SECTION l.O SUMMARY l
TABLE OF CONTENTS Page SECTION l.O SUMMARY l SECTION 2.0 INTRODUCTION SECTION 3.0 DESIGN OVERVIEW 3.1 Turbine Intermediate Caseand Low-Pressure Turbine Concepts 3.1.1 Turbine Intermediate Case 3.1.2 Low-Pressure Turbine 3.2 Predicted Performance SECTION 4.0 INTERMEDIATE CASE ANDLOW-PRESSURE TURBINE AERODYNAMIC DESIGNII 4.1 AerodynamicDesign Parameters II 4.2 Flowpath and Airfoil Cross Section Definition 12 4.2.1 Intermediate CaseFlowpath 12 4.2.2 Intermediate CaseStrut Fairing 14 4.2.3 Low-PressureTurbine Flowpath 18 4.2.4 Turbine Airfoil Definition 20 4.3 Supporting Technology Programs 60 4.3.1 Introduction 60 4.3.2 Transition Duct Model Test Program 60 4.3.3 Low-Pressure Turbine Subsonic Cascade Technology Program 64 4.3.4 Low-Pressure Turbine Boundary Layer Program 68 4.4 Performance Status and Adjustment 81 SECTON 5.0 TURBINE THERMAL-MECHANICAL DESIGN 5.1 Mechanical Design Objectives and Goals 5.2 Turbine Mechanical Configuration 5.2.1 Turbine Rotor Assembly 5.2.1.I Blades 88 5.2.1.I.I Mechanical Design Features 5.2. l. l. 2 Structural Analysis I03 5.2.1.2 Disk and HubAssembly I03 5.2.1.2.1 Mechanical Design Features I04 5.2.1.2.2 Structural Analysis 5.2.1.3 Thrust Balance Seal 5.2.1.3.1 Mechanical Design Features I09 5.2.1.3.2 Structural Analysis I09 5.2.1.4 Inner Cavity Knife-Edge Seals I09 5.2.1.4.1 Mechanical Design Features I09 5.2.1.4.2 Structural Analysis I09 5.2.1.5 Rotor Shaft Assembly I09 5.2.1.5.1 Mechanical Design Features ll2 5.2.1.5.2 Structural Analysis iii TABLE OF CONTENTS (Cont'd) 5.2.2 Turbine Vaneand CaseAssembly ll4 5.2.2.1 Vanes ll4 5.2.2.1.I Mechanical Design Features ll4 5.2.2.1.2 Structural Analysis 120 5.2.2.2 Vane Cases 5.2.2.2.1 Mechanical Design Features 5.2.2.2.2 Structural Analysis 5.2.2.3 Turbine Exhaust Case 5.2.2.3.1 Mechanical Design Features 5.2.2.3.2 Structural Analysis 5.2.2.4 Low-Pressure Turbine Active Clearance Control System 134 5.2.2.4.1 Design Approach 5.2.2.4.2 Mechanical Design Features 5.2.2.4.3 Clearance Analysis 5.2.3 Turbine Intermediate Case 5.2.3.1 Mechanical Design Features 5.2.3.1.I Structural Struts 5.2.3.1.2 Strut Fairings 5.2.3.1.3 Outer Caseand Outer Case Heatshield 144 5.2.3.1.4 Engine Mount Lugs 5.2.3.1.5 Oil Supply and ScavengeLines 5.2.3.1.6 Thrust Balance Seal Lands 148 5.2.3.2 Structural Analysis 5.2.3.2.1 Primary Support Structure Stress and Life Analysis 5.2.3.2.2 Strut Fai ring DurabiI i ty Analysis 5.2.3.2.3 Thrust Balance Seal Vibration Analysis 5.2.4 Low Rotor Critical SpeedAnalysis 5.2.5 Low-Pressure Turbine Secondary Flow System 5.2.5.1 System Description 5.2.5.2 Low-Pressure Turbine Rotor 167 5.2.5.3 Low-Pressure Turbine Case 175 5.2.5.4 Turbine Intermediate Case 180 5.2.6 Turbine SystemWeight Summary SECTION 6.0 CONCLUDING REMARKS iv
APPENDIX A NASTRAN Stress Analysis
TABLE OF CONTENTS (Cont'd) APPENDIXES APPENDIX A NASTRAN Stress Analysis APPENDIX B Turbine Intermediate Case Strut Fairing Platform and Airfoil Deflection Contours APPENDIX C Turbine Intermediate Case Strut Fairing Airfoil Coordinates APPENDIX D Low-Pressure Turbine Vane and Blade Airfoil Coordinates APPENDIX E Low-Pressure Turbine Exit Guide Vane Airfoil Coordinates LIST OF SYMBOLS REFERENCES DISTRIBUTION LIST
LIST OF ILLUSTRATIONS
Number Ti tl e
Page
2-I
ProgramSchedule for Energy Efficient Engine Turbine
Intermediate Case and Low-Pressure Turbine Design Efforts
3.1 -I
Major Components in the Energy Efficient Engine Turbine
Section
3.1 .l -l Cross-Sectional View of Turbine Intermediate Case
3.1.2-I Cross-Sectional
View of Low-Pressure Turbine Component
4.2.1 -l
Turbine Intermediate Case Flowpath for the Integrated Core/Low Spool and the Preliminary Design Scaled to a 160,135 N (36,000 Ibs) Thrust Size 4.2.1-2 Intermediate Case Wall Pressure Distribution Profiles, Showing the High Loading of Integrated Core/Low Spool Design
4.2.2-I
Incidence Range Capability as a Function of Pressure Loss, Showing Greater Range With 400 Series Foil
4.2.2-2 Definition for Section l --
Fairing Aerodynamic Spool Intermediate Case Integrated Core/Low
4.2.2-3 Definition for Section 2 --
Fairing Aerodynamic Integrated Core/Low Spool Intermediate Case
4.2.2-4 Definition for Section 3 --
Fairing Aerodynamic Spool Intermediate Case Integrated Core/Low Definition for Section 4 --
4.2.2-5
Fai ri ng Aerodynamic Spool Intermediate Case Integrated Core/Low
4.2.2-6
Strut Fairing Stacki ng Arrangement
4.2.3-I
Low-Pressure Turbine FLowpath
4.2.4-I Effects of Controlled Vortexing Compared to Free
Vortexi ng
4.2.4-2 Turbine Inlet Guide Vane Mean Section Aerodynamic
Definition and Pressure Distribution vi LIST OF ILLUSTRATIONS (Cont'd) Number Ti tle
Page
4.2.4-3 Turbine Inlet Guide Vane Root and Tip Section Aerodynamic Definitions and Pressure Distributions 25 4.2.4-4 Turbine Second Stage Blade Root and One-Quarter Root Section Aerodynamic Definitions and Pressure Di stri buti ons 26 4.2.4-5 Turbine Second Stage Blade Mean Section Aerodynamic Definition and Pressure Distribution 4.2.4-6 Turbine Second Stage Blade Tip and One-Quarter Tip Section Aerodynamic Definitions and Pressure Di stri buti ons 4.2.4-7 Turbine Third Stage Vane Root and One-Quarter Root Section Aerodynamic Definitions and Pressure Di stri buti ons 29 4.2.4-8 Turbine Third Stage Vane Mean Section Aerodynamic Definition and Pressure Distribution 4.2.4-9 Turbine Third Stage Vane Tip and One-Quarter Tip Section Aerodynamic Definitions and Pressure Distributions 4.2.4-I0 Turbine Third Stage Blade Root and One-Quarter Root Section Aerodynamic Definitions and Pressure Di stri butions 32 4.2.4-II Turbine Third Stage Blade Mean Section Aerodynamic Definition and Pressure Distribution 33 4.2.4-12 Turbine Third Stage Blade Tip and One-Quarter Tip Section Aerodynamic Defi nitions and Pressure Distributions 4.2.4-I 3 Turbine Fourth Stage Vane Root and One-Quarter Root Section Aerodynamic Definitions and Pressure Di stri buti ons 4.2.4-14 Turbine Fourth Stage Vane Mean Section Aerodynamic Definition and Pressure Distribution 4.2.4-15 Turbine Fourth Stage Vane Tip and One-Quarter Tip Section Aerodynamic Definitions and Pressure Distributions vii
LIST OF ILLUSTRATIONS (Cont'd)
Number
Ti tl e
Page
4.2.4-16
Turbine Fourth Stage Blade Root and One-Quarter Root
Section Aerodynamic Definitions and Pressure
Distributions
4.2.4-17
Turbine Fourth Stage Blade MeanSection Aerodynamic
Definition and Pressure Distribution 39
4.2.4-18
Turbine Fourth Stage Blade Tip and One-Quarter Tip
Section AerodynamicDefinitions and Pressure
Distributions
4.2.4-19
Turbine Fifth Stage Vane Root and One-Quarter Root
Section AerodynamicDefinitions and Pressure
Distributi ons
4.2.4-20
Turbine Fifth Stage VaneMeanSection Aerodynamic
Definition and Pressure Distribution
4.2.4-21
Tu_-bineFifth Stage VaneTip and One-Quarter Tip Section
AerodynamicDefinitions and Pressure Distributions
4.2.4-22
Turbine Fifth Stage Blade Root and One-Quarter Root
Section Aerodynamic Definitions and Pressure
Distri buti ons
4.2.4-23
Turbine Fifth Stage Blade Mean Section Aerodynamic Definition and Pressure Distribution
4.2.4-24
Turbine Fifth Stage Blade Tip and One-Quarter Tip Section Aerodynamic Definitions and Pressure Di stri buti ons
4.2.4-25
Second Stage Turbine Vane Stacking Arrangement
4.2.4-26
Second Stage Turbine Blade Stacking Arrangement
4.2.4-27
Third Stage Turbine Vane Stacking Arrangement
4.2.4-28
Third Stage Turbine Blade Stacking Arrangement 48
4.2.4-29
Fourth Stage Turbine Vane Stacking Arrangement
4.2.4-30
Fourth Stage Turbine Blade Stacking Arrangement
4.2.4-31
Fifth Stage Turbine Vane Stacking Arrangement viii LIST OF ILLUSTRATIONS (Cont'd) Number Ti tl e
Page
4.2.4-32 50 Fifth Stage Turbine Blade Stacking Arrangement 4.2.4-33 Turbine Exit Guide Vane Flowpath for the Integrated Core/Low Spool 4.2.4-34 Turbine Exit Guide Vane Cascade Performance Test Results Under a Navy-Sponsored Program (Contract NO OOl9-77-C-0546 ).
4.2.4-35 Turbine Exit Guide Vane Cascade Performance Test Results Under a Navy-Sponsored Program (Contract NO001 9-77-C-0546). 54 4.2.4-36 Turbine Exit Guide Vane Root Section Aerodynamic Definition and Pressure Distribution 55 4.2.4-37 Turbine Exit Guide Vane 25-Percent Span Section Aerodynamic Definition and Pressure Distribution 4.2.4-38 Turbine Exit Guide Vane Mean Section Aerodynamic Definition and Pressure Distribution 57 4.2.4-39 Turbine Exit Guide Vane 75-Percent Span Section Aerodynamic Definition and Pressure Distribution 4.2.4-40 Turbine Exit Guide Vane Tip Section Aerodynamic Definition and Pressure Distribution 4.2.4-41 Turbine Exit Guide Vane Stacking Arrangement 4.3.2-I Rig Inlet Mass-Averaged Air Angle Spanwise Data 4.3.2-2 Strut Exit Mass-Averaged Air Angle Spanwise Data 4.3.2-3 Low-Pressure Turbine Inlet Guide Vane Exit Mass-Averaged Air Angle Spanwise Data 4.3.2-4 Outer Wall Loading Profile 4.3.2-5 Inner Wall Loading Profile 4.3.3-I Loss Versus Incidence for the Low Camber Vane 4.3.3-2 Aft-Loaded Transonic Versus Squared-Off Subsonic Airfoil Profiles and Pressure Distributions ix LIST OF ILLUSTRATIONS (Cont'd) Number Ti tl e
Page
4.3.3-3 Aft-Loaded Transonic Versus Squared-Off Subsonic Airfoils Predicted Versus Measured Profile Loss Data 67 4.3.3-4 Aft-Loaded Transonic Versus Squared-Off Subsonic Airfoils Predicted Versus Mass-Averaged, Mixed-Out, Mi d-Span Losses. 67 4.3.3-5 Aft-Loaded, Heavyweight, Lightweight -- Predicted Versus Profile Loss Over Range of Incidence, Fixed Exit Mach Number (0.72) 4.3.3-6 Aft-Loaded, Heavyweight, Lightweight -- Predicted Versus Measured Mass-Averaged Mid-Span Profile Loss at Design Point Incidence, Variable Exit Mach Numbers 69 4.3.4-I Turbulent Boundary Layer Velocity Profiles Using Dimensionless Parameters 4.3.4-2 Comparison of Measured Integral Parameters for the Squared-Off Test Section With Theoretical Predictions 72 4.3.4-3 Comparison of Measured Mean Velocity Profile data (Squared-Off) With Theoretical Predictions 73 4.3.4-4 Comparison of Measured Integral Parameters for the Aft-Loaded Test Section With Theoretical Predictions 74 4.3.4-5 Comparison of Measured Mean Velocity Profile data (Aft-Loaded) With Theoretical Predictions 4.3.4-6 Distribution of Boundary Layer Momentum Loss Thickness (e), Shape Factor (H), and Skin Friction (Cf) 4.3.4-7 Comparison of Measured Total Turbulence Intensity Profiles with Flat Plate Data 4.3.4-8 Distribution of Normalized Turbulence Intensity Components in the Fully Turbulent Region of the Squared-Off Test Configuration 4.3.4-9 Distribution of Normalized Turbulence Intensity Components in the Transitional Region of the Squared-Off Test Configuration LIST OF ILLUSTRATIONS (Cont'd) Number Ti tl e
Page
4.3.4-10 Growth of Turbulence Intensity in the Laminar Region of 8O Each Boundary Layer Low-Pressure Turbine Component Mechanical Configuration Energy Efficient Engine Low-Pressure Turbine Rotor Assembly 5.2.1-2 Low-Pressure Turbine Blade General Characteristics 5.2.1 -3 Blade-to-Disk Attachment Design Detail s 5.2.1-4 Tip Shroud Design Details 5.2.1-5 Low-Pressure Turbine Second Stage Rotor Resonance Diagram 5.2.1-6 93 Low-Pressure Turbine Third Stage Rotor Resonance Diagram 5.2.1 -7 94 Low-Pressure Turbine Fourth Stage Rotor Resonance Diagram 5.2.1-8 Low-Pressure Turbine Fifth Stage Rotor Resonance Diagram 5.2.1-9 Low-Pressure Turbine Fifth Stage Rotor Resonance Diagram Ti tani um-Al umi num All oy Bl ade 5.2.1 -lO 95 Shrouded Turbine Blade Flutter Analysis 5.2.1-II Second Stage Blade Durability Design Conditions and Calculated Stress (Blade material: PWA 1447, Coating: PWA 73) 5.2.1 -12 Third Stage Blade Durability Design Conditions and Calculated Stress (Blade material: PWA 655; No coating) 5.2.1 -13 Fourth Stage Blade Durability Design Conditions and Calculated Stress (Blade material: PWA 655, No coating) 5.2.1 -14 Fifth Stage Blade Durability Design Conditions and lO0 Calculated Stress (Blade material: PWA 655, No coating) 5.2.1 -15 Energy Efficient Engine Second Blade Transient Strains I02 (Flight Propulsion System) xi LIST OF ILLUSTRATIONS (Cont'd) Number Ti tl e 5.2.1 -16 Energy Efficient Engine Third Blade Transient Strains (Flight Propulsion System) I02 5.2.1-17 I03 Low-Pressure Turbine Disk and Hub Assembly 5.2.1-18 Rotor Metal Temperature Distribution Used in Stress and Life Analysis I05 5.2.1-19 I06 Rotor Stress Summary 5.2.1-20 I07 Rotor Low Cycle Fatigue Life Summary 5.2.1-21 1 08 Low-Pressure Turbine Thrust Balance Seal Assembly 5.2.1-22 Low-Pressure Turbine Interstage Inner Cavity Knife-Edge Seal s llO 5.2.1-23 Ill Low Rotor Shaft Assembly 5.2. 1-24 ll3 Low Rotor Shaft Front Temperature and Stress Summary 5.2.1-25 ll3 Low Rotor Shaft Rear Temperature and Stress Summary 5.2.2-I ll5 Low-Pressure Turbine Vane and Case Assembly 5.2.2-2 Low-Pressure Turbine Vane General Characteristics ll6 5.2.2-3 Attachment Scheme Utilized for Third, Fourth, and Fifth Stage Vanes 117 Second Vane Support Structure 117 Location of Second Vane Chordal Cuts Employed for lib Attachment Leakage Control 5.2.2-6 ll9 Low-Pressure Turbine Second Stage Vane Cluster 5.2.2-7 ll9 Typical Third, Fourth, and Fifth Stage Vane Cluster 5.2.2-8 Second Stage Vane Durability Design Conditions and Calculated Stress (Vane material: PWA 1480, Coating: PWA 73) 121 5.2.2-9 Third Stage Vane Durability Design Conditions and Calculated Stress (Vane material: PWA 1455, No coating) xii
LIST OF ILLUSTRATIONS (Cont'd)
Number Ti tl e
Page
5.2.2-I0 Fourth Stage Vane Durability Design Condtions and
Calculated Stress (Vane material: PWA 655, No coating)
5.2.2-II Fifth Stage VaneDurability Design Conditions and
Calculated Stress (Vanematerial: PWA 655, No coating)
5.2.2-12 Energy Efficient Engine SecondVaneTransient Strains
5.2.2-13 Energy Efficient Engine Third Vane Transient Strains
5.2.2-14 Low-Pressure Turbine CaseAssembly
5.2.2-15 Typical Outer Airseal Configuration
5.2.2-16 Low-Pressure Turbine Case, Baffle and Cooling Manifold
Material s Summary
5.2.2-17 Low-Pressure Turbine CaseTemperature Summary at Sea
Level Takeoff Hot Day Conditions
5.2.2-18 Low-Pressure Turbine CaseStress Summary
5.2.2-19 Flight Propulsion System Turbine Exhaust CaseAssembly -
Preliminary Design
5.2.2-20 Integrated Core/LowSpool Turbine Exhaust Case A_sembly
5.2.2-21 Low-Pressure Turbine CaseActive Clearance Control
Features
5.2.2-22 Low-Pressure Turbine CaseModifications for Flight
Propulsion SystemActive Clearance Control System
5.2.2-23 Typical Low-Pressure Turbine Rotor and CaseRadial
Growth Causedby Centrifugal Force, Thermals, and
Pressure
5.2.3-1 Turbine Intermediate Case
5.2.3-2 Turbine Intermediate CaseStructural Strut Details 142
5.2.3-3 Bearing Compartment Axial and Radial Maneuverand
Inbal ance Loads
5.2.3-4 Structural Strut-to-Torque Ring Weldment
xiii
LIST OF ILLUSTRATIONS (Cont'd)
Number Ti tle
Page
5.2.3-5 Turbine Intermediate Case Strut Fairing 145 5.2.3-6 Turbine Intermediate Case Outer Case and Outer Case Heatshi el d Detai Is 5.2.3-7 Engine Rear Mount Details 5.2.3-8 Routing for Nos. 4-5 Bearing Compartment Oil Supply, Scavenge and Drain Lines 147 5.2.3-9 Turbine Intermediate Case Front and Rear Thrust Balance Seal Lands 5.2.3-I0 Turbine Intermediate Case Structure Illustrating Complex Geometry 149 5.2.3-II NASTRAN Model Used to Analyze Nonaxisymetric Structural Loads 5.2.3-12 NASTRAN Model Used to Analyze Axisymetric Loads 152 5.2.3-13 Bearing Load Diagram and Spring Rates Resulting from NASTRAN Analysis 153 5.2.3-14 Radial and Circumferential Deflections Caused by Thermal Gradients Between Strut and Cases 5.2.3-15 Axial Deflections Caused by Thermal Loads and Thrust Balance Pressure Loads 5.2.3-16 Location of Minimum Clearance Between Structural Strut and Aerodynamic Fai ring 154 5.2.3-17 Structural Strut Stress and Life Summary for Flight Propulsion System and Intergrated Core/Low Spool 156 5.2.3-18 NASTRAN Model Utilized to Conduct Turbine Intermediate Case Strut Fairing Durability Analysis 5.2.3-19 Hot-Spot Temperature Profiles Utilized in Strut Fairing Durability Analysis 158 5.2.3-20 Flight Propulsion System Structural Analysis Summary 5.2.3-21 Intergrated Core/Low Spool Structural Analysis Summary xiv
LIST OF ILLUSTRATIONS (Cont'd)
Number Ti tl e
Page
5.2.3-22 Illustration of Flight Propulsion System Strut Fairing Transient Strain Calculation Procedure 160 5.2.3-23 Summary of Flight Propulsion System Strut Fairing Transient Strains at Maximum Stress Locations 161 5.2.4-I Low Rotor Bearing Locations 5.2.4-2 Rotor-Frame Model Used for Critical Speed Analysis 5.2.4-3 Strain Energy Results from Component Preliminary Design Critical Speed Analysis 5.2.4-4 Increase in Low-Pressure Turbine Mode Strain Energy as Low Rotor Design Evolved 5.2.4-5 Effects on Fan, Low-Pressure Compressor and Shaft Mode Responses Due to the Addition of an Oil-Film Damper to the Number 5 Bearing 5.2.4-6 Illustration Showing Effectiveness of Damped Number 5 Bearing in Reducing Low-Pressure Turbine Mode Imbalance Response 5.2.4-7 Sumary of Critical Speed Analysis Results for Integrated Core/Low Spool and Flight Propulsion System 5.2.4-8 Mode Shape for Integrated Core/Low Spool Low-Pressure Turbine Mode -- Damped Number 5 Bearings 5.2,4-9 Mode Shape for Integrated Core/Low Spool Fan Mode -- Damped Number 5 Bearing 5.2.4-I0 Mode Shape for Integrated Core/Low Spool Shaft Mode -- Damped Number 5 Bearing 5.2.5-I Low-Pressure Turbine Secondary Flow System 5.2.5-2 Pressure-Balanced Cooling Air Distribution System 5.2.5-3 A-Frame Rotor Construction Showing Turbine Rim Cooling Air Flow Di stri buti on 5.2.5-4 Disk Rim and 'Wing' Support Structure Cooling Scheme XV LIST OF ILLUSTRATIONS (Cont'd) Number Ti tle 5.2.5-5 Rotor Finite Element Model Showing Temperature Distribution at Steady-State Sea Level Takeoff Hot Day Engine Operating Conditions 174 5.2.5-6 Temperature Versus Time History at Selected Second Stage Low-Pressure Turbine Disk Locations During Severe Engine Power Excursions 5.2.5-7 Temperature Versus Time History at Selected Fourth Stage Low-Pressure Turbine Disk Locations During Severe Engine Power Excursions 176 5.2.5-8 Low-Pressure Turbine Case Cooling Flow Distribution System 177 5.2.5-9 Details of Case Hook Cooling Configuration 177 5.2.5-I0 Cooling Flow Film Coefficient Variation as a Function of Inner Case-To-Baffle Gap Width 178 5.2.5-II Comparison of Integrated Core/Low Spool and Flight Propulsion System Low-Pressure Turbine Case Cooling Ai rflow Di stri buti ons 5.2.5-12 Design Features Utilized to Thermally Isolate the Inner Case and Case Hooks from the Hot Gaspath 5.2.5-I 3 Case Finite Element Model Showing Temperature Distribution at Steady-State Sea Level Takeoff Hot Day Operating Conditions 181 5.2.5-14 Temperature Distribution at Low-Pressure Turbine Third Vane Attachment at Steady-State Sea Level Takeoff Hot Day Operating Conditions 5.2.5-15 Temperature Versus Time History at Selected Second Stage Low-Pressure Turbine Vane/Outer Airseal Locations During Severe Engine Power Excursions 182 5.2.5-16 Temperature Versus Time History at Selected Fourth Stage Low-Pressure Turbine Vane/Outer Airseal Locations During Severe Engine Power Excursions 183 5.2.5-17 Low-Pressure Turbine Second Stage Outer Airseal Radial Deflection with Active Clearance Control in Operation xvi
LIST OF ILLUSTRATIONS (Cont'd)
Number Ti tl e
Page
5.2.5-18 Low-Pressure Turbine Fourth Stage Outer Airseal Radial
Deflection with Active Clearance Control in Operation
5.2.5-19 Turbine Intermediate CaseSecondary Flow System
5.2.5-20 Integrated Core/Low Spool Turbine Intermediate Case
Temperature Mapat Sea Level Takeoff Engine Operating
Conditi ons 185
5.2.5-21 Temperature-TimeHistory for Turbine Intermediate Case
Torque Box-To-Strut Case Support Structure During
Temperature Excursions
xvii
LIST OF TABLES
Ti tl e
Table Pa e
3.2-1 Goal AerodynamicEfficiency
ll
4.1-I Turbine General Aerodynamics
4.2.1 -I Intermediate CaseGeometry Comparison
4.2.2-I
Integrated Core/Low Spool Design
4.2..3-I Airfoil Quantity by Stage
4.2.4-I Low-Pressure Turbine Airfoil Design Summary
4.2.4-II
Gas Triangle Summary
Exit Guide VaneCharacteristics
4.2.4-III
4.2.4-IV
Comparisonof Navy Foil to Energy Efficient Engine
Goal Performance for Low-Pressure Turbine
4.5-I
4.5-II Low-Pressure Turbine Efficiency Update
4.5-III
Detail Design Report Status Performance
5.l-I Life Requirements for Major Turbine Subassemblies
5.2.1-I Blade Attachment Stress Summary
5.2.1 -II
Tip Shroud Stress Summary
lOl
5.2.1 -III Blade Airfoil Creep Strength Margins
lOl
5.2.1 -IV Summary of Predicted Blade Lives
I04
5.2.1 -V Disk Structural Summary
5.2.2-I Airfoil Stress Summary
5.2.2-II VaneAirfoil Creep Strength Margins
5.2.2.-III Summary of Predicted Vane Lives
5.2.2-IV CaseThickness Requirements for Blade Containment
Predicted Tailplug Resonant Frequencies
5.2.2-V
xviii
LIST OF TABLES (Cont'd)
Table Titl e
Page
5.2.2-VI
Active Clearance Control SystemCooling Flow Management
5.2.2-Vli
Energy Efficient Engine Low-PressureTurbine
Blade Tip Clearance Summary
5.2.2-VIII
Energy Efficient Engine Low-Pressure Turbine
Interstage Seal Clearance Summary
5.2.2-IX
Energy Efficient Engine Low-Pressure Turbine Blade Tip Clearance Summary
5.2.2-X
Energy Efficient Engine Low-Pressure Turbine Interstage Seal Clearance Summary
5.2.3-I 157
Strut Bolt Stress Summary
5.2.3-II Gas Environment Definition 158
5.2.3-III Predicted Lives 1 61
5.2.6-I
Preliminary Weight Summary for Integrated Core/Low Spool Turbine Intermediate Case and Low-Pressure Turbine Component xix
SECTION 1.0
SECTION 1.0 SUMMARY A four stage low-pressure turbine component has been designed to power the fan and low-pressure compressor system for the Energy Efficient Engine. In con- junction with the low-pressure turbine definition, designs for a turbine in- termediate case and an exit guide vane assembly have been also established.
The low-pressure turbine aeromechanical design incorporates numerous advanced features to enhance efficiency, durability and performance retention. The aerodynamic definition is aggressive, considering the large expansion ratio and high inlet swirl. To attain a successful aerodynamic design, several fea- tures were incorporated including low loss airfoil designs, counterrotating spools, and a low ratio of through flow to wheel speed (Cx/U).
Additional performance gains are achieved by minimizing endwall losses with the use of flow guides on airfoil platforms, improved interstage sealing and active clearance control. Turbine airfoils are fabricated from high strength/ high temperature materials to meet life and durability requirements. The use of these materials negates the requirement for cooling, which contributes substantially to a higher level of performance.
The combined effect of these technology features results in a status component efficiency of 91.3 percent for the flight propulsion system and 90.3 percent for the integrated core/low spool. This represents a 2.3 percent improvement for the flight propulsion system and a 1.3 percent improvement for the inte- grated low spool relative to the low-pressure turbine component in the refer- ence base engine. This prediction for the flight propulsion system is slightly below the efficiency goal of 91.5 percent mainly because of maneuver deflec- tions, which are calculated to require greater blade tip clearances than goal values. Calculated life of major parts meets or exceeds design goals.
Separate from the turbine module is the exit guide vane assembly. The exit guide vanes are designed to present the mixer with a low Mach number, zero swirl gas stream. This component is designed for low loss performance with a predicted total loss of 0.9 percent APT/P T.
The turbine intermediate case is a compact configuration, containing eleven structural struts to support the number 4 and 5 bearing compartment. The struts are encased by aerodynamically-shaped fairings that turn the gas flow 5 degrees to provide the low-pressure turbine inlet with the proper flow field.
A pressure loss of 0.7 percent Z_PT/PT, which is consistent with the goal, is predicted for the intermediate case designed for a future flight propulsion system. However, duct pressure loss is expected to be 1.5 percent APT/P T for the integrated core/low spool because of geometry differences to maintain aerodynamic compatibility with the low-pressure turbine.
The low-pressure turbine, exit guide vane and intermediate case designs are a positive step towards improving engine fuel efficiency on a component level.
Many of the technology advancements in the areas of aerodynamics, structures and materials are universally applicable to any engine of the next generation.
Technology verification is the next step, and this will be accomplished during the scheduled integrated core/low spool test program.
SECTION 2.0
SECTION 2.0
INTRODUCTION
The Energy Efficient Engine Component Development and Integration Program,
sponsored by the National Aeronautics and Space Administration, is directed
towards demonstrating the technology required to achieve greater fuel effi-
ciency for future commercial gas-turbine engines. The overall program goals
include a reduction in fuel consumption of at least 12 percent and a reduction
in direct operating cost of at least 5 percent relative to a Pratt & Whitney
Aircraft JT9D-7A base engine. To demonstrate the technology required to
achieve these goals, the program currently consists of three active tasks.
These include:
Task 1
Flight Propulsion System Analysis, Design and Integration
Task 2
Component Analysis, Design and Development
Task 4
Integrated Core/Low Spool Design, Fabrication and Test
Under Task 2 of the program, two major accomplishments have been the design of
the turbine intermediate case and the low-pressure turbine component. These
components utilize technology advancements in the areas of aerodynamics
materials and structure-mechanics to meet the requirements for the integrated
core/low spool as well as a future flight propulsion system. The program
schedule for the design of the turbine intermediate case and low-pressure
turbine system, including supporting technology programs, is shown in Figure
2-I.
This report presents a comprehensive description of the aerodynamic and
thermal-mechanical designs of the intermediate case and low-pressure turbine
components. The following section, Section 3, provides an overview of the
component designs and the predicted performance. Section 4 contains the
details of the aerodynamic designs, and Section 5 presents a discussion of the
thermal-mechanical designs. Concluding remarks are presented in Section 6.
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COMPONENT ANALYSIS AN00ESXGN I PRELIMINARY DESIGNS I I EARLY FASRICAT ION I TO SUPPORTING TECH- :NOLCGY PROGRAMS REOUIREIENTS I I v COMPONENT FABRICATION I I 91 ) I I TO IC/LS I
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SUPPORTING TECHNOLOGY ASSEMBLY I I I I I I 81_JNOARY LAYER I I I I J AIRFOIL SELECTION I I I SUBSONIC CASCADE 9 *0 V i I I I TRANSITION DUCT I
I
LPT DEFINITION I I TO IC/LS OETAILEO I DESIGN 1 I I I I I I I I I TASK 4 I
Ii '
I I I: PDR OUR IC/LS - ANALYSISANO DESIGN I" I i I EARLY PROCUREMENT I I I INITIATED
I I
i, IC/LS - FABRICATION V v I ; TRR TRR I ? V _Vl IC/LS - ASSEMBLY 9 TEST Z I TEST 1 I IC/LS - TEST 9----9 V--.V I I 9----V 9----V [C/LS - POST-TEST ANALYSIS "14OENOTES f_OR MILESTONE *O DENOTES KEY OEC, ISIOI| POINT
Figure 2-I
Program Schedule for Energy Efficient Engine Turbine Inter-
mediate Case and Low-Pressure Turbine Design Efforts
SECTION 3.O
SECTION 3.O DESIGN OVERVIEW 3.1 TURBINE INTERMEDIATE CASE AND LOW-PRESSURE TURBINE CONCEPTS A cross-sectional view of the turbine intermediate case, low-pressure turbine and exit guide vane, as integrated in the turbine section of the Pratt & Whitney Aircraft Energy Efficient Engine, is shown in Figure 3.l-l. The princi- pal design goals include a turbine efficiency of 91.5 percent, an intermediate case pressure loss of 0.7 percent APT/P T and an exit guide vane pressure loss of 0.9 percent APT/P T.
I INTERMEDIATE CASE LJ HIGH LOW-PRESSURE EXIT GUIDE TURBINE PRESSURE VANE TURBINE ROTOR Major Components in the Energy Efficient Engine Turbine Figure 3.l-l Section 3.l.l Turbine Intermediate Case In the Energy Efficient Engine design, the turbine intermediate case provides the gaspath transition between the high-pressure turbine exit and the low- pressure turbine inlet. It also provides a frame for the turbine rotor sup- ports and rear engine mounts. This arrangement permits use of a straddle mounted, or simply-supported, high-pressure rotor system for improved clear- ance control. Another advantage is that the number 4 and 5 bearing compartment can be located after the high-pressure turbine disk in a less severe thermal envi ronment.
However, the intermediate case introduces certain technical challenges that have been addressed during the design process. These include: (1) high gaspath temperatures which can cause excessive thermal stresses and distortion resul- ting in limited life; (2) strut fairing airfoil size constraints that limit the internal passage volume available for the structural struts and oil ser- vice lines; and (3) high thrust balance loads on the support structure, causing high bending stresses.
A cross-sectional view of the turbine intermediate case is shown in Figure 3.1.l-l. The design is compact and minimizes pressure loss resulting from excessive boundary layer formation and local separation.
Structural support for the number 4 and 5 bearing compartment is provided by eleven support struts. To minimize blockage and pressure loss, the struts are encased by aerodynamically-shaped fairings. 0il supply and scavenge tubes to the bearing compartment are routed through hollow fairings in the cavity just forward of the strut leading edge. These lines are fully insulated for protec- tion from the high temperature environment. Also, the struts are cooled with a small percentage of high-pressure compressor bleed air.
¢
#4 BEARING HOUSING HPT REAR THRUST / BALANCE SEAL STRUCTURAL STRUTS (11) STRUT FAIRINGS (ii) OUTER CASE _RETENTION BOLT OIL SERVICE TUBES (NOT SHOWN) Figure 3.1.l-I Cross-Sectional View of Turbine Intermediate Case
3.1.2 Low-Pressure Turbine
The low-pressure turbine is a highly loaded four-stage design. The turbine for
a future flight propulsion system is shownin Figure 3.1.2-I and has several
features to minimize leakage, improve aerodynamics and reduce weight. Someof
these features are:
o Double wall case construction with case-tied outer air seals and an
internal active clearance control system for controlling blade tip
clearances
o
Accomplishing the required expansion ratio (5.5) in four stages
rather than five, thereby reducing componentweight
o
Increased velocity ratio (Vm = 0.468) comparedto the reference
base engine coupled with a low ratio of through flow velocity to
wheel speed (Cx/U) for improved aerodynamic performance
o Low loss aft-loaded airfoils with elliptical leading edges
o Controlled vortexing to reduce root section aerodynamic losses
o
Stepped labyrinth inner air seals in the third through fifth stages,
and case-tied knife-edge inner air seals (thrust balance seals) in
the second stage to control leakage
o
Blade leading and trailing edge flow guides to minimize cavity
recirculation losses
In addition, the low-pressure turbine is counter rotating relative to the
high-pressure turbine to enhance aerodynamicperformance by reducing turning
losses in the second stage vane. Performance is further improved by the elimi-
nation of cooling in the second stage through the use of high strength and
high temperature capability materials for the airfoils. The combinedeffect of
all these features translates into a low-pressure turbine efficiency benefit
of 2.3 percent for the Energy Efficient Engine flight propulsion system
relative to the reference base engine.
Separate from the turbine module is the turbine exit guide vane assembly,
which is designed to present the mixer with a low Machnumber, zero swirl gas
stream. The vanes are a hollow, welded construction in the flight propulsion
system. Solid castings are used, however, in the integrated core/low spool as
a cost saving feature.
