interaction between the leakage
vortex and the passage vortex was evident from the improved definition of the passage
vortex. A reduction in tip gap height caused the leakage vortex to move towards the blade
suction-side.
Coolant injection from the tip trench was successful in filling in the total pressure
defect originally resulting from the leakage vortex without injection. Coolant mass flow
rates of 0.41%, 0.52%, 0.63%, and 0.72% were investigated for coolant injection from a
single blade with a gap height of t/h = 1.40%.
179
At all injection rates, the reduction in total pressure defect is very similar to that
when the gap height t/h = 1.40% is reduced by about half to 0.72%. This result shows
that directed tip coolant jets can be effective in creating beneficial “sealing” effects
previously observed only from small clearances.
Injection at Minj = 0.41% core flow was the most effective in reducing the total
pressure loss in the leakage flow of the test blade. This was observed at a radius near the
core of the tip vortex. However, it appears that 0.63% injection is the most effective from
a global point of view, as shown by the passage averaged pressure coefficient obtained in
the last 25 % of the blade height.
The tip leakage vortex moves closer to the tip in a radially outward direction,
especially at the higher injection rates. The cross section of the new tip leakage vortex,
with coolant injection, is smaller and some of the total pressure defect is eliminated by
the injection process. The upper passage vortex is better defined when the tip leakage
vortex cross section is smaller and located nearer the casing.
The relatively high radial position of the leakage vortex resulting from Minj values
0.63% and 0.72% implies that the interaction with the passage vortex is less pronounced.
Thus, tip injection is capable of reducing losses, especially those due to leakage
flow/passage vortex/blade wake interaction.
At the two highest injection rates the leakage vortices of adjacent blades (to the
right of the test blade) are affected by the injection. The high momentum associated with
these jets moves the tip vortices in adjacent channels against the direction of rotation.
This might be due to the alteration of the near casing flow between the pressure side
corner of the test blade and the adjacent blade. Since the tip jets are faced towards the
180
pressure side corner of the test blade, especially at high blowing rates, the static pressure
distribution near the outer casing in the adjacent passage is expected to be altered.
The location of the coolant injection holes was also studied, by individual
injection from each hole and combinations thereof. The coolant supply pressure was
maintained constant for all tests in this series. Individual injection results show that
coolant injection from 61%, 71%, and 81% chord locations reduce the leakage vortex
size at the measurement location. This is attributed to a reduction in normal momentum
exiting the tip gap.
Injection from 81% chord is the most successful in filling the total pressure defect
in the vortex core. Thus it appears that leakage flow responsible for the greatest total
pressure deficit occurs around 80%.
The injection location at 91%, with the largest hole diameter of 1.8 mm, does not
have a significant effect on the leakage flow. This hole is very close to the trailing edge
of the blade. However, the tip-side passage flow shows an increase in the total pressure
drop coefficient.
Combined injection in general shows better desensitization. There is observable
movement of the leakage vortex towards the suction-side of the test blade. Combined
injection from H1-H3 is found to be almost as effective as injection from all locations.
Coolant injection from all five cooled blades was studied at coolant mass flow
rates of 0.43%, 0.62%, and 0.72%. In all cases the effect of injecting from more than one
blade caused a greater total pressure recovery. Increasing coolant mass flow rate leads to
greater defect reduction, although it appears that 0.62% might be an inflection point in
the performance benefit curve. Multiple blade injection is believed to cause a strong near
181
casing flow in a direction opposing over tip leakage and might be expected to change the
static pressure distribution on the casing wall. Tip passage vortex appears to be more
energized.
Casing surface roughness was investigated by applying smooth plastic shim and
two grades of sandpaper to the casing surface with double-sided adhesive tape. Casing
surface roughness greatly reduced the leakage vortex defect measured downstream of the
rotor exit. The area occupied by the tip leakage vortex was also substantially smaller. The
increase in turbulent kinetic energy within the tip gap is believed to reduce the normal
momentum exiting the tip gap, leading to a smaller leakage vortex. The leakage vortex is
also contained within the blade wake and hence it can be concluded that the leakage flow
is turned more efficiently. The tip-side passage vortex was found to be more energized
especially with the coarse grade of sandpaper causing the tip-side passage vortex to move
towards the hub and away from the blade wake. This is attributed to the reduced
interaction with the tip leakage vortex, which also means that the mixing losses due to
this interaction must be lower.
8.2 Conclusions
The following conclusions are drawn form the results,
1. Surface flow visualization is an extremely effective tool in understanding the flow
physics in rotating turbine blade passages.
2. Rotor endwall flow features observed in the rotating frame of reference compare
well with those observed in previous stationary cascade facilities.
182
3. Low momentum fluid from the nozzle suction surface propagates through the
rotor passage causing under-turning of flow at rotor exit.
4. Chord-wise flow on the tip surface occurs in a region between the separated flow
at blade pressure-side corner and re-circulating flow within the tip gap.
5. The location of the reattachment line, measured from the pressure-side corner
increases linearly with gap height.
6. The distance of the reattachment line measured from the pressure-side corner is
about 2*gap height, at all gap heights.
7. Tip gap flow is fully separated over the last 5% axial chord length of the blade,
where the driving pressure differential across the tip gap is minimal.
8. The tip trench causes the gap vortex to be weakened, by preventing it from
remaining attached to the tip surface.
9. Coolant injection from the test blade causes the leakage vortex to decrease in size
and total pressure defect is reduced to levels observed for blades with half the gap
height. Tip surface coolant injection could serve as a highly effective tip leakage
sealing system.
10. Coolant injection closer to the trailing edge, where the blade tip profile is thin,
caused a much greater reduction in total pressure defect. Coolant injection closer
to the leading edge on the other hand is more effective in reducing the area
occupied by the tip leakage vortex.
11. Heat transfer benefits might arise from the elimination of reattachment and re-
circulation of gap flow on the tip surface and the blockage presented by the
coolant jets.
183
12. Coolant injection from multiple blades increases the reduction in total pressure
defect, by about 50%. Leakage flow of un-cooled blades also appears to be
affected beneficially.
13. Increasing the surface roughness of the turbine casing leads to the tip-side passage
vortex to appear at a lower radius and the tip leakage vortex to be greatly reduced
in both size and total pressure defect.
14. Tip leakage reduction methods studied in this thesis are highly applicable to many
axial flow turbomachinery systems. The tip injection and casing roughness based
OTL control are candidates for use in modern counter-rotating ducted fan systems
built into many unmanned flight vehicles.
8.3 Recommendations for Future Work
This experimental study has shown that tip surface coolant injection, with coolant
jets directed towards the pressure-side corner, can serve as an effective tip leakage
sealing strategy and can also improve the thermal performance of turbine blade tips.
There are however certain parameters that cannot be addressed economically in a rotating
rig. These may however be studied either in stationary cascades or through numerical
investigations. The influence of tip gap height on the exit flow field may be used to
“calibrate” such investigations. Parametric studies of the location of the coolant injection,
from the pressure-side corner can be conducted relatively inexpensively. Similarly, the
potential thermal benefits of tip surface coolant injection may be easier to quantify,
184
particularly in stationary cascades. The effect of casing surface roughness may also be
simulated numerically and compared to available experimental results.
185
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51. Allen, H. W., and Kosfskey, M. G., Visualization Study of Secondary Flows in Turbine Rotor Tip Regions. NACA TN-3519
52. Dring, R. P., and Joslyn, H. D., 1981, “Measurement of Turbine Rotor Blade Flows,” ASME J. Engineering for Power, 103, pp. 400-405.
53. Gregory-Smith, D. G., Graves, C. P., Walsh, J. A., 1988, “Growth of Secondary Losses and Vorticity in an Axial Turbine Cascade,” ASME J. Turbomachinery, 110, pp. 1-8.
54. Xiao, X., McCarter, A., Lakshminarayana, B., “Tip Clearance Effects in a Turbine Rotor: Part 1- Pressure Field and Loss,” ASME Paper No. 2000-GT-476.