THRUST BALANCE SEAL 3B 4V 48 5V 5B TURBINE EXIT VANE INNER CASE AiR SEALS CASE TIED *LPT INLET GUIDE VANE (IGV} IS 2ND STAGE VANE OF COMPLETE ENGINE TURBINE SECTION Figure 3.1.2-I Cross-Sectional View of Low-Pressure Turbine Component 3.2 PREDICTED AERODYNAMIC PERFORMANCE Aerodynamic studies conducted prior to detailed mechanical design confirmed that the initially established goal of 0.7 percent APT/P T loss for the turbine intermediate case in the flight propulsion system could be attainable.
Similar studies conducted for the integrated core/low spool predicted that loss would be 1.5 percent APT/P T because of geometry differences necessary to maintain aerodynamic compatibility with the low-pressure turbine. Thisloss level is consistent with the integrated core/low spool goal of 1.5 percent APT/P T.
The turbine intermediate case structural struts are designed for a 15,000 hour life for the flight propulsion system and 50 hours of hot time for the inte- grated core/low spool.
The goal performance of the low-pressure turbine componentis summarizedin
Table 3.2-I. The items of improvement relative to the reference base engine
are indicated. Studies conducted prior to completion of the mechanical design
indicated that flight propulsion system aerodynamic efficiency would surpass
the goal if the clearances assumed could be maintained. An updated performance
prediction (status) after completion of the mechanical design is presented in
Section 4.5, Performance Status and Adjustment.
Turbine life predictions for the flight propulsion system are consistent with
the goals of 15,000 hours or 3300 missions for the airfoils and 30,000 hours
or 20,000 missions for the disks, hubs, shaft and static structure. For the
integrated core/low spool, life estimates exceed the goal of at least 50 hours
of hot time operation or lO00 cycles.
TABLE 3.2-I
GOAL AERODYNAMIC EFFICIENCY
(AerodynamicDesign Point)
89.0%
Reference Base Engine
+0.3
Increased Velocity Ratio
+l .0
Improved AerodynamicBenefits
+l. 2
ReducedClearance and Flow Guides
TOTAL 91.5%
Assumptions:
(1) 0.051 cm (0.020 in) Clearance for Inner Seals
(2) 0.051 cm (0.020 in) Clearance for Outer Blade Tip Seals
(3) 0.152 cm (0.060 in) Clearance on Flow Guides
(4) No Instrumentation
SECTION 4.0
SECTION 4.0
INTERMEDIATE CASE ANDLOW-PRESSURE TURBINE AERODYNAMIC DESIGN
4.1 AERODYNAMIC DESIGN PARAMETERS
The aerodynamic definition of the intermediate case and low-pressure turbine
componentswas based on a series of analyses to establish the flowpath, air-
foil contours and flow characteristics to achieve the highest level of per-
formance within the constraints of the basic mechanical configuration. In
addition, the final designs were influenced largely by the results obtained
from supporting technology programs which were in progress concurrent with the
design process.
The general parameters governing the aerodynamic design of the intermediate
case and low-pressure turbine are presented in Table 4.l-I. The values listed
in this table are for the aerodynamic design point of the integrated core/low
spool.
TABLE 4.l-I
TURBINE GENERAL AERODYNAMICS
(At AerodynamicDesign Point)
INTERMEDIATE CASE
Annulus Area Ratio
l .57
Effective Area Ratio
l .42
a2 - e2 Free Vortex, (deg)
1.5
APT/P T, (%)
LOW-PRESSURE TURBI NE
Stages
Rotation
Counter
3902.0
Speed (rpm)
Inlet Total Pressure, Pa (psia)
319,229.2 (46.3)
ll61 (2090)
Inlet Total Temperature, K.__(°R)
Inlet Corrected Flow, W _/T/Pt 69.342
323.17
Exit Corrected Flow, W Vrf_Pt
Pressure Ratio
5.51
A H, _Btu/sec )
12760.0
0.468
Mean Velocity Ratio
Work Factor, (Ah/u 2) 2.28
0.76
Average Flow Coefficient, Cx/u
Work Split, %
23/24/26/27
Mean Reaction
.45/.45/.45/.46
Goal Clearances, cm(in)
o.o51 (0.020)
Goal Efficiency Split, %
90/88.9/90.4/90.9
91.5
Goal Overall Efficiency, %
]!
4.2 FLOWPATH AND AIRFOIL CROSS SECTION DEFINITION
4.2.1 Intermediate Case Flowpath
The basic flowpath of the turbine intermediate case for the Energy Efficient
Engine required certain modifications in the integrated core/low spool to
maintain aerodynamic compatibility with the low-pressure turbine component.
Figure 4.2.1-I shows a comparison of the final flowpath for the integrated
core/low spool and the preliminary flowpath established during the earlier
part of the program, but scaled to the current thrust size of 160,135 N
(36,000 Ib). A further comparison of the geometric differences of the two
designs is provided in Table 4.2.1-I. As indicated, several major differences
are apparent because of the requirements of the integrated/core low spool
design.
(19) INTEGRATED CORE/LOW SPOOL (IC/LS) .... PRELIMINARY DESIGN REVIEW (PDR) (SCALED) (18) (17) r- IC/LS INLET E
u_ 40
(16) a PDR INLET <( nr (15) (14)
m_ J
7.6 10.1 1 2.7 1 5.2 17.7 20.3 0 2.5 5.0 (3) (4) (5) (6) (7) (8) (0) (1) (2)
AXIAL DISTANCE, cm (in)
Turbine Intermediate Case Flowpath for the Integrated Core/
Figure 4.2.1-I
Low Spool and the Preliminary Design Scaled to a 160,135 N
(36,000 Ib) Thrust Size
The increase in corrected flow of the integrated core/low spool cycle compared
to the flight propulsion system cycle dictated a 5 percent increase in the
intermediate case exit annulus area. Also, the intermediate case support strut
was changed from a nonworking to a working airfoil, turning the gas flow 5
degrees back towards axial. These modifications provide the proper inlet flow-
field to the low-pressure turbine in the integrated core/low spool. In addi-
tion, to obtain proper case stiffness," the posture of the support strut was
]2
TABLE 4.2.1 -I
INTERMEDIATE CASE GEOMETRY COMPARISON
Integrated Core/
DUCT Preliminary Design Low Spool
21.0 (8.2) (7.7 scaled) 19.8 (7.8)
Length, cm (in)
3.0
3.0
L/H
I. 57
l .50
Ann. Area Ratio
l .42
l .26
Eff. Area Ratio
Duct Exit Gas Angle (deg)
3O 35
Hub
Mi dspan
Tip
Duct Exit Mach No.
0.40
Hub O. 50
0.39
0.44
Midspan
0.41 0.30
Ti p
NONWORKING WORKING
STRUT
No. of Foils 14 II
Type 65C/A 4OO
2 -_2 Free Vortex, deg 0 5
changed by canting it tangentially approximately II degrees opposite rotation
and shifting the sections axially rearward. Finally, the fairing cross section
was changed from a 65 circular arc to a 400 series airfoil to increase the
incidence range. The width of the fairing was set by the area required for
adequate oil scavenging from the number 4 and 5 bearing compartment.
For high-pressure turbine blade structural considerations, the f!owpath length
was increased 0.2 cm (O.l in) for a final duct length of 19.8 cm (7.8 in).
Moreover, the number of fairings was reduced from 14 to II to preclude vibra-
tory excitation of the high-pressure turbine blade precipitated by a passing
frequency.
Overall, the net effect of these changes is a more aggressive flowpath design, as indicated in Figure 4.2.1-2 by the considerably higher level of diffusion,
with an attendant higher system pressure loss. This loading is achievable, as
demonstrated by the results of a Transition Duct Model Test Program (Ref. l)
(Section 4.3.2). The estimated total duct pressure loss for the integrated
core/low spool is 1.5 percent APT/PT.
For the flight propulsion system, however, the intermediate case flowpath
retains many of the features in the original definition to achieve the goal
pressure loss of 0.7 percent. These include maintaining an exit annulus area
ratio of 1.50, which is 5 percent lower than that established for the inte-
grated core/low spool, and a free vortex (nonworking) type fai ri ng.
i 0.6 ,,,,="'_N T EG 0594 RATED CORE/ TiP WALL ' _-- 1 LOW SPOOL / TRAILI/NG EDGE
/ /
0.4
- /
0.
k- a.
Cq
>
I I
O_ a.
TRAILING EDGE 'x _ r%._......- _- 0.415 o_ 0.4 Q.
B_ADING EDGE _ 0.2 U
,,=..v I I I " I I
I I !
O,Ocm 2.5cm 5.0cm 7.6cm 10.1 cm 12.7 cm 15.2 cm 17.7 cm 20.3 cm (0.0 in) (1.0 inl (2.0 in) (3.0 in) (4.0 in) (5.0 in) {6.0 in) (7.0 in) (8.0 in) AXIAL DISTANCE, cm (in)
Figure 4.2.1-2
Intermediate Case Wall Pressure Distribution Profiles,
Showing the High Loading of Integrated Core/Low Spool Design
4.2.2 Intermediate Case Strut Fairing
The fairing, which encases the structural strut, is a standard 400 series air-
foil. The definition was established on the basis of airfoil thickness, length
and leading and trailing edge radii. Also, as shown in Figure 4.2.2,1, a 400
series airfoil offers a larger incidence range with only a slight compromise
in pressure loss in comparison with other airfoil series.
The main criteria governing the fairing aerodynamic design were thickness and
flow turning. To accommodate the structural strut, the thickness was estab-
lished at 2.54 cm (l.O in). Flow turning, as discussed earlier, was determined
at 5 degrees to satisfy low-pressure turbine aerodynamics in the integrated
core/low spool. Axial spacing as well as the canted posture of the fairing
were dictated by structural requirements.
A summary of the final fairing definition is contained in Table 4.2.2-I. Also,
Figures 4.2.2-2 through 4.2.2-5 present corresponding pressure distributions
and section aerodynamics for each of the four radial sections identified in
Table 4.2.2-I. Figure 4.2.2-6 shows the stacking arrangement of th_ fairing.
]4
0.01 0 400 SERIES [] 65C/A SERIES _C/A
I I I I I I
-4 -2 0 +2 +4 +6 INCIDENCE, DEG
Figure 4.2.2-I
Incidence Range Capability as a Function of Pressure Loss,
Showing Greater Range With 400 Series Foil
TABLE 4.2.2-I INTEGRATED CORE/LOW SPOOL DESIGN Transition Duct Strut Fairing Design Section (Planar) l 2 3 4 Radius, cm(in) 35.003(13.781) 39.187(15.428) 42.735(16.825) 47.444(18.679) Foil Type 400 400 400 400 Number of Foils 11 ll 11 11 Axial Chord, cm(in) 10.97(4.32) 11.25(4.43) 11. S0(4.53) 11.78(4.64) Actual Chord, cm(in) 20.39(8.03) 17.29(6.81 ) 15.59(6.14) 14.93(5.88)
Maximum Thickness, cm(in) 2.5(1.0} 2.5(1.0) 2.s(1.0) 2.5(l .0)
Gap/Chord 0.98 1.29 1.59 1.82 Inlet Metal Angle, 150.2 144.5 138.7 131.6 Inlet Gas angle, 148.5 142.5 141.8 148.1 Inlet Mach Number 0.51 0.50 0.48 0.43 33.4 44.1 51 54.9 Exit Metal Angle, Exit Gas Angle, 31.9 35.3 38.4 42.4 Exit Mach Number 0.37 0.40 0.40 0.31
]5
1.0 0.2 0.9 0,4 0.8 0.6 0.7 0,8 X B 0.6 (_ 1.0 >- 0.5- 1.2 0.4 B 1.4
o.3 I I I I I I
I I I I I
1.6 0.20 0.40 0.60 0.80 1.00 -0.20 0.00 0 0.2 0.4 0.6 0.8 1.0 X/BX, AIRFOIL X/BX, AIRFOIL
Figure 4.2.2-2 Fairing Aerodynamic Definition for Section 1 -- Integrated
Core/Low Spool Intermediate Case
1.0 0.9 0.8 0.7_ ).- o_ X 0.6 0.6 >..
0.8 0.5 1.0 0.4 I I I I I
, I I I I I 0.3 I
1.2 -0.20 0.00 0.20 0.40 0.60 0.80 1.00 0 0.2 0.4 06 0.8 1.0 X/BX, AIRFOIL X/BX, AIRFOIL
Figure 4.2.2-3 Fairing Aerodynamic Definition for Section 2 -- Integrated
Core/Low Spool Intermediate Case
1.0 i 0 m 0.9 0.8 0.7 M- D.
X _m O6 0.6 )- 0.8 0.5 1.0 0.4
121, I I I I l O3J
I I I I I I 0 0.2 0.4 0.6 0.8 1.0 - 0.20 0.00 0.20 0.40 0.60 0.80 1.00 X/BX, AIRFOIL X/BX, AIRFOIL
Fi gu re 4.2.2-4
Fairing Aerodynamic Definition for Section 3 -- Integrated
Core/Low Spool Intermediate Case
1.o m o.9_ 0.2 0.8-- 0.4 0.7-- {- o_ X 0.6 0.6-- 0.5- 1.0 0.4 -- 08 f
1.2 I I I I I
0.3 I I I I I I 0 0.2 0.4 0.6 0.8 1.0 0.00 0.20 0.40 0.60 0.80 1.00 X/BX, AIRFOIL X/BX, AIRFOIL
Figure 4.2.2-5
Fairing Aerodynamic Definition for Section 4 -- Integrated
Core/Low Spool Intermediate Case
]?
ROOT TIP
Figure 4.2.2-6 Strut Fairing Stacking Arrangement
4.2.3 Low-Pressure Turbine Flowpath
The low-pressure turbine flowpath evolved from a series of iterative design
analyses. The final definition, as presented in Figure 4.2.3-I, is sized for a
thrust of 160,135 N (36,000 Ib), consistent with the integrated core/low spool.
As shown in Figure 4.2.3-I, the turbine flowpath, from the inlet to the exit
of the last or fifth-stage blade, is 61.23 cm (24.12 in) in length with a
maximum elevation of 65.93 cm (25.96 in). The axial gap between stages is
sufficient for the incorporation of flow guides and allows adequate margin to
preclude clashing that may result from pressure load deflections on the air-
foils, vibratory deflections and mechanical tolerances. Flow guides are exten-
sions of the airfoil platforms and serve to minimize cavity recirculation
losses.
The four turbine stages contain a total of 756 airfoils. A delineation of the
number of airfoils for each stage is presented in Table 4.2.3-I. All airfoils
are a high aspect ratio and designed for low loss. Turbine blades in all
stages are tip shrouded.
]8
38.1 (15.0) I 43.1 m (17.0) I 48.2 m
E
(19.0) m 53.3 I (21.0) < 5V 5B B rr EGV 58.4 m (23.0) m 63.5 n (25.0)
68.5 I I I I I I I I I I I
I I I I I I I
(27.0) 20.3 25.4 30.4 35.5 40,6 45.7 50.8 55.8 60.9 66.0 71.1 76.2 81.2 (8.0) (10.0) (12.0) (14.0) (16.0) (18.0) (20.0) (22.0) (24.0) (26.0) (28.0) (30.0) (32.0) AXIAL DISTANCE, cm (in)
Figure 4.2.3-I
Low-Pressure Turbine FLowpath
TABLE 4.2.3-I
AIRFOIL QUANTITY BY STAGE
Second Stage*
Vanes 54
Blades 120
Third Stage
Vanes 72
Blades 96
Fourth Stage
Vanes 84
Blades lO0
Fifth Stage
Vanes 108
Blades 122
Total 756
* The low-pressure turbine inlet guide vane is the same as the second-
stage vane in the complete Energy Efficient Engine turbine section.
In terms of component aerodynamics, the turbine is characterized by a high
expansion ratio (5.5) compared to the number of stages, a low mean velocity
ratio (0.468) and a low ratio of throughflow to wheel speed (Cx/U) (0.76).
Also, the low-pressure turbine is counterrotating in relation to the single-
stage high-pressure turbine. This results in a low camber inlet guide vane for
improved performance. Airfoils inthe remaining stages have a high level of
turning. The rotating stages are defined with a reaction level of 45, 45, 45
and 46 percent to enhance efficiency, and the turbine stage work split is
defined at 23, 24, 26 and 27 percent, respectively. These levels provide both
an optimum efficiency and exit guide vane loss level.
]9
The addition of the turbine exit guide vane increases the flowl)ath length by
21.59 cm (8.5 in) for a total length of 82.85 cm (32.62 in). The radius is
increased by 2.74 cm (I.08 in) for a total height of 68.68 cm (27.04 in).
There is a total of 30 turbine exit guide vanes. Endwall contours are defined
for compatibility with the internal exhaust mixer. The airfoil is also con-
toured to present the mixer with an acceptable inlet Mach number and zero
swirl.
The adaptation of the turbine flowpath designed for the integrated core/low
spool to the flight propulsion system would conceptually require only one
minor modification. This would entail a 5 percent reduction in the turbine
inlet annulus area to accommodate the lower inlet corrected flow of the flight
propulsion system cycle. The revised flowpath would also necessitate a slight
change to the airfoils.
4.2.4 Turbine Airfoil Definition
The turbine airfoils are comprised of individually designed sections. Contours
and performance characteristics were defined with the use of Pratt & Whitney
Aircraft's analytical techniques used specifically for designing turbine air-
foils. Basically, a streamline computer design simulation generated the radial
aerodynamic environment. This information served as input to the interactive
airfoil design system for definition of the external contour. Analyses were
then performed to ascertain pressure distribution and boundary layer character-
istics. On the basis of these results, iterations of the airfoil shape were
made to optimize performance. In addition, boundary layer and plane cascade
tests, conducted as part of the Low-Pressure Turbine Subsonic Cascade
Technology Program (Ref.2) and Boundary Layer Technology Program (Ref.3), were
instrumental in substantiating the selected airfoil contours. Salient results
from these programs are summarized in Sections 4.3.3 and 4.3.4 of this report.
Pertinent characteristics of the final airfoil designs are presented in Table
4.2.4-I. Velocity triangle data are summarized in Table 4.2.4-II. The velocity
triangles were established using a controlled vortex philosophy. In this man-
ner, vane trailing edge root sections were opened and the trailing edge tip
sections were closed. This maintains the meanline aerodynamics, while reducing
the turning done by the blade root sections and increasing the blade root
reactions. The change in root section aerodynamics produces a reduction in
endwall loss generation, thus increasing efficiency. Figure 4.2.4-I shows an
example of the effects produced by controlled vortexing in comparison to a
free vortex design for the turbine fourth stage. Fifth-stage blade vortexing
was set to achieve reasonable load factors for the turbine exit guide vane.
Besides controlled vortexing, the airfoils are designed with an elliptical
leading edge to provide a uniform pressure distribution.
Of particular importance, the Low-Pressure Turbine Subsonic Cascade Technology
Program verified the choice of blade and vane contours needed to obtain low
loss characteristics as well as defined the characteristics of the low camber
inlet guide vane. Results showed that an airfoil that continuously accelerates
the flow on the suction surface to the region of the gaging point can provide
an approximate 18 percent reduction in profile loss compared to an airfoil
which achieves maximum velocity further forward on the suction surface. These
lower loss airfoils, termed "aft loaded" because of the shape of the pressure
distribution curve, were selected for the turbine design. Furthermore, the
Subsonic CascadeProgramevaluated the performance and loss characteristics of
the unique low camber inlet guide vane over a range of inlet angles. At design
conditions, measuredprofile losses and secondary losses were 0.54 and 0.56
percent APT/P T, respectively. The vanes also showed a negative incidence
range of 8 degrees, as defined by the point where the loss level is 50 percent
above the design point loss.
Figures 4.2.4-2 through 4.2.4-24 present airfoil contours of different vane
and blade sections in each stage. These figures also summarize the section
aerodynamics and contain the static pressure distribution. Figures 4.2.4-25
through 4.2.4-32 show the stacking arrangements of the airfoil in each stage.
TABLE 4.2.4-I
LOW-PRESSURE TURBINE AIRFOIL DESIGN SUMMARY
Second-Stage Vane
8.76(3.45)
Trailing Edge Span, cm (in)
0.826
Trailing Edge Hub/Tip Ratio
2.203
Aspect Ratio
Second-Stage Blade
9.45 (3.724)
Trailing Edge Span, cm (in)
0.820
Trailing Edge Hub/Tip Ratio
3.545
Aspect Ratio
Thi rd-Stage Vane
Trailing Edge Span, cm (in) II.282(4.442)
0.796
Trailing Edge Hub/Tip Ratio
2.759
Aspect Ratio
Thi rd-Stage Blade
13.314(5.242)
Trailing Edge Span, cm (in)
O. 770
Trailing Edge Hub/Tip Ratio
4.339
Aspect Ratio
Fourth-Stage Vane
Trailing Edge Span, cm (in) 15. 946(6.278)
O. 737
Trailing Edge Hub/Tip Ratio
Aspect Ratio 4.282
Fourth-Stage Blade
Trailing Edge Span, cm (in) 18.224(7.175)
O.711
Trailing Edge Hub/Tip Ratio
Aspect Ratio 5.400
Fi fth-Stage Vane
Trailing Edge Span, cm (in) 19.913(7.840)
0.693
Trai Iing Edge Hub/Ti p Ratio
Aspect Ratio 5.433
Fi fth-Stage Blade
20.863(8.214)
Trailing Edge Span, cm (in)
O. 684
Trail ing Edge Hub/Ti p Ratio
6.285
Aspect Ratio
2]
TABLE 4.2.4-11
GAS TRIANGLE SUMMARY
(Aerodynamic Design Point)
Inlet Exit Inlet Exit Gas Vexi t/ Angle Angle Mach No. Mach No. Turning Conve rgence Vinlet IGV R 144.3" 26.6" 0.396 0.666 9.1" 1.43 l .65 M 141.3" 24.2" 0.394 0.658 14.5" 1.47 1.64 T 139.4" 21.7" 0.306 0.534 19.0" 1.53 1.73 1.27 1.40 43.9" B2 R 23.0" 0.433 0.614 113.2" 24.2" 0.422 0.638 114.4" 1.34 I .48 41.4" I/4R 25.9" 0.394 0.620 113.3" 1.42 1.58 40.7" M 27.6" 0.335 0.580 109.5" 1.51 1.70 42.0" l/4T 50.2" 28.7" 0.256 0.543 101.2" 1.78 2.09 T V3 R 36.0" 30.3" 0.350 0.601 113.6" 1.44 1.69 1.48 1.70 1/4R 41.7" 24.2" 0.373 0.648 114.0" M 47.6" 22.8" 0.350 0.656 109.6" 1.61 ] .85 1.80 2.00 1/4T 54.4" 22.2" 0.307 0.631 103.3" T 59.3" 21.9" 0.266 0.590 98.8" 1.90 2.17 B3 R 53.5" 23.5" 0.380 0.645 102.9" I .42 1.67 42.4" 23.2" 0.383 0.679 114.5" 1.47 l .73 I/4R l .78 M 41.7" 22.7" 0.365 0.665 I15.7" I .56 44.7" 22.1" 0.324 0.629 113.3" l .72 l.90 I/4T 21.7" 0.278 0.599 108.5" I .83 2.12 T 4g.g" 37.5" 28.9" 0.368 0.647 113.6" l .41 1.72 V4 R 22.4" 0.391 0.700 117.5" 1.48 1.79 40.1" I/4R 42.9" M 21.2" 0.357 0.689 115.9" l .62 1.88 20.3" 0.309 0.644 112.2" 1.75 2.03 47.5" I/4T 52.4" T 19.6" 0.279 0.598 108.0" 1.88 2.10 49.2" B4 R 22.1" 0.397 0.692 108.7" l .40 1.70 l .79 38.9" 22.1" 0.401 0.738 118.9" l .49 I/4R 40.9" M 22.2" 0.359 0.689 116.9" l .64 l .87 2.03 2.21 46.7" 22.4" 0.286 0.649 llO.9" I/4T 2.28 2.76 55.5" T 22.5" 0.228 0.643 lOI.9" l .36 l .63 37.2" V5 R 32.4" 0.409 0.685 110.5" 39.4" 27.3" 0.426 0.728 113.3" l .42 l .66 I/4R 22.6" 0.348 0.736 I11.8" 1.70 2.05 45.6" M 53.9" 22.6" 0.297 0.704 103.5" l .89 2.29 I/4T 22.6" 0.289 0.648 99.7" l .87 2.19 57.8" T 37.3" 0.450 0.784 87.5" 1.35 1.69 55.3" B5 R 48.9" 28.8" 0.439 0.818 102.3" 1.43 l .79 I/4R 44.7" 25.5" 0.385 0.791 108.9" l .60 1.98 M 52.0" 27.0" 0.325 0.770 I01.0" l .86 2.28 1/4T 62.7" T 29.5" 0.263 0.741 87.8 ° 2.0 2.72 EGV R 55.7" 90.0" 0.515 0.396 34.3" 1/4R 47.9" 90.0" 0.507 0.373 42.1" M 50.1" 90.0" 0.441 0.350 39.9" I/4T 58.I" 90.0" 0.405 0.366 31.9" T 67.9" 90.0" 0.415 0.390 22.1"
,2°r
m 32 _E3 _ 116__hl_ll_'_,_ < _ , m :::D 108'-- 24 _ I- t- < 1.1.1 20 104
lOO I I I I I
0 2O 40 60 80 100 0 20 40 60 80 100 PERCENT SPAN PERCENT SPAN FOURTH VANE 56r=- a -- 30 a 40 18 _ 1001 I
I I I
0 20 40 60 80 1O0 0 20 40 60 80 1O0 PERCENT SPAN PERCENT SPAN FOURTH BLADE FREE VORTEX ENERGY EFFICIENT ENGINE DESIGN
Figure 4.2.4-I Effects of Controlled Vortexing Compared to Free Vortexing
X m o.6 >- 0 0.2 0.4 0.6 0.8 1.0 X/BX, AIRFOIL 1.0 RADIUS cm (in) 46.012 (18.115) INLET METAL ANGLE 137.5 ° EXIT METAL ANGLE 24.4 ° 0.9 AXIAL CHORD cm'(in) 3.975 (1.565) FOI LS 54.
SOLIDITY (bx) .743 0.8 ELLIPSE RATIO 4/1 CHORD cm (in) 7.244 (2.852) 0.7 INLET GAS ANGLE 140.5 ° EXIT GAS ANGLE 24.2 ° INLET MACH NO. .378 0.6 EXIT MACH NO. .657 INCIDENCE -2.95 ° MMAX .753 Ap/Q T.E. .229 !
0.5
L
0.4
! I ! I I I
0 0.2 0.4 0.6 0.8 1.0 X/BX, AIRFOIL
Fi gure 4.2.4-2 Turbine Inlet Guide Vane Mean Section Aerodynamic Definition
and Pressure Distribution
fl I-- I-- Q3 °e_ o0 EE | I -o ._ 0
=$
I-- g0 I u
o
u_
_4
n'- I-- i rn m d x °r,-
I
I
I
t
o d 6 d d 0 m 0.2 n 0.4 X £D 0.6 >- 0.8 - 1.0 B
| I I I ! !
1.2 0 0.2 0.4 0.6 0.8 1.0 X/BX, AIRFOIL m 1.0 RADIUS cm (in) 47.759 (18.803) INLET METAL ANGLE 36.4 ° EXIT METAL ANGLE 25.2 ° 0.9 AXIAL CHORD cm (in) 2.66 (1.05) FOILS 120.
SOLIDITY (bx) 1.07 0.8 ELLIPSE RATIO 4/1 CHORD cm (in) 2.943 (1.1 59) INLET GAS ANGLE 41.0 ° i:_ 0.7 EXIT GAS ANGLE 25.7 ° INLET Mn .378 EXIT Mn .624 0.6 _ INCIDENCE B4.6 ° MMAX .834 /_ P/Q T .E. .418 0.5-
0.4, I ,, I, I I, I I
0 0.2 0.4 0.6 0.8 1.0
X/BX AIRFOIL
Fi gure 4.2.4-5
Turbine Second Stage Blade Mean Section Aerodynamic Defini-
tion and Pressure Distribution
r- ,_oo__ _,:_ o _._,.. _ _ " "_od _, o o_"_
m 0..
I I I I I t
I I I I I i
_ e- _ . (_ d d d -- °; ,¢ d _ d X d
d x
°]
I I I I I i
L I I I I I
0 _ ao o c_L ,_" ¢.D o d o d d d d d d .Ld/d XB/A I-.
nm o
I
I I I I I I I ,
I I I
o i
d
d
I _0 .--I
d_
u-
,g
o
d
o
rr- X
d x
d I
I I I I _
I I I I
o 0 o_ co _ co to ._.
O0 ",.,0 '_" cN 0
d d o d d d
o d o d
XBIA
-Ld/d
1.2 1.0 0.8 X m 0.6 >- 0.4 0.2 0 0.2 0.4 0.6 0.8 1.0 X/BX, AIRFOIL 1.o RADIUS cm (in) 49.68 (19,56) INLET METAL ANGLE 41.7 ° EXIT METAL ANGLE 21.0 ° 0.9 AXIAL CHORD cm (in) .4.08 (1.61) FOILS 72.
SOLIDITY (bx) .942 0.8 _ ELLIPSE RATIO 4/1 CHORD cm (in) 4.831 (1.902) i.- INLET GAS ANGLE 49.5 ° 0.. 0.7- EXIT GAS ANGLE 22.9 ° INLET Mn .337 EXIT Mn ,657 0.6- INCIDENCE --7.9 ° MMAX .912 Ap/Q T .E. .466 0.5 - 0.4
I I ! I I I
0 0.2 0.4 0.6 0.8 1.0 X/BX, AIRFOIL
Figure 4.2°4-8
Turbine Third Stage Vane Mean Section Aerodynamic Definition
and Pressure Distribution
,_ _._ CN 1.0 0 C_C_I _, CO ... co _do _
_'_'_'g_o_._o_ • • _ _ _,_ _o
A e_ o _.1 I.IJ v e- I---4-.)
_. :._._ oo_ °
o o m i go d -o _0 ._I _J II n_ o.
I-- o 0_ ° d X -o x v_g- o m 0 I I I I " I I I I I I q o .,_ 0 .
cO I-'- ,_ -o d _D d -- -o -- !
a.
LI.
I-
_d
d
,4
x"
en d × " d X o_ o -- O !
I I I I I _ I _ I I (N d d d d - d d o d d XB/A .Ld/d e=, o o c_ o e- 0_i c-4_ _. ,e- _ru_ e- l.
e- _/) I- o_- o 0: e" eU "_ ,.. e- _3 0 I_ *r m •I_ "r- edo u,- (U
I I I I I
I I I I I I
S.- °rm (_) e'- .I_- o m CO _0 d
d
F.-,_ (o
°o 0
I-
"7
o
d
" d
o
ed
,g
d x
d X
I 0
! I I I I I I I I I I I
0 _ _" "..0 oO o d d d d _ d 0 o d o c5 XBIA
id/d
0.2 0.4 X o3 0.6 >- 0.8 1.0
• I I I I I I
1.2 0 0.2 0.4 0.6 0.8 1.0 X/BX, AIRFOIL 1.0 RADIUS cm (in) 51.191 (20.1 54) INLET METAL ANGLE 35.2 ° EXIT METAL ANGLE 20.6 ° 0.9 AXlALCHORDcm (in) 3.068 (1.208) FOI LS 96.
SOLIDITY (bx). .916 0.8 ELLIPSE RATIO 3/1 CHORD cm (in) 3.550 (1.398) INLET GAS ANGLE 42.2 °- EXIT GAS ANGLE 22.7 ° INLET MN .355 EXIT MN .666 0.6 INCIDENCE -7.1 ° MMAX .918 A P/Q T.E. .465
I ! I I I I
0 0.2 0.4 0.6 0.8 1.0 X/BX, AIRFOIL
Figure 4.2.4-II Turbine Third Stage Blade Mean Section Aerodynamic Defini-
tion and Pressure Distribution
e- O °_ • r- O °r..- ,L °r-" f,,,. f,.
°,-.
o (- _/) _)._ oi-= c.- 0.
o_ ._-- ._.- o C_)4-
I I I I I i
o *r.- CO QO s.
d
o (0 ..1 o
o
u_ ii
'T
o_ D2 d <_ N o
x
,g
en ED X d X o .p..
o I o
I
I I I I I I I I I I
o CN - 0 d d d d c5
6 6 6 o
±d/d XBI,A.
I- nr T-- 0 or-- C
I I I I I I I I I I I I
°p- ._ o 0 . 0 _. _..
_0 O0
" d
d
",D ..I
B
U.
cc I-
o
" d
d .
o
cr m X °_-
d x - d
-- 0
I I I I I I I I I I I
0 _ _ 0 tZ_ d d o o d d d d d 0
XQ/A
_d/d
3S
1,0
12 f
0.8 X r_ 0.6 >- 0.4 0.2 0 0.2 0.4 0.6 0.8 1.0 X/BX, AIRFOIL 1.0 RADIUS cm (in) 52.699 (20.748) INLET METAL ANGLE 38.7 ° EXIT METAL ANGLE 18.9 ° 0.9 AXIAL CHORD cm (in) 3.73 (1.47) FOILS 84.
SOLIDITY (bx) .9445 0.8 ELLIPSE RATIO 4/1 CHORD cm (in) 4.328 (1.704) INLET GAS ANGLE 43.6 ° 0.7 EXIT GAS ANGLE 21.2 ° P, INLET Mn .362 EXIT Mn .690 0.6 INCIDENCE --4.9 ° MMAX .925 Ap/Q T.E. .428 0.5
0.4 ! I ! I I I
0 0.2 0.4 0.6 0.8 1.0 X/BX, AIRFOIL
Turbine Fourth Stage Vane Mean Section Aerodynamic Defini-
Figure 4.2.4-14
tion and Pressure Distribution
.0 ¢.1 ¢,¢1 .p.- ¢1.1_ °r.= L l- "4 t,n ¢-- £= °_ ¢=.
¢- 4_ r- 4.- 4._ e-_ %.
0 .e- .._ 0
"T
LI_ XS/A ±did I-- _l_ °_ _ _-: _: ....
_._ _ _ m _ o o I o d c- _ 00_ _ _ 0 0 I C) C) o .la A 4.a o o e.-.
o L. °_,.- _4a s.. ,_
_ _-_° o
N5
_ °_.- 0 e'- ¢'-_ O _0 d d d --
d 6
°o
I- rt- c5 n,- o O
x"
m t_ °r,...
c_ X d X I O 0r,.- t a I I I I
I I I I I I
0r",- cO o0 d d _D _o -J
"T
n.- n..
d .
tw d ¢.q d × d X O
L I I I I I
L I I I I I
d d d d '- - d d d c_ 0 d XB/A J'd/d 0.2 -- 0.4 - X m 0.6 -- >- 0.8 -- 1.0 m 1.2 -
i t _ I I I
-0.2 0 0.2 0.4 0.6 0.8 1.0 X/BX, AIRFOIL 1.0 RADIUS crn(in) 53.962 (21.245) INLET METAL ANGLE 33.8 ° EXIT METAL ANGLE 22.1 ° 0.9 AXIAL CHORD cm (in) 3.37 (1.33) FOILS 100.
SOLIDITY (bx) .9954 0.8 ELLIPSE RATIO 4/1 CHORD cm_n) 3.934 (1.544J_ ' "INLET GAS ANGLE 41.5 ° 0.7 EXIT GAS ANGLE 22.2 ° INLET Mn .350 EXIT Mn .696 0.6 INCIDENCE --7.6 ° MMAX. .902 AP/O T.E. .386 0.5 i
0.4 I I I I I I
-0.2 0 0.2 0.4 0.6 0.8 1.0 X/BX, AIRFOIL
Figure 4.2.4-17
Turbine Fourth Stage Blade Mean Section Aerodynamic Defini-
tion and Pressure Distribution
4O
il
I- n,- I-- IZ: I I
I I I I I I I I I I
e,i 0 QO d d d d _ d d d d d o
xs/A "a/a
1.2 i 1.0 -- 0.8- X n_ 0.6 -- >- 0.4 - 0.2 -- 0 .
0 0.2 0.4 0.6 0.8 1.0 X/BX, AIRFOIL 1.0 RADIUS cm (in) 54.91 (21.62) INLET METAL ANGLE 42.0 ° EXIT METAL ANGLE 22.6 ° 0.9 AXIAL CHORD cm (in) 3.65 (1.44) FOILS 108.
SOLIDITY (bx) 1.148 0.8 ELLIPSE RATIO 4/1 CHORD cm (in) 4.155 (1.636)'.
INLET GAS ANGLE 45.4 ° I:_ 0.7 EXIT GAS ANGLE 22.7 °
a-
INLET Mn .350 EXIT Mn .736 m 0.6 INCIDENCE --3.4 ° MMAX .919 /%P/Q T.E. .336 0.5
0.4 I I I I I I
0 0.2 0.4 0.6 0.8 1.0 X/BX, AIRFOIL
Figure 4.2.4-20
Turbine Fifth Stage Vane Mean Section Aerodynamic Definition
and Pressure Distribution
I-- a.