55. Yang W., 1989, Handbook of Flow Visualization, Hemisphere Publishing Corporation.
56. Prasad, A., and Wagner, J. H., 2000, “Unsteady Effects in Turbine Tip Clearance Flows,” ASME Paper No. 2000-GT-0444.
57. Kwak, J. S., Ahn, J., Han, J. C., Lee, C. P., Bunker, R. S., Boyle, R., and Gaugler, R., 2003, “Heat Transfer Coefficients on The Squealer Tip and Near Tip Regions of a Gas Turbine Blade With Single or Double Squealer,” ASME Paper No. GT-2003-38907.
58. Prasad, D., and Hendricks, G. J., 2000, “A Numerical Study of Secondary Flow in Axial Turbines With Application to Radial Transport of Hot Streaks,” ASME Paper No. 2000-GT-0448.
59. Ristic, D., Lakshminarayana, B., Chu, S., 1998, “Three-Dimensional Flow Field Downstream of an Axial Flow Turbine Rotor,” AIAA paper 98-3572, presented at the 34th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, Cleveland, Ohio.
190
Appendix A
Total Pressure Probe Characteristics
The sensitivity of the total pressure coefficient to probe incidence was tested at
two radial locations in the AFTRF and the results are presented Section A.1. The
frequency content of the total pressure field in the measurement plane and in the region
of influence of the tip leakage vortex was measured to decide on the low pass filter cut-
off frequency. The frequency spectrum is discussed in Section A.2.
A.1 Angular Sensitivity
Measurement sensitivity to exit flow angle was measured in the AFTRF by
rotating the probe in increments of 5°, on either side of α3 = 25.4°. This was done at two
radial locations, 0.93h (near tip region) and 0.49h (mid span). Counterclockwise (CCW)
rotation of the probe is positive and makes the probe more tangential with every
increment. The rotor averaged, and passage averaged total pressure coefficient for blade
B21, are shown in Figure A.1. The absolute velocity vector at mid span is at an angle of
29.19° from axial, corresponding to an incidence angle of 3.79°. The mean value of Cpt,P
at mid-span in a ±15° range is -3.916 and all values in this range lie within the
uncertainty band of δCpt = ± 0.024. Furthermore, the passage averaged coefficient for
blade B21 indicates that the blade passage behaves almost identically as the rest of the
191
rotor. At 0.93h however, the difference between the rotor averaged and passage averaged
values is considerable, due to the large leakage vortex. Passage averaged coefficient in
the range -25° to 10° are within uncertainty limits of the mean value (Cpt,P = -4.22),
computed for the range ±15°. The value at 15° is just outside the uncertainty band.
McCarter, et al. [16] measured lower relative tangential velocity in the region dominated
by the leakage vortex. This means that the leakage vortex approaches the probe at a
negative incidence. Thus, from measurements in the range of -25° to 0° it is possible to
conclude that error in the measurement of total pressure associated with the tip leakage
vortex of the test blade is within uncertainty bounds.
A.2 Frequency Spectrum of Flow Field
The frequency content of the rotor exit flow field, 30% chord length downstream
of the rotor, was obtained by feeding the signal to a frequency analyzer and computing
the spectrum shown in Figure A.2. The peaks identified are those at the blade passing
frequency (BPF) and its harmonics. The 2nd harmonic of the BPF is not as distinct as the
fundamental and 1st harmonic. Furthermore there is at least a 10 dB drop in signal power
between the fundamental and 2nd harmonic. The drop off is around 30 dB at a frequency
of 12 kHz and the spectrum appears to be leveling out. Based on this it was decided that
using a low-pass filter cut-off frequency of 20 kHz would not remove any content from
the signal. Furthermore, the full range of the sensor frequency response would be utilized.
192
Figure A.1: Probe Response to Incidence. (Squares denote rotor averaged Cpt and circles
denote passage averaged Cpt.).