I-- • r=.
l.:- i I I I I I I I I I I 0 00 ',D ,_- _ 0 0 _ CO r_ co _ ,_ o d d d _ d d o d o d XS/A ld/d I- ¢¢ i=-
I , I I I I
I I I I I
O0
d
to _J I-
d
d .
n"
d ×
d x _
o
, I I I I I
I I I I I a
0 o_ co _ ¢o _ _l- 0 t'_ _ CO _0 0 o,,I d d o d d o d o d d
XB/,_
ld/d 0 ¸ 0.2 - 0.4 -- X 0.6 - nn >- 0.8 " 1.0 "-
I I I I I I
1.2 0 0.20 0.40 0.60 0.80 1.00 X/BX AIRFOIL 1.0 RADIUS cm {in) 55.501 (21.851) INLET METAL ANGLE 40.2 ° EXIT METAL ANGLE 25.5 ° 0.9 AXIAL CHORD cm (in) 3.32 (1,31) FOILS 122.
SOLIDITY (bx) 1.161 0.8 ELLIPSE RATIO 4/1 CHORD cm (in) 3,759 (1.480) INLET GAS ANGLE 44.9 ° 0.7 EXiT GAS ANGLE 25.7 ° INLET Mn .384 EXIT Mn ,795 0.6 INCIDENCE --4.7 ° M MAX .986 _/Q T.E. ,322 0.5
0.4 I I I I I I
0 0.20 0.40 0.60 0.80 1.00 X/BX AIRFOIL
Figure 4.2.4-23
Turbine Fifth Stage Blade Mean Section Aerodynamic Defini-
tion and Pressure Distribution
o °1,- (.)
(/1 0_m,,, 0 I ,=- ,m-" _4-J 0 5- _r", c n _o i/i I-- ,,,,m o e- ¢,._ *_m °r" t_. N I I I I I I I I I I I I o o gO d d I 0,.
rr" nr I-- d .
d .
e_ X d X d or,- o o m m m I I I I I I I t 0 _ _0 r.. _D _ ,_t 0 IN _ _ID gO 0 d o o d d o d o d d _ ×B/k J'dld ROOT TIP
Figure 4.2.4-25 Second Stage Turbine Vane Stacking Arrangement
ROOT Ti
Figure 4.2.4-26 Second Stage Turbine Blade Stacking Arrangement
TIP ROOT
Figure 4.2.4-27 Third Stage Turbine Vane Stacking Arrangement
ROOT TIP
Figure 4.2.4-28 Third Stage Turbine Blade Stacking Arrangement
TIP ROOT
Figure 4.2.4-29 Fourth Stage Turbine Vane Stacking Arrangement
ROOT TIP
Figure 4.2.4-30 Fourth Stage Turbine Blade Stacking Arrangement
TIP ROOT
Figure 4.2.4-31 Fifth Stage Turbine Vane Stacking Arrangement
ROOT TIP
Figure 4.2.4-32 Fifth Stage Turbine Blade Stacking Arrangement
The turbine exit guide vane is designed to present the exhaust mixer with a
low Mach number and zero swirl flow field. Figure 4.2.4-33 shows the aero-
dynamic definition of the exit guide vane, and Table 4.2.4-III presents a
summary of the design characteristics at the aerodynamic design point.
16.1378 cm (6.3535 in) -0.9748 cm (-0.3838 in) '-'=='' 66 68.6892 (27.043 in)R (26) T" 66.497 cm 4 J (26.180 in)R J J (24) -- I A f C J 3 I E I J c} I -: 55 f _1221 -- I ,<
J
J
(20) --
i--
49.4436 cm (19.4660 in)R (18) -- 16.1378 (6.3535 in)
J
45.2704 cm (17,8230 in)R 40 10.0) I 0.0 (16) I I I I I -5.0 0 5.0 10.1 15.2 20.3 (-2) (2.0) (4.0) (6.0) (S.O) AXIAL DISTANCE, cm (in)
Figure 4.2.4-33
Turbine Exit Guide Vane Flowpath for the Integrated Core/Low
Spool
TABLE 4.2.4-III EXIT GUIDE VANE AERODYNAMIC CHARACTERISTICS (Planar Sections) Integrated Core/Low Spool - Aerodynamic Design Point I0,668 m (35,000 ft) 0.8 Mn 1 2 3 4 5 Radius LE, cm (in) 45.270 (17.823) 50.505 (19.884) 55.740 (21.945) 60.972 (24.005) 66.207 (26.066) Radius TE, cm(in) 49.443 (19.466) 54.254 (21.360) 59.067 (23.255) 63.878 (25.149) 68.689 (27.043) Inlet Gas Angle, deg 55.67 47.6 49.5 54.4 66.8 Exit Gas Angle, deg 90.0 90.0 90.0 90.0 90.0 Inlet M N 0.516 0.503 0.452 0.410 0.412 Exit MN 0.395 0.372 0.350 0.366 0.386 Incidence, deg -5.5 -4.5 -4.5 -4.0 -3.5 Deviation, deg 0 3.3 5.9 6.9 5.8 Gap/chord 0.598 0.644 0.701 0.762 0.826 Diffusion Factor* 0.33 0.455 0.465 0.390 0.250 REBact @ inlet 2.86 X lO 5 3.0 X lO 5 2.88 X IO5 2.65 X lO5 2.Sl X lO5 Chord, cm (in) 17.310 (6.815) 17.650 (6.949) 17.642 (6.946) 17.559 (6.913) 17.421 (6.859) *"D" Factor = 1 - C2 + Z_Cu . /bact
-CT-
To establish the definition of the exit guide vane, a controlled diffusion
design approachwas used to ensure an attached boundary layer and attain the
desired gas exit angle. This approach was successfully verified during cascade
testing conducted under a Navy-sponsoredprogram (Contract N00019-77-C-0546).
Table 4.2.4-IV showsthe similarity in design parameters between the Energy
Efficient Engine and Navy exit guide vane configurations. Figures 4.2.4-34 and
-35 present test results acquired under this program and showthe excellent
agreement betweenpredicted and measuredpressure distribution and gas exit
angle. It is also noteworthy to point out the large incidence range that
exists at the inlet Machnumberlevels (0.4 to 0.6) of the Energy Efficient
Engine.
The final aerodynamic definition of the exit guide vane is shown in Figures
4.2.4-36 through -40. These figures showairfoil contours and pressure distri-
butions for the five sections identified in Table 4.2.4-III. The stacking
arrangement of the exit guide vane is presented in Figure 4.2.4-41.
TABLE 4.2.4-IV
CO_ARISON OF NAVY FOIL TO ENERGY EFFICIENTENGINE
Energy Efficient Engine
Navy Radius TE = 54.254 (21.360 in)
(Design Point) (Design Point)
Inlet Mach 0.763 0.503
Exit Mach 0.529 0.372
Inlet Gas Angle 46.8" 47.6 °
Exit Gas Angle 90.0 ° 90.0 °
Gap Chord Ratio 0.70 0.62
Gas Tunning 43.2 ° 42.4 °
L_ ; Z Z Z _ ']"IONV INV] EI.I.SN_OC]
_' '9NINUN.L
DESIGN CONDITIONS 1,4 I M 1 = 0.763 M 2 = 0.529 /31 = 46.8 DEGREES /32 = 90.0 DEGREES /_ TURNING = 43.2 DEGREES 1.2 1.0 n- LU rn D 0.8 Z "1- <
DO
U,I 0.6 O , < U,, ri- D 0.4 SOLID: DESIGN 0.2 SYMBOL: TEST DATA FOR POINT 72
I I
I I I
0.6 0.8 1.0 0.0 0.2 0.4
X/A'XlAL CHORD
Turbine Exit Guide Vane Cascade Performance Test Results
Figure 4.2.4-35
Under a Navy-Sponsored Program (Contract N00019-77-C-0546)
0.2 0.4 X m 0.6 m >- 0.8-- 1.0 --
I I I I I
1.2 0 0.2 0.4 0.6 0.8 1.0 X/BX, AIRFOIL 1.2 1.0 0.8
(5
,_P/Q TE = 0.61 Z "r (J < 0.6_ 0.4 0'2 t
I I I I I
0 0.2 0.4 0.6 0.8 1.0 X/BX, AIRFOIL
Figure 4.2.4-36 Turbine Exit Guide Vane Root Section Aerodynamic Definition
and Pressure Distribution
SS
0.2 0.4 X ¢n 0.6 >- 0.8 m 1.0 --
I I I I t
1.2 k 0 0.2 0.4 0.6 0.8 1.0 X/BX, AIRFOIL 1.20 -- 1.00 AP/QTE = 0.70 0.80
d
Z 0.60_
0 7 "
0.20/0_ I I I I I
0 0.2 0.4 0.6 0.8 1.0 X/BX, AIRFOIL
Figure 4.2.4-37
Turbine Exit Guide Vane One-Quarter Section Aerodynamic
Definition and Pressure Distribution
0.2 0.4 X m 0.6 >- 0.8 n 1.0m
1.2!L I I I I I
0.2 0.4 0.6 0.8 1.0
X/BX, AIRFOIL
4P/Q TE = 0.63 OZ 0.6 0.4 0.2
I I I I I
0 _ 0.2 0.4 0.6 0.8 1.0
X/BX, AIRFOIL
Figure 4.2.4-38
Turbine Exit Guide Vane Mean Section Aerodynamic Definition
and Pressure Distribution
0.2 0.4 X n-_ 0.6 >- 0.8 m 1.0 --
I I I I I
1.2- 0.2 0.4 0.6 0.8 1.0
X/BX, AIRFOIL
m 0.6 0.5 0.4 "I- 0.3 m (,J ,&P/QTE = 0.54 < 0.2 0.1 m
I I I I I
0 0.2 0.4 0.6 0.8 1.0
X/BX, AIRFOIL
Figure 4.2.4-39
Turbine Exit Guide Vane Three-Quarter Section Aerodynamic
Definition and Pressure Distribution
X rm 0.4 >- 0.6 m
I I I I I
0.8 0.2 0.4 0.6 0.8 1.0 X/BX, AIRFOIL 0.5' 0.4 zO 0.3 "1" (.3 < 0.2 0.1 m
I I I I I
0.2 0.4 0.6 0.8 1.0
X/BX, AIRFOIL
Fi gure 4.2.4-40 Turbine Exit Guide Vane Tip Section Aerodynamic Definition
and Pressure Distribution
TIP ROOT
Turbine Exit Guide Vane Stacking Arrangement
Figure 4.2.4-41
4.3 SUPPORTING TECHNOLOGY PROGRAMS
4.3.1 Introduction
Three supporting technology programs were conducted during the design process
to assess critical aspects of the turbine intermediate case and low-pressure
turbine designs. Results acquired from these efforts provided technical guid-
ance and insight to ensure a viable aerodynamic design. A Transition Duct
Model Test Program was completed to verify duct performance. Also, a Low-
Pressure Turbine Subsonic Cascade Technology Program demonstrated that the
loss and incidence range of the low camber second-stage vane are reasonable as
well as verifying the benefits of an aft-loaded pressure distribution. The
Low-Pressure Turbine Boundary Layer Program corroborated that an aft-loaded
pressure distribution was reasonable for low-loss airfoils.
The significant results from these programs are summarized in the following
sections. Complete documentation of results is contained in the following NASA
technical reports: Energy Efficient Engine Transition Duct Model Test Program
(pending publication), Energy Efficient Engine Low-Pressure Turbine Boundary
Layer Program (CR-165338) and Energy Efficient Engine Low-Pressure Turbine
Subsonic Cascade Technology Program (CR-165592).
4.3.2 Transition Duct Model Test Program
This technology program was structured to verify the aerodynamic design of the
turbine intermediate case. The effort consisted of a two phase test, using a
model of the intermediate case, which included the fairing and low-pressure
turbine i nlet gui de vane.
In the first phase, the intermediate case, designed on a preliminary basis for
a 182,376 N (41,000 Ib) thrust engine, was evaluated. At design flow condi-
tions, the total pressure loss for the flight propulsion system transition
duct design was 0.7 percent APT/PT, which was consistent with the goal.
The pressure loss increased to 1.3 percent APT/P T at off-design swirl
conditions up to 5 degrees. Also, the pressure coefficient along the outer
duct wall indicated that the desired diffusion was attained across the strut.
Strut airfoil pressure distribution data verified that the design conditions
were achieved with no flow separation.
Inlet, strut exit and low-pressure turbine inlet guide vane exit air angle
data were in good agreement with the design prediction. With a 5 degree off-
design inlet swirl angle, the struts returned the unturned flow to within one
degree of the design point swirl.
In the second phase of the program, several modifications were made to the
test model to reflect the intermediate case design for the integrated core/low
spool. The changes consisted of: (1) an increased area ratio from 1.50 to
1.57, (2) a reduction in the number of structural support struts from 14 to
II, (3) tangential canting of the strut, and (4) revising the fairing from a
65 circular arc series nonworking airfoil to a 400 series working airfoil with
a flow turning capability of 5 degrees.
In Figures 4.3.2-I through 3, mass-averagedair angle spanwise data from the
design and off-design test are comparedto the design prediction. Data are
presented for the rig inlet, strut exit and turbine inlet guide vane exit,
respectively. Rig inlet data in Figure 4.3.2-I showthat the measuredoff-
design swirl was approximately 6.7 degrees less than the measureddesign point
inlet swirl. This impact on the exit swirl angle is shownin Figure 4.3.2-2,
in which the measuredstrut exit swirl is approximately 2.8 degrees less than
that represented by design point data.
Figure 4.3.2-3 indicates that the measuredinlet guide vane exit swirl at the
rig off-design inlet conditions was within 0.75 degree of design point data.
This suggests that the high-pressure turbine off-design operation will have
little effect on the angle of the flow entering the low-pressure turbine
first-stage rotor.
Profiles of outer and inner wall loadings are presented in Figure 4.3.2-4 and
5, respectively. Data trends are comparedto the design predictions, and
results showa reduction in diffusion through the strut because of the change
in inlet angle. The flow then accelerates through the low-pressure turbine
inlet guide vane. These loading results indicate that the flow is separation
free.
u_ LU LU mr -- DESIGN PREDICTION L9 W DESIGN POINT DATA m OFF-DESIGN DATA Z < rr U.I u.I > < <
. I I I I I I I I I I
10 20 30 40 50 60 70 80 90 100 PERCENT SPAN
Figure 4.3.2-I Rig Inlet Mass-Averaged Air Angle Spanwise Data
6]
50 J DESIGN PREDICTION UJ LU 48 -- '--C)'-- DESIGN POINT DATA nr OFF-DESIGN DATA LU r_ 46--
w"
_J 44-- Z < 42--
<
tn w (3 40 ,< it,,.
u.i > oo < 34 -
I I I I I I I I I I
10 20 30 40 50 60 70 80 90 100
PERCENT SPAN
Figure 4.3.2-2 Strut Exit Mass-Averaged Air Angle Spanwise Data
30 In -- DESIGN PREDICTION OFF-DESIGN DATA UJ V ('/) L "'O"'- DESIGN POINT DATA (,.9 28
_ 26
<
@
oO 20 oO <
:E
I I I I I I I I I I
0 10 20 30 40 50 60 70 80 90 100
PERCENT SPAN
Figure 4.3.2-3 Low-Pressure Turbine Inlet Guide Vane Exit Mass-Averaged Air
• Angle Spanwise Data
q _ PREDICTION 0.6 -- ----O---DESIGN POINT DATA o.
--O--OFF-DESIGN D_ a. Q 0.4 -- / I
>
0.2 II r_ o o.
Z uJ _J -0.2 m u.
u.
uJ w - 0.4 -- ffl I I I 1 STRUT STRUT IGV tGV n --0,6 I LE[ I TE LE I TE I -5.0 O 5.0 10.1 15.2 20.3 ( - 2.0) (0) (2.0) (4.0) (6.0) (8.0) AXIAL DISTANCE, cm (in) Figure 4.3.2-4 Outer Wall Loading Profile ------ PREDICTION DESIGN POINT DATA ---O--OFF-DESIGN DATA I I[ _IGV STRUT STRUT IGV TE LE TE LE ) I I I 20.3 0.0 5.0 10.1 15.2 (8.0) (0.0) (2.0) (4.0) (6.0) AXIAL DISTANCE, cm (in) Figure 4.3.2-5 Inner Wall Loading Profile
4.3.3 Low-Pressure Turbine Subsonic Cascade Technology Program
The use of counterrotating turbines and the requirement for highly loaded,
"low loss" airfoils results in airfoil geometries that are a substantial
departure from conventional designs. The objective of the Subsonic Cascade
Technology Program was to investigate the performance characteristics of a low
camber inlet guide vane and to evaluate the fore and aft loading type loadings
used to optimize the low-pressure turbine vane and blade aerodynamic defini-
tions. This was accomplished through three airfoil cascade tests conducted at
varying pressure ratios and inlet incidence angles.
To assess the performance of the low camber inlet guide vane, pressure loss
data were generated for both design and off-design conditions in the first
test series. Secondary loss data were acquired at the design incidence angle,
and were found to be in good agreement with the Pratt & Whitney Aircraft
cascade correlation. The secondary losses were 0.56 percent of the inlet total
pressure and the prediction was 0.59 percent of the inlet total pressure. Pro-
file losses, measured as a function of incidence angle, were found to be in
good agreement with the Pratt & Whitney Aircraft boundary layer prediction
method, as shown in Figure 4.3.3-I.
1.0 m 0•_ m _ PREDICTIONS /_ 0.6 LOSS (_.p.__T APT "_ %
)
0.4 B 0.2 -- DESIGN INCIDENCE
I
0.0 I I I I
-15 -'lO -5 0 5 10 INCIDENCE ANGLE (DEGREES)
Figure 4.3.3-I Loss Versus Incidence for the Low Camber Vane
The results presented in Figure 4.3.3-I show that losses increase more rapidly
for the negative incidence side of the loss bucket compared to the positive
side. Consequently, the shape of the low camber airfoil loss bucket is differ-
ent from that of a typical turbine airfoil. The measured off-design incidence
performance at a constant exit Mach number showed the vane section to have a
negative incidence range of 12.4 degrees and a positive incidence range of 7
degrees, as defined by the point where the loss level is 50 percent above the
design loss point. Results showed that an improved incidence range could be
achieved by revising the vane incidence angle from -5 to -3 degrees.
In the second part of the program, tests were completed with two airfoil cas-
cades to evaluate the aerodynamic performance difference between an airfoil
that accelerated the flow on the suction surface near the gaging point (aft
loaded) and an airfoil that accelerated the flow more rapidly in the forward
region of the suction surface (squared-off).
Most of the airfoil profile losses are generated on the suction surface be-
cause the average velocity on this surface is high. Furthermore, the flow on
this surface of the airfoil accelerates from the leading edge to some high
value and then diffuses to the trailing edge velocity. Two factors affect the
suction side loss, the nature of the boundary layer, i.e. laminar, transition-
al, or turbulent in the accelerating portion of flow and the maintenance of
attached flow in the diffusion region. The latter is characterized by the
diffusion parameter Ap/Q, the ratio of exit static pressure minus static
pressure at the maximum velocity point to the maximum dynamic head.
For a given Zweifel load coefficient, aft-loaded airfoils generally have a
higher diffusion parameter, maximum Mach number and thickness to chord ratio
than the "squared-off" airfoil. In addition, the squared-off concept offers a
potential for lower weight. Figure 4.3.3-2 shows profiles and pressure distri-
butions of two airfoils that have the same load coefficient and exit Mach
number.
Test results are compared to theoretical predictions in Figure 4.3.3-3. As
shown, the performance of the aft-loaded airfoil is better than or equal to
the alternate configuration over the range of incidence angles tested. Similar
results were obtained with variations in Mach number (Figure 4.3.3-4). Theor-
etical predictions were found to be in good agreement with measured data over
the entire range of Mach number and incidence angles. Secondary loss measure-
ments showed that the aft-loaded airfoil generated approximately II percent
lower losses than the squared-off airfoil at the design incidence and Mach
number conditions.
In a third series of tests, three airfoil sections, representing the mean
section of the fourth-stage blade, were evaluated to determine surface static
pressure distribution and profile loss characteristics for a range of inlet
gas angles and Mach numbers. One airfoil section with an aft-loaded pressure
distribution was used as a baseline. The other two, which had a squared-off
type pressure distribution and different airfoil thickness distributions, were
compared to the baseline. One squared-off airfoil was distinguished as a heavy-
weight configuration because of its thick sections, and was designed with the
same leading and trailing edge wedge angles as the aft-loaded airfoil. The
other configuration, referred to as a lightweight airfoil, had almost the same
static pressure distribution as the heavyweight airfoil. However, it was
designed for low inlet and exit wedge angles to enable a thin foil geometry.
d or,- A o q..
Q s- _J d _ o .X
, r//-/r
_ X ..
d q- I (_1 °r-" o m ,lm '_4.a ,..o °_-- • I o q _J o u'l c o bE d _o _J .r- I q.- <n.
o< --J3
.$
_-0 0 _ <: u_
I
Z
<
I I
I o ¢j
/'
J
X o x
/
0I'D I.
0 _ °' 0 0 c I,.- DATA I PREDICTIONS SQUARED-OFF-SUBSONIC _] AFT-LOADED TRANSONIC (_ F-
T
Q.
0.5 Q.
<3 POSITIVE DESIGN NEGATIVE INCIDENCE INCIDENCE --= INCIDENCE 39.5 44.5 49.5 54.5 59.5 14.5
INLET GAS ANGLE, DEGREES
Figure 4.3.3-3 Aft-Loaded Transonic Versus Squared-Off Subsonic Airfoils
Predicted Versus Measured Profile Loss Data
1.5 J
i/.]__
J 1.0 0.
a.
0.5 DATA PREDICTIONS SUBSONIC SQUARED-OFF TRANSONIC AFT-LOADED 0.7 0.8 0.9 1.0
EXIT MACH NUMBER
Figure 4.3.3-4 Aft-Loaded Transonic Versus Squared-Off Subsonic Airfoils
Predicted Versus Mass-Averaged, Mixed-Out, Mid-Span Losses.
6?
In general, surface static pressure distributions and profile losses were in
good agreement with Pratt & Whitney Aircraft prediction methods over the
entire range of Mach numbers and inlet gas angles. Cascade results showed the
aft-loaded airfoil produced lower profile losses than the other designs. Per-
formance trends from data and theoretical predictions are presented in Figures
4.3.3-5 and 4.3.3-6. Compared to the aft-loaded airfoil at the design point,
the heavyweight and lightweight airfoils had 34 and 21.3 percent higher
profile losses, respectively. The data indicate that the profile loss for the
aft loaded airfoil is lower than either the heavyweight or lightweight airfoil
for positive incidence, whereas all airfoil sections show almost equal losses
for negative inci dence.
A measurement of secondary loss was also obtained for the aft-loaded and light-
weight airfoils at three separate incidence angles. The aft-loaded airfoil
demonstrated lower secondary losses for the entire range of inlet gas angles
and Mach numbers.
In summary, the cascade program verified that the loss and incidence range of
the low camber inlet guide vane was reasonable. Also, the benefits of an aft-
loaded pressure distribution for low loss airfoils were verified.
4.3.4 Low-Pressure Turbine Boundary Layer Program
The Low-Pressure Turbine Boundary Layer Program was aimed at investigating the
probability of further improving turbine efficiency by reducing airfoil pro-
file losses. The pressure distributions of aft-loaded and squared-off airfoils
were assessed in terms of losses resulting from boundary layer development on
the airfoil suction surface. Velocity distributions were simulated on flat
plates in a low speed, high aspect ratio wind tunnel designed specifically for
these investigations. Measurements of boundary layer mean velocity and turbu-
lence intensity were acquired for_an inlet turbulence level of 2.4 percent and
an exit Reynolds number of 8 x lO _. Flush-mounted hot film probes were used
to identify boundary layer transition regimes that were found to be located in
the adverse pressure gradient regions for both test simulations.
Mean Velocity Profiles
A total of twenty mean velocity profiles was measured, ten for each of the two
pressure distributions. Nine of the profiles were located in the laminar flow
region, four in the transitional flow region, and the remaining seven in the
fully turbulent flow region. A comparison of mean velocity profile data in the
transitional and turbulent regimes with well established semi-empirical formu-
lations is as follows.
All turbulent boundary layer data have a universal region where Equation l is
val id.
U+ = l In y+ + B (Eq. l)
k
The fully turbulent boundary layer mean velocity profile data from the two
configurations are presented in Figure 4.3.4-I, using the dimensionless para-
meters of Equation I.
I DATA PREDICTION AFT LOADED { 1.5 _- HEAVYWEIGHT E_ LIGHTWEIGHT { ------.!-%....
....2 • 0.5 POSITIVE DESIGN INCIDENCE -',,,Q_ POINT _- NEGATIVE 39.5 44.5 49.5 54.5 59.5 64.5
INLET GAS ANGLE, DEGREES
4.3.3-5
Figure
Aft-Loaded, Heavyweight, Lightweight-- Predicted Versus
Profile Loss Over Range of Incidence, Fixed Exit Mach Number
(0.72)
1.5
f-
I
f
I
DESIGN I
J
I
I
1.0
a.
DATA PREDICTIONS
I
0.5
AFT LOADED . {
I
I HEAVYWEIGHT
I
LIGHTWEIGHT { SOLID SYMBOLS INDICATE LOSS DATA OBTAINED WITH THE TURBULENCE GRID INSTALLED AT INLET TO AIR TEST SECTION
I ! I
0.6 0.7 0.8 0.9
1.0
MACH NUMBER
4.3.3-6
Figure
Aft-Loaded, Heavyweight, Lightweight -- Predicted Versus
Measured Mass-Averaged Mid-Span Profile
Loss at Design Point
Incidence, Variable Exit Mach Numbers
20 -- _= #__SQ-OF F $ _0.836 15 -- ""_1_'_" -- .,._[_,,_ SQ-O F F 15 _ '-- ... _0_ l_-_s =0.925 5 _ 0 -- _Q-OF F ' _._ s =0.93 d 1150 _ f S _ SURFACE DISTANCE FRO M 5 _ LEADING EDGE IN METERS (M) 0 _ _.__FT-LOAD E D , _ s =0.721 FT-LOADED
o__ ,_o_o
5 0 . A_FT'LOADED 0. _Z_"-"
1ol-
:F T..OAO O
_l °°'" I I
1 10 1 O0 1000 y+
Figure 4.3.4-I
Turbulent Boundary Layer Velocity Profiles Using Dimension-
less Parameters
?0
Squared-Off Pressure Distribution
Experimental data for the integral parameters of boundary layer momentum loss
thickness Reynolds number(ReO), shape factor (H) and skin friction (Cf),
from the squared-off design are shownin Figure 4.3.4-2, along with analyti-
cally predicted values. In general, the predictions are in good agreementwith
test data. Figure 4.3.4-3 showsdetailed meanvelocity profile data compared
to the predictions. The predictions are in good agreement with laminar and
turbulent velocity profiles, but in somewhat poor agreementwith transitional
profile data.
Aft-l oaded Pressure Di stri buti on
A comparison of test data and predictions for the aft-loaded test is presented
in Figure 4.3.4-4. The agreement is good for the accelerating part of the
flow, but someflow separation in the diffusing part is indicated. The calcu-
lations were repeated and the boundary layer was artifically madetransitional
at a distance of two boundary layer thicknesses upstream of the expected sepa-
ration point to obtain theoretical predictions. Although it is difficult to
ascertain whether the boundary layer actually separated at the location that
was predicted, the calculated separation point was slightly upstream of the
transition region identified by the hot film probes. If separation occurred,
however, it did not influence the behavior of the intermittency factor in the
transitional region. Velocity profile data and predictions are presented in
Figure 4.3.4-5. Again, measuredand predicted profiles are in good agreement
for the laminar and turbulent regions.
Profile Loss Assessment
Relative magnitudes of profile losses associated with aft-loaded and squared-
off pressure distributions can be achieved by a direct comparison of the
integral parameters obtained from test data. Distribution of the three inte-
gral parameters, boundary layer momentum thickness (e), shape factor (H) and
skin friction (Cf), is shownin Figure 4.3.4-6.
Momentum Loss Thickness Results indicated that the momentum loss thickness
for the aft-loaded airfoil section was 8 percent larger than for the squared-
off section, as shown in Figure 4.3.4-6(a). If the exit velocity for the two
airfoil simulations was identical, a squared-off airfoil of the same design as
the aft-loaded foil would have an 8 percent lower loss on the suction side.
However, the exit velocity for the squared-off airfoil was about 4 percent
higher than than the aft-loaded airfoil. This implies that the loss (U2 e)
for the two tests would be almost identical.
Shape Factor The shape factor associated with the aft-loaded test increased
at a location before the transition region, thereby indicating a possibility
of separation before transition. The distribution of boundary layer shape
factor is shown in Figure 4.3.4-6(b). The distributions of each design are
similar in the laminar (H = 2.3) and the turbulent (H = 1.4) regions.
Skin Friction The distribution of laminar and turbulent skin friction showed
similar behavior. The aft-loaded skin friction rate of decrease with stream
distance was high, indicating that laminar separation may have occurred.
?!
/_ DATA 3.0 m
O
PREDICTIONS
[]
[] [] []
2.0 0.006 1.0 0.005
A
0.004 0.003 0.002 0.001
1,oo z
O
t
'_ 800 20£ TRANSITION REGION (MEASUREMENT)
I
I I I I T'////z'////f//Vl I
1.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 STREAMWISE DISTANCE FROM THE LEADING EDGE (M)
Fi gure 4.3.4-2
Comparison of Measured Integral Parameters for the Squared-
Off Test Section With Theoretical Predictions
?2
Q w,.: CO !
c_ _0 _O Z c_ © F- ,,_ C_ c_ O O O.
o1..- O e- O _" O 0r-- U co O C_ LU OJ_" L_ o_..
m.l-_ O a o Z o c- F-- c- o e- W O • i..- .r.- © n" LL LU OI4" (--)O Z O I O In n- Il O or.-
I I L
Q o_ _O O
...,-.- 6/x
3.0 DATA
O
t
OREDICTIONS n-
[3
[] []
V!
P 2.0
u u __
0.O_ B 1.o I 0.005 m 0.004
t
0.003 0.002 0.001
Z
14ooL _ 1000 m
t
¢r 600 --
• O O
400 m GION __REDICTION) I (MEASUREMENT)
I
t t I I I T '_'ZT_ I t t
1.0 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 STREAMWISE DISTANCE FROM THE LEADING EDGE (M)
Figure 4.3.4-4
Comparison of Measured Integral Parameters for the
Aft-Loaded Test Section With Theoretical Predictions
°t- O e- _ 0 ¢'_ °F,- 0r-- "_ 0 S- r-,. a.
e.- U #_I .r- S,,.
S,,. e- _ .,1-) _'.r- 4-- L._.I I d II X I S,.
• r- I.L B 1.4 (A) i 1.2
O
1.0
0[]
i 0.8
E
v m
• o.6
0.006 -- (C) 0.4 m 0.2 _ 0.005
-
O
0 0.004 0 0.2 0.4 0.6 0.8 1.0
O
- O
X b 0.003
O o
0.002
- C! O
0.001 2.6 ¸ (B)
O
I I i I I
n 2.4 0.2 0.4 0.6 0.8 1.0
O
O
X m 2.2 2.0
O
3:
O O AFT-LOADED
1.8 O SQUARED-OFF 1.6 1.4 I I I I I 1.2 0 0.2 0.4 0.6 0.8 1.0 X
Figure 4.3.4-6
Distribution of Boundary Layer Momentum Loss Thickness (e),
Shape Factor (H), and Skin Friction (Cf)
?6
Boundary Layer Turbulence Intensity Profiles
Boundary layer turbulence intensity profiles for the turbulent, transitional
and laminar flow regions are described in the following paragraphs.
Turbulent Region Turbulent2int_nsityRrofiles were obtained from the three
components of tu'rbulence (u , v and w_) measured at the exit plane of
the two airfoil designs. Test data are in good agreement with the flat plate
da_a, as indicated in Figure 4.3.4-7. Relative magnitudes of the streamwise
(uL) and normal (v2) components are shown for the squared-off configura-
tion in Figure 4.3.4-8. The streamwise and normal components contain about 50
and 20 percent, respectively, of the total turbulence intensity. These results
are consistent with the Pratt & Whitney Aircraft prediction system.
Transitional Region Relative magnitudes of turbulence intensity components
for the transitional region of the squared-off design are shown in Figure
4.3.4-9. Data show that the streamwise and normal components contain approxi-
mately 80 and lO percent, respectively, of the total turbulence intensity.
This means that the turbulence in transitional boundary layers is more non-
isotropic than in fully turbulent boundary layers.
Laminar Re_ion Systematic growth of the streamwise component of turbulence
intensity was observed in the laminar region of both airfoil designs. Dimen-
sionless turbulent intensity (u+) data in the laminar boundary region for both
airfoil designs were presented as functions of y+ in Figure 4.3.4-I0. The data
suggest two important features. First, turbulence intensity profiles in the
laminar region had a maximum value at approximately y+ = 25. Second, maximum
turbulence intensity (u+ aty+ = 25) increased in the downstream direction as
the onset of transition was approached.
Subsonic Cascade Test Results Versus Boundary Layer Test Results
Results from the boundary layer tests indicate that the Reynolds number based
on momentum loss thickness for the two simulated airfoils was about the same
at the exit plane. This suggests that the suction surfaces of the two airfoils
generate about the same losses. However, loss data acquired from the Subsonic
Cascade Program indicate that the heavyweight airfoil generates about 17 per-
cent higher losses than the aft-loaded airfoil. This apparent paradox can be
explained by examining the pressure distributions on the suction surfaces of
the two airfoils and the simulated pressure distributions.
The Boundary Layer Program did not simulate the entire airfoil suction sur-
face, omitting the leading edge region. It is in this region of the airfoil
that the potential flow analysis indicates a difference between the two air-
foils with the squared off pressure distribution predicting a larger velocity
defect. This may be the explanation for the relative loss difference measured
in the cascade. This theory tends to be supported by the off-design data which
show that the loss difference between the two airfoils decreases as they
operate at higher inlet angles, as previously indicated by Figure 4.3.3-4. At
the higher negative incidence, the leading edge velocity distributions become
nearly identical.
7?
1.0 &
O.8
k TM
0 _ • 'SOuAR_O-OF_"
_k 0.6
_ O@
Y O.4 0..2 1.0 2.0 3.0 4.0 .5.0 6.0 q--_/2 v,2
Figure 4.3.4-7 Comparison of Measured Total Turbulence Intensity Profiles
with Flat Plate Data
OJl 0 u2/q2
I-I v21q2
0.7
A w_/q2
Z ud Z 0.8 n_
0 0 0 0 0 0 0 0 0
_ O_
o o
Z
_z
uJ 0.4 U Z Lu .J m O: O.3
z_ zx _ A A _ A _ _ 13
w N
[] 0
I
0 0 0 0 0 0 0 0 0
Z 0.1
I I I I I
0.2 0.4 0.8 (3.1 1,0 Y/6
Figure 4.3.4-8 Distribution of Normalized Turbulence Intensity Components
in the Fully Turbulent Region of the Squared-Off Test
Confi gu ration
?8
_8
_0 O0 0
0 u2/q2
0 v2/q_
0.7 0.6
|
)- I- 0.5 Z u.i I,- Z g,I 0.4 Z IM ..I m 0.3 I.-
&
N mm ..I '_ 0.2
I
m o z
o
0.1 _o ooo .
I I I I !
0 0.2 0.4 0.6 0.8 I .O
Y/6
Figure 4.3.4-9
Distribution of Normalized Turbulence Intensity Components
in the Transitional Region of the Squared-Off Test Configu-
ration
?9
U ILl r- e'- ÷ 1.- -- >.
e-" I ....J ..¢: ..I LO Cn 0 d _.
e" it d " '1 II O_ O_
I I I I I I I I 1 I I
e- e-- 4- U e.-
,-g
o_
3c S. 0 !
o
,4
S,.
I 1 I I I I I I I I I
c_ o o c; 4-
From the above discussion, it could be concluded that when the leading edge
velocity defect is accounted for, the cascade test and large scale boundary
test results are consistent. Further studies directed towards understanding
the influence of airfoil leading edge velocity profile on the performance of
turbine cascades are required before some definite conclusions can be formu-
lated with respect to its application in turbine designs.
In summary, the results of the Low-Pressure Turbine Boundary Layer Program
corroborated that an aft-loaded pressure distribution was reasonable for low-
loss airfoils in the low-pressure turbine of the Energy Efficient Engine.