193
A.3 Uncertainty Analysis
A sample calculation of the uncertainty analysis is shown in this section. The
propagation of uncertainty is calculated by the method derived by Kline and McClintock
[45]. The expression for the total pressure coefficient is as shown in Equation A-1. The
uncertainty in the total pressure coefficient is obtained from Equation A-2, which is
-6.00E+01
-5.00E+01
-4.00E+01
-3.00E+01
-2.00E+01
-1.00E+01
0.00E+00
1.00E+01
2.00E+01
3.00E+01
0.00E+00 2.00E+03 4.00E+03 6.00E+03 8.00E+03 1.00E+04 1.20E+04 1.40E+04
Frequency, (kHz)
Mag
nitu
de, (
dB)
BPF Harmonics
Figure A.2: Frequency Spectrum of Rotor Exit Flow Near Rotor Tip.
194
obtained by differentiating Equation A-1 and dividing the resulting expression by
Equation A-1.
where W[] is the uncertainty associated with the parameter in the square brackets.
The individual uncertainty and nominal value of each measured parameter (used
in the denominator) is shown in Table A-1. These in turn are obtained from manufacturer
specifications of the precision associated with the measurement. Uncertainty values for
air density (ρ) and blade mean wheel speed (Um) are calculated in a similar fashion. The
values are listed in Table A-2, without presenting the sample calculations.
2
0
21
),(m
ptU
PjiC
ρ
∆= , where 01030 ),( PjiPP −=∆
(A-1)
222
0
20
⎟⎟⎠
⎞⎜⎜⎝
⎛+⎟
⎟⎠
⎞⎜⎜⎝
⎛+⎟
⎟⎠
⎞⎜⎜⎝
⎛
∆=
∆
m
UP
pt
C
U
WW
P
W
C
Wmpt
ρρ
, (A-2)
195
Table A-1: Uncertainty and Nominal Values in Measured Parameters
Parameter Precision Error Uncertainty Nominal Value
∆P0 0.1% of 34.474*103
(kPa) ±34.474 (Pa) 7471.387 (Pa)
Tamb ±0.5 (K) ±0.5 (K) 300 (K)
Pamb 0.1% ±100 (Pa) 98700 (Pa)
N ±1 (rpm) ±1 (rpm) 1328 (rpm)
Table A-2: Uncertainty in Derived Parameters
Parameter Uncertainty
pt
C
C
Wpt ±0.0058
0
0
PW P
∆∆ ±0.00461
ρρW
±0.0035
m
U
U
Wm ±7.53*10-4
196
Appendix B
AFTRF Tip Clearance Distribution
Tip clearance measurements in three regions of the blade chord are presented in
tabulated form and graphically. Data is presented for two of the four clearance
distributions reported in this manuscript. The clearance distributions TCL1 and TCL3 as
defined in Chapter 2 are presented in Table B-1 and Table B-2, respectively. The
graphical representation of the variation of measured gap height along blade axial chord
length is shown in Figure B.1. Only the variation in distribution TCL1 is shown.
197
Table B-1: Gap Height Variation Along Blade Axial Chord For Clearance Distribution TCL1
Blade #
Clearance Average Clearance
LE-33% 33-66% 66-TE mm t/h% mm t/h% mm t/h% mm t/h%
1 1.02 0.83% 0.99 0.81% 1.02 0.83% 1.01 0.82% 2 0.97 0.79% 0.99 0.81% 1.02 0.83% 0.99 0.81% 3 0.99 0.81% 0.99 0.81% 0.97 0.79% 0.98 0.80% 4 0.97 0.79% 1.02 0.83% 0.97 0.79% 0.98 0.80% 5 0.94 0.76% 0.91 0.74% 0.94 0.76% 0.93 0.76% 6 0.91 0.74% 0.91 0.74% 0.86 0.70% 0.90 0.73% 7 0.89 0.72% 0.86 0.70% 0.86 0.70% 0.87 0.71% 8 0.91 0.74% 0.86 0.70% 0.84 0.68% 0.87 0.71% 9 0.91 0.74% 0.86 0.70% 0.86 0.70% 0.88 0.72%
10 0.89 0.72% 0.89 0.72% 0.89 0.72% 0.89 0.72% 11 0.94 0.76% 0.94 0.76% 0.94 0.76% 0.94 0.76% 12 0.94 0.76% 0.94 0.76% 0.91 0.74% 0.93 0.76% 13 0.94 0.76% 0.91 0.74% 0.91 0.74% 0.92 0.75% 14 0.94 0.76% 0.86 0.70% 0.86 0.70% 0.89 0.72% 15 0.89 0.72% 0.86 0.70% 0.84 0.68% 0.86 0.70% 16 0.91 0.74% 0.89 0.72% 0.86 0.70% 0.89 0.72% 17 0.99 0.81% 1.04 0.85% 1.09 0.89% 1.04 0.85% 18 0.76 0.62% 1.14 0.93% 1.09 0.89% 1.00 0.81% 19 1.07 0.87% 1.12 0.91% 1.22 0.99% 1.13 0.92% 20 0.89 0.72% 0.97 0.79% 0.99 0.81% 0.95 0.77% 21 1.63 1.32% 1.70 1.38% 1.83 1.49% 1.72 1.40% 22 1.02 0.83% 1.02 0.83% 1.02 0.83% 1.02 0.83% 23 0.91 0.74% 0.97 0.79% 0.97 0.79% 0.95 0.77% 24 0.97 0.79% 0.97 0.79% 1.04 0.85% 0.99 0.81% 25 1.02 0.83% 1.02 0.83% 1.04 0.85% 1.02 0.83% 26 1.02 0.83% 1.07 0.87% 1.09 0.89% 1.06 0.86% 27 1.09 0.89% 1.02 0.83% 1.02 0.83% 1.04 0.85% 28 0.97 0.79% 0.99 0.81% 1.02 0.83% 0.99 0.81% 29 0.97 0.79% 0.99 0.81% 0.97 0.79% 0.98 0.79%
198
Table B-2: Gap Height Variation Along Blade Axial Chord For Clearance Distribution TCL3
Blade # Clearance Average Clearance
LE-33% 33-66% 66-TE
mm t/h% mm t/h% mm t/h% mm t/h% 1 1.02 0.83% 0.99 0.81% 1.02 0.83% 1.01 0.82% 2 0.97 0.79% 0.99 0.81% 1.02 0.83% 0.99 0.81% 3 0.99 0.81% 0.99 0.81% 0.97 0.79% 0.98 0.80% 4 0.97 0.79% 1.02 0.83% 0.97 0.79% 0.98 0.80% 5 0.94 0.76% 0.91 0.74% 0.94 0.76% 0.93 0.76% 6 0.91 0.74% 0.91 0.74% 0.86 0.70% 0.90 0.73% 7 0.89 0.72% 0.86 0.70% 0.86 0.70% 0.87 0.71% 8 0.91 0.74% 0.86 0.70% 0.84 0.68% 0.87 0.71% 9 0.91 0.74% 0.86 0.70% 0.86 0.70% 0.88 0.72%
10 0.89 0.72% 0.89 0.72% 0.89 0.72% 0.89 0.72% 11 0.94 0.76% 0.94 0.76% 0.94 0.76% 0.94 0.76% 12 0.94 0.76% 0.94 0.76% 0.91 0.74% 0.93 0.76% 13 0.94 0.76% 0.91 0.74% 0.91 0.74% 0.92 0.75% 14 0.94 0.76% 0.86 0.70% 0.86 0.70% 0.89 0.72% 15 0.89 0.72% 0.86 0.70% 0.84 0.68% 0.86 0.70% 16 0.91 0.74% 0.89 0.72% 0.86 0.70% 0.89 0.72% 17 1.78 1.45% 1.78 1.45% 1.78 1.45% 1.78 1.45% 18 1.68 1.36% 1.70 1.38% 1.73 1.41% 1.70 1.38% 19 1.80 1.47% 1.98 1.61% 1.91 1.55% 1.90 1.54% 20 1.70 1.38% 1.70 1.38% 1.73 1.41% 1.71 1.39% 21 1.63 1.32% 1.70 1.38% 1.83 1.49% 1.72 1.40% 22 1.02 0.83% 1.02 0.83% 1.02 0.83% 1.02 0.83% 23 0.91 0.74% 0.97 0.79% 0.97 0.79% 0.95 0.77% 24 0.97 0.79% 0.97 0.79% 1.04 0.85% 0.99 0.81% 25 1.02 0.83% 1.02 0.83% 1.04 0.85% 1.02 0.83% 26 1.02 0.83% 1.07 0.87% 1.09 0.89% 1.06 0.86% 27 1.09 0.89% 1.02 0.83% 1.02 0.83% 1.04 0.85% 28 0.97 0.79% 0.99 0.81% 1.02 0.83% 0.99 0.81% 29 0.97 0.79% 0.99 0.81% 0.97 0.79% 0.98 0.79%
199
0.40%
0.60%
0.80%
1.00%
1.20%
1.40%
1.60%
0 5 10 15 20 25 30
Blade Number
Cle
aran
ce G
ap H
eigh
t, t/h
(%)
LE-33%33%-66%66%-TEAverage
Figure B.1: Clearance Gap Variation Along Blade Axial Chord Length For TCL1.