4.4 PERFORMANCE STATUS AND ADJUSTMENT
The performance history of the low-pressure turbine component, intermediate
case, and exit guide vane can be summarized in three categories: l • Goal efficiency established early in the program.
2.
Efficiency predictions made prior to completion of the mechanical
design.
.
Efficiency prediction update made after completion of the mechanical
design.
The efficiencies in the first two categories for the low-pressure turbine were
made with assumed blade tip and air seal clearances, while that in the third
category reflects the clearances resulting from design analysis.
The goal efficiencies established early in the program and presented at the
Low-Pressure Turbine Preliminary Design Review are as shown in Table 4.5-I.
When work progressed to the point where aerodynamic studies were nearly com-
plete but the mechanical design was still in process, a review of predicted
performance showed that component efficiency had the potential to exceed the
goal in Table 4.5-I. The improvement was attributed to the aerodynamics of the
airfoil contours. The updated performance levels are summarized in Table
4.5-II.
On completion of the mechanical design the performance was again updated. The
mechanical design that resulted had an impact on performance because of final
blade tip clearances. Rotor deflection studies showed that the goal 0.050 cm
(0.020 in) blade tip clearances would not be sufficient for aircraft maneuvers.
In addition, it was found that greater thermal closure of the active clearance
control system (reduced blade tip clearance) was possible with eighth stage
high-pressure compressor bleed air than with originally selected tenth stage
air. Since considerable integrated core/low spool external part design work
had been completed with the tenth stage system, it was retained for the inte-
grated core/low spool design. To take advantage of the improved efficiency
with reduced tip clearance, the flight propulsion system adopted the eighth
stage high-pressure compressor bleed air system, partially off-setting the
efficiency penalty resulting from increased tip clearance required for
maneuver deflections.
TABLE 4.5-I GOAL PERFORMANCE FOR LOW-PRESSURE TURBINE (Intermediate Case and Exit Guide Vane at Aerodynamic Design Point) Flight Propulsion Integrated Core/ System Low Spool Low-Pressure Turbine Component Base Efficiency, % 89.0 89.0 Energy Efficient Engine Changes Improved Aerodynamics +I.0 +0.5 +0.3 +0.3 Increased Velocity Ratio +I. 2 +I.0 Reduced Leakage Integrated Core/Low Spool Effects -0.2 Off-Design Operation Part Quality -0.2 Increased Clearances -0.3 and Leakage 91.5 90.l Low-Pressure Turbine Total, % Intermediate Case Pressure Loss,APT/PT, % 0.7 1.5 Exit Guide Vane Pressure Loss,APT/PT, % 0.9 0.9 Assumptions: (1) 0.051 cm (0.020 in) clearance for interstage air seals (2) 0.051 cm (0.020 in) clearance for outer blade tip air seals (3) 0.152 cm (0.060 in) clearance on flow guides (4) No loss included for gaspath instrumentation TABLE 4.5-II LOW-PRESSURETURBINE EFFICIENCY UPDATE (At Aerodynamic Design Point -- Prior to Completion of Mechanical Design) Flight Propulsion Integrated Core/ System Low Spool Low-Pressure Turbine Component 89.0 89.0 Base Efficiency, % Energy Efficient Engine Changes +1.3 +1.3 Improved Aerodynamics Increased Velocity Ratio +0.3 +0.3 Reduced Clearance +0.7 +0.7 and Leakage Flow Guides +0.5 +0.2 Integrated Core/Low Spool Effects -0.I Off-Design Operation Part Quality -0.I Increased Clearances and -0.3 Leakage 91.8 91.0 Low-Pressure Turbine Total, % Intermediate Case Pressure 0.7 1.5 Loss,APT/P T, % Exit Guide Vane Pressure 0.9 0.9 Loss,Z_PT/P T, % Assumptions: (I) 0.051 cm (0.020 in) clearance for interstage air seals (2) 0.051 cm (0.020 in) clearance for outer blade tip air seals (3) 0.152 cm (0.060 in) clearance on flow guides (4) No loss included for gaspath instrumentation
The resulting status efficiencies are _1.3 and 90.3 percent for the flight
propulsion system and integrated core/low spool, respectively, as shown in
Table 4.5-III. Reasons for the lower integrated core/low spool efficiency are
presented below.
Leakage levels in the integrated core/low spool are expected to be
greater than design values for the flight propulsion system because
the quality of experimental hardware is not as likely to he as high as
that specified for the flight hardware.
Clearance levels for the integrated core/low spool are expected to be
greater for the same reason, plus the fact that the integrated core/
low spool utilizes tenth stage high-pressure compressor bleed air for
active clearance control, which is less effective than the eighth
stage air used in the flight propulsion system. Eighth stage high-
pressure compressor bleed air ,_as chosen for the flight propulsion
system because of a 0.19 percent efficiency advantage and 0.28 percent
thrust specific fuel consumption advantage. Tenth stage compressor
bleed air was selected for the integrated core/low spool to avoid
design expense and additional hardware costs associated with external
plumbing to the eighth stage.
Experimental hardware is not expected to achieve design-point
operating conditions without more testing than is currently planned
for the integrated core/low spool.
The flight propulsion system status efficiency of 91.3 percent falls short of
the 91.5 percent goal because maneu,,er deflections produce higher than ex-
pected design blade tip clearances, particularly in the rear stages of the
turbine. The resultant clearance gaps were sufficiently large so that closure
to goal clearance levels could not be achieved with the active clearance con-
trol system at cruise conditions even by utilizing an optimized eighth stage
high-pressure compressor bleed schedule. If the goal clearance of 0.050 cm
(0.020 in) could be maintained, the calculated efficiency is 91.8 percent, as
indicated in Table 4.5-II.
Reducing the impact of maneuver deflections would require basic changes in the
rotor support system which translates into a major redesign effort. It was
decided that this would not be prudent, since the status efficiency is so near
the goal.
TABLE 4.5-III
DETAIL DESIGN REPORT STATUS PERFORMANCE
(At Aerodynamic Design Point -- After Completion of Mechanical Design )
Flight Propulsion Integrated Core/
System
Low Spool
91.8 91.0
Low-Pressure Turbine Component
Mechanical Effects
Reduction due to Tip Clearance
Increase 0.051 to 0.086 cm
(0.020 to 0.033 in) -0.4
Speed and Thermal Effect:
with Eighth Stage Air
-0.3
with Tenth Stage Air
Low-Pressure Turbine
91.3 90.3
Detailed Design Report Status, %
Intermediate Case Pressure
0.7 1.5
Loss,APT/PT, %
Exit Guide Vane Pressure
0.9 0.9
Loss,APT/PT, %
Assumptions:
(1) Interstage seal clearances as shown on page 139
(2) Outer blade tip seal clearances as shown on page 138
(3) 0.152 cm (0.060 in) clearance on flow guides
(4) No loss included for gaspath instrumentation
SECTION 5.0
SECTION 5.0
TURBINE THERMAL-MECHANICAL DESIGN
5.1 MECHANICAL DESIGN OBJECTIVES ANDGOALS
The primary objective of the mechanical design effort was to provide an inter-
mediate case and low-pressure turbine configuration with acceptable life that
would meet or exceed the componentaerodynamic and performance goals. A secon-
dary objective was to achieve this with the lightest possible structure at the
lowest possible cost.
The mechanical definition of the low-pressure turbine componentevolved
through an iterative process based on the results of the aerodynamic analyses
and various supporting technology programs. Wherenecessary, the turbine
design for the flight propulsion system was modified to meet specific require-
ments associated with the integrated core/low spool. Life requirements for the
major subassemblies are summarized in Table 5.l-I.
TABLE 5.1-I LIFE REQUIREMENTS FOR MAJOR TURBINE SUBASSEMBLIES Subassembly/ Requi red Life Remarks Turbine Intermediate Case: 15,000 hours (1)/ Strut Fairings Oxidation life is limiting failure mode.
3,300 missions Case/Strut Frame 30,000 hours (1)/ Structural limitation is 20,000 missions the radial load capacity required to prevent cata- strophic failure in the event of fourth stage turbine blade failure.
Rotor: Hub, disk, shaft 30,000 hours/ 20,000 missions (2) Blaries 15,000 hours (l)/ 3,300 missions Vane and Case: Case 30,000 hours/ 20,000 missions Vanes 15,000 hours (1)/ 3300 mi ssions Turbine Exhaust Case 30,000 hours(l )/ Materials and configuration 20,000 missions set by plug tuning design requirements.
Notes: 50 hot hours fo (I) integrated core/low spool Greater than lO _ (2) cycles for integrated core/low spool
5.2 TURBINE MECHANICAL CONFIGURATION
The mechanical configuration of the low-pressure turbine componentis illus-
trated in Figure 5.2-I. The major subassemblies of this system are: (1) the
turbine intermediate case; and (2) rotor, shaft, vanes and cases, and turbine
exit guide vane. These and the active clearance control system are described
in more detail in the following sections of this report. Significant design
features and advanced technology concepts that enhance performance, perfor-
manceretention, and durability are described where appropriate.
HUB THRUST BALANCE SEAL TURBINE EXHAUST CASE INNER CASE AIR SEALS- CASE TIED
'
Figure 5.2-I Low-Pressure Turbine Component Mechanical Configuration
5.2.1 Turbine Rotor Assembly
The short, stiff, high-spool rotor of the Energy Efficient Engine is straddle-
mounted by locating its front bearing in the compressor intermediate case and
its aft bearing in the turbine intermediate case forward of the low-pressure
turbine. This eliminates the overhung high turbine configuration of other
engine designs that can create large radial rotor deflections during flight
maneuvers. Centralizing the rotor mass betweenand near the support structure
causes the cases and rotors to deflect in a more similar fashion under nor-
mally encountered flight loads.
The low-spool rotor design takes advantage of this concept with a three bear-
ing support. The two front bearings, located at the fan intermediate case,
provide momentrestraint for the overhung fan/low-pressure assembly to mini-
mize maneuverdeflections. The low-pressure turbine rotor is cantilevered off
a rear bearing that is axially positioned near the plane of the front stage of
the low-pressure turbine to minimize rotor tip deflection where clearances
have a more significant effect on efficiency than with the longer rear stages.
An analysis of rotor system critical speed is discussed in Section 5.2.4 of
this report.
The rotor assembly for the low-pressure turbine is illustrated in Figure
5.2.1-I. Its primary elements are the blades, winged disks and nonintegral
hubs, a four-lip thrust balance seal, separate knife-edge seals for the inner
cavities, and the shaft. Design details pertaining to these componentsare
discussed in the following sections.
_o 4LIPTHRUST _ " I1[ BALANCE SEAL _ --i _ WINGED DISKS EDGESEALS _'_ I SEPARATE KNIFE " ' _ Figure 5.2.1-I Energy Efficient Engine Low-Pressure Turbine Rotor Assembly
5.2.1.I Blades
5.2.1.I.I Mechanical Design Features
The major criterion guiding the mechanical design of the turbine blades was
the elimination of the requirement for blade cooling, which eliminated the
need for complex internal cooling flow passages and permitted fabrication of
the blades from solid castings. Since the low-pressure turbine inlet tempera-
ture of ll61K (2090°R) is appreciably higher than current gas-turbine engines,
this approach requires the use of high-strength, high-temperature capability
materials and coatings on the front stages to provide the desired life.
The general characteristics of the turbine blades are illustrated in Figure
5.2.1-2. Airfoil geometry was determined by aerodynamic analysis and is
described in Section 4.3.2.2.2 of this report.
The total numberof blades is 438. The second stage blade is cast of PWA 1447
(MAR-M-247) and coated with PWA 270 (vapor deposition NiCoCrAly) for the flight
propulsion system in order to meet its life requirements. A conventional over-
lay coating, PWA 73, (diffused aluminide) is adequate to meet life requirements
of the integrated core/low spool. For both the integrated core/low spool and
flight propulsion system the third and fourth stage blades are cast of PWA 655
(Inconel 713) material. The third stage blade in the flight propulsion system
requires a PWA 73 coating, but this coating is not necessary for this stage in
the integrated core/low spool. The fifth stage blade in the integrated core/
low spool is also cast of this material. However, the fifth stage blade in the
flight propulsion system is fabricated from titanium-aluminum alloy because of
its lower weight property.
Fourth and fifth stage blade platforms incorporate leading edge and trailing
edge flow guides to reduce rim cavity flow ingestion and improve efficiency.
Blade platform weight is minimized by providing cast, conical, constant
thickness surfaces at the flowpath and cast pockets at the underside.
Thickness of the pocket and walls was set to minimize leakage through the
attachment. To balance out gas bending momentsat the blade root, all blades
incorporate a slight amountof built-in tilt.
Details of the blade-to-disk attachment are shown in Figure 5.2.1-3. To save
time and minimize program cost, an existing two-tooth broach design was uti- lized. The selected design provided the lightest weight while still meeting design requirements. This broach was utilized on all four stages and no attempt was made to optimize the attachment. A conventional blade retention tang is used on the second, third, and fourth stages, while the fifth stage has a shear lock arrangement. The shear lock arrangement was chosen in order to reduce a fifth stage disk stress concentration factor and improve disk rim low-cycle fatigue life. This shear lock is the same configuration used in the seventh and eighth stage blades of the Energy Efficient Engine high-pressure compressor.
Tip shrouds are included on the blades to enchance aerodynamic efficiency and control blade vibration. Details of these shrouds are shown in Figure 5.2.1-4 Pretwist and bearing surface angles were set to achieve the required tight- ness, while maintaining low bearing stresses in order to prevent excessive wear on the bearing surfaces. Forward and aft surface angles were set to main- tain curling stress within limits.
R1 _ a.-_ I* 'H j_I R1 _ a_ _--._L R 1 Wp ,'-I I R2 R3 _-- C -P R 4 =-Cc.
R5 I -_R 3 -.- Cc-- ,--- CC,- _ R3 STAGE 2nd 3rd 4th 5th NO. OF BLADES 120 96 100 122 BLADE MATERIAL PWA 1447 PWA 655 PWA 655 PWA 655 (3) COATING MATERIAL PWA 270 (1) PWA 73 (2) NOT REQ'D NOT REQ'D ATTACHMENT AXIAL LENGTH, (a), cm (in) 2.41 (0.95) 2.28 (0.90) 2.54 (1.00) 2.41 (0.95) AVERAGE AIRFOIL CHORD, (C), cm (in) 2.66 (1.05) 3.07 (1.21) 3.37 (1.33) 3.42 (1.35) 40.1 (15.8) 41.6 (16.4) 42.1 (16.6) 42.6 (16.8) R 1 , cm (in) 42.4 (16.7) 44.1 (17.4) 44.4 (17.5) 44.7 (17.6) R 2, cm (in) R 3, cm (in) 51.5 (20.3) 56.6 (22.3) 61.9 (24.4) 65.2 (25.7) PLATFORM AXIAL WIDTH, Wp INCLUDING FLOW GUIDE, cm (in) 4.77 (1.88) 5.61 (2.21) 5.91 (2.33) 5.46 (2.15) (1) COATING REQUIRED FOR FLIGHT PROPULSION SYSTEM. INTEGRATED CORE/LOW SPOOL WILL USE PWA 73 (2) NOT REQUIRED FOR INTEGRATED CORE/LOW SPOOL (3) MERL 101 (Ti-AI) IS BEING CONSIDERED FOR THE FLIGHT PROPULSION SYSTEM Figure 5.2.1-2 Low-Pressure Turbine Blade General Characteristics LpG LI STAGE 2rid 3rd 4th 5th 11 o 7o30 , 15 ° 0 BROACH ANGLE 14o45 ' 25 ° 15 ° 27o15 ' PLATFORM GAP ANGLE, _G 0.030 (0.012) 0.030 (0.012) 0,025 (0.O10) 0.030 (0,012) PLATFORM GAP, LpG cm (in) 22 ° 22 ° 22 ° 22 ° TAPER ANGLE, _T + 4030 , ± 3 ° ± 3 ° ± 3 ° RESTAGGER CAPACITY 0.78 (0.31) 0.78 (0.31) 0.78 (O.31) 0,78 (0.31) L 1 cm (in) 0.81 (0.32) 0.81 (0.32) 0,81 (0.32) 0.81 (0.32) L2, cm (in) B1ade-to-Di sk Attachment Desi gn Detail s Figure 5.2.1-3 9O # I / / L4 L 3' LSG STAGE / 2nd 3rd 4th 5th 1 ° lO15 ' 2o30 ' 0045 ' PRETWIST ANGLE, (_p 20 ° 30 ° 35 ° 25 ° BEARING SURFACE ANGLE, 0 B 37o30 ' 33 ° 35 ° 42 ° FORWARD, SURFACE ANGLE, 0 F 46030 ' 42 ° 35 ° 47030 ' AFT SURFACE ANGLE, 0 A 3.35 (1.32) 3.73 (1.47) 3.68 (1.45) 3.04 (1.20) AXIAL WIDTH, (W) cm (in) 2.00 (0.79) 1.93 (0.76) 1.85 (0.73) 1.44 (0.57) L 1 ~ cm (in) 1.27 (0.50) 1.95 (0.77) 1.93 (0.76) 1.14 (0.45) L2 _ cm (in) 0.81 (0.32) 1.14 (0.45) 1.11 (0.44) 0.73 (0.29) L 3 - cm (in) 1.44 (0.57) 1.47 (0.58) 1.21 (0.48) 0.83 (0.33) L4 - cm (in) 0.030 (0.012) 0.038 (0.015) 0.030 0.012 0.38 0.015) THRUST AIR, LSG - cm (in) Figure 5.2.1-4 Tip Shroud Design Details 5.2.1.I.2 Structural Analysis Structural analysis of the blades involved an analysis of blade vibration and flutter characteristics, attachment stresses, tip shroud stresses, and blade airfoil durability.
Stage Vibration and Flutter Analysis Individual rotor stage frequencies were selected to avoid engine orders known from experience to produce high stresses. These are the second and third orders for all stages, the eleventh (twenty-second order) excited by the up- stream struts for the first two stages and the thirtieth order excited by the exit struts for the last stage.
For the second stage rotor, liE and 22E resonances were of concern because of the proximity of the II upstream intermediate case struts. Figure 5.2.1-5 shows that for the liE first mode, a frequency margin of 9.8 percent at the maximum rotor speed is predicted. For 22E first mode, a margin of -I0.7 percent at the minimum cruise speed is calculated. All critical second mode resonances occur well outside of the operating range.
The third stage design has a -38 percent 22E and a -16 percent liE first mode frequency margin at the minimum cruise speed, with the 22E resonance occurring away from the idle range. The 3E first mode frequency margin is 14 percent and the 22E second mode frequency is also 14 percent at the maximum rotor speed (Figure 5.2.1-6).
Critical resonances for stage four were limited to the low orders. The final design resulted in margins of 52 percent 2E and 18 percent 3E for the first mode at the maximum rotor speed (Figure 5.2.1-7). Avoidance of the liE and 22E resonances was not required.
The fifth stage blade design in the integrated core/low spool incorporates a nickel-base alloy and has a 9 percent first mode 3E resonance margin at maxi- mum rotor speed. The 30E first and second mode resonances are predicted to occur at 1400 and 1760 revolutions/minute, respectively. These were considered acceptable for the integrated core/low spool (Figure 5.2.1-8).
Titanium alloy blades are being considered for the flight propulsion system to reduce weight. Vibration analysis with this material indicated that the fifth stage has an a_le first mode frequency margin above the second and third engine orders at maximum rotor speed. The thirtieth order resonance with the downstream struts will be well below minimum cruise speed for the first and second modes. Testing with strain gages will be required to demonstrate that stresses are low enough for flight engine safety. The resonance diagram is shown in Figure 5.2.1-9.
Figure 5.2.1-I0 presents the results of a flutter analysis on the four low- pressure turbine rotors. The Energy Efficient Engine designs are compared to previous Pratt & Whitney Aircraftdesigns in terms of aerodynamic damping and tip exit reduced velocity parameters. The results of t_is analysis indicate that all four rotors operate well above the unstable limit and are therefore not expected to encounter flutter problems.
2500 -- R EDLII_,_ 30E / 25E / 22E FIRST MODE - BLADE + MIN-CRUISE_ , / /
o
v > Z w D IDLE O
,c s
_5E _4E _-3E _2E 1 I 100_ 2000 3000 4000 5000 6000 7000 ROTOR SPEED (RPM) Figure 5.2.1-5 Low-Pressure Turbine Second Stage Rotor Resonance Diagrmn 2500 F I O FIRSTMODE 30E 22E BLADE+ DISK COUPLED REDLINE 1'T SECOND MODE MIN-CRUISE / / i 15E 14% > 111E Z IC/LS D IDLE O 10001 -.------- 2E 14% I I I l 0 1000 2000 3000 4000 5000 6000 7000 ROTOR SPEED (RPM) Figure 5.2.1-6 Low-Pressure Turbine Third Stage Rotor Resonance Diagram 2500-- REDLINE/ 30E / 25E " / 20E O FIRST MODE O SECOND MODE
i
BLADE + DISK COUPLED / / 2000i A ICLS Z IDLE O w 1000
/
_ _ _5E _ 4E I________ ___.__ _ _3E _------_ - 18% /2E I I I I I 1000 2000 3000 4000 5000 6000 7000 ROTOR SPEED (RPM) Figure 5.2.1-7 Low-Pressure Turbine Fourth Stage Rotor Resonance Diagram 25OO O FIRST MODE BLADE+ DISK I"t SECOND MODE 25E COUPLED REDLINE/30E
, -cr7
ICLS 0E IDLE 5E 4E 2E L I J _J 1000 2000 3000 4000 5000 6000 7000 ROTOR SPEED (RPM) Figure 5.2.1-8 Low-Pressure Turbine Fifth Stage Rotor Resonance Diagram MIN-CRUISE RED LINE [3 1ST BENDING MODE O 2ND BENDING MODE 30E 20E fl 5E ICLS IDLE U3 Q.
O >.- IOE Z LU D O UJ 5O0 t_ u.
2E I 0 1000 2000 3000 5000 ROTOR SPEED, RPM Figure 5.2.1-9 Low-Pressure Turbine Fifth Stage Rotor Resonance Diagram Ti tani um-Al umi num Al loy Bl ade 0.02 R2 R 3 (9 EEE LPT PREDICTION + D - ®% <> R4 R5® r_ or_ 0 -0.02 0 0 uJ < [9 × Oo E> _D D _ A t.,.'3 -0.04 Q o C>D Z LOW STABLE TURBINE UNSTABLE LIMIT ,"t O -0.06 .d I,,U < I I I I I I I I 4O 80 120 160 200 240 280 320 TIP EXIT REDUCED VELOCITY'_ 24 W2/b w Figure 5.2.1-I0 Shrouded Turbine Blade Flutter Analysis Attachment Stress Analysis Blade attachment stresses were calculated with consideration given to axial gas loads, airfoil and shroud residual moments, blade centrifugal pulls, and platform offset moments. Table 5.2.1-I summarizes the results of these anal- yses and indicates that the attachments for all stages have adequate stress margin. The margins shown represent the maximum stress levels.
TABLE 5.2.1-I BLADE ATTACHMENT STRESS SUMMARY (Margins in Percent) Second Blade Third Blade Fourth Blade Fifth Blade Blade Material PWA 1447 PWA 655 PWA 655 PWA 655 (*) Neck Tension 173 247 167 188 Tooth Shear 130 51 38 27 Tooth Bending 349 232 1 97 172 Tooth Bearing 108 48 16 29 NOTE (*): A Ti-Al alloy (MERL lOl) is being considered for the blade in the flight propulsion system. This would have even higher margins than those shown for PWA 655.
Tip Shroud Stress Analysis Bearing stresses and curling stresses were analyzed for the tip shrouds.
Results of this analyses indicated that adequate stress margins were achieved for all stages. These results are summarized in Table 5.2.1-II.
TABLE 5.2.1-II TIP SHROUD STRESS SUMMARY (Margins in Percent) Second Blade Third Blade Fourth Blade Fifth Blade Bearing Stress 15 24 3 8 Curl ing Stress 27 54 98 39 Blade Airfoil Durability Analysis Blade airfoil durability objectives were to provide 15,000 hours of service life (or 3300 flight missions) in the flight propulsion system and 50 hours of hot life (28°C (84°F) day at sea level takeoff power) in the integrated core/ low spool. The durability design analysis consisted of first defining metal temperatures and stresses at the design condition and then evaluating expected life relative to the durability objectives. Materials and coatings utilized in this analysis are listed in Figure 5.2.1-2.
Gas temperature profiles and resultant spanwise calculated stresses for the blade airfoils are shown in Figures 5.2.1-II through 5.2.1-14. Blade stresses were calculated using average profile temperatures. Airfoil creep strength margins were calculated using the relationship: % Margin = (Allowed Stress - Actual Stress) x lO0 Al lowed Stress A (1800) LL o n- 871 (1600) I-- < LLI rt TURBINE INLET TEMPERATURE = 1015°C(1859°F) LU 760 I-- 5TH STATOR EXIT TEMPERATURE = 651 °C(1205°F) (1400)
I I I I I
20 40 60 80 1 O0 PERCENT SPAN 206,844 (30) -- OO 137,896 (20) --, u_ oO uJ 68,948 rr I.- (10) B (/)
I I I I I
0 20 40 60 80 1 O0 PERCENT SPAN Fi gure 5.2. l-l I Second Stage Blade Durability Design Conditions and Calcu- lated Stress (Blade Material: PWA 1447, Coating: PWA 73) o_ (1600) Z) (1400) 1-- ,< r,-" I.,g TURBINE INLET TEMPERATURE = 1015°C(1859°F) r, 5TH ROTOR EXIT TEMPERATURE = 652°C(1206°F) I"- (1200)
I I I I I
20 40 60 80 1 O0 PERCENT SPAN 206,644 B (30) 137,896 ,,z (20) Q..
v; U3 LU 68,948 n" (10) I-- U'J I I I I I 0 20 40 60 80 1 O0 PERCENT SPAN Third Stage Blade Durability Design Conditions and Calcu- Figure 5.2.1-12 lated Stress (Blade Material: PWA 655; No Coating) (1600) o tO o i.u" (1400) rr t'.- < tr iii I:L TURBINE INLET TEMPERATURE = 1015°C(1859°F) m 5TH ROTOR EXIT TEMPERATURE = 652°C(1206°F) Iii (1200) I-
I I I I I
20 40 6O 8O 1 O0 PERCENT SPAN 206,844 __ (30) 1"37,896 (20) 68,948 (10) u3
I I I I I
0 20 40 60 80 1 O0 PERCENT SPAN Fourth Stage Blade Durability Design Conditions and Calcu- Figure 5.2.1-13 lated Stress (Blade Material: PWA 655, No Coating) (16OO) TURBINE INLET TEMPERATURE = 1015°C(1859°F) 5TH ROTOR EXIT TEMPERATURE = 651 °C(1205°F) A II v oO 760 (1400) # (1200) F-
I I I I I
20 40 60 80 1 O0 PERCENT SPAN 206.844 (30) ' _. _ PWA 655 _ ----- MERLI01 137.896 B _ (20) -._ 68.948__ "_.
-_ _
o I I I I --1
0 20 80 100 40 60 PERCENT SPAN Fi gure 5.2.1 -I 4 Fifth Stage Blade Durability Design Conditions and Calcu- lated Stress (Blade Material: PWA 655, No Coating) where allowed stresses were based on the desired life and l.O percent material creep properties.
The resultant airfoil creep strength margins are summarized in Table 5.2. l-III for both the flight propulsion system and the integrated core/low spool. All airfoils were determined to have adequate creep strength.
TABLE 5.2.l-Ill BLADE AIRFOIL CREEP STRENGTH MARGINS (Margins in Percent at Limiting Span) Flight Propulsion Integrated Core/ System Low Spool Second Blade 14 62 Third Blade 21 52 Fourth Bl ade 50 66 Fifth Blade 33 75 Since the second and third-stage blades operate in a high temperature environ- ment, a transient thermal strain analysis was conducted to determine thermal fatigue life. Engine operating conditions in a typical flight cycle were utilized in the analysis. These included startup and idle, acceleration to takeoff power, takeoff, climb, cruise power, deceleration to flight idle, approach power, and thrust reverse. Two strain cycles are experienced during each typical flight. The resultant strain histories for the second and third blades are shown in Figures 5.2.1-15 and -16. As indicated, the maximum strain ranges (i.e., strain cycle amplitudes) for these two airfoils were equal to or less than 0.24 percent.
Airfoil lives were subsequently calculated and are presented in Table 5.2.l-IV.
Life-limiting conditions of concern were cracking and metal surface oxidation.
As noted in this table, life for both the flight propulsion system and the integrated core/low spool exceeded the goals. Oxidation was the only concern for the integrated core/low spool because the integrated core/low spool low- pressure turbine will not accumulate enough running time to initiate cracks.
TABLE 5.2.l-IV SUMMARY OF PREDICTED BLADE LIVES Integrated Core/ Flight Propulsion System Low Spool Oxidation (hrs) Cracking (hrs) Oxidation (hrs) lO0 20,000 Second Blades 20,000 lO0 20,000 Third Blades 20,000 Fourth & Fifth Blades (a) (a) (a) Note: (a) Not Life Limiting 0.3 m 0.2 DECEL 0.1 Z _L_GH T ID_ TAKEOFF o
....,s'-i /
o.
-0.1 -0.2 -0.3 I I I I I I I I I - 17.7 93 204 3i 5 426 537 648 760 871 982 (0) (200) (400) (600) (8OO) " (1000) (1200) (1400) (1600) (1800) METAL TEMPERATURE, °C (°F) Figure 5.2.1-15 Energy Efficient Engine Second Blade Transient Strains (Flight Propulsion System) 0,3 B 0.2-- DECEL 0.1 -- Z ,( STRAIN RANGE TAKEOFF mr u_ F- ,,z
"'''_'------'"_,CCEL
mr -0.1 -0.2 -- -0.3 , I I ! ! I I I I -17.7 93 204 315 426 537 648 760 871 (0) (200) (400) "(600) (800) (1000) (1200) (1400} (1600) METAL TEMPERATURE, °C (OF) Figure 5.2.1-16 Energy Efficient Engine Third Blade Transient Strains (Flight Propulsion System)
5.2.1.2 Disk and HubAssembly
5.2.1.2.1 Mechanical Design Features
The low-pressure turbine disk and hub assembly shownin Figure 5.2.1-17 incor-
porates an "A-frame" construction technique, designed to provide the stiffness
required to control deflections caused by maneuver loads. The second and fifth
disks are bolted to the legs of the "A-frame", and the third and fourth stage
winged disks are bolted to the second and fifth stages to form the base of the
"A-frame". Separate knife-edge inner air seals also shield the base of the
"A-frame" from hot gaspath air and, in the third and fourth stages, provide
cooling air passages to the disk attachments. (Cooling air from inside the
drum is directed to the disk rim front and passes through the blade root gaps
to cool the disk rim and blade roots.)
(5.6 in) 14.2 cm ]_ 23.3 cm (9.2 in)R 32.2 cm 12.7 DISK RIM COOLING AIR 37.5 cm PASSAGES (14.8 in)R 39.6 (15 KNIFE-EDGE INNER AIRSEALS Figure 5.2.1-17 Low-Pressure Turbine Disk and Hub Assembly For both the flight propulsion system and integrated core/low spool, the hub and inner airseal materials were selected on the basis of availability, cost, and structural/radial clearance requirements. Hubs are made from PWA I003 (Incoloy 901). Inner air seals are made of PWA I099 (modified Inconel lO0 or MERL 76) to match the disk material, which is also PWA I099.
Existing tierod bolts are used for the integrated core/low spool. Spacers are utilized to accommodate the existing bolt lengths. The number of bolts employ- ed was established to satisfy the blade loss design criteria, except for the bolted joint between the fourth and fifth stage disks. Here, the number of bolts was doubled to improve flange sealing.
5.2.1.2.2 Structural Analysis Structural analysis of the hub assembly was conducted using the results of the transient thermal analysis described in Section 5.2.5 of this report. Rotor metal temperatures used in the analysis are shown in Figure 5.2.1-18. They were calculated based on sea level takeoff hot day conditions. The resultant stresses, summarized in Figure 5.2.1-19, were calculated based on this temper- ature distribution and an assumed 60-second acceleration to takeoff power.
Stress concentration factors for the disk rim locations were established by performing a finite element analysis on the fifth stage and correcting those values for the remaining stages. This was permissible because the same attach- ment geometry existed for all stages. All stresses were within allowable margins.
A subsequent life analysis indicated that all areas exceeded lO0,O00 cycles with the exception of those noted in Figure 5.2.1-20, which just meet flight propulsion system requirement of 20,000 cycles in the areas noted. Margin could be added here by a slight "tuning" of the design. Disk burst margins, creep life, and average tangential rim stresses are summarized in Table 5.2.1-V. All stresses and lives are adequate to meet integrated core/low spool requirements.
TABLE 5.2.1-V DISK STRUCTURAL SUMMARY Rotor 2 Rotor 3 Rotor 4 Rotor 5 Burst Margin 1.66 l .43 l l
1o4 6 8
Creep Life, hrs. lO4 Avg. Tangential Stress, MPa (ksi) 434,372 584,679 652,248 549,51 5 (63) (84.8) (94.6) (79.7) ]04 (939) (937) (949) (947) 506 (952) (981) (1001) (989) (1 O32) 563 574 1940) (1047) (1066) 568.
(1056) 11131) 586 718 l (1087) 11325) 602 667 591 571 (1116) (1233) (1057) (1097) 11060) TEMPERATURE, °C (OF) Figure 5.2.1-18 Rotor Metal Temperature Distribution Used in Stress and Life Analysis a B = BENDING STRESS, MPa (KSI) a H = NOOPSTRESS, MPa (KSI) aB a B _61,900 (96) a H = 431,614 (62.6 a B = 723,264 (104.9) a H = 150,306 (21.8) a B = 472,983 a B = 206,844 o B = 377,145 (68.6) (30) (54.7) a H = 160,648 a H = 127,553 o H = 366,803 (23.3) (18.5) (53.2) o B = 606,052 (87.9) o B = 293,029 o" H = 199,949 (42.5)K (29) I a B = 57,916 a H = 339,224 (8.4) (49.2)K a H = 275,792 (401 (95.5) _ a B = 658,453 o H = 130,311 a B = 83,427 (18.9) (12.1) a B = 896,324 a B = 51"7,110 o 8 = 713,611 a B = 259,933 (130) (75) (103.5) (37.7) a H = 122,037 o H = 368,182 o H = 57,916 a H = 26,889 (17.7) (53.4) (8.4) (3.9) Figure 5.2.1-19 Rotor Stress Summary NOTE: ALL AREAS EXCEED 100,000 CYCLES EXCEPT THOSE NOTED 38,000 CYCLES 20,000 CYCLES \ 48,000 CYCLES 20,000 CYCLES Fi gure 5.2.1-20 Rotor Low Cycl e Fati gue Li fe Summary 5.2.1.3 Thrust Balance Seal Assembly 5.2.1.3.1 Mechanical Design Features The configuration selected for the second stage thrust balance airseal is shown in Figure 5.2.1-21. It features a single knife-edge outer seal and a three knife-edge inner seal. Resonance and coincidence analyses indicated the need for a damper on the three knife-edge inner seal. Interlocking segmented sideplates on the front of the second disk protect the disk rim from hot gas- path air. The material for the rotating seals is PWA 1O03 and PWA 655 for the sidepl ates.
31.7 cm STAGE _.. SECOND (12+5 inla SK BALANCE SEAL INNER THRUST BALANCE SEAL 37.5 cm (14.8 in)R SEGMENTED SIDEPLATES (20) Figure 5.2.1-21 Low-Pressure Turbine Thrust Balance Seal Assembly 5.2.1.3.2 Structural Analysis Stress analysis of the thrust balance seal was included as part of the rotor hub stress analysis, and results are shown in Figure 5.2.1-19. Stresses shown in the figure are well within allowables. In addition to stress analysis, a vibration analysis was conducted to ensure that no resonance or coincidence problems existed within the operating range (ll03 to 3902 rpm). Resonance margins were 23 percent for the stator and lO0 percent for the rotor. These results indicated that both the inner and outer rotating seals were free of resonance. However, a damper, as shown in Figure 5.2.1-21, was required on the inner seal to avoid coincidence with the static seal lands.
5.2.1.4 Inner Cavity Knife-Edge Seals 5.2.1.4.1 Mechanical Design Features The interstage inner cavity knife-edge seals are shown in Figure 5.2.1-22. As mentioned in Section 5.2.1.2.1, the seals between stages two-three and three- four are multifunctional. They not only provide gaspath sealing between stages, but shield the disk flanges from hot gaspath air and provide a passage for cooling air to the disk rim and airfoil attachments, as shown in the figure. All of the inner air seal knife-edge axial stations were established to run axially over the seal lands at sea level takeoff and aerodynamic design point steady state conditions. Two of the three knife edges for the second- third and third-fourth stage seals and one of the two fourth-fifth stage knife edges will run over the seal lands during flight propulsion system engine transient conditions. (Seal clearance control is discussed in more detail in Section 5.2.3.6 of this report.) Seal material is PWA I099 for thermal com- patibility with the disk material.
5.2.1.4.2 Structural Analysis Stress analysis of the inner cavity knife-edge seals was included as part of the rotor hub stress analysis and results are shown in Figure 5.2.1-19. The stresses shown in the figure are well within allowables. Vibration analysis of these seals indicated that resonance margins were 85 percent for the second- third stage seal, 74 percent for the third-fourth stage seal, and greater than lO0 percent for the fourth-fifth stage seal. All were free of coincidence problems.
5.2.1.5 Rotor Shaft Assembly 5.2.1.5.1 Mechanical Design Features The low rotor shaft assembly is illustrated in Figure 5.2.1-23. The low rotor shaft connects the fan and low-pressure compressor to the low-pressure turbine.
It is supported at the front end by the Nos. l and 2 bearings and at the rear by the No. 5 bearing. The front end of the shaft connects to the fan and low- pressure compressor stub shaft through a spline. The low-pressure turbine rotor is mounted on the rear of the shaft and is cantilevered off the No. 5 bearing, which is damped to desensitize the case to rotor vibration. A center vent seal is located at the aft end of the shaft. In the No. 5 bearing compartment area, oil drain holes are provided to remove any oil that may accumulate in the low shaft inner diameter.
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/ I _4 / / / cN _ / _ v / The shaft is sized to satisfy stress requirements under the combined loadings of gyroscopic limit maneuver loads, maximum design torque, and shaft thrust.
Adequate intershaft clearance has been provided to prevent intershaft contact under limit load conditions. The relative radial deflection has been predicted to be 0.317 cm (0.125 in) under a one radian/second gyro, plus 3 g's, plus 3 radians/second acceleration. With the design "zero-load" radial clearance of 0.825 cm (0.325 in), this results in a minimum running clearance under load of 0.508 cm (0.2 in). The maximum deflection occurs approximately 64.7 cm (25.5 in) forward of the No. 5 bearing location.
The intershaft seal is a constant diameter seal designed to eliminate the potential for a self-destructive rub mechanism occurring. Seal geometry permits an axial translation of 0.330 cm (0.130 in) between knife edges and seal land during engine transient excursions. The aft center vent seal inner diameter was sized by the flow area required for the No. l bearing compartment deoiler.
Seal geometry was established by allowing 1.651 cm (0.65 in) rearward and l.Ol6 cm (0.40 in) forward translation of the rotating hardware relative to the static structure in order to satisfy intermesh criteria and engine tran- sient excursions. Shaft material for the flight propulsion system is PWA 733 (low alloy steel). To save cost for the integrated core/low spool, AMS 6304 has been selected.
5.2.1.5.2 Structural Analysis Structural analysis of the shaft was conducted using sea level takeoff hot day conditions. All notched locations on the shaft were evaluated for high cycle fatigue life and bearing stresses were evaluated at the front and rear spline locations. In addition, shaft ovalization, flutter and low cycle fatigue life were assessed.
Figure 5.2.1-24 summarizes the temperatures and calculated maximum stresses at critical stress locations on the shaft front. All stresses are within allow- able limits. A water quench heat treat of the front of the shaft at the spine area is required to develop the high tensile and fatigue properties needed for adequate vibratory shear stress margin.
Low cycle fatigue lives in all locations exceeded the flight propulsion system goal of 20,000 cycles, with the exception of the front spline, which had a life of 1300 cycles. This deficiency can be resolved by increasing the pitch diameter by 0.381 cm (0.15 in), increasing the root fillet radius, and crown- ing the spline to better distribute the loads. The intershaft seal has ade- quate burst margin and no vibration analysis of this seal was required.
Temperatures and calculated stresses at critical stress locations on the shaft rear are summarized in Figure 5.2.1-25. All stresses are within allowables and low cycle fatigue lives at all locations exceed the design goal of 20,000 cycles. The center vent seal has adequate burst margin and vibration analysis established its resonance margin at greater than lO0 percent and its coinci- dence margin at 45 percent.
High cycle fatigue lives and ovalization flutter are well within allowables for the shaft design.
OrC = 588,126 MPa (85.3 ksi) --7 O" C = 562,615 MPa (81.6 ksi) T = 176°C (350OF) / T = 121°C (250OF) OrA = 84,116 MPa (12.2 ksi) / / T = 1760C (350OF) / / MATER_AL = PWA 733 V--ETBRG = 141'343 MPa '20"5 ksi) / _ I- T = 176°C (350OF) / / _-_-_- .............. ,I--_- --_/_ _ _[ _34-,- ' Figure 5.2.1-24 Low Rotor Shaft Front Temperature and Stress Summary O" C = 416,445 MPa (60.4 ksi) / T = 510°C (950OF) / / or C = 539,173.3MPa(78.2ksi) / O'BRG = 110,316MPa(16ksi) T = 204°C (400OF) / T = 482oc (900OF) or C = 422,651 MPa (61.3 ksi) T = 510°C (950OF) Figure 5.2.1-25 Low Rotor Shaft Rear Temperature and Stress Summary 5.2.2 Turbine Vane and Case Assembly The low-pressure turbine vane and case design is based on a requirement for tight clearance control and uncooled vane airfoils; both aimed at achieving a high level of component efficiency. In addition, the case structure must be adequate for blade containment. Major elements of this assembly, illustrated in Figure 5.2.2-I, are the vanes, the inner and outer cases, and the turbine exhaust case/exit guide vane. Design details pertaining to these components are discussed in the following sections.
5.2.2.1 Vanes 5.2.2.1.I Mechanical Design Features The major criterion for the mechanical design of the vanes was the same as that for the blades -- elimination of the requirement for cooling. As with the blades, advanced high-strength, high-temperature capability materials and coatings are required to provide the desired life.
The general characteristics of the turbine vanes are illustrated in Figure 5.2.2-2. Airfoil geometry was determined by aerodynamic analysis and is described in Section 4.3.2.2.2 of this report.
There are 318 vanes in the four low-pressure turbine stages. To achieve flight propulsion system life goals without the use of air cooling, the second vane (inlet guide vane) is cast in single crystal SC 2000 (advanced high nickel alloy single crystal) with a PWA 286 (advanced NiCoCrAly vapor deposition) coating and the third vane is cast in PWA 1447 with a PWA 73 coating. Life requirements for the integrated core/low spool second vane can be met with cast single crystal PWA 1480 and PWA 73 coating. Third vane life requirements for the integrated core/low spool can be met with cast PWA 1455 and PWA 73 coating. Fourth and fifth stage vanes for both applications are cast PWA 655.
Coatings are not required for these airfoils.
All vane airfoils are stacked on a radial line passing through the center of gravity at the root section and incorporate a slight amount of tangential tilt to balance out root bending stresses. All have elliptical leading edges to enhance aerodynamic performance. The inner diameter platforms of stages three, four, and five incorporate flow guides for compatibility with the adjacent blades. The inner airseals on stages three, four, and five incorporate 0.158 cm (0.0625 in) cell honeycomb made from 0.005 cm (0.002 in) thick Hastelloy X foil and are brazed to the vane inner diameter seal lands.
The third, fourth, and fifth stage vanes are supported in the conventional manner, cantilevered from the vane case and attached with vane feet and case hooks as illustrated in Figure 5.2.2-3. The rear vane feet for stages three and four are notched to reduce surface contact with the case and minimize conductive heat flow into the case structure. The second vane support struc- ture is illustrated in Figure 5.2.2-4. The outer diameter support, made of AMS 5707 Waspaloy, is a simple hook arrangement that permits the vane to move radially, but restrains axial motion. The inner diameter support made of AMS 5666 incorporates a double hook arrangement, in conjunction with an anti- rotation pin, to provide the required axial, radial, and torque restraint. The inner support is bolted to the turbine intermediate case inner torque box.
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S" _4 or- STAGE 2nd 3rd 4th 5th NO, OF VANES 54 72 84 108 VANE MATERIAL SC2000 (1) PWA 1447 (2) PWA 655 PWA 655 COATING MATERIAL PWA 286 (3) PWA 73 NOT REQ'D NOT REQ'D WIDTH (W), cm (in) 4.97 (1.96) 5.18 (2.04) 4.87 (1.92) 5.02 (1.98) AVERAGE AIRFOIL CHORD, (C), cm (in) 3.96 (1.56) 4.08 (1.61) 3.73 (1.47) 3.68 (1.45) R1, cm(in) 37.5(14.8) 41.4(16.3) 41.9(16.5) 42.9(16.9) R2, cm (in) 40.3(15.9) 43.6(17.2) 44.4(17.5) 44.7(17.6) R 3, cm (in) 49.0 (19.3) 54.1 (21.3) 59.4 (23.4) 64.0 (25.2) (1) INTEGRATED CORE/LOW SPOOL WILL USE PWA 1480 (2) INTEGRATED CORE/LOW SPOOL WILL USE PWA 1455 (3) INTEGRATED CORE/LOW SPOOL WILL USE PWA 73 Figure 5.2.2-2 Low-Pressure Turbine Vane General Characteristics 3v I I I Figure 5.2.2-3 Attachment Scheme Utilized for Third, Fourth, and Fifth Stage Vanes OUAL HOOKS ANTI-ROTATION PIN _ _ ":'_ "-_ f CASE HOOK AMS 5707 Figure 5.2.2-4 Second Vane Support Structure This configuration provides for proper mating of the turbine flowpath with that of the turbine intermediate case strut fairings and minimizes steps in the flowpath during engine operation. The second vane support structure per- mits the vane to tilt slightly fore and aft under thermal and gas loads. To minimize attachment leakage caused by this tilt action, chordal sealing sur- faces are machined on the vane and attachment mating surfaces at the locations shown in Figure 5.2.2-5. The fourth stage vane incorporates a slight amount of rearward tilt to satisfy a blade meshing criteria in the event of a shaft failure.
Individual semi-machined vanes are Transient Liquid Phase (TLI_) bonded together into clusters for improved strength and reduced leakage. Second stage vane cluster size (Figure 5.2.2-6) is limited to two vanes. This is the maxi- mum cluster size that will permit straight chordal seal faces to be cut in the radial space available. Each second stage vane cluster incorporates one anti- rotatiomE,pin. As shown in Figure 5.2.2-7, the third through fifth stage vanes are TLP_Jbonded in clusters of three. Each vane in the cluster incorporates an anti-rotation pin.
CUT Figure 5.2.2-5 Location of Second Vane Chordal Cuts Employed for Attachment Leakage Control BONO JOINTS % CHORDAL SEALING SURFACE Figure 5.2.2-6 Low-Pressure Turbine Second Stage Vane Cluster •BOND JOINTS BOND JOINT, ( Figure 5.2.2-7 Typical Third, Fourth, Cluster and Fifth Stage Vane 5.2.2.1.2 Structural Analysis Structural analysis of the vanes comprised analysis of stresses in the vane airfoils, vane hooks, and anti-rotating pins, plus an assessment of vane air- foil durability.
Vane Stress Analysis A vane stress analysis was conducted assuming sea level takeoff operating con- ditions. Hot streak temperatures were considered where appropriate. Limits were based on one percent creep and 0.2 percent yield strength for 50 hours (anticipated integrated core/low spool running time). Calculated stresses are listed in Table 5.2.2-I. All are well within allowables with the highest being a combined stress of 226,149 MPa (32.8 ksi) at the second vane inner diameter hook and a bending stress of 226,838 MPa (32.9 ksi) at the third stage anti- rotation pins.
TABLE 5.2.2-I AIRFOIL STRESS SUMMARY (Stresses in MPa (ksi) and Temperatures oc (OF) 2nd 3rd 4th 5th stage stage Stage stage Ai rfoiI : Bending Stress 20,684 (3.0) 29,647 (4.3) 62,742 (9.I) 60,674 (8.8) Hot Streak Temperature I176 (2150) I048 (1920) 915 (1680) 787 (14SO) Vane Hook : Bending Stress .... I08,248 (15.7) 200,638 (29.1) 153,064 (22.2) Max. Combined Stress 226,149 (32.8) ............
Temperature .... 732 (1351) 677 (1251) 671 (1240) Hot Streak Temperature 940 (1725) ............
Anti-Rotation Pins: Shear Stress 15,858 (2.3) 70,326 (I0.2) 61,363 (8.9) 43,437 (6.3) Bending Stress 19994 (2.9) 226,838 (32.9) 197,191 (28.6) 146,169 (21.2) Temperature .... 623 (I154) 597 (If07) 615 (If40) Hot Streak Temperature 1,037 (1900) ............
Vane Airfoil Durability Analysis Durability analysis of the vanes was conducted in a manner similar to that of the blades described earlier. Life objectives for the vanes were identical to those for the blades. Materials and coatings utilized in the vane durability analysis are listed in Figure 5.2.2-2.
Gas temperature profiles and resultant spanwise calculated stresses for the clustered vane airfoils are shown in Figures 5.2.2-8 through 5.2.2-II. Since the vanes are stationary, the worst case is when a vane is exposed to a hot streak. Consequently, full hot streak temperatures were used in the calcul- ations.
The resultant airfoil creep strength margins are summarized in Table 5.2.2-II for both the flight propulsion system and the integrated core/low spool. All airfoils were determined to have adequate creep strength.
TURBINEINLETTEMPERATURE = IOISeCiIBS9OFI 5TH ROTOR EXIT TEMPERATURE - 652 °C! 1206°F) o.. 1.o93 f,J (2000( o LG n..
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£RAGE f 1BC_)) I I I I I 20 40 60 80 100 PERCENT SPAN 137.896 (20) -- U_ 68.948 (loi _4 __ 0 -68.948 (-I0) i w I I I 20 40 60 80 IO0 PERCENT SPAN Figure 5.2.2-8 Second Stage Vane Durability Design Conditions and Calcu- lated Stress (Vane Material: PWA 1480, Coating: PWA 73) HOT STREAK ,.p 982 (18OO) O o u/ n- :_ 871 F- ,< (16001 rr 5TH ROTOR EXIT TEMPERATURE - 651 eCl1205°F) F- (1400) I I i I I 0 20 40 60 80 100 PERCENT SPAN 137,896 (20) v. 68.948 (10) g.
m _ 0 -68.948 I I I I I (-10) 0 20 40 60 80 100 PERCENT SPAN Figure 5.2.2-9 Third Stage Vane Durability Design Conditions and Calculated Stress (Vane Material: PWA 1455, No Coating) HOTSTREAK
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_. _ TURBINE INLET TEMPERATURE - 1015°C11859°F) J 5TH ROTOR EXIT TEMPERATURE - 692eC(1206eFI _- 8,8 1 I I ,,200, I I I 0 20 40 60 80 100 PERCENT SPAN 137,896 (20I I _ o 0 20 40 60 80 100 PERCENT SPAN Figure 5.2.2-I0 Fourth Stage Vane Durability Design Conditions and Calcu- lated Stress (Vane Material: PWA 655, No Coating) HOT STREAK u. 760 o_ 11400)
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-68,848 I I I I I 1-10) 0 20 40 6O 80 100 PERCENT SPAN Figure 5.2.2-11 Fifth Stage Vane Durability Design Conditions and Calculated Stress (Vane Haterial: PWA 655, No Coating) TABLE 5.2.2-II VANE AIRFOIL CREEP STRENGTH MARGINS (Margins in Percent at Limiting Span) Flight Propulsion Integrated Core/ System Low Spool Second Vane 28 45 Third Vane 34 55 Fourth Vane 44 59 Fifth Vane 75 82 Since the second and third vanes operate in a high temperature environment, a transient thermal strain analysis was conducted to determine their thermal fatigue life. Engine operating conditions were the same as those used in the blade durability analysis described earlier. The resultant strain histories for the two airfoils are shown in Figures 5.2.2-12 and 5.2.2-13. As indicated, the maximum strain range for the second vane was 0.51 percent and for the third vane was 0.49 percent. Strain range for fourth vane was 0.42 percent.
The fifth vane is not life limited as a result of transient strain. The strain range for the vanes was higher than for the blades because of the effects of clustering the vanes.
03 m (FPSi 0.2 0._ FLIGHT / z m TAKEOFF o IDLE I1_ U_ -0.1 STRAIN RANGE 0,2
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-0.3 -0.4 I I I I I I I I I 93 204 315 426 537 648 760 871 982 (200) (400) (600) (800) 11000) (1200) 11400) (1600) (1800) METAL TEMPERATURE, °C (°F) Figure 5.2.2-12 Energy Efficient Engine Second Vane Transient Strains 0.3 0.2 DECEL 0.1 FLIGHT _z IDLE TAKEOFF _ o ,# START _ -0.1 _ STRAIN RANGE -0.2 _ ACCEL_ -0.3 _ I I I I I I I I - 17.7 93 204 315 426 537 648 760 871 1400] (600) 18001 ( 10001 (1200) ( 140OI (1600) (O) (200) METAL TEMPERATURE. °C (°F) Figure 5.2.2-13 Energy Efficient Engine Third Vane Transient Strains Airfoil lives were calculated and are presented in Table 5.2.2-111. Life- limiting conditions of concern were cracking and metal surface oxidation. As noted, life goals for the integrated core/low spool were exceeded for all vanes. Goals for the flight propulsion system were met or exceeded for all vanes, with the exception of the second vane whose life was limited to 9300 hours with a single coating. Analysis indicated that one recoating of this vane would extend the life beyond 15,000 hours.
TABLE 5.2.2-III SUMMARY OF PREDICTED VANE LIVES Integrated Core/ Fli_ht Propulsion System Low Spool Oxidation (hrs) Cracking (hrs) Oxidation (hrs) Second Vanes 20,000 9,300 (1) lO0 Third Vanes 20,000 15,000 lO0 Fourth Vanes 20,000 20,000 lO0 Fifth Vanes (2) (2) (2) Notes: (1) Life 15,000 hrs. with one recoating.
(2) Not life limiting.
5.2.2.2 Vane Cases 5.2.2.2.1 Mechanical Design Features The basic design objectives for the low-pressure turbine case were to: (1) provide blade containment, (2) support the vanes, (3) provide active clearance control, and (4) meet life requirements.
The case design evolved from preliminary design studies that compared bolted vane support and hooked vane support designs (both with external active clear- ance control) with a hooked vane support design incorporating a double wall and internal active clearance control. The latter design approach was selected as being the lightest, least expensive, and best for active clearance control.
The concept was refined in the detail design process and the resultant design is shown in Figure 5.2.2-14. The main feature of this design is the double wall, with the space between the walls serving as a passage for cooling air.
This cooling air provides a dual service in that active clearance control air is also used for cooling the vane feet and case hooks via drilled holes that meter the air into the cooling passage manifolds. This system of metering makes cooling flow control independent of part-to-part tolerances. The active clearance control baffle-to-case gap can be tailored for each stage and cooling air supply temperature adjusted, through the use of mixing valves, to achieve optimum case thermal response characteristics. (Active clearance control is discussed in more detail in Section 5.2.2.4 of this report.)
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Airseals are fabricated from forged rings by electron beam cutting. The honey- comb rubstrip is fabricated from Hastelloy X, has a 0.157 cm (0.062 in) cell size with a 0.005 cm (0.002 in) foil thickness and is nickel brazed to the outer airseal. Anti-torque pins serve to position the outer airseals so as to cover the gaps between vane clusters. The heatshields are also fabricated from Hastelloy X in segments of eight per stage. Heatshields under the vanes are assembled with an axial pinch for vibration damping. Those under the outer airseals incorporate standoffs formed into the sheetmetal to minimize heat transfer from the outer airseals and heatshields to the cooling manifold.
Materials selected for the case, baffles, and cooling manifolds are summarized in Figure 5.2.2-16. Inconel 718 was selected for the turbine inner case be- cause of its inherent strength at elevated temperatures. The outer case material, A 286 was chosen for its higher thermal expansion characteristics to minimize thermal growth differences between the two cases, since the outer case is always cooler than the inner case. The internal cooling air manifolds were to be initially fabricated from Inconel 718 for all six locations but stress analysis dictated a change to Waspaloy for the three manifolds adjacent to the front feet of the vanes. The forward set of baffle panels absorb the axial seal load from the outer case and are made from Inconel 718. The rear two sets of baffle panels are not load-carrying and are made from stainless •steel. All the panels incorporate circumferential sheet metal seals between panels to minimize leakage.
5.2.2.2.2 Structural Analysis Structural analysis of the case comprised an evaluation of blade containment and low cycle fatigue life. In the containment analysis, no credit was taken for any containment capability in the segmented outer airseals, sheet metal heatshields, and active clearance control intercase flow guides. The inner and outer cases were assumed to provide total blade containment. The thickness variation required was provided only in the inner structural case so that the outer case thickness was maintained constant. The resultant wall thickness distribution required to provide adequate containment capability is shown in Table 5.2.2-IV.
TABLE 5.2.2-IV CASE THICKNESS REQUIREMENTS FOR BLADE CONTAINMENT Inner Case Outer Case Blade Stage Thickness, cm (in) Thickness, cm (in) Second 0.203 (0.080) 0.I14 (0.045) Third 0.266 (0.I05) 0.114 (0.045) Fourth 0.393 (0.155) 0.114 (0.045) Fifth 0.558 (0.220) SEGMENTED OUTER AIRSEAL (18 PER STAGE) HONEYCOMB ABRADABLE ANTI-TORQUE PIN HEATSHIELD (8 PER STAGE) MANIFOLD COOLING AIR PASSAGE INNER CASE Figure 5.2.2-15 Typical Outer Airseal Configuration Life analysis considered the combined effects of stresses resulting from case buckling loads and stresses due to thermal transients. The four types of loading that contribute to low-pressure turbine case buckling and used in the analysis are external pressure, bending, axial compression, and shear. Also, a 6-g limit maneuver load for the flight propulsion system was included in the analysi s.
Thermal stresses were calculated based on the transient thermal strain analysis used for the blade and vane durability analyses. Sea level takeoff hot day conditions were assumed. The case temperature distribution for these conditions is shown in Figure 5.2.2-17. Buckling stresses (predominantly due to maneuver loads) were superimposed over those caused by thermal transients and the resultant stresses are summarized in Figure 5.2.2-18. Case life based on this distribution was determined to be in excess of the required 20,000 cycles for the flight propulsion system.
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5.2.2.3 Turbine Exhaust Case 5.2.2.3.1 Mechanical Design Features The turbine exhaust case includes a turbine exit guide vane to remove swirl from the low-pressure turbine exit flow. It is an integral ring-strut-ring, welded structure, with the 30 struts doubling as airfoils for the exit guide vanes. Design requirements for this structure were to meet design life and provide radial and trunion springrates in the ring-strut-ring structure that would tune exhaust plug pitch and vibratory modes outside the excitation frequency. The flight propulsion system preliminary design which meets these requirements is illustrated in Figure 5.2.2-19. It is a lightweight structure with hollow vanes/struts fabricated from sheet Ti-Al. The inner and outer rings are forged from the same material and welded to the struts. The inner ring supports the exhaust plug as shown. A flange at the rear of the outer ring provides support for the exhaust/mixer and serves as the engine-to-fan duct interface for cowl load sharing. The front outer support flange transfers axial loads from the turbine exhaust case forward to the low-pressure turbine case and provides containment for the fifth stage rotor blade.
In order to reduce cost for the integrated core/low spool, a simpler design was established that would provide the same structural and aerodynamic charac- teristics as the flight design, but was heavier and less costly. The design is illustrated in Figure 5.2.2-20. Here, the vanes/struts are solid castings o'f Greek ASCOLOY material. The inner and outer rings and flanges are forged from the same material. All are welded together into a single assembly.
5.2.2.3.2 Structural Analysis The structural feasibility of the turbine exhaust case designs for the flight propulsion system and the integrated core/low spool is based on experience with similar structures fabricated from current, lower strength material and operate in more severe thermal environments. The maximum metal temperatures expected for the integrated core/low spool turbine exhaust case assembly are 526°C (980°F) for the inner case ring, 648°C (1200°F) for the vanes/struts, and 604°C (ll20°F) for the outer duct case ring.
A three-dimensional NASTRAN analysis of the flight propulsion system ring- strut-ring structure was performed to determine trunion and radial springrates.
These springrates were used to determine the vibratory resonances of the tail- plug. The turbine exhaust case structure was resized as needed to tune the tailplug pitch and bounce vibratory modes out of the operating ranges. The predicted tailplug resonant frequencies are shown in Table 5.2.2-V. They meet or exceed the requirements for the modes indicated.
TABLE 5.2.2-V PREDICTED TAILPLUG RESONANT FREQUENCIES Vi bratory Mode Resonant Frequency Pitch 4,800 rpm Bounce 20,000 rpm INNER O4AMETER RING BOX STRUCTURE
\
FRONT FLANGE/ FLOW GUIDE EXHAUST RLUG SUPPORT EDGE VANE/STRUT OUTER DIAMETER DUCT CASE RING FRONT OUTER FAN DUCT/EXHAUST MIXER SUPPORT FLANGE SUPPORT FLANGE Figure 5.2.2-19 Flight Propulsion System Turbine Exhaust Case Assembly - Prel iminary Design INNER RING SUPPORT FLANGE FRONT ._ FLANGE _ 43.9 cm 44.4 cm .L [////'_ DUCT CASE RING iNNER DIAMETER _l_m)R II_//_ 48.7 ¢m (19 R .
15.7 cm .._ _ -- (6.2 in)
.7 EXHAUST PLUG
LEADING _FLANGE EDGE VANE/STRUT • EDGE I/ '_k_"_TRAIUNG 66..2 cm (26.1 inIR | | , _. 16.7 cm ._._ _6.6 ,nlR, j 7 1 f 6.3 cm " DUCT CASE R,NG 4_Y c28_ inIR / OUTER SUPPORT FAN DUCT / OUTER RING FLANGE SUPPORT FLANGE REAR SUPPORT FLANGE Figure 5.2.2-20 Integrated Core/Low Spool Turbine Exhaust Case Assembly 5.2.2.4 Low-Pressure Turbine Active Clearance Control System 5.2.2.4.1 Design Approach To achieve high efficiency in the low-pressure turbine, it is desirable to maintain tight clearances at the blade tips and inner airseals throughout the engine operating envelope and especially at the design point. Causes for these clearances (or gaps) fall into two major categories: (1) those associated with the relative difference in radial growth between the rotor and case resulting from centrifugal forces and thermals; and (2) those associated with maneuver deflections, rotor whirl, case ovalization, tolerances, and eccentricities.
The latter, which were included in the structural design criteria for the component, are essentially uncontrollable once the design has been fixed. Gaps caused by the former can be controlled, to a certain extent, by modulating case thermal expansion or contraction through the application of controlled temperature cooling air. It is the function of the active clearance control system to provide the case movement required to minimize clearances and avoid rotor-to-case "pinch", while maintaining case temperature within acceptable Iimits.
To minimize deterioration, a "no-rub" philosophy was adopted. In conjunction with this, a clearance goal of 0.050 cm (0.020 in) was established for the blade tips and inner airseal at the aerodynamic design point.
5.2.2.4.2 Mechanical Design Features The major features of the low-pressure turbine case active clearance control system are shown in Figure 5.2.2-21. Cooling air is pumped to the turbine case, enters a manifold, and is directed aft between segmented sheet metal flow guides and the inner turbine case wall. These flow guides are attached to the inner case with rivets and spacers to provide a controlled gap dimension at each stage, tailored to maintain high thermal coefficients for optimum case response. Outer airseals are case-tied to ensure no relative motion between them and the case. A portion of the cooling flow is metered through holes in the inner turbine case wall to internal manifolds adjacent to the vane feet.
This cooling air flows around the vane feet areas and reduces heat transfer from the vanes and outer airseals into the case.
Outer airseal radial growth for active clearance control is regulated through the use of a combination of tenth and fifteenth stage high-pressure compressor bleed air; maximum growth occurring through the use of all fifteenth stage air and minimum growth through the use of all tenth stage air. However, use of all fifteenth stage air at sea level takeoff power produces case hook temperatures beyond the capability of the Inconel 718 case material. Therefore, the active clearance control system proposed for the integrated core/low spool uses a combination of tenth and fifteenth stage high-pressure compressor bleed air.
The pressure in the outer manifold of the case is 30 percent of the total pressure at the high-pressure compressor exit, and O.l percent of core engine airflow is used to cool each hook. Total active clearance control air in this case is l.O percent of core engine airflow.
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In contrast, the flight propulsion system uses a combination of eighth and fifteenth stage high-pressure compressor bleed air. Use of eighth stage air increases case closure capability over tenth stage air, but use of all eighth stage air reduces the supply pressure in the active clearance control manifold below the inlet guide vane gaspath static pressure. Therefore, cooling holes are not drilled in the second stage vane support hook area or the third vane front support hook area. Cooling passages in these areas are filled with insulation. In this case, pressure in the outer manifold of the case is 16 percent of the total pressure at the high-pressure exit and total active clearance control air is reduced to 0.8 percent of core engine airflow. The case design, as modified for the flight propulsion system, is illustrated in Figure 5.2.2-22.
5.2.2.4.3 Clearance Analysis To establish the relative growth between the case and rotor as a result of rotor centrifugal forces, thermals, and pressure, shell analyses were con- ducted over the full range of engine operating conditions. The resultant radial growth of the case and rotor were plotted versus time and superimposed to establish "pinch" points. As shown in Figure 5.2.2-23, the "pinch" point occurred during a snap acceleration from idle to takeoff power. Clearance deviations associated with maneuver loads, rotor whirl, ovalization, toler- ances, and eccentricities were assumed to occur at the pinch point to ensure a "no rub" condition. Combining the effects of these deviations with clearances determined from the rotor and case growth analyses resulted in clearances sub- stantially in excess of the 0.050 cm (0.020 in) goal. Thermal matching with the active clearance control system yielded excellent clearance results at the aerodynamic design point and sea level takeoff conditions, but these were negated by large maneuver deflections.
In an effort to achieve clearances closer to the aerodynamic design point, a requirement was added that the blade tip knife edges be ground at assembly to + 0.005 cm (+ 0.002 in). tolerance on the diameter. In addition, any toler- ances that cause deviation from nominal would be allowed to rub in at the pinch point and in effect "custom machine" blade tips and seals to a line-on- line or nominal condition. While this alters the "no-rub" philosophy, the rub- bing is not considered deterioration since the parts, in effect, are rubbed to nominal dimensions. Approximately 0.012 cm (0.005 in) tip clearance reduction is gained through the tip grind and rub in philosophy.
The cooling flow management scheme that resulted from these analyses is shown in Table 5.2.2-VI.
Eighth stage compressor bleed air was chosen for the flight propulsion system because of a 0.19 percent efficiency advantage and 0.28 percent thrust speci- fic fuel consumption advantage. Tenth stage bleed air was selected for the integrated core/low spool to avoid design expense and additional hardware costs involved with external bleed plumbing to the eighth stage.
Blade tip and interstage airseal clearances resulting from these analyses for the integrated core/low spool are listed in Tables 5.2.2-VII and -VIII. Those for the flight propulsion system are listed in Tables 5.2.2-IX and -X.
[I rl ADDED COOLING AIR INSULATION HOLES ELIMINATED Low-Pressure Turbine Case Modifications for Flight Propul- Figure 5.2.2-22 sion System Active Clearance Control System LINE-ON-LINE f_. AT ADP Z -/ X CASE /i h- (J ill .J LL iii
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// _, I
(( r_
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(( fr I ACCEL PINCH I I / I / I IDLE SNAP ACCEL _- TAKEOFF .'=-.. SNAP DECEL--- IDLE ADP TIME Typical Low-Pressure Turbine Rotor and Case Radial Growth Figure 5.2.2-23 Caused by Centrifugal Force, Thermals, and Pressure TABLE 5.2.2-VI ACTIVE CLEARANCE CONTROL SYSTEM COOLING FLOW MANAGEMENT Integrated Core/Low Spool Flight Propulsion System Power Setti n9 Idle No cooling air No cooling air Mixed 8th & 15th Accel-Sea Level Mixed lOth & 15th Takeoff stage compressor stage compressor bl eed ai r bleed air All 8th stage Aerodynamic Design All lOth stage Point compressor bleed air compressor bleed air No cooling air Decel No cooling air TABLE 5.2.2-VII ENERGY EFFICIENT ENGINE LOW-PRESSURE TURBINE BLADE TIP CLEARANCE SUMMARY (Clearances are in cm (in)) TENTH STAGE ACTIVE CLEARANCE CONTROL SYSTEM AIR Rotor S Gap Component Rotor 2 Rotor 3 Rotor 4 Concentricity (I/2 Max.) O.OIS (0.006) 0.015 (0.006) 0.017 (0.007) 0.017 (0.007) Maneuvers 0.020 (0.008) 0.027 (O.Oll) 0.048 (0.019) 0.071 (0.028) Whirl 0.010 (0.004) 0.012 (0.005) 0.015 (0.006) 0.015 (0.006) Bearing Clearance 0.005 (0.002) 0.007 (0.003) 0.007 (0.003) 0.007 (0.003) Ovallzation & Cowl Loads 0.005 (0.002) 0.005 (0.002) 0.005 (0.002) o.oos (0.002) Total Mechanical 0.055 (0.022) 0.068 (0.027) 0.093 (0.037) 0.116 (0.046) Speed & Thermals @ ADP 0.010 (0.004) 0.020 (0.008) 0.020 (0.008) 0.027 (0.011) Speed & Thermals @ SLTO 0 0 0 0.033 (0.013) ADP Clearance 0.066 (0.026) 0.088 (0.035) 0.114 (0.045) 0.144 (0.057) SLTO Clearance 0.055 (0.022) 0.068 (0.027) 0.093 (0.037) 0.149 (O.OSg) TABLE 5.2.2-VIII ENERGY EFFICIENT ENGINE LOW-PRESSURE TURBINE INTERSTAGE SEAL CLEARANCE SUMMARY (Clearances are in cm (in)) TENTH STAGE ACTIVE CLEARJkNCE CONTROL SYSTEM AIR Rotor 2-3 Rotor 3-4 Rotor 4-5 Outer Inner Middle Middle Front Thrust Thrust Tooth Tooth Tooth Balance Balance Gap Component 0.012 (0.005) 0.012 (0.005) 0.012 (0.005) 0.012 (0.005) 0.012 (0.005) Concentricity (I/2 Max.)