200
Appendix C
Total Pressure Coefficient Contour Map
The total pressure contours presented and discussed previously showed a sector of
five rotor blades, including the cooled, test blade B21. A complete map of the total
pressure field is shown in Figure C.1. As noted in Section 4.2.1, the total pressure probe
is in the stationary frame and the data acquisition system acquires 6000 data points per
revolution of the rotor. Thus, in the total pressure map of the entire rotor 29 blade wakes,
passage cores, and tip leakage vortices are seen as shown in Figure C.1. The variations in
the tip leakage vortex structure with tip gap height are clearly observed.
201
Figure C.1: Total Pressure Contour Map Of the Entire Rotor Exit Flow Field In the
Measurement Plane.
ΩBlade Wake
Passage Core
Tip Leakage Vortex
VITA
Nikhil Molahally Rao
Nikhil Rao received his Bachelor of Engineering degree in Mechanical
Engineering, from the University of Mysore in February 1994. His career in
turbomachinery engineering started in May 1994, at Turbotech Precision Engineering (P)
Ltd., in Bangalore, India. The author left his position in May 1997 to commence his
graduate study at The Virginia Polytechnic Institute and State University and received his
M.S. degree in Mechanical Engineering in May 1999. Subsequently, he entered the
graduate program in Aerospace Engineering at The Pennsylvania State University and
successfully defended his doctoral thesis on May 02, 2005. The author will be graduating
with a Ph.D. in Aerospace Engineering in August 2005 and has accepted the position of
Engineer in the Turbine Engineering Group, at The Elliott Company in Jeannette,
Pennsylvania, starting in June 2005.
Refereed Papers
Rao, N., and Camci, C., 2005, “Visualization of Rotor Endwall, Tip Gap, and Outer Casing Surface Flows In a Rotating Axial Turbine,” ASME Paper GT2005-68264, ASME Turbo Expo 2005, Reno, Nevada, June 2005.
Rao, N., and Camci, C., 2004, “A Flow Visualization Study of Axial Turbine Tip Desensitization by Coolant Injection From a Tip Platform Trench,” ASME Paper IMECE2004-60943, 2004 ASME International Mechanical Engineering Conf., Anaheim, California, November 2004.
Publications
Rao, N. M., Feng, J., Burdisso, R. A., and Ng, W. F., 2001, “Experimental Demonstration of Active Flow Control to Reduce Unsteady Stator-Rotor Interaction,” AIAA Journal, Vol.39, No.3, March 2001, pp.458-464.
Presentations
Rao, N., “Axial Flow Turbine Tip Desensitization by Injection From a Tip Trench. Part 1- Effect of Injection Mass Flow Rate,” Presented at ASME Turbo Expo 2004, Vienna, Austria, June 14, 2004.
Rao, N., “Axial Flow Turbine Tip Desensitization by Injection From a Tip Trench. Part 2- Leakage Flow Sensitivity to Injection Location,” Presented at ASME Turbo Expo 2004, Vienna, Austria, June 14, 2004.