0.022 (0.009) 0.022 (0.009) Maneuvers 0.022 (0.009) 0.038 (0.015) 0.058 (0.023) O.OlO (0.004) O.OlO (0.004) Wh irl 0.012 (0.005) 0.012 (0.005) 0.015 (0.006) o. 005 0.007 (0.003) 0.005 (0.002) (0.002) 0.007 (0.003) 0.OO7 (0.003) Bearing Clearance 0.005(0.002) 0.o05 (0.002) 0.005 (0.002) Ovalization & Cowl Loads 0.005 (0.002) O.O05 (0.002) 0.099 (0.039) 0.055 (0.022) 0.055 (0.022) Total Mechanical 0.060 (0.024) 0.076 (0.030) 0.002 (0.001) 0.040 (0.016) 0.038 (0.015) 0.007 (0.003) 0.045 (0.018) Speed & Thermals @ ADP 0 0 0 0 0.025 (O.OlO) Speed & Thermals @ SLTO 0.058 (0.023) O.ll6 (0.046) 0.137 (0.054) 0.063 (0.025) ADP Clearance 0.106 (0.042) 0.055 (0.022) 0.055 (0.022) SLTO Clearance 0.060 (0.024) 0.076 (0.030) 0.124 (0.049) TABLE 5.2.2-IX ENERGY EFFICIENT ENGINE LOW-PRESSURE TURBINE BLADE TIP CLEARANCE SUP_4ARY (Clearances are in cm (in)) EIGHTH STAGE ACTIVE CLEARJ_NCE CONTROL SYSTEM AIR Gap Component Rotor 2 Rotor 3 Rotor 4 Rotor 5 Concentricity (I/2 Max.) 0.015 (0.006) 0.015 (0.006) 0.017 (0.007) 0.017 (0.007) Maneuvers 0.020 (0.008) 0.027 (O.Oll) 0.048 (0.019) 0.071 (0.028) Whirl 0.010 (0.004) 0.012 (0.005) 0.015 (0.006) 0.015 (0.006) Bearing Clearance 0.005 (0.002) 0.007 (0.003) 0.007 (0.003) 0.007 (0.003) Ovalization & Cowl Loads 0.005 (0.002) 0.005 (0.002) 0.005 (0.002) 0.005 (0.002) Total Mechanical 0.055 (0.022) 0.068 (0.027) 0.093 (0.037) 0.I16 (0.046) Speed & Thermals @ ADP - - 0.017 (0.007) Speed & Thermals @ SLTO 0.068 (0.027) 0.063 (0.025) 0.066 (0.026) 0.055 (0.022) ADP Clearance 0.055 (0.022) 0.068 (0.027) 0.093 (0.037) 0.134 (0.053) SLTO Clearance 0.124 (0.049) 0.132 (0.052) 0.160 (0.063) 0.172 (0.068) TABLE 5.2.2-X ENERGY EFFICIENT ENGINE LOW-PRESSURE TURBINE INTERSTAGE SEAL CLEARANCE SUMMARY (Clearances are in cm (in)) EIGHTH STAGE ACTIVE CLEARANCE CONTROL SYSTEM AIR Rotor 2-3 Rotor 3-4 Rotor 4-5 Outer Inner Middle Middle Front Thrust Thrust Tooth Tooth Tooth Balance Balance Gap Component o.o12 (o.oo5) o.o12 (0.005) 0.012 (0.005) 0.012 {0.005) Concentricity (I/2 Max.) 0.012 (0.005) 0.022 (0.009 0.022 (0.009) Maneuvers 0.022 (0.009) 0.038 0.015) 0.058 (0.023) O.OlO (0.004 O.OlO (0.004) Whirl 0.012 (0.005 0.012 0.005) 0.015 (0.006) 0.005 (0.002 0.007 0.003) 0.007 (0.003) 0.005 (0.002) Bearing Clearance 0.007 (0.003 0.005 (0.002) 0.005 (0.002) 0.005 (0.002 0.005 0.002) Ovalization & Cowl Loads 0.005 (0.002 0.055 (0.022 0.055 (0.022) Total Mechanical 0.060 (0.024 0.076 0.030) 0.099 (0.039) O.OO2 (O.OOl 0.007 (0.003) 0.007 0.003) 0.033 (0.013) Speed & Thermals @ ADP 0.068 (0.027 0.083 (0.033) 0.055 (0.022) Speed & Thermals @ SLTO 0.]04 (0.041 0.058 (0.023) 0.083 (0.033) 0.132 (0.052) 0.063 (0.025) ADP Clearance O.12g (0.051 0.055 (0.022) 0.055 (0.022) SLTO Clearance 0.165 (0.065 0.132 (0.052) 0.182 (0.072) 5.2.3 Turbine Intermediate Case 5.2.3.1 Mechanical Designs Features The turbine intermediate case provides for gaspath transition between the high-pressure turbine exit and the inlet of the low-pressure turbine. It also provides a frame for rotor supports and the rear engine mounts. A basic aero- dynamic requirement is to duct gasflow between the turbines without separation and with minimal loss.
Figure 5.2.3-I identifies the major design features of the turbine inter- mediate case assembly. The assembly comprises the high-pressure turbine outer case, the high-pressure turbine blade tip seal, eleven structural struts that traverse the gaspath and are shielded by aerodynamic fairings, an inner ring torque box that forms an interface between the structural struts and the rear bearing support structure, and second stage turbine vane inner support, and the front and rear secondary air seal lands. Engine mount and ground handling attachment lugs are located on the outer case between the pads where tiebolts and dowels secure the bearing support structure to the case. The strut and its associated support structure serve to maintain structural integrity of the bearing support frame in the event of a turbine failure, to minimize case ovalization caused by engine mount loads, and to provide a route for oil service lines to the No. 4-5 bearing compartment. Flight propulsion system design life goals are 15,000 hours/3300 missions for the fairing and 30,000 hours/20,O00 missions for the structure. Integrated core/low spool life goal for all hardware are 50 hour's of hot time and lO00 cycles. The aerodynamic pressure loss goal for the transition duct flowpath is 1.5 percent /kPT/P T.
THRUST BALANCE SEAL LAND BEARING TORQUE COMPARTMENT RING SUPPORT SUPPORT STRUCTURE 30.7 cm WELDMENT (12.1 in) R FAIRII_ STRUCTURAL STRUT 37.3 cm (14.7 in)R THRUST BALANCE SEAL LANDS (REAR) (7.3 in) 47.7 cm (18.8 in)R 49.5 cm (19.5 in)R HEATSH( OUTER CASE 53.8 cm (21.2 in)R TIEBOLT Figure 5.2.3-I Turbine Intermediate Case 5.2.3.1.I Structural Struts The structural struts, as shown in Figure 5.2.3-2 are forged from Inconel 718 material and include drilled cooling air passages. The cooling hole exits in the integrated core/low spool strut are circular for ease of manufacturing while they are elliptical in the flight propulsion system for a lower stress concentration factor and thus greater life. Each strut is flame sprayed with an aluminum oxide insulation coating to reduce radiation heating from the fairings and local conductive heating from possible temporary contact with the fairing during high structural deflections. The struts are canted rearward from the outer case tiebolt connection to carry a rearward thrust balance load of approximately III,205 N (25,000 Ib), while reducing bending stresses to an acceptable level. Strut axial loads as well as the radial loads imposed on the bearing compartment from maneuver and imbalance are shown in Figure 5.2.3-3.
Each strut is electron beam welded to the torque ring at the inner diameter, as shown in Figure 5.2.3-4, and fastened to the outer case with a single, high-strength tiebolt. Externally removable dowels are installed on either side of each tiebolt to help the struts resist the tightening torque, absorb twisting moments at the end of the strut arising from blowoff and thermal loads, and prevent potential shear loads on the tiebolt at the strut-to-outer case joint surface. Minimum clearance between the strut and the strut fairing, due to relative motion between the two, was calculated to be 0.076 cm (0.03 in), as shown in Figure 5.2.3-2.
ALUMINUM OXIDE INSULATING COATING COOLING AIR • PASSAGES REMOVABLE /_ DOWELS / I 1 D"==--', _J(. --_ _ SPACE. // _.._""_ L _.._'_ TIEBOLT 0-,_ (0.030 inl MIN J_ 0.076 crn STRUT COOLING AIR INLET LOW-PRESSURE TURBINE OUTER CASE SECTION D-D WITH AIRFOIL FAIRING Figure 5.2.3-2 Turbine Intermediate Case Structural Strut Details MANEUVER AND IMBALANCE 6761 N 10.675 N (1520 LB) (2400 LB) \
\
32.0 CM (12.6 IN) R = 47.54 CM (18.72 IN) 11 STRUTS 315.5°C (600°F) IC/LS 1 •106 CM (0.4375 IN) DIA, MP159 551,584 Pa _80 PSII OR FPS 1.27 CM (0.500 IN) DIA, INCO 718 Figure 5.2.3-3 Bearing Compartment Axial and Radial Haneuver and Imbalance Loads .___ GUSSET I TORQUE RING " ,,, ._. ......o: _.,oo _ -_i_ Y", \.
I _' Figure 5.2.3-4 Structural Strut-to-Torque Ri ng Weldment 5.2.3.1.2 Strut Fairings The strut fairings, illustrated in Figure 5.2.3-5, are cast in eleven cir- cumferential segments that are sealed at the inner and outer platform gaps through the use of Haynes 188 cobalt alloy feather seals. To meet life requirements, SC 2000 single crystal alloy was selected for the flight pro- pulsion system strut fairings and MAR-M-509 cobalt alloy for the integrated core/low spool. Segment edges are set at the angles of the gas flow stream- lines at the inner and outer shrouds (platforms) to minimize boundary layer flow disturbances. Shroud edges are thickened to provide straight grooves for the feather seals. The inner shroud seals are retained axially by tabs at the front seal support and at the segment rear flange by a split ring. The outer shroud seals are trapped at the front end by a fairing support ring and at the rear by the second stage vane outer shroud. The fairing inner front shroud is supported by a hook arrangement on a flange which, in turn, is supported by the torque ring. This inner support arrangement reduces transient thermal variations in the radial gap between the high-pressure turbine disk rear sideplate seal lip and the front end inner shroud inner diameter. The fairing inner diameter rear shroud is supported through a hook arrangement with the second vane inner diameter support flange.
5.2.3.1.3 Outer Case and Outer Case Heatshield The outer case for the flight propulsion system is a lightweight design that features a polygonal cross section at the strut connection plane with flat plates joined at the tiebolt bosses. This has been modified for the integrated core/low spool to a continuous ring section between tiebolts in order to reduce fabrication cost and time. These rings are machined from Inconel 718 forgings. The outer case incorporates a heatshield, shown in Figure 5.2.3-6, which channels hot discharge air from the high-pressure turbine blade tip seal along the case inner wall. This flow raises the case temperature and reduces strut thermal compression experienced during engine acceleration when radia- tion from the strut fairings results in strut temperatures appreciably hotter than outer case and torque ring temperatures.
The high-pressure turbine blade tip seal exhaust air flows out of the heat- shield channel rear end into the cavity between the outer case and the outer shroud. Some of this flow leaks past the fairing segment edge seals, some through the flexible finger seals at the front and rear ends of the outer shroud, and some passes through the fairing segment airfoils into the inner cavity and its exit orifice holes. The heatshield is retained in position by bolts at each strut pad location, where flat heatshield inserts are in contact with the case. Six of these bolts, spaced at alternate pads, are inserted through bushings engaging holes both in the case pad and the heatshield to locate the heatshi el d more precisely. The Hastel loy S heatshield material is selected to provide a relatively constant gap between its outer diameter and the outer case inner diameter for the slow accelerations of the integrated core/low spool test program. The heatshield temperature is higher than that of the case, but its lower coefficient of expansion keeps its thermal growth closely matched to the outer case. For the flight engine, with fast accelera- tions the heatshield grows faster than the outer case, with the possibility of buckling. This problem can be eliminated by adding slip joints in the heat- shield between each strut location to permit circumferential expansion.
0.025 cm (0.010 in)L SEGMENT EDGE SEAL 0.254 cm (0.100 in) SECTION W-W Figure 5.2.3-5 Turbine Intermediate Case Strut Fairing HIGH-PRESSURE TURBINE BLADE TIP EXHAUST AIR 48 DIMPLES PER ROW HASTELLOY S HEATSHIELD INCONEL 718 OUTER CASE Figure 5.2.3-6 Turbine Intermediate Case Outer Case and Outer Case Heat- shield Details 5.2.3.1.4 Engine Mount Lugs The engine rear mount comprises three sets of double lugs that are located between struts on the upper half of the outer case as shown in Figure 5.2.3-7.
For the integrated core/low spool, these lugs are designed only for test stand load conditions and are electron beam welded to the outer case to save cost.
Flight propulsion system lugs would be integrally forged with adequate strength to meet the flight loads indicated in the figure. Single lugs for ground hand- ling are also welded to the outer case, one in each bottom quadrant at the aft end of the case strut ring, leaving the forward location free for instrumenta- tion bosses.
2.23 cm (0.88 in) !
(53.._iCn_nR-- __ TYPICAL LIMIT LUG REACTIONS: _L MOUNT R 1 = 75,619 N(17,000 LB) R 2 = 104,532 N(23,500 LB) UG S = 185,489 N(41,700 LB) _.,._ STRUT FLIGHT CONDITION 1.5G + THRUST FWD + W_DEJ _R1 1GDOWN + 2GSIDE+ 1 RAD/SEC YAW + .1 RAD/SEC 2 YAW MOUNT DIAGRAM Figure 5.2.3-7 Engine Rear Mount Details 5.2.3.1.5 Oil Supply and Scavenge Lines Oil supply and scavenge lines for the No. 4-5 bearing compartment are routed through the cavity formed between the structural strut and the strut fairing leading edge as shown in Figure 5.2.3-8. These lines are fully insulated to protect the oil flow from the high temperature environment within the strut fai ring.
OIL SUPPLY LINE \ STRUT
"\
/ 4-s /
,_ BEARING COMPARTMENT SCUPPER DRAIN LINE I OIL SCAVENGE FRONTVIEW LINES Figure 5.2.3-8 Routing for Nos. 4-5 Bearing Compartment Oil Supply, Scavenge and Drain Lines 5.2.3.1.6 Thrust Balance Seal Lands The front and rear thrust balance seal lands are shown in Figure 5.2.3-9. The front seal land accommodates the three knife-edge seal bolted to the rear of the high-pressure turbine disk. It is bolted to a ring spacer which, in turn, is bolted to the front face of the inner torque ring. Support for the front inner diameter strut fairing segements is provided through a hook arrangement, shown in the figure. The rear seal accommodates the single knife-edge outer seal and three knife-edge inner seals that one bolts to the front face of the first stage low-pressure turbine disk. This land is bolted to the rear face of the inner torque ring. Material for both is AMS 5671. Rubstrip material bonded to the front seal land is PWA 24-3 fiber metal. Material for the rear seal rubstrip is AMS 5536 honeycomb.
BALANCE SEAL LAND FRONT THRUST T I TORQUE RING RING SPACER ATTACHMENT HOOK REAR THRUST BALANCE SEAL LANDS Figure 5.2.3-9 Turbine Intermediate Case Front and Rear Thrust Balance Seal Lands ]48
5.2.3.2 Structural Analysis
Major concerns addressed in the structural analysis of the intermediate case
included the high temperature environment IOlO°C (1850°F) average and I176°C
(2150°F) maximum hot spot). This temperature environment can produce high
thermal stresses resulting from temperature gradients betweenthe structural
struts and inner and outer cases. The differential pressure across the inner
bearing support results in high axial thrust balance loads that can cause high
bending stresses. Rotor imbalance and maneuverloads transmitted to the
turbine intermediate case through the bearings and bearing support structure
add to the stresses. Life requirements for the primary elements of the turbine
intermediate case assembly are presented in Table 5.l-I. Structural analysis
of the primary support structure included calculation of strut stresses and
lives, and strut bolt loads. Durability of the strut fairing was analyzed in
detail because of its exposure to a high-temperature environment.
5.2.3.2.1 Primary Support Structure Stress and Life Analysis
Becauseof the complex loading to which the turbine intermediate case is sub-
jected and the complex geometry of the case itself (Figure 5.2.3-I0) with both
axial and tangential tilt in the structural struts, a three-dimensional analy-
sis was required to determine load paths, loading and deflections in the
various turbine intermediate case components. The structural deflections were
required to determine the No. 4 and 5 bearing support spring rates and for
OUTERCASE STRUTS TILTED AXIALLY LLY tG SUPPORT \ Figure 5.2.3-I 0 Turbine Intermediate Case Structure Illustrating Complex Geometry calculating clearances between the strut and aerodynamic fairing. Two separate NASTRAN models were used. A full 360-degree model, shown in Figure 5.2.3-II, was used to analyze nonaxisymetric loads. This model comprises 900 plate and bar elements. A cyclic symmetry model (modeling only a segment of the struc- ture), shown in Figure 5.2.3-12, was used to analyze axisymetric loads such as thermal loading.
Thermal and pressure loads were applied for acceleration, steady-state takeoff and deceleration conditions. The acceleration and deceleration time points chosen for analysis were those having maximum potential for ring-strut-ring thermal interference. The acceleration thermals produced the largest thermal loads.
Bearing support springrates were established for the No. 4 bearing (rear high rotor) and No. 5 bearing (rear low rotor) from the NASTRAN analysis and were used in the critical speed analyses discussed in Section 5.2.4 of this report.
The springrates shown in Figure 5.2.3-13 for both the No. 4 and 5 bearing supports satisfied critical speed requirements.
Circumferential and axial deflections resulting from the NASTRAN analysis are summarized in Figures 5.2.3-14 and 5.2.3-15. The circumferential deflection of the structural strut was found from the NASTRAN analysis to be almost entirely due to the thermal gradient between the strut and the inner and outer cases.
Approximately 40 percent of the axial deflection resulted from the thermal gradient between the strut and the inner and outer cases, and 60 percent from the large axial pressure load. Deflection analysis of the structural strut and the aerodynamic fairing determined that the minimum operating clearance between the fairing and the strut was 0.076 cm (0.030 in) at a section near the strut outer diameter, as shown in Figure 5.2.3-16. This represents the worst possible dimensional tolerance stack-up and occurs during a snap accel- eration. The minimum strut fairing clearance shown will actually be increased by fairing deflection due to gas loading. To provide an added margin of safety for the integrated core/low spool, the structural strut are coated with 0.025 cm (O.Ol in) thick ceramic to eliminate any possibility of metal-to-metal contact between the fairing and the strut.
There was some initial concern that local distortions from the eleven struc- tural struts would be "felt" by the case over the tips of the high-pressure turbine blades and thus adversely affect high-pressure turbine tip clearances.
The NASTRAN models did not extend as far foreward as the plane of the high- pressure turbine; the case was modeled only to a plane which is approximately 3.810 cm (I.5 in) aft of the plane of the high-pressure turbine. At the fore- ward edge of the model, the maximum distortions were less than 0.005 cm (0.002 in) and occurred during a normal flight propulsion system maneuver. The dis- tortions at the plane of the high-pressure turbine would be even less than 0.005 cm (0.002 in). It was judged that these distortions are reasonably small and that they would have minimal influence on high-pressure turbine tip clear- ances for the flight propulsion system. The integrated core/low spool is not subjected to maneuver loads and local distortions are even smaller than 0.005 cm (0.002 in).
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N >,, i.,..= ¢-- ,,=¢: Z "T _4 or,- Figure 5.2.3-12 NASTRAN Model used to Analyze Axisymetric Loads t14 BEARING SUPPORT SPRING RATE K - 1.7X I06LB/IN #5 BEARING SUPPORT SPRING RATE K - 5.8X 105 LB/IN Figure 5.2.3-13 Bearing Load Diagram and Spring Rates Resulting from NASTRAN Analysis 0.175cm (0.069 in)
i
+ VIEW: AXIALLY FORE'WARD Fi gure 5.2.3-I 4 Radial and Circumferential Deflections Caused by Thermal Gradients Between Strut and Cases O.175 cm A 0.053 cm (O.069 inl _1 in) / K 0.142 cm (0.056 in) 0.302 cm (O. 119in) 0.027 cm (0.011 in( 0.325 AXIALLY REARWARD n) Figure 5.2.3-15 Axial Deflections Caused by Thermal Loads and Thrust Balance Pressure Loads MOTION OF FAIRING RELATIVE TO STRUT DURING ACCEL APPROX. 0.076 cm (0.030 in)' CLEARANCE AXIAL AND TANGENTIAL DURING ACCEL. AT SECTION K-K Location of Minimum Clearance Between Structural Strut and Fi gure 5.2.3-I 6 Aerodynamic Fairi ng Loads in the structural strut, determined from the NASTRAN analysis, were used to determine stresses in the strut. These stresses, together with appropriate stress concentration factors, are summarized in Figure 5.2.3-17. Flight pro- pulsion system structural strut lives at all strut sections are greater than 20,000 cycles and have greater than 30,000 hour capability to O.l percent creep. All integrated core/low spool strut lives were found to be greater than the design requirements of lO00 cycles or 50 hours at maximum temperature to O.l percent creep.
The struts for the flight propulsion system are bolted to the outer case with 1.2 cm (0.5 in) Inconel 718 bolts, where as the integrated core/low spool uses l.ll2 cm (0.438 in) MP 159 bolts because Inconel 718 bolts are not presently available. The stresses in these tiebolts are summarized in Table 5.2.3-I. The preloads shown were set at a level to prevent separation of the strut from the outer case during "worst-case" normal loads, which occur during engine acceler- ation.
Low-pressure and high-pressure turbine blade loss analyses indicated that the structural strut and tiebolts could successfully resist the imbalance loads caused by the loss of either a high-pressure turbine blade or a fifth-stage low-pressure turbine blade.
5.2.3.2.2 Strut Fairing Durability Analysis In the durability analysis, the strut fairing was considered in the analysis an airfoil and, therefore, low-pressure turbine airfoil criteria were used.
Durability goals for the fairing are 15,000 hours service life (3300 flight missions) in the flight propulsion system and 50 hours of hot life at 28vC (84°F) day sea level takeoff conditions in the integrated core/low spool.
The analysis was conducted using a simplified NASTRAN model, where the fairing airfoil and platforms were assumed to be plate elements, as shown in Figure 5.2.3-18. The loads on the fairing considered in the analysis included: (1) those caused by the differential in pressure distribution between the gaspath and cooling air sides of the fairing, (2) those caused by differential radial growth between adjacent fairing segments, resulting from hot spot thermal loadings on one of the fairings, and (3) those caused by nonuniform tempera- ture distributions on the fairing surface during transient engine operating conditions. The later would be of concern only in the flight propulsion system, since transient operation is not planned for the integrated core/low spool.
Gaspath parameters assumed for the analysis are summarized in Table 5.2.3-II.
The assumed temperature profiles for the integrated core/low spool and flight propulsion system are shown in Figure 5.2.3-19. These profiles reflect combus- tor exit "hot-spot" conditions as defined in the component preliminary design phase. Figure 5.2.3-20 summarizes the calculated maximum steady stresses for the flight propulsion system strut airfoil fairing. All stresses shown are primarily bending stresses, with the high stress locations predictably at the fillet areas; the inner diameter leading edge fillet (location 2) being high- est. Maximum steady stresses calculated for the integrated core/low spool are shown in Figure 5.2.3-21. As with the flight propulsion system, the highest stress is at the inner diameter leading edge fillet. The material used for the integrated core/low spool strut fairing (PWA 647) is expected to undergo local creep deformation from these stresses at hot spot temperature conditions, but this will not be an inhibiting factor in the integrated core/low spool test program.
ELLIPTICAL COOLING HOLE EXIT K -18 INCO 718 BOLT _A --JK _ _J _ T- • _'____ _D -- --'_'J_ INCO 718 %_._,899 MPa (77 KSI) "__ WELD 37_,214 MPa (55 KSI) _A SECT J-J SECT K-K SECT L-L K T = 1.9 K T = 1 K T = 1 393,003 MPa (57 KSI) 689,480 MPa (100 KSI) 965,272 MPa (140 KSI) FLIGHT PROPULSION SYSTEM CIRCULAR COOLING HOLE EXIT _/ K T -- 2.5 / Cr\455,056 MPa (66 KSII F_--_L INCO 718 O MP159 BOLT '_ i_.. A _4"_L J_L'Y FOR_GED /
o WE,D
._ 324,055 MPa (47 KSl) SECT J-J SECT K-K SECT L-L K T = 1.9 K T = 1 K T = 1 310,266 MPa (45 KSI) 565,373 MPa (82 KSI) 799,796 MPa (116 KSI) INTEGRATED CORE/LOW SPOOL Propul- Structural Strut Stress and Life Summary for Flight 5.2.3-17 Figure sion System and Integrated Core/Low Spool
TABLE 5.2.3-I
STRUT BOLTSTRESS SUMMARY
Flight Propulsion System
Integrated Core/Low Spool
Inconel 718 Bolt
MP159 Bolt
Cold Assembly
Max Direct Tensile
903,218 _a (131 ksi) l ,241,064 MPa(180 ksi)
Max Shear 592,952 MPa(86 ksi)
827,376 MPa(120 ksi)
1,027,325 MPa(149 ksi) l ,434,118 MPa(208 ksi)
MaxPrincipal
Yiel d Strength 1,137,642 MPa(165 ksi) 1,654,752 MPa(240 ksi)
Ultimate Strength l ,378,960 MPa(200 ksi) 1,896,070 _a (275 ksi)
Steady State
MaxDirect Tensile
772,217 MPa(ll2 ksi)
1,006,640 MPa(146 ksi)
Max Shear
503,320 MPa(73 ksi) 579,163 MPa(84 ksi)
882,534 MPa(128 ksi)
Max Pri nci pal 1,082,483 MPa(157 ksi)
Yield Strength 999,746 MPa(145 ksi) l ,378,960 _a (200 ksi)
1,275,538 I_a (185 ksi)
Ultimate Strength 1,551,330 MPa(225 ksi)
Assembly Prel oad
6,667 kg (14,700 Ibs) 5,896 kg (13,000 Ibs)
INNER DIAMETER OUTER DIAMETER Figure 5.2.3-18 NASTRAN Model Utilized to Conduct Turbine Intermediate Case Strut Fairing Durability Analysis TABLE 5.2.3-II GAS ENVIRONMENT DEFINITION Fli ght Propulsion System* Integrated Core/Low Spool Combustor Exit Average Temperature, oc (OF) 1500 (2733) 1519 (2767) Design Pattern 0.42 Factor 0.42 Strut Falri ng 1004 (1840) I015 (185g) Average Temperature, oc (OF) Design Pattern 0.37 0.37 Factor Hot Streak I162 (2125) ll81 (2158) Temperature, oc (OF) Note: *Includes Deterioration and Tolerances INTEGRATED CORE/ (2200) LOW SPOOL 1148 - F_L (2100) o, 0 ' LM rr • 1093
2 (10001
FLIGHTPROP CC Lid SYSTEM LU I-- 1.037 (1900 I I I I 0 20 40 .... 160 80 1O0 PERCENT SPAN Figure 5.2.3-19 Hot-Spot Temperature Profiles Utilized in Strut Fairing Durabil ity Analysis CALCULATED STRESS (ELASTIC) LOCATION PRESSURE LOADS, HOT SPOT, CREEP AT 15,0OO HRS MPa (k$i) MPa (ks)) (NO RELAXATION) 1. AIRFOIL 35,163 (5.1) 8,963 (1.3) 2% 2. ID FILLET 149,617 (21.7) 36.542 (5.3) 2.5% 3. ID FILLET 47,574 (6.91 14.479 (2.1) 0.2% 4. OD FILLET 106.669 (15.5) 12,410 (1.8) 1% 5. ID PLATFORM 42,058 (6.1} 17,237 (2.5) 0.1% 5.2.3-20 Fi gure Flight Propulsion System Structural Analysi s Summary CALCULATED STRESS (ELASTIC) LOCATION PRESSURE LOADS, HOT SPOT, COMMENTS MPa (ks)) MPa (ks)) 1. AIRFOIL 36,542 (5.3) 8,963 (1.3) -- LOCAL DEFORMATION AT HOT SPOTS 2. ID FILLET 157,201 (22.8) 36,542 (5.3) 3. ID FILLET 49,642 (7.2) 14,479 (2.1) - MAX CREEP < 1.3% 4. OD FILLET 112,385 (16.3) 12,410 (1.8) - LOCAL DEFORMATION AT HOT SPOTS 5. ID PLATFORM 44,126 (6.4) 17,237 (2.5) - MAX. CREEP < 1.0%
5 rTj --
5.2.3-21 Figu re Integrated Core/Low Spool Structural Analysi s Summary The procedure used to calculate transient strains in the flight propulsion system strut fairing is illustrated in Figure 5.2.3.22. Here, the strain cycle for a typical snap acceleration-deceleration transient was calculated using a standard beam analysis. Results of this analysis were then adjusted to match strain results from the more sophisticated NASTRAN analysis. This matching was done at the most critical location (inner diameter leading-edge fillet) in order to obtain a conservative transient strain range estimate. The adjusted beam analysis procedure could then be utilized for the remainder of the analyses. The results of this analysis are summarized in Figure 5.2.3-23.
Predicted lives, based on these analyses, are summarized in Table 5.2.3-III, which indicates that durability goals can all be met with one recoating of the flight propulsion system fairing. Operation of the integrated core/l ow spool parts without a recoating does not present a failure risk, but as noted earlier, some local deformation can be expected in hot spot regions.
Sample stress and deflection contours for the strut fairing platform and air- foil sections, which were calculated as part of the NASTRAN analysis, are shown in Appendixes A and B.
0"3 F 0.2 0.56% ;TANDARD BEAM ANALYSIS \ / \ / \ / -0.3 _ NASTRAN ANALYSIS I I I I I I -0.4 METAL TEMPERATURE Illustration of Flight Propulsion System Strut Fairing Figure 5.2.3-22 Transient Strain Calculation Procedure m LOCATION STRAIN RANGE 1. AIRFOIL 0.27% 2. ID FILLET 0.56% 3. OD FILLET 0.39% 4. ID PLATFORM 0.53% Summary of Flight Propulsion System Strut Fairing Transient Figure 5.2.3-23 Strains at Maximum Stress Locations TABLE 5.2.3-111 PREDICTED LIVES FLIGHT PROPULSION SYSTEM 23,000 Hours (5100 Flights) Cracking 9300 Hours (Life 15,000 hours Coating Oxidation with one recoating) INTEGRATED CORE/LOW SPOOL 90 hours Coati ng Oxidation 5.2.3.2.3 Thrust Balance Seal Vibration Analysis Frequency response analysis of the front thrust balance seal assembly was con- ducted as part of the high-pressure turbine component design and analysis effort. Results of that analysis indicated that margins for resonance and coincidence met or exceeded commercial requirements and that operation would be flutter-free. Results for the rear thrust balance seal assembly are sum- merized in Section 5.2.1.3.2.
5.2.4 Low Rotor Critical Speed Analysis The low-pressure rotor comprises the fan, low-pressure compressor, low-pressure turbine, and the shaft connecting these components. The rotor is supported by three bearings, as illustrated in Figure 5.2.4-I. The two front bearings, located at the fan intermediate case, provide moment restraint for the over- hung fan/low-pressure compressor assembly to minimize maneuver deflections.
The low-pressure turbine is cantilevered off a rear bearing, which is damped to control the low-pressure turbine mode. For analysis purposes, the entire engine rotor and case structure was modeled as a system of beams that simula- ted the high and low rotor, the core cases, and the inner and outer fan ducts.
These were connected by a series of linear and torsional springs which duplicated the stiffness of the bearings and bearing support structures. In this way, the rotor-frame model, illustrated in Figure 5.2.4-2, was able to account for vibratory interaction between components. Critical speed design goals for the low rotor were to position the low strain energy fan and low- pressure turbine modes below minimum cruise speed, and the high strain energy shaft bending mode above redline speed to provide a mode-free running range.
During the component preliminary design effort, critical speed analysis indi- cated that the fan and low-pressure turbine modes were below minimum cruise speed and that both had low strain energy (less than 15 percent). The shaft bending mode, although indicating high strain energy, occurred well above maximum low rotor speed (67 percent margin). All low rotor bearings at this time were undamped. Results of this early analysis are summarized in Figure 5.2.4-3.
As the low rotor design evolved and components weights and structural spring- rates became better defined, strain energy in the low-pressure turbine mode increased, as shown in Figure 5.2.4-4. This raised a concern because of the sensitivity of this mode to rotor imbalance. A viscous oil-film damper was subsequently incorporated into the number 5 bearing design to control the low-pressure turbine mode response. Addition of this damper has the following effects, which are illustrated in Figure 5.2.4-5.
l .
Shaft bending strain remained at its initial level, but the bending mode margin above maximum low rotor speed dropped from 67 percent to 25 percent -- still an acceptable margin.
.
Fan mode strain energy dropped to 7 percent, while fan mode speed increased.
.
Low-pressure turbine mode speed was reduced and strain energy was reduced to 14 percent.
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0 I I I 0 2000 3000 4000 5000 6000 7000 LOW ROTOR SPEED, RPM Effects on Fan, Low-Pressure Compressor and Shaft Mode Figure 5.2.4-5 Responses Due to the Addition of an Oil-Film Damper to the Number 5 Bearing The effectiveness of this damper in controlling low-pressure turbine mode response is also shown in Figure 5.2.4-6, where the relative amplitude is based on an arbitrary amount of induced rotor imbalance.
Results of the critical speed analysis for the integrated core/low spool and flight propulsion system are summarized in Figure 5.2.4-7. No high strain energy modes are anticipated in the range of operating speeds. The fan and low-pressure turbine modes have low strain energy and occur below minimum cruise speed. The high strain energy shaft bending mode occurs above maximum low rotor speed with an acceptable margin. Resultant mode shapes for the integrated core/low spool are presented in Figures 5.2.4-8, -9 and -lO.
5.2.5 Low-Pressure Turbine Secondary Flow System 5.2.5.1 System Description The low-pressure turbine secondary flow system shown in Figure 5.2.5-I, is designed to provide maximum utilization of secondary airflow required for cooling and active clearance control. Major sources of low-pressure turbine cooling air are: l .
1.07 percent of core engine airflow from high-pressure turbine disk bore cooling; (source: twelfth stage high-pressure compressor and fifteenth stage high-pressure compressor inner bleeds) .
0.22 percent of core engine airflow from high-pressure turbine active clearance control system; (source: tenth stage high-pressure compressor/fifteenth stage high-pressure compressor mixture) o 0.35 percent of core engine airflow from turbine intermediate case strut cooling; (source: tenth stage high-pressure compressor outer bleed) o l.O0 percent of core engine airflow from low-pressure turbine case cooling and active clearance control; (source: tenth stage high-pressure compressor mixture) TOTAL = 2.64 percent airflow The rotor, case and turbine intermediate case incorporate specific features to best utilize this cooling air. They are described in the following sections.
5.2.5.2 Low-Pressure Turbine Rotor The major features of the rotor cooling air distribution system are: A pressure-balanced system independent of rim seal clearances.
0 A-frame rotor construction that provides uniform cooling flow to the disk rims.
Individually metered cooling air to disk rims.
Rim shields to thermally isolate disk-blade attachments from rim cavities.
Flow guides to minimize hot gas ingestion and recirculation.
LOW SHAFT IDLE MAX N 1 <{ mr LPT UNDAMPED #5 II LPT I I FAN DAMPED #5 i'_r _ J_ , ] I I00( 2000 30O0 4000 5000 LOW ROTOR SPEI_D,, RPM Figure 5.2.4-6 Illustration Showing Effectiveness of Damped Number 5 Bearing in Reducing Low-Pressure Turbine Mode Imbalance Response IO0
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MODE I °l , I : I I I O, 3OOO 4O0O 50(0) 6000 7000 LOW ROTOR SPEED, RPM Summary of Critical Speed Analysis Results for Integrated Fi gure 5.2.4-7 Core/Low Spool and Flight Propulsion System UJ RPM = 2409 a F- ro TURBINE O.
LOW-PRESSURE < UJ SHAFT > u F- < -J LU nr •AXIAL LENGTH STRAIN ENERGY 13.8% Mode Shape for Integrated Core/Low Spool Low-Pressure Fi gure 5.2.4-8 Turbine Mode -- Damped Number 5 Bearings UJ RPM = 3130 C3 0- < I I LU > < UJ tr STRAIN ENERGY 6.9% Figure 5.2.4-9 Mode Shape for Integrated Core/Low Spool Fan Mode -- Damped Number 5 Bearing LU RPM = 4900 a m .J (}- LOW-PRESSURE TURBINE < I I I i , UJ > m < "_"-_'_'-' '/ ' ,_ LU mr AXIAL LENGTH HIGH STRAIN ENERGY, 72.7% FREE-FREE BENDING MODE Figure 5.2.4-I 0 Mode Shape for Integrated Core/Low Spool Shaft Mode -- Damped Number 5 Bearing E o u_ u ,L I'-- f,.
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o ._1 "T ed _4 °_,,, The major elements of the pressure-balanced system are shown in Figure 5.2.5-2. Holes at the rear of the strut accurately meter 0.83 percent of core engine airflow for cooling flow into the cavity bounded by the strut, the rear thrust balance seal, and the second stage turbine disk rim. 0.44 percent of this air leaks past the thrust balance seal and is used to cool the low- pressure turbine bore. 0.26 percent of the remaining air passes through holes in the disk rim shield and serves to cool the blade attachment before being spent in the disk rim rear cavity, where it effectively reduces the tempera- ture in the cavity to less than 732°C (1350°F). The disk rim shield protects the disk rim from gaspath temperatures and incorporates a "fishmouth" seal, which prevents hot gaspath ingestion back into the inner cavity. A small amount of flow (0.13 percent of core engine airflow) is permitted to leak past this seal to maintain the pressure between the gaspath and the inner cavity.
One of the advantages of this system of flow control is that extremely well controlled rim seal clearances are not required.
A portion of the low-pressure turbine bore cooling airflow is introduced at the base of the A-frame and is metered to the third and fourth stage disk rims to cool the attachments, as shown in Figure 5.2.5-3. As the air flows radially outward toward the metering holes, a free-vortex flow field is established that efficiently maintains a positive pressure gradient across the metering holes. The split between the metered flows reflects the temperature exposure of the third and fourth stage turbine disks. As shown in the figure, spent attachment cooling air is exhausted into the downstream rim cavities and serves to reduce cavity temperatures.
For structural and life purposes, it was desirable to maintain the third and fourth stage disk 'wing' support structures at the lowest possible temperature.
This was accomplished through the use of spent disk rim attachment cooling air, disk rim shields, and disk rim flow guides, as shown in Figure 5.2.5-4.
The mount of hot gas ingested into the rim cavity is directly proportional to the amount of leakage past the inner airseal. Therefore, a three knife-edge, stepped seal was incorporated to minimize this leakage. Spent rim attachment cooling air from the upstream disk mixes with the ingested gas to effectively reduce its temperature. The knife-edge support structure functions as a temperature shield between the rim cavity and disk rim support structure.
Steady-state and transient thermal analyses of the rotor were conducted to verify the effectiveness of the cooling flow di stri buti on system in control - ling rotor structure temperature. A steady-state thermal analysis of the flight propulsion system rotor at sea level takeoff hot day +29°C (+84°F) engine operating conditions indicated a high degree of uniformity in disk and spacer temperatures as well as effective shielding of the rim cavities from gaspath temperatures. This can be seen in the finite element model results shown in Figure 5.2.5-5. Temperatures shown would be approximately +17°C (+30°F) higher for the integrated core/low spool.
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I ta_ S.- oe,- kl_ I I/SEAL SHIELDS DISK 'WING" I I I_ ROTOR BACKBONE SUPPORT STRUCTURE I i I I FROM RIM CAVITY I I I I I " _ • / 3 KNIFE EDGE SPENT COOLING AIR I I I I \ / AIR SEA,S + LOWERS FOLLOWING RIM CAVITY TEMP_ FLOW GUIDES MINIMIZE / GASPATH RIECIRCULAII1ON e RIM SHIELOS ISOLATE OISKS FROM HOT RiM CA(/ITIES Figure 5.2.5-4 Disk Rim and 'Wing' Support Structure Cooling Scheme 783 _ J-_29- _ Te70-'lf 1206)°'< _-_-I I I I '_'_ 592_J i I / c_'3"3_) 720 ._JJl 802 ....... 667 -J./)l ,C_C?dnoeal _..Ckq97.k__ ...... I 57111 =a(,o6,), 832 ' -- / _X/ _ 1,566 ii 610,J _ -'_ /" 'L_--_,]_--"-'L---- (10481 -------]_" _ 4 543 I H Ittl I_''_°,_" ,o, tto_';) _ )_ ,,o,,o, " H )JI _- 19401 525 r__'U---"---'=-_ II \X <1 I\ .t--t // / so7 \ X X/ (945) \_ J5_02 _ 494 _\ (95 _ (922) NOTE: IC/LS TEMPERATURES _._ = 17°C (30°F) HOTTER X_ I \ 491 (917) '_ 482 . "/ \ 4 7 5.2.5-5 _igure Rotor Finite Element Model Showing Temperature Distribution at Steady-State Sea Level Takeoff Hot Day Engine Operating Conditions
A transient thermal response analysis was conducted for the second and fourth
stage disks to determine the thermal gradients imposed oh the disk during
severe power excursions (i.e., 6 seconds from steady state idle to steady
state sea level takeoff power and then back to idle). The second stage disk
was selected because it is exposed to the highest temperatures. The fourth
stage disk was selected to assess the transient response of a disk within the
A-frame. Results of this analysis are presented in the temperature-time
histories shownin Figures 5.2.5-6 and-7, where the temperature gradients at
the noted disk locations are represented by the slopes of the curves.
Differences between curve shapes are an indication of the thermal "fight" going on at the noted disk locations with the exception of the snap decel- eration region the curves exhibit similar characteristics. Even in this region, the thermal gradients were within design limits. The analysis, therefore, verfied that the thermal gradients in the second and fourth stage disks were well within design limit. It also verified the benefits of the free vortex flow field generated by the A-frame in minimizing fourth stage disk thermal gradient.
5.2.5.3 Low-Pressure Turbine Case Major features of the case cooling air distribution system are: A combined cooling air and active clearance Control system approach to minimize total flow requirements.
o Thermal isolation of the structural case from the hot gaspath.
O Cool ing ai rflows metered through drilled holes - independent of part-to-part tolerances.
O Circumferential staggering of vane and airseal segment gaps to ensure minimum Ieakage.
O An active clearance control shield-to-case flow passage tailored to each stage for optimum thermal response.
O Temperature modulation of active clearance control air to provide minimum blade tip clearance.
The primary function of the case cooling air distribution system, shown schematically in Figure 5.2.5-8, is to provide case temperature modulation for blade tip active clearance control. Adjustment to the blade tip clearances are accomplished by controlling the temperatures of the inner case hooks. A double wall case construction facilitates this inasmuch as the outer case functions as a cooling air manifold from which air is taken to cool the individual hooks. Referring to Figure 5.2.5-8, high-pressure bleed air enters the case manifold and is channeled through a passage created by the inner case and a segmented baffle. This air has a dual function. It first cools the inner case structure and is then fed through metering holes in the case to cool the individual case hooks, as shown in Figure 5.2.5-9. Through this metering technique, the cooling airflow distribution becomes independent of part-to- part tolerances.
(1200) (1000) A LI.
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-_-- SNAP ACCEL SNAP DECEL -D (200) 6500 6700 6900 7100 7300 7500 7700 7900 8100 TIME, SEC.
Figure 5.2.5-6 Temperature Versus Time History at Selected Second Stage Low-Pressure Turbine Disk Locations During Severe Engine Power Excursions (1200) (1000) v U o 426 Lu (800) tr :::) I-- ///13OLT CIRCLE 1 _"(_" < _ // / BORE _1 _"9 tr 315 LLI o_ (600)
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Fi gure 5.2.5-7 Temperature Versus Time History at Selected Fourth Stage Low-Pressure Turbine Disk Locations During Severe Engine Power Excursions
Figure 5.2.5-8 Low-Pressure Turbine Case Cooling Flow Distribution System
CONTROLLED SHIELD GAP (VIA VARIOUS STAND-OFF THICKNESS) "TUNE" ACC FILM COEFFICIENTS CASE METERING HOLE " INNER CASE BAFFLE Figure 5.2.5-9 Details of Case Hook Cooling Configuration Temperature variation in the inner case and case hooks is accomplished in two ways. First, by mixing bleed air from different stages in the high-pressure compressor and second, by controlling the flow gap between the inner case and baffle. For the integrated core/low spool, a mixture of tenth and fifteenth stage high-pressure compressor bleed air was found acceptable. For the flight propulsion system, a mixture of eighth stage and fifteenth stage air was selected, as discussed in Section 5.2.2.4 of this report. By adjusting the inner case-to-baffle gap, the cooling air film coefficient can be regulated to provide the optimum case temperature at each stage. The range of film coeffi- cients available for this "tailoring" process is shown in Figure 5.2.5-I0. The distribution of case cooling airflow for both the integrated core/low spool and the flight propulsion system is shown in Figure 5.2.5-II. The flight propulsion system uses less total cooling airflow than the integrated core/low spool because it has two less case hook metering holes. The reasons for this are described in Section 5.2.2.4 of this report. To maintain optimum control of the outer airseal radial position, the inner case and inner case hooks must be thermally isolated from gaspath temperatures as much as possible. This thermal isolation is accomplished by the use of the design features illustrated in Figure 5.2.5-12. First, the gaps between segments of the outer airseals and vane platforms are staggered so that the airseal segment gap does not line up with the adjacent vane platform gap. This effectively reduces a potential leak path from the gaspath to the case. Second, the case is further isolated from the gaspath wall through the use of heatshields at all outer airseal and vane platform locations indicated in the figure. These heatshields act as radiation barriers as well as diverters of any hot gaspath air that might leak into the cavities. Finally, the cooling air manifolds provided at each case hook loca- tion, while directing O.lO percent of core engine flow for cooling flow through the case hooks and vane feet, act as an additional shield between the gaspath and the inner case structure.
i IO0 LL WACC ' 80 .J 40 ¢.
,< I I I I I I o I I I I I I I I I 0.254 cm 0.304 cm 0.355 cm Oom 0.050 om 0.101 cm 0.152 cm 0,203 cm 10.100 in) 10.120 in) (0.1401hi lo.o _I qo.o2o k_l Io.o4o _ io.o6o in) 1o.o8o _ CASE O.D.ACC GAP. cm (in) Figure 5.2.5-I0 Cooling Flow Film Coefficient Variation as a Function of Inner Case-To-Baffle Gap Width / O%PT3 FROM HPC ,OTN,,ST. STAGES INTEGRATED CORE/ t LOW- SPOOL L,
-.---'--_;_ F"-_._ I
FROM HPC 8TH/1 STH STAGES FUGHT PROPULSION :L_ ' _.LL+, SYSTEM Figure 5.2.5-I 1 Comparison of Integrated Core/Low Spool and Flight Propul- sion System Low-Pressure Turbine Case Cooling Airflow Distributions SEGMENT GAPS SPACED CIRCUMFERENTIALLY /-..
SEGMENTED OUTER _ (I : I ' IA_' AIRSEAL J _"-.,L"_ r .... _ _ | LCOOUNG A,R MANIFOLD COOUN_'A,R "_'-___ ( / II I *_"" MANIFOLD _ED HEATSHIELD _. I_E_ ___ L Figure 5.2.5-I 2 Design Features Utilized to Thermally Isolate the Inner Case and Case Hooks from the Hot Gaspath A steady-state thermal analysis of the flight propulsion system case at sea level takeoff operating conditions indicated good isolation of vane-to-case hooks from the hot gaspath flow. This can be seen in the finite element model results shown in Figures 5.2.5-13 and 5.2.5-14. Leading edge hooks are a nominal 26°C (80°F) cooler than trailing edge hooks because of the greater degree of gaspath isolation possible at the front of each vane. Temperatures shown would be approximately 17°C (30°F) higher for the integrated core/low spool.
A transient thermal response analysis of the case was conducted to assess the thermal behavior of the second stage vane and fourth stage vane outer airseal case hooks during snap acceleration/deceleration operation. The second and fourth stage vane locations were selected in order to be compatible with the locations selected for the rotor thermal analysis. The initial analysis was conducted without the use of active clearance control. Results of this analysis are shown in Figures 5.2.5-15 and -16. These figures show the degree to which the case hooks were effectively shielded from hot gaspath tempera- tures and also indicate the desired comparable thermal response at the front and rear hook locations. The second analysis was conducted to determine the degree of outer airseal radial deflection possible by exercising the active clearance control system during the same engine transient operation. Results of this analysis are shown in Figures 5.2.5-17 and -18. The deflection resulting from use of fifteenth stage high-pressure compressor bleed air only defines the maximum deflection possible, whereas the deflections resulting from use of eighth or tenth stage high-pressure compressor bleed air only define the minimum deflection possible. Deflections within these ranges can be accomplished by mixing the bleed air. It should be noted that use of all fifteenth stage bleed air is not possible at sea level takeoff conditons since this would elevate the temperatures at the case hooks above the capability of the Inconel 718 case material. Use of eighth stage bleed air provides a small improvement in deflection range at the second stage vane location and a significantly larger improvement at the fourth stage vane location.
5.2.5.4 Turbine Intermediate Case Major features of the turbine intermediate case cooling air distribution system shown in Figure 5.2.5-19. These include: Feather seals that bridge the inner and outer diameter chordal gaps between strut fairing segments and limit leakage to 0.25 percent of core engine flow.
Struts cooled internally by 0.35 percent of core engine flow from the tenth stage high-pressure compressor outer diameter bleed. (Because of the large surface area and high film coefficients involved, it is not practical to cool the strut fairing.)
An outer case with temperature controlled by spent air from the high-pressure turbine active clearance control system. The outer case has been thermally matched with the strut inner torque box with no additional cooling air required.
e- !
o °r- °r,- _ 4-_ c- 0r-- ,r.- "_o o e-_..
• r'- 0 V_F- r..- r-,- o m ILl "_" • r'- {_ _n 0 4_.,- c,') L_ "-s °_...
i, "_ 378 _ -515 _)_11 _lb (713)_ _'_(959) t--"f'--I'_, 564"-I 762 408 _(971 )"_-'_i'--_ 1126)al I- _774 _(768)_ 450 _",,,,J'_-_,.583"T"_"_ 1(1426L NOTE: ,C/LS TEMPERATURES "_-_'_) _ "_5064241 _ i\ _ " 1 7°C (30°F)HOTTER _,,_(7425 _(1590531 L._ TEMPERATURES °C (°F) L",_'-_ ""_'_/"_ Thi rd Vane Figure 5.2.5-14 Temperature Distribution at Low-Pressure Turbine Attachment at Steady-State Sea Level Takeoff Hot Day Operati ng Conditi ons 1,204 - (2200) m GAS PATH (1800)
t
ii O
t
760 _ o (1400) i.u"
, /
n-" FRONT HOOK__ I-- REAR HOOK ,_ 537 n,- (1000) iii I:1.
LU e_CO_LI N -- "_ G AIR/t I-- 315 m (600) O b D (/) (I) (200) --__;NAP _)ECEI]_ I I I I'__---- _NAP _CCE_ 0 I I I 6700 6900 71 O0 7300 7500 7700 7900 8100 TIME, SEC.
Figure 5.2.5-15 Temperature Versus Time History at Selected Second Stage Low-Pressure Turbine Vane/Outer Airseal Locations During Severe Engine Power Excursions (1400) r-- -" /"
I
GAS PATH
E
f
648 I A (1200) LL o m CJ O ,,._ 11530701 - FRONT HOOK ___ - _ HOOK l- < LU
I
a. 426 (3 m :; (800) U) UJ t- (6OO) n 2O4 SNAP DECEL-- u} SNAP ACCEL (400) ' I I I I I I I I I I I I 6500 71 O0 7300 7500 7700 7900 6700 6900 8100 TIME, SEC.
Figure 5.2.5-16 Temperature Versus Time Hi story at Selected Fourth Stage Low-Pressure Turbine Vane/Outer Airseal Locations During Severe Engine Power Excursions
0.508 I
0.114cm 15TH STAGE ACC (0.200) (0.045 in)
I
IU.
MAX. ALLOWABLE DEFLECTION I 0.406 I (0.160) E J J o p- CJ Iii .-J LL 8TH STAGE ACC Iii a .J
I I
<
I
(0.080) r_ < mr
I I
I
0.101 I (0.040) ._ O I U) 8001 8003 6700 6900 71 O0 7300 7500 7700 TIME, SE_.
Figure 5.2.5-17 Low-Pressure Turbine Second Stage Outer Airseal Radial Deflection with Active Clearance Control in Operation 0.508 I ,sT.ST*OE*CC L I (0.20OI / _n r_o L,h _ I _ '_ 0EFLEC'nON 0.406 (0.160) i'.
E u
/ ./'-__--+---uT',,>-,L_.,,__-_I.,- .-'
0.304 _o F- (0.120) .J _L H STAGE ACC Ill 0.203 10.0801 < m
I
(3 0.101 (0.0401
__1 _1 I _1
+1 I I
I I l I I I I I 6700 6900 7100 7300 7500 7700 8001 8003 TIME, SEC.
Low-Pressure Turbine Fourth Stage Outer Airseal Radial Figure 5.2.5-18 Deflection with Active Clearance Control in Operation RATE FLOW l% WAI ) HPT DISK 8ORE • _ C0OUNG AIR /_ LEAKAGE FROM HPT ACC EXIT FROM 10TH HPC 00 BLEED Figure 5.2.5-19 Turbine Intermediate Case Secondary Flow System
The temperature distribution in the integrated core/low spool turbine inter-
mediate case structure was assessed at sea level takeoff engine operating
conditions. Results of this assessmentare shown in Figure 5.2.5-20 and
indicate that the load carrying struts, inner diameter and outer diameter ring
steady-state temperatures are all maintained at, or below, 551°C (I025:F).
(900) ( 1 OO3) (890) (1775) (1730) (1715) (967) NOTE: STRUT TEMPERATURES 17°C (30°F) COOLER FOR FPS CONDITIONS oc (°F) / Fi gure 5.2.5-20 Integrated Core/Low Spool Turbine Intermediate Case Tempera- ture Map at Sea Level Takeoff Engine Operating Conditions The torque box-to-strut case structure represents a classic ring-strut-ring thermal stress problem. To minimize thermal stresses in this support structure during temperature excursions, it is desirable to thermally match the inner ring torque box and the outer case. This was accomplished by ducting 0.25 percent of core engine airflow from the high-pressure turbine active clearance control system along the outer case wall, as shown in Figure 5.2.5-19, and not cooling the torque box structure. The resultant time-temperature history in Figure 5.2.5-21 shows that the thermal response rates for the two rings were equalized, indicating that the desired thermal matching was achieved. The faster responding strut is permitted to expand by circumferential deflection.
of the rings relative to each other, utilizing the circumferential tilt of the struts to advantage (see Section 5.2.3.2.1 of the report for further discussion).
(1000) STRUT (900) (800) _ASE A U.
O v 371 (700) TORQUEBOX O w" rr D (600) < mr W O- LU F- (500) (400)
(300) I I I I
TIME, SEC.
Figure 5.2.5-21 Temperature-Time History for Turbine Intermediate Case Torque Box-To-Strut Case Support Structure During Tempera- ture Excursions 5.2.6 Turbine System Weight Summary Preliminary weight analyses were conducted for the turbine intermediate case and low-pressure turbine as configured for the integrated core/low spool.
Results of these analyses are presented in Table 5.2.6-I. A detailed weight assessment will not be performed until the final flight propulsion system preliminary design update.
TABLE 5.2.6- I PRELIMINARY WEIGHT SUMMARY FOR INTEGRATED CORE/LOW SPOOL TURBINE INTERMEDIATE CASE AND LOW-PRESSURE IURBINE COMPONENT WEIGHT
Kg--TT )
Turbine Intermediate Case 139.4 (307) Low-Pressure Turbine Component Rotor 381.8 (841) Case and Vanes 302.8 (667) Exhaust Case 323.7 (713) Shaft 114.9 (253) TOTAL 1262.6 (2781)
SECTION 6.0
SECTION 6.0 CONCLUDING REMARKS The intermediate case and low-pressure turbine designs represent a positive step towards improving engine fuel efficiency on a component level through advances in aerodynamics, structure and materials. The turbine component, in particular, introduces several technical features to enhance performance and durability, while offering a substantial reduction in weight. Also, the designs are sensitive to the variables of cost and maintainability, which are fundamental requirements of the commercial market.
Overall, much of the technology is universally applicable to any engine of the next generation since it reflects refinements in aerodynamic and thermodynamic design techniques. Included are concepts for reducing gaspath leakage, active clearance control and high strength/high temperature capability materials.
Completion of the intermediate case, low-pressure turbine and turbine exit guide vane designs is a major accomplishment under Task 2 of the Energy Efficient Engine Program. The supporting technology programs have been successful in providing pertinent design information as well as substantiating the benefits of various advanced concepts. Technology verification on a system level is the next step, and this will be accomplished during the scheduled integrated core/low spool test program.
i_¸ APPENDIXES I_ !%
APPENDIX A
APPENDIX A NASTRAN stress results for turbine intermediate case strut fairing.
A-I : Platform and airfoil stress contours caused by pressure loads A-2: Platform and airfoil stress contours caused by hot spot thermals A-3: Platform and airfoil stress contours caused by transient thermals
APPENDIX A
APPENDIX A NASTRAN STRESS ANALYSIS
APPENDIX A
APPENDIX A
NASTRAN stress results for turbine intermediate case strut fairing.
A-I: Platform and airfoil stresss contours caused by pressure loads
A-2: Platform and airfoil stress contours caused by hot spot thermals
A-3: Platform and airfoil stress contours caused by transient thermals
A-l
PLATFORM AND AIRFOIL STRESS CONTOURS CAUSED BY PRESSURE LOADS
PLATFORM AND AIRFOIL STRESS CONTOURS CAUSED BY PRESSURE LOADS
SYMBOL VALUE 1 - 7.142109E + 03 2 -3.817291E+03 3 -4.924727E+02 4 2.832346E + 03 5 6.157164E + 03 6 9.481920E+03 7 1.280680E+04 8 1.613161E+04 9 1.945643E+04 10 2.278126E+04
Inner Diameter Platform Stress Results - Pressure Load
(Integrated Core/Low Spool )
PLATFORM AND AIRFOIL STRESS CONTOURS CAUSED BY PRESSURE LOADS
SYMBOL VALUE 1 -1.633836E+04 6 2 -1,349928E+04 3 -1.066021E+04 4 -7.821133E+03 5 -4,982059E+03 6 -2.142984E +03 7 6.960898E+02 8 3.535164E+03 9 6.374238E+03 10 9.213312E+03
Outer Diameter Platform Stress Results - Pressure Load
(Integrated Core/Low Spool
PLATFORM AND AIRFOIL STRESS CONTOURS CAUSED BY PRESSURE LOADS
SYMBOL VALUE 1 - 2.172700E + 04 2 -1,898413E+04 -1.624126E+04 -1.349839E+04 5 -1.075552E+04 6 -8.012645E+03 7 - 5.269773E + 03 8 -2.526904E+03 9 2.159648E + 02 10 2.958826E+03
Airfoil Suction Wall Stress Results - Pressure Load
(Integrated Core/Low Spool)
PLATFORM AND AIRFOIL STRESS CONTOURS CAUSED BY PRESSURE LOADS
SYMBOL VALUE 1 - 9.281922E + 03 2 - 7.454176E + 03 3 - 5.626430E + 03 4 -3.798684E+03 5 -1.970937E+03 6 -1.431914E+02 7 1.684555E+03 8 3.512301E+03 9 5.340047E+03 10 7.167793E + 03
Airfoil Pressure Wall Stress Results - Pressure Load
(Integrated Core/Low Spool)
A-2
PLATFORM AND AIRFOIL STRESS CONTOURS CAUSED BY HOT _OT THER_LS
PLATFORM AND AIRFOIL STRESS CONTOURS CAUSED BY HOT SPOT THERMALS
SYMBOL VALUE 1 -2.224739E+03 2 -1.192309E+03 3 -1.598794E+02 4 8.725503E+02 5 1.904980E+03 B 2.937410E+ 03 7 3.969839E+03 8 5.002266E+03 9 6.034695E+03 10 7.067129E+03
Inner Diameter Platform Stress Results - Hot Spot Thermals
(Integrated Core/Low Spool)
PLATFORM AND AIRFOIL STRESS CONTOURS CAUSED BY HOT SPOT THERMALS
/
SYMBOL VALUE 1 - 6.139207E + 03 2 - 5.222969E + 03 3 - 4.306730E +03 4 - 3.390494E + 03 5 - 2.474258E + 03
/
6 - 1.558021E+03 7 - 6,417852E +02 8 2.744512E + 02 9 1.190687E + 03 10 2.106922E+03
Outer Diameter Platform Stress Results - Hot Spot Thermals
(Integrated Core/Low Spool )
PLATFORM AND AIRFOIL STRESS CONTOURS CAUSED BY HOT SPOT THERMALS
SYMBOL VALUE 1 - 5.293922E+03 2 -4.625516E +03 3 - 3.957111E+03 4 - 3.288706E + 03 5 - 2.620301E+03 6 - 1.951896E÷03 7 - 1.283492E + 03 8 - 6.150869E + 02 9 5.331787E+01 10 7.217253E +02
Airfoil Suction Wall Stress Results - Hot Spot Thermals
(Integrated Core/Low Spool)
A-3
PLATFORM AND AIRFOIL STRESS CONTOURS CAUSED BY T_NSIENT THER_LS
PLATFORM AND AIRFOIL STRESS CONTOURS CAUSED BY TRANSIENT THERMALS
8 SYMBOL VALUE 1 - 2.548554E + 03 2 - 2.150909E + 03 3 - 1.753265E+03 4 - 1.355620E+03 5 -9.579751E+02 6 - 5.603303E + 02 7 - 1.626855E+02 8 2.349592E +02 9 6.326040E + 02 10 1,030250E + 03
Airfoil Pressure Wall Stress Results - Hot Spot Thermals
(Integrated Core/Low Spool )
PLATFORM AND AIRFOIL STRESS CONTOURS CAUSED BY TRANSIENT THERMALS
SYMBOL VALUE 1 -5.807805E+03 2 -3.497070E+02 3 5.108391E+03 4 1.056649E+04 5 1.602459E+04 6 2.148268E+04 7 2.694078E+04 8 3.239888E+04 9 3.785698E+04 10 4.331508E+04
Inner Diameter Platform Stress Results - Transient Thermal
(Integrated Core/Low Spool )
PLATFORM AND AIRFOIL STRESS CONTOURS CAUSED BY TRANSIENT THERMALS
SYMBOL VALUE 1 -3.584471E+03 2 -1.034071E+03 3 1.516329E+03 4 4.066729E+03 5 6.617129E+03 6 9.167527E+03 7 1.171793E+04 8 1.426832E+04 9 1.681872E+04 10 1.936913E+04
Outer Diameter Platform Stress Results - Transient Thermal
(Integrated Core/Low Spool)
PLATFORM AND AIRFOIL STRESS CONTOURS CAUSED BY TRANSIENT THERMALS
SYMBOL - 1.2_ +04 "_ - 9.607574E + 03 - 6.794836E + 03 4 - 3.982099E + 03 - 1.169361E+03 1.643376E + 03 4.456113E+03 8 7.268848E + 03 1.008158E + 04 1.289432E + 04
Airfoil Suction Wall Stress Results - Transient Thermal
(Integrated Core/Low Spool)
PLATFORM AND AIRFOIL STRESS CONTOURS CAUSED BY TRANSIENT THERMALS
SYMBOL VALUE 1 - 3.177446E + 04 2 - 2.764677E + 04 3 - 2.351907E + 04 4 - 1.939137E+04 5 - 1.526367E+04 6 - 1.113597E+04 7 - 7.008270E + 03 8 - 2.880570E +03 9 1.247129E + 03 10 5.374855E + 03
Airfoil
Pressure Wall Stress Results - Transient Thermal
(Integrated Core/Low Spool )
APPENDIX B
APPENDIX B
TURBINE INTERMEDIATE CASE STRUT FAIRINGPLATFORM
ANDAIRFOILDEFLECTION CONTOURS CAUSED BYPRESSURE LOADS
TURBINE INTERMEDIATE CASE STRUT FAIRING PLATFORM
AND AIRFOIL DEFLECTION CONTOURS CAUSED BY PRESSURE LOADS
SYMBOL VALUE 1 -7.503985E-04 2 -2.536122E-04 3 2.431741E -04 4 7.399605E-04 5 1.236747E-03 6 1.733533E-03 7 2.230319E-03 8 2.727106E -03 9 3.223892E -03 10 3.720682E -03
Inner Diameter Platform Radial Deflections - Pressure Load
(Integrated Core/Low Spool )
TURBINE INTERMEDIATE CASE STRUT FAIRING PLATFORM
AND AIRFOIL DEFLECTION CONTOURS CAUSED BY PRESSURE LOADS
SYMBOL VALUE 1 -3.091799E-02 2 -2.786454E-02 3 -2.481109E -02 4 -2.175764E -02 5 -1.870418E-02 6 -1.565073E -02 7 -1.259728E -02 8 -9.543825E-03 9 -6.490372E.-03 10 -3.436932E--03
Outer Diameter Platform Radial Deflections - Pressure Load
(Integrated Core/Low Spool )
TURBINE INTERMEDIATE CASE STRUT FAIRING PLATFORM
AND AIRFOIL DEFLECTION CONTOURS CAUSED BY PRESSURE LOADS
_10 SYMBOL VALUE -2.755137E-02 -2.407670E-02 -2.060202E-02 4 -1.712735E-02 -1.365267E-02 -1.017800E-02 -6.703321E -03 -3.228648E -03 2.460242E -04 3.720682E -03
Airfoil Suction Wall Transverse Deflections - Pressure Load
(Integrated Core/Low Spool )
i)_i:Lrr..F., _,.""
2 - _'098_g E O_ r° -I ,B424% OE _u_ g _,0_' 2041E -OS _0
pressure Load
_a _ "T_ '_ns_erse Def%ecti°ns "
_%_tegrat, ed coreltOm SP°°il
_irfOi_ presSU_'e
APPENDIX C
APPENDIX C TURBINE INTERMEDIATE CASE STRUT FAIRING AIRFOIL COORDINATES STRUT FAIRING 400 SERI ES (Section I) SUCTION PRESSURE SIDE SIDE X/BX Y/BX Y/BX l .00000 l .49436 l .49436 0.99000 l .50131 l .47229 0.98000 l .49121 l .45161 0.97000 l .47953 l .43120 0.96000 l .46802 l .41086 0.94000 l .44469 l .37039 0.92n00 l .421 21 l .33022 0.88000 l .37374 l .25066 0.84000 l .32558 l .17166 0.80000 l .27682 l .09275 0.76000 l .22760 l .O1407 0.72000 l. 17804 0.93597 0.68000 l .12812 0.8 5870 0.64000 l.07774 0.78248 0.60000 l .02677 O. 70747 0.56000 0.97514 0.63384 O. 52000 0.92274 0.56172 0.48000 0.86947 0.49125 0.44000 0.81524 0.42254 0.40000 0.75998 0.35574 O. 36000 0.70350 0.29104 0.32000 0.64558 0.2287 0 O. 28000 0.58590 O. 16909 0.24000 0.52391 O.11288 0.2 0000 0.45907 O. 06048 0.16000 0.39076 O. O1330 0.12000 0.31792 -0.02835 0.08000 0.23837 -0.05686 0.06000 O. 19460 -0.06494 0.04000 0.14702 -0.06583 O. 03000 0.12077 -0.06286 0.02000 0.09091 -0.05704 O.OlO00 0.05806 -0.04654 0.0 0.0 0.0 STRUT FAIRING 400 SERIES (Section 2) SUCTION PRESSURE SIDE SIDE XIBX Y/BX Y/BX I.00000 l .lOl08 1.10108 0.99O00 l .I0790 1.08457 0.98000 l .I0073 1.0701 5 0.97000 l .09321 1.05592 O. 96000 l .08568 1.04170 0.94000 l .07043 1.01338 0.92000 l .05502 O. 98520 0.88000 l .02369 0.92923 0.84000 0.99167 0.87347 0.80000 0.95902 0.81753 0.76000 0.92588 0.76136 0.72000 0.89235 0.70524 O.68000 0.85842 0.64935 0.64000 0.82399 0.59389 0.60000 0.78897 O. 53898 0.56000 0.75327 0.4847 4 O. 52000 0.71682 0.43131 0.48000 0.67954 0.37898 0.44000 O.64136 O. 3 2744 0.40000 0.60216 0.27693 0.36000 0.56184 O. 2 2758 0.32000 0.52025 0.17971 0.28000 0.47699 O. 13378 0.24000 0.43152 0.09022 O.20000 0.38331 0.04961 O. 16000 0.33179 O. O1316 O. 12000 0.27586 -0.01819 0.08000 0.21359 -0.03993 0.06000 O. 17900 -0.04759 0.04000 0.14022 -0.04955 0.03000 O.11861 -0.04713 0.02000 0.09506 -0.04199 O. O1000 0.06674 -0.03271 0.0 0.0 -0.0 STRUT FAIRING 400 SERIES (Section 3) SUCTION PRESSU RE SIDE SIDE XIBX YIBX Y/BX l.O0000 0.88068 0.88068 0.99000 0.88727 0.86652 O. 98000 0.88177 O. 8 5490 0.97000 0.87625 0.84343 O. 96000 0.87065 0.831 96 0.94000 0.85932 0.80914 0.92000 0.84782 0.78643 0.88000 0.82435 0.74136 0.84000 0.80021 0.69649 0.80000 0.77547 0.65145 0.7 6000 0.75028 0.6061 9 0.72000 0.72473 0.56091 0.68000 O. 6 9880 0.51578 0.64000 0.67238 0.47096 0.60000 0.64540 0.42656 0.56000 0.61777 0.38268 O. 5 2000 O. 58945 O. 33942 0.48000 0.56034 0.29700 0.44000 O. 53039 0.25537 0.40000 0.49950 0.21449 O. 3 6000 0.46758 O. 17453 0.32000 0.43445 0.13575 O. 28000 0.39974 0.09852 0.24000 0.36926 0.06328 0.2 0000 0.32370 0.0 3044 O.16000 0.28144 O.OOl 15 O.12000 0.23519 -0.02408 0.08000 0.18335 -0.04066 0.06000 0.15432 -0.04537 O.04000 0.12156 -0.04700 0.03000 O. 10341 -0.04471 0.02000 0.08349 -0.03971 O. O1000 O. 05988 -0.03059 0.0 0.0 0.0 STRUT FAIRING 400 SERIES (Section 4) SUCTION PRESSURE SIDE SIDE X/BX Y/BX Y/BX I.00000 0.73531 0.73531 0.99000 0.74174 0.72222 O. 98000 0.73716 0.71188 0.97000 0.73265 0.70170 O. 96000 0.72805 0.69153 0.94000 0.71874 0.67132 0.92000 0.70929 0.65128 0.88000 0.68993 0.61164 O.84000 0.66999 0.57236 0.80000 0.64955 0.53310 0.76000 0.62876 0.49379 0.72000 0.60771 0.45462 O. 68000 0.58635 0.41577 0.64000 0.56459 0.37734 O. 60000 0.54234 0.33945 O.56000 0.51954 0.30217 O.52000 0.49614 0.26560 0.48000 0.47207 0.22988 0.44000 0.44726 O. 19505 0.40000 0.42164 0.16115 O. 36000 0.39511 O. 12832 0.32000 0.36748 0.09677 O.28000 O.33843 0.06680 0.24000 0.30764 0.03878 O. 20000 0.27477 O. O1300 O. 16000 0.23935 -0.00965 0.12000 0.20047 -0.02865 0.08000 0.15683 -0.04074 0.06000 0.13213 -0.04318 0.04000 O.10434 -0.04340 O.03000 0.08900 -0.04131 0.02000 0.07163 -0.03646 O. O1000 0.04930 -0.02750 0.0 0.0 0.0
APPENDIX D
APPENDIX D LOW-PRESSURE TURBINE VANE AND BLADE AIRFOIL COORDINATES SECOND-STAGE VANE COORDINATES ROOT SECTION (HOT RADIUS = 16.39000) PERCEN T X X Y TOP Y BOT 0.0 0.00209 2.29910 2.29910 O.Ol 0 0.01781 2.31407 2.25376 0.020 0.03354 2. 30879 2.22621 0.030 0.04926 2.29923 2.20238 0.040 0.06498 2. 28774 2. 17851 0.050 0.08071 2. 27433 2.15455 0.060 0.09643 2. 26020 2.13065 0.070 O.l l 21 6 2.24533 2.10674 0.080 O. 12788 2. 23046 2. 08285 0.090 O.l 4360 2.21553 2.05890 O. lO0 O. 15933 2. 20054 2.03505 0.125 O.19863 2.16276 l .97535 O.150 0.23794 2.12453 l .91 567 0.175 0.27725 2.08583 l .85607 0.200 0.31656 2.04661 1.79639 0.225 0.35587 2.00683 l .73680 0.250 0.3951 8 l .96644 l .67723 0.275 0.43449 l .92538 l. 61772 0.300 0.47380 1.8835B 1.55824 0.325 0.51310 1.84097 1.49879 0.350 O. 55241 l .7 9747 l .43937 0.37 5 0.59172 l .75300 l .38000 0.400 0.631 03 l .70744 l .32066 0.425 0.67034 l .66068 l .26136 0.450 0.70965 1.61257 1.20210 0.47 5 0.74896 l .56296 l. 14289 0.500 0.78827 l .51 167 l .08370 0.525 0.82757 l .45850 l .02463 O. 550 0.86688 l .40320 O. 96556 0.575 0.90619 1.34548 0.90649 0.600 0.94550 l .28508 0.84756 0.625 0.98481 1.22178 0.78866 0.650 l .0241 2 l .15545 0.72977 0.67 5 l .06343 l .08607 0.67104 0.700 l .10274 l .Of 373 0.61 235 0.725 I .14204 0.93862 0.55369 0.750 l. 181 35 0.861 Ol 0.49515 0.77 5 l .22066 0.78117 0.43676 0.800 l .25997 0.69940 0.37852 0.825 l .29928 O. 61597 0.32042 0.850 l .33859 0.53114 0.26245 0.87 5 I .37790 0.4451 3 0.20471 0.900 l .41 721 0.3581 l O.14734 0.91 0 l .43293 0.32307 O. 12452 0.920 l .44865 O. 28789 O. lOl 78 0.930 l .46438 0.25259 0.07917 0.940 l .48010 0.21 718 0.05669 0.950 l .49582 O. 18167 0.03441 0.960 l .51 155 0.14607 O.Ol 247 0.970 l .52727 O. l I037 -0.00902 0.980 l .54299 O. 07460 -0.02337 0.990 l .55872 0.03873 -0.02382 I .000 l .57444 '0.00202 -0.00202 SECOND-STAGE VANE COORDINATES MEAN SECTION (HOT RADIUS : 18.11501) PERCENT X X Y TOP Y BOT 0.0 0.00652 2.36065 2.36065 0.010 0.02218 2.37622 2.31705 2.29072 0.020 0.03783 2.37197 2.26651 0.030 0.05349 2.36463 0.040 0.06915 2.35471 2.24219 2.21787 0.050 0.08481 2.34311 2.19355 0.060 0.10047 2.33083 2.16912 0.070 0.11613 2.31861 2.14472 0.080 0.13178 2.30620 0.090 0.14744 2.29371 2.12046 0.I00 0.16310 2.28111 2.09612 0.125 0.20225 2.24912 2.03537 0.150 0.24139 2.21642 l .97443 0.175 0.28054 2.18293 l .91366 l .85280 0.200 0.31968 2.14862 0.225 0.35883 2.11341 l .79198 0.250 0.39797 2.07725 l .73125 0.275 0.43712 2.04005 l .67052 0.300 0.47626 2.00174 l .60982 0.325 0.51541 1.96225 l .54914 0.350 0.55455 1 .92146 l .48848 0.3?5 0.59370 1.87928 l .42?84 0.400 0.63284 1 .83561 l .36723 0.425 0.67199 1 .79031 l .30664 0,450 0.71113 1.74327 l .24607 0.475 0.75028 1 .69434 l .18553 0.500 0.78942 1 .64339 l .12502 0.525 0.82857 1 .59024 l .06458 0.550 0.86772 1 .53473 l .00418 0.575 0.90686 1 .47670 0.94368 0.600 O. 94601 1 .41 595 0.88329 0.625 0.98515 1 .35232 0.82300 0.650 1.02430 1 .28566 0.76258 0.675 1.06344 1 .21580 0.70231 0.700 1.10259 1 .14263 0.64216 0.725 1.14173 1.06605 0.58200 0.750 1.18088 0.98599 0.52185 0.775 1 .22002 0.90241 0.46171 0.800 1.25917 0.81529 0.40164 0.34171 0.825 1 .29831 0.72468 0.28200 0.850 1 .33746 0.63065 0.22242 0.875 1 .37660 0.53329 0.900 1.41575 0.43273 0.16290 0.910 1 .43141 0.39164 0.13918 0.920 1 .44707 0.35006 0.I1550 0.930 1 ,46272 0.30802 0.09188 0.06838 0.940 1.47838 0.26552 0.04503 0.950 1 .49404 0.22256 0.02179 0.960 1 .50970 0.17918 0.00088 0.970 1 .52536 0.13538 -0.02006 0.980 1.54101 0.09115 -0.02222 0.990 1.55667 0.04652 -0.00202 1 .000 1.57233 -0.00202 SECOND-STAGE VANE COORDINATES TIP SECTION (HOT RADIUS = 19.84001) PERCEN T X X Y TOP Y BOT 0.0 O.Ol Ol 9 2.5671 0 2.56710 O.Ol 0 0.02580 2.58257 2.52332 0.020 0.04141 2. 57929 2.49703 0.030 0.05702 2.5721 l 2.47173 0.040 0.07264 2. 56276 2. 44679 0.050 0.08825 2.55330 2.42165 0.060 O.l 0386 2. 54263 2. 39654 0.070 O.l l948 2.53186 2.37142 O. 080 O. 13509 2. 52094 2. 34628 0.090 O. 15070 2. 50988 2.3 2112 O. lO0 0.16631 2.49868 2. 29595 O.125 0.20535 2.47006 2. 23298 O. 150 O. 24438 2.44048 2.16990 0.175 0.28341 2. 40988 2.10674 0.200 O. 32244 2. 37816 2. 04348 0.225 0.36147 2.34526 l .9801 l 0.250 0.40051 2.31 I09 l .91 663 0.275 0.43954 2.27553 l .85304 0.300 0.47857 2.23850 l .78934 0.325 O. 51760 2.19988 l .72551 0.350 0.55664 2.15955 l .661 55 0.375 0.59567 2.1 1739 l .59746 0.400 O. 63470 2.07326 l. 53323 0.425 0.67373 2.02704 l .46884 0.450 O. 71 276 l .97858 l .40431 0.475 0.75180 l .92774 l .33960 O. 500 0.79083 I .87437 l .27474 0.525 0.82986 I. 8]833 l .20969 O. 550 0.86889 l .7 5949 l .14444 0.575 0.90792 1.69768 1.07900 0.600 0.94696 I .63277 l .Of 334 0.625 0.98599 1.56460 0.94746 0.650 l .02502 I .49298 0.881 34 0.675 l .06405 l .41775 0.81496 0.700 l. 10308 l .33871 0.74832 0.725 I. 1421 2 l .25564 0.68138 0.750 l .18115 I .16832 0.6141 5 0.775 l .2201 8 l .07653 0.54660 0.800 l .25921 O. 98001 0.47869 0.825 l .29824 0.87849 0.41041 0.850 l .33728 0.771 68 O. 34172 0.875 l .37631 0.65925 0.27261 0.900 l .41 534 0. 54088 0.20302 0.91 0 l .43095 0.49180 O. 17506 0.920 l .44657 0.441 69 O.l 4700 0.930 l .4621 8 0.39050 O.11886 O. 940 l .47779 O. 33827 0.09064 0.950 l .49341 0.28491 0.06232 0.960 l .50902 O. 23042 0.03391 0.970 l .52463 0.17478 0.00540 0.980 l. 54024 O. 11794 -O.Ol 855 0.990 l. 55586 0.05989 -0.021 50 l .000 l .571 47 -O.OOl 98 -O.OOl 98
SECOND-STAGE BLADE COORDINATES
ROOT SECTION (HOT RADIUS = 16.94200)
PERCENTX X Y TOP Y BOT
0.0 0.00171 0.43803 0.43803
O.Ol 0 O. 01220 0.471 l 5 0.42735
O. 020 0.02269 0.49032 0.43075
0.030 0.03318 0.50825 0.43689
0.040 0.04367 O. 52556 0.44465
0.050 0.0541 5 0.54228 0.45362
0.060 0.06464 O. 55843 0.46294
0.070 0.07513 O. 57401 0.47180
0.080 0.08562 O. 58906 0.48017
0.090 0.0961 l 0.60357 0.48808
O.lO0 O. 10660 0.61759 0.49557
0.125 O. 13282 0.65048 O. 51260
O.150 O. 15905 0.68042 0.52740
0.175 O. 18527 0.70758 0.5401 8
0.200 0.21149 0.73207 0.55114
0.225 0.23771 0.7 5400 0.56039
O. 250 O. 26394 0.77344 O. 56802
0.27 5 0.2901 6 0.79045 0.5741 2
0.300 0.31638 0.8051 0 O. 57875
0.325 0.34261 0.81741 O. 58193
0.350 O. 36883 0.8 2739 O. 58371
0.375 0.39505 0.83508 0.58410
0.400 0.421 27 0.84047 O. 58309
0.425 0.44750 0.84353 O. 58069
0.450 0.47372 0.84427 O. 57687
0.475 0.49994 0.84265 0.571 61
O. 500 O. 52617 0.83861 O. 56483
O. 525 0. 55239 0.83209 0.55649
O. 5 50 O. 57861 0.82303 0.54650
0.575 0.60483 0.81 l 31 0.53475
0.600 0.631 06 0.7 9683 0.521 09
0.625 0.65728 0.77945 0.50533
O. 650 O. 68350 0.7 5899 0.48725
0.67 5 0.7097 3 0.73525 0.46649
0.7 O0 0.7 3595 O. 70796 0.44264
0.725 0.76217 0.67681 0.41 517
0.7 50 0.78840 0.641 40 0.38383
0.775 0.81462 0.60130 0.34895
0.800 0.84084 0.55615 0.31132
0.825 0.86706 0.50568 0.27156
O. 850 0.89329 0.44978 O. 23014
0.87 5 O. 91951 0.38850 0.18737
0.900 O. 94573 0.321 98 O. 14342
O. 910 0.95622 0.29400 O. 12555
O. 920 0.96671 0.26524 O. 10752
0.930 0.97720 0.23575 0.08936
O. 940 0.98769 O. 20556 0.07106
0.950 0.9981 8 0.17469 0.05262
0.960 1.00867 0.1431 6 0.03406
0.970 l .O191 6 O.l1102 O.O1537
O. 980 l .02964 0.07828 -O.OOl 19
0.990 1.04013 0.04499 -0.00361
l. 000 l .05062 -0.00920 -0.00920
SECOND-STAGE BLADE COORDINATES QTR ROOT SECTION (HOT RADIUS = 17.87300) PERCEN T X X Y TOP Y BOT 0.0 0.00941 0.48439 0.48439 O.Ol 0 O.O1982 0.51 936 0.4731 0 0.020 0.03023 0.53968 0.47671 0.030 0.04064 0.55879 0.48324 0.040 0.05106 O. 57707 0.491 51 0.050 0.06147 0.59457 0.50109 0.060 0.07188 0.61132 0.51 lO0 0.070 0.08229 0.62738 O. 52038 0.080 O. 09271 O. 64275 O. 52922 0.090 O. I031 2 0.65750 0.53755 O.lO0 O.11353 0.67162 O. 54542 0.125 O. l 3956 0.70438 0.56321 O. 150 O. l 6559 0.73376 O. 57859 O. 175 O. l 9162 0.75999 O. 59183 O. 200 O. 21765 O. 78329 O. 6031 2 0.225 O. 24368 0.80382 O. 61261 0.250 0.26971 0.82170 0.62042 0.275 0.29574 0.83705 0.62662 0.300 0.32177 0.84994 0.63127 0.325 0.34780 0.86045 0.63444 O. 350 0.37383 0.86862 O. 6361 6 0.375 0.39986 0.87448 0.63641 0.400 0.42589 O. 87804 0.63523 0.425 0.45192 0.87933 0.63259 0.450 0.47795 0.87832 0.62848 0.475 0.50398 0.87499 O. 62284 0.500 0.53002 0.86931 0.61 562 0.525 0.55605 0.86122 0.60676 O. 550 O. 58208 0.85066 O. 59613 0.575 0.6081 l 0.83754 0.58362 0.600 0.6341 4 0.821 73 O. 56903 0.625 0.66017 0.8031 2 0.55215 0.650 0.68620 0.781 51 O. 53264 0.67 5 0.71223 0.75670 O. 51008 0.700 0.73826 0.72842 0.48383 0.725 0.76429 0.69632 0.45323 0.750 0.79032 0.66003 0.41 857 0.775 0.81635 O. 61921 0.38107 0.800 0.84238 O. 57364 0.3417 5 0.825 0.86841 0.52325 0.301 l 2 0.850 0.89444 O. 46813 O. 25948 0.875 0.92047 0.40850 0.21701 0.900 0.94650 O. 34471 O. 17385 0.91 0 0.95691 O. 31 81 l 0.I 5641 O. 920 0.96733 O. 29094 O. 13887 0.930 0.97774 0.26322 0.I 2126 O. 940 O. 9881 5 O. 23500 O. 10354 0.950 0.99856 0.20628 0.08577 0.960 l .00897 O. 17709 0.06792 0.970 l .O1939 O. 14746 0.04997 0.980 l .02980 O.11743 0.03914 0.990 l .04021 0.08700 0.03905 l .000 l .05062 -0.05368 -0.05368
SECOND-STAGE BLADE COORDINATES
MEAN SECTION (HOT RADIUS = 18.80299)
PERCENTX X Y TOP Y BOT
0.0 0.00941 O. 55491 0.55491
O. Ol 0 O. O1982 O. 59002 0.54357
0.020 0.03024 0.61 044 O. 54721
0.030 0.04065 0.62949 0.55377
0.040 O. 05106 0.64758 O. 56209
0.050 0.06147 0.66477 0.57172
0.060 0.07188 0.68113 0.58168
0.070 0.08230 0.69671 0.59105
0.080 0.09271 0.71154 0.59983
0.090 0.I0312 0.72567 0.60809
O.lO0 O.11353 O. 73914 0.61 585
0.125 0.13956 0.77008 0.63328
O.l 50 O. 16559 0.7 9743 0.64822
0.175 0.19162 0.82153 0.66090
0.200 0.21765 0.84260 0.671 57
0.225 O. 24368 0.86084 0.68036
0.250 0.26971 0.87641 O. 68740
0.275 0.29574 0.88942 0.69276
0.300 O. 321 77 0.89999 0.69650
O. 325 0.34781 0.90816 O. 69869
O. 350 O. 37384 O. 91402 O. 69932
0.375 0.39987 0.91759 0.69841
0.400 O. 42590 O. 91889 O. 69596
0.425 0.451 93 0.91793 0.69192
0.450 0.47796 O. 91471 O. 68627
0.475 O. 50399 0.90919 0.67893
O. 500 O. 53002 0.901 34 0.66981
0.525 0.55605 0.891 l l 0.65880
O. 550 O. 58208 0.87841 0.64572
0.575 O. 6081 l 0.8631 4 0.63037
O. 600 O. 63414 0.8451 8 0.61 245
0.625 0.66017 0.82437 O. 59153
O. 650 O. 68620 0.80051 0.56704
0.67 5 O. 71223 0.77335 0.53821
0.700 0.73826 0.74257 0.5051 4
0.725 0.76429 0.70779 0.46916
O. 7 50 O. 7 9032 O. 66870 0.431 42
0.775 0.81635 0.6251 3 0.39248
0.800 0.84238 0.5771 l 0.35260
0.825 0.86841 O. 52482 0.31200
O. 8 50 0.89444 0.46859 0.27075
0.875 0.92047 0.40882 0.22899
O. 900 0.94650 0. 34596 0.18676
O. 910 0.95691 0.32005 0.16974
O. 920 O. 96733 0.29374 O. 15266
0.930 0.97774 0.26707 O. 13552 O. 940 0.9881 5 0.24004 O. 11832 0.950 0.99856 O. 21270 O. 1 0106 O. 960 l .00897 0.18505 0.08375 0.970 l .O1939 O. l 571 2 0.06637 O. 980 l .02980 0.12892 0.05504 O. 990 l .04021 O. I0049 0.05473 l .000 l .05062 O. 0691 5 O. 0691 5 SECOND-STAGE BLADE COORDINATES QTR TIP COORDINATES (HOT RADIUS = 19.73500) PERCENT X X Y TOP Y BOT 0.0 -0.00037 0.65325 0.65325 O.OlO O. Ol Ol 4 0.6851 2 0.64294 0.020 0.02065 0.70346 0.64614 0.030 0.031 l 6 0.72025 0.65197 0.040 0.041 67 0.7361 8 0.65934 0.050 0.0521 8 0.7 51 31 0.66786 0.060 0.06269 0.76568 0.67671 0.070 0.07320 0.77935 0.68506 0.080 0.08371 0.79234 0.69287 0.090 0.09422 0.80469 0.70018 O.lO0 O. I0473 0.81 643 0.70704 0.125 O.l 3101 0.84331 0.72233 0.150 O. 15728 0.86688 0.73520 0.175 0.18356 0.88744 0.74592 0.200 0.20983 0.9051 8 0.75465 0.225 0.23611 0.92028 0.76150 0.250 0.26238 O. 93287 O. 76658 0.275 0.28866 0.94304 0.76995 0.300 0.31 493 0.95089 0.771 65 0.325 0.341 21 0.95646 0.771 71 0.350 O. 36748 0.95979 0.7701 2 0.375 0.39375 0.96093 0.76686 0.400 0.42003 O. 95984 O. 7 61 90 0.425 0.44630 0.95655 0.75517 0.450 0.47258 0.951 02 0.74657 0.475 0.49885 0.94320 0.73599 0.500 0.5251 3 0.93304 0.72326 0.525 0.55140 0.92046 0.70815 0.550 0.57768 O. 90534 0.69033 0.575 0.60395 0.88756 0.66935 0.600 O. 63023 0.86694 0.64456 0.625 0.65650 0.84325 0.61534 0.650 O. 68278 O. 81 622 O. 58227 0.675 0.70905 0.78549 0.54685 0.700 0.7 3533 0.7 5064 0.50993 0.725 0.76160 0.71139 0.47199 0.750 0.78788 0.66768 0.43323 0.775 0.81 41 5 O. 61969 0.39382 0.800 0.84042 O. 56777 0.35383 0.825 0.86670 0.51242 0.31338 0.850 0.89297 0.4541 4 O. 27250 0.875 O. 91925 0. 39345 0.23124 0.900 0.94552 O. 33075 O. 1 8962 0.910 0.95603 0.30521 O. 17288 0.920 O. 96654 O. 27941 O. 15610 0.930 0.97705 0.25341 O.l 3928 0.940 O. 98756 0.2271 9 O. 12240 0.950 0.99807 0.20079 0.I0548 0.960 l .00858 O. 17422 0.08853 0.970 l .Ol 909 O. 14749 0.07l 51 0.980 l .02960 O. 12061 0.05520 0.990 l .0401 l 0.09359 0.05186 l .000 l .05062 O. 0641 2 0.0641 2 SECOND-STAGE BLADE COORDINATES TIP SECTION (HOT RADIUS = 20.66600) PERCENT X X Y TOP Y BOT 0.0 -0.02208 0.79950 0.79950 O.Ol 0 -O.O1135 0.82495 0.79054 0.020 -0.00062 O. 83950 0.7 9250 0.030 O.Ol Oll 0.85206 0.79652 0.040 0.02083 0.86393 O. 801 74 0.050 0.031 56 0.87516 0.80784 0.060 0.04229 0.88580 0.81426 0.070 0.05301 0.89586 0.82032 0.080 0.06374 0.90539 O. 82595 0.090 0.07447 0.91439 0.831 l 9 O.l O0 0.08519 O. 92291 0.83606 O.125 O.11201 0.9421 8 0.84667 O.l 50 O.13883 0.95875 0.8551 8 0.17 5 O.16565 0.97283 0.86174 0.200 O.19246 0.98457 0.86644 0.225 O. 21928 0.9941 l 0.86933 0.250 0.2461 0 l .OOl 50 0.87045 0.275 0.27292 l .00685 0.86983 0.300 0.29973 l .01017 0.86746 0.325 0.32655 l .Of l 52 0.86330 O. 350 0.35337 l .Of 091 0.85729 0.375 0.3801 9 l .00832 0.84936 0.400 0.40700 l .00372 0.83938 0.425 0.43382 0.99709 O. 8271 9 0.450 0.46064 O. 98837 O. 81255 0.47 5 0.48746 0.97746 0.7951 5 O. 500 O. 51 427 0.96429 0.77454 0.525 0.541 09 0.94868 0.75006 O. 550 O. 56791 0.93046 0.721 l 5 0.575 0.59473 0.90941 0.68854 0.600 0.621 54 0.88522 0.65373 0.625 0.64836 0.85746 0.61755 O. 650 O. 6751 8 0.82559 O. 58038 0.67 5 0.70200 0.78903 0.54244 0.700 0.72881 0.74754 O. 50389 0.725 0.75563 0.70130 0.46481 0.7 50 0.7 8245 O. 65085 0.42527 0.775 0.80927 O. 59692 0.38533 0.800 0.83608 O. 54023 0.34503 0.825 0.86290 0.48141 0.30439 0.850 O. 88972 0.42096 0.26346 0.875 0.91654 0.35923 0.22224 0.900 O. 94335 0.29654 O. 1 8076 0.91 0 0.95408 0.27124 O. 1641 0 0.920 O. 96481 0.24583 O. 14741 0.930 0.97553 0.22033 O.13066 0.940 0.98626 O.l 9473 O.11390 0.950 0.99699 O. 16904 0.09709 0.960 l .00771 O.l 4328 0.08026 0.970 l .01844 O.11745 0.06338 O. 980 l .02917 0.091 54 0.04649 0.990 l .03989 0.06561 0.03157 l .000 l .05062 O. 0381 9 0.0381 9 THIRD-STAGE VANE COORDINATES ROOT SECTION (HOT RADIUS = 17.35500) PERCENT X X Y TOP Y BOT 0.0 -0.00041 O. 65793 0.65793 O.OlO O.O1570 0.71300 0.64681 0.020 0.031 81 0.7 5378 0.65750 0.030 0.047 91 0.79133 0.67258 0.040 0.06402 O. 82610 0.68799 0.050 0.0801 3 0.85845 0.70308 0.060 0.09624 0.88864 0.71782 0.070 O.l 1234 O. 91690 0.73221 0.080 O.12845 0.94342 0.74625 0.090 O. 14456 O. 96835 0.75991 O.lO0 O.16067 0.991 82 0.7731 9 0.125 0.20094 1.04471 0.80463 0.150 0.24121 l .09036 0.83342 0.175 0.28147 l. 12975 0.85937 0.321 74 l .16354 0.88228 0.200 0.225 0.36201 l. 19225 0.90199 0.250 0.40228 l .21 631 0.91 832 0.275 0.44255 1.23601 0.93111 0.300 0.48282 l .251 62 O. 94023 0.325 0.52309 l. 26331 0.94558 0.350 O. 56336 l .27123 0.94707 0.375 0.60363 l .27550 0.94468 0.400 0.64390 l .2761 8 0.93839 0.425 0.6841 7 l .27331 0.92825 0.450 O. 72443 l .26694 O. 91430 0.475 0.76470 1.25702 0.89666 0.500 0.80497 l .24354 0.87545 0.525 0.84524 l .22642 0.85083 0.550 0.88551 l .20557 0.82295 0.575 0.92578 l .18086 0.79201 0.600 O. 96605 l . 15209 0.7581 8 0.625 l .00632 l .I1904 0.72166 0.650 l .04659 l .08142 0.68262 0.675 l .08685 l .03884 0.64125 0.700 l .1271 2 O. 99086 O. 59773 0.725 l. 16739 O. 93700 0.55222 l .20766 0.87699 0.50487 0.750 0.775 l .24793 O. 81077 0.45583 l .28820 0.73850 0.40526 0.800 0.825 1.32847 0.66054 0.35324 0.850 l .36874 O. 57738 0.29993 0.875 l .40901 0.48960 0.24540 0.900 l .44928 0.39781 O. 1 8977 0.910 l .46538 0.3601 0 O. 16723 l .481 49 O. 321 86 O.l 4453 0.920 0.930 1.49760 O. 2831 4 O. l 2167 0.940 l .51 371 0.24396 O. 09867 0.950 l .52981 0.20436 0.07553 0.960 l. 54592 O. l 6436 0.05224 0.970 l .56203 O. 12400 0.02883 0.980 l .57814 0.08327 0.00528 0.990 l .59424 0.04223 -O.Ol 343 l .61035 -0. OOl 46 -0.00146 l .000
THIRD-STAGE VANE COORDINATES
QTR ROOTSECTION (HOT RADIUS = 18.46500)
PERCENTX X Y TOP Y BOT
0.0 -0. 00007 n.83670 0.83670
O.Ol 0 0.01603 0.88372 0.82705
0.020 0.03214 0.91 825 0.83617
0.030 0.04824 0.95080 0.84887
0.040 0.06434 0.981 59 0.86175
0.050 0.08045 l .01077 0.87434
0.060 0.09655 l .03845 0.88662
0.070 O.11266 l .06475 0.89858
0.080 O. 12876 l .08975 0.91023
0.090 O. 14487 l. 11355 O. 92156
O.lO0 O. 16097 l . 13622 0.93256
O. 125 0.20123 l . 18828 O. 95859
O. 150 0.241 49 l .23440 O. 98240
O. 175 0.28175 l .2751 4 l .00390
0.200 O. 32201 l .31 lOl l .02298
0.225 0.36227 l .34231 l .03952
0.250 0.40253 l .36936 l .05341
0.275 0.44279 l .39235 l .06454
0.300 0.48306 l .41145 l .07279
0.325 0.52332 l .4267 9 l .07806
0.350 O. 56358 l .43847 l .08025
0.375 0.60384 l .44655 1.07924
0.400 0.6441 0 l .451 04 l .07495
0.425 0.68436 l .45197 l .06730
0.450 0.72462 l .44931 l .05622
0.475 0.76488 1 .44300 l .04162
0.500 0.80514 l .43297 l .02346
0.525 0.84540 l .41909 l .OO172
0.550 0.88566 1.40122 0.97634
0.57 5 0.92592 l .3791 5 0.947 31
0.600 O. 9661 8 l .35264 O. 91466
0.625 l .00644 l .32135 0.87838
0.650 l .04670 l .28487 0.83852
0.67 5 l .08696 l .24264 0.79509
0.700 l. 12722 l . 19394 0.7481 8
0.725 l. 16748 l . 13787 0.69784
0.750 l .20775 l .07356 0.6441 5
0.775 l .24801 l .00035 O. 58719
0.800 l .28827 0.91799 O. 52706
0.825 l .32853 0.82663 0.46386
0.850 l .36879 0.72679 0.39769
0.875 l .40905 O. 61933 0.32867
0.900 l .44931 0.50520 0.25692
0.91 0 l .46541 0.45789 0.22746
O. 920 l .481 52 0.40974 O. l 9760
0.930 l .49762 0.36081 O. 16733
O. 940 l .51373 0.31 114 O. 13667
0.950 l .52983 0.26078 O. I0561
O. 960 1.54593 0.20979 0.0741 6
O. 970 l .56204 O. 15822 0.04234
O. 980 l .57814 O. 1061 l 0.0101 5
0.990 1.59425 0.05348 -0.01343
l. 000 l .61035 -0. 00147 -O.OOl 47
THIRD-STAGE VANE COORDINATES MEAN SECTION (HOT RADIUS : 19.57600) PERCENT X X Y TOP Y BOT 0.0 0.00028 1.02438 1.02438 0.010 0.01638 1.06534 1 .01 555 0.020 0.03248 1.09460 1 .02322 0.030 0.04858 1.12231 1 .03391 0.040 0.06468 1.14861 1 .04470 0.050 0.08078 1.17360 1 .05530 1 .06567 0.060 0.09688 1.19735 0.070 0.I1299 1.21996 1 .07583 0,080 0.12909 l .24148 1 .08576 0.090 0.1451 9 l .26199 1.09546 0.100 1.10493 0.16129 l .281 51 0.125 0.20154 1.32636 1.12745 0.150 0.24179 1.36600 1 .14827 0.175 0.28204 1.40089 1.16722 0.200 0.32229 1.43138 1.18414 0.225 0.36255 1.45774 1.19886 0,250 0.40280 1.48021 1 .21118 0.275 0.44305 1.49893 1.22091 0.300 0.48330 1.51404 1 .22784 0.325 0.'52355 1.52563 1 .23177 0.350 0.56381 1.53376 1.23246 0.375 0.60406 1.53848 1 .22972 0.400 0.64431 1.53979 1.22335 0.425 0.68456 1.53766 1 .21314 0.450 0.72481 1.53206 1 .19894 0.475 0.76506 1.52292 1.18061 0.500 0.80532 1.51012 1.15803 0.525 0.84557 1.49354 1.13115 0.550 0.88582 1.47298 1 .09992 0.575 1 .06437 0.92607 1.44822 0.600 0.96632 1.41895 1 .02455 0.625 1.00657 1.38478 0.98056 0.650 1.04683 1.34523 0.93254 0.675 0.88063 1.08708 1.29963 0.700 0.82502 1.12733 1.24712 0.725 1.16758 1.18669 0.76591 0.750 1.20783 I.I1754 0.70352 0.775 1.24809 1.03910 0.63805 0.800 1.28834 0.95133 0.56972 0.825 1.32859 0.85457 0.49873 0.850 1.36884 0.74956 0.42529 0.875 1.40909 0.63726 0.34958 0.900 1.44934 0.51872 0.27180 0.910 1.46544 0.46977 0.24014 1.48155 0.42005 0.920 0.20819 0.930 1.49765 0.36962 0.17595 0.940 1.51375 0.31850 0.14343 0.950 1.52985 0.26678 0.11064 0,960 1.54595 0.21448 0.07761 0.970 1.56205 0.16165 0.04432 0.980 1.57815 0.I0833 0.01079 0.990 1.59425 0.05455 -0.01343 1.000 1.61035 -0.00146 -0.00146 THIRD-STAGE VANE COORDINATES QTR TIP SECTION (HOT RADIUS : 20.68600) PERCENT X X Y TOP Y BOT 0.0 0.00065 1 .22253 l .22253 O.Ol 0 0.01675 I .257 21 l . 21416 0.020 O. 03285 I .28096 l .22001 0.030 0.04894 1.30357 1.22831 0.040 0.06504 1.32510 1.23666 0.050 0.08114 I .34562 l .24485 0.060 0.09724 l .36517 l .25289 0.070 O.l 1333 1 .38379 l .2607 5 0.080 0.12943 1.40156 1.26846 0.090 0.14553 1.41849 1.27597 O.lO0 0.16162 1.43461 1.28330 0.125 0.20187 1.47162 1.30074 0.150 0.24211 1 .50421 1.31680 0.175 0.28235 1.53270 I .331 31 0.200 O. 32259 l .55734 1 .3441 l 0.225 O. 36284 l .57830 l .35497 0.2 50 0.40308 l .59577 I .36368 0.275 0.44332 I .60985 1.36997 0.300 0.48356 l .62061 I .37359 0.325 0.52381 l .62813 l .37422 0.350 0.56405 l .63243 l .371 57 0.375 0.60429 1.63352 1.36532 0.400 0.64453 l .631 39 l .35517 0.425 0.68478 l .62599 l .34080 0.450 0.72502 l .61728 l .321 97 0.475 0.76526 l .60515 l .29848 0.500 0.80550 l .58948 l .2701 8 0.525 0.84574 l .57011 l .23698 O. 550 0.88599 l .54683 l . 19891 0.575 0.92623 1.51939 1.15608 0.600 O. 96647 l .48745 1 . I0861 0.625 l .00671 l .45057 I .05677 0.650 l .04696 l .40823 l .00081 0.675 1.08720 1.35966 0.94104 0.700 l. 12744 l .30392 0.87776 0.725 l. 16768 l .24001 0.81130 0.7 50 l .20793 l .16714 0.741 94 0.775 l .24817 1 .08479 0.66999 0.800 l .28841 0.99289 0.59571 0.825 l .32865 0.89179 O. 51934 0.850 l .36890 0.78221 0.441 l l 0.875 1.40914 0.66510 0.36120 0.900 l .44938 O. 541 47 0.27982 O. 91 0 l .46548 0.49042 0.24688 0.920 l .481 58 0.43855 0.21374 0.930 l .49?6? 0.38592 0.18040 0.940 l .51377 0.33258 0.I 4688 0.950 l .52987 0.27858 0.I 131 9 0.960 l .54596 0.22398 0.07932 0.970 l .56206 0.16879 0.04528 0.980 l .5781 6 O.11309 O.Ol 109 0.990 l .59425 0.05691 -O.O1343 l .000 l .61 035 -0.00146 -0.00146 THIRD-STAGE VANE COORDINATES TIP SECTIOM (HOT RADIUS = 21.79700) PERCENT X X Y TOP Y BOT 0.0 O.OOl 05 l .43534 l .43534 O.Ol 0 O.O171 5 l .46634 l .42696 0.020 0.03324 l .48687 l .431 53 0.030 0.04933 l .50634 l .43818 0.040 0.06543 l .52483 l .44482 0.050 0.081 52 l .54236 l .45132 0.060 0.09761 l .55900 l .45767 0.070 O.11370 l .57479 l .46387 0.080 O. 12980 l .58975 l .46992 0.090 0.14589 l .60394 l .47579 O.lO0 O.161 98 1 .61739 l .48149 0.125 0.20222 l .64788 1 .49492 O. 150 0.24245 l .67422 l .50702 0.17 5 0.28268 l .69668 1.51764 0.200 0.32291 l .71 548 l .52652 0.225 0.3631 5 l .73082 l .53346 0.250 0.40338 l .74282 l .53817 0.27 5 0.44361 l .75158 l .54032 0.300 0.48384 l .7571 8 l .53958 0.325 0.52408 l .75966 l .53557 0.350 0.56431 l .75907 l .52788 0.375 0.60454 l .75539 l. 51 61 l 0.400 0.64477 l .74862 l .49988 0.425 0.68501 l .73870 l .47881 0.450 0.72524 1.72556 1.45263 0.475 0.76547 l .7091 2 l .42117 O. 500 0.80570 l .68925 l .38434 0.525 0.84594 1.66578 1.34221 0.550 0.88617 l .63851 l .29495 0.575 0.92640 l .6071 7 l .24284 0.600 0.96663 1.57142 1.18622 0.625 l .00686 l .53082 l. 12548 0.650 I .0471 0 l .48482 l .06104 0.675 I .08733 l .43265 0.99329 0.700 l .12756 l .37336 0.92261 0.725 I. 16779 l .30606 0.84935 0.750 I .20803 l .22990 0.77382 0.775 I .24826 l. 14428 0.69632 0.800 I .28849 l .04895 0.6171 0 0.825 I. 3287 2 0.94400 0.53634 0.850 1.36896 0.82988 0.45425 0.875 I .4091 9 0.70735 0.37100 0.900 I .44942 O. 57729 0.28672 0.910 1.46551 0.52337 0.25274 O. 920 I .48161 0.46845 O. 218_3 0.930 1.49770 0.41 262 0.18439 0.940 1.51379 0.35590 O.15001 0.950 I. 52989 0.29836 O.l 1552 0. 960 I. 54598 O. 24006 0.08092 0.970 I. 56207 O.l 8105 0.04621 0.980 I .57817 O. 121 36 O.Ol 138 0.990 I. 59426 O. 061 05 -0. O1343 l .000 I .61035 -0.00146 -O.OO146
THIRD-STAGE BLADE COORDINATES
ROOTSECTION (HOT RADIUS = 17.53300)
PERCENTX X Y TOP Y BOT
0.0 -0.00048 0.67741 0.67741
O.Ol 0 O. 01282 0.7087 2 0.66855
0.020 0.0261 l 0.72689 0.67276
0.030 0.03940 0.74362 0.67956
0.040 0.05269 O. 76008 O. 68794
0.050 0.06598 0.77625 0.69717
0.060 0.07928 0.7921 3 0.70608
0.070 0.09257 0.80770 0.71455
0.080 0.10586 0.82297 0.7 2261
0.090 O.ll 915 0.83791 0.73026
O.lO0 O.13245 0.85253 0.7 3751
O. 125 O.16568 0.88765 0.75402
O. 150 O. 19891 0.92063 0.76832
0.175 0.23214 0.95141 0.78057
O. 200 0.26537 0.97992 0.79087
0.225 0.29860 l .00608 0.79932
0.2 50 O. 331 83 I .02983 0.80598
0.275 0.36506 1 .05111 0.81092
0.300 0.39829 1.06984 0.81 415
0.325 0.43152 1 .08595 0.81573
0.350 0.46475 I .09935 0.81562
0.375 0.49799 I. 10997 0.81387
0.400 O. 53122 I . 11770 0.81 044
0.425 O. 56445 l. 12247 0.80530
0.450 0.59768 I . 1241 4 0.79843
0.475 0.63091 1 .I 2260 0.78974
0.500 O. 66414 1.11773 0.77919
0.525 0.69737 l. 10938 0.76666
0.550 O. 73060 1 .09738 0.75205
0.57 5 0.76383 l .08157 0.73521
0.600 0.79706 1 .06174 0.71 594 0.625 0.83029 1 .03765 0.69396 0.650 0.86353 1 .00906 0.66881 0.675 0.89676 0.97568 0.64006 0.700 0.92999 O. 93716 0.60741 0.725 0.96322 0.89 313 O. 57020 0.750 O. 99645 0.8431 3 0.52776 0.775 l .02968 0.78671 0.48043 0.800 1.06291 0.72352 0.42925 0.825 l .09614 0.65342 0.37 511 0.850 l. 12937 O. 57654 0.31 852 0.87 5 l .l 6260 0.49320 0.25994 O. 900 l. 19583 0.40390 O. l 9970 O. 91 0 l .2091 2 0.36664 O. 17517 O. 920 l .22242 0.32857 O.l 5038 0.930 l .23571 0.28971 0.12544 O. 940 l .24900 0.2501 4 0.10026 0.950 l .26229 0.20986 0.07489 O. 960 l .27559 O. 16893 0.04931 0.970 l .28888 O.l 2738 0.02357 0.980 l .30217 0.08526 -0.00236 0.990 l .31546 0.04259 -0. Ol 495 l .OOO l .32876 -0.00248 -0.00248 THIRD-STAGE BLADE COORDINATES QTR ROOT SECTION (HOT RADIUS = 18.84399) PERCENT X X Y TOP Y BOT 0.0 -0.00046 0.68076 0.68076 O.OlO 0.01283 0.72322 0.66959 0.020 0.02612 0.74885 0.67625 0.03941 0.77326 0.68661 0.03O 0.040 0.05271 0.79659 O. 69933 0.050 0.06600 0.81891 0.71302 0.060 0.07929 0.84028 0.?2594 0.070 0.09258 0.86073 0.73806 0.080 0.I0587 0.88033 0.74943 0.090 0.I1917 0.89909 0.76013 O.lO0 0.13246 0.91707 0.77019 0.16569 0.95872 0.79285 0.125 0.19892 0.99600 0.81230 0.150 0.175 0.23215 1.02923 0.82892 0.200 0.26538 1.05865 0.84297 0.225 0.29861 1.08448 0.85464 0.250 0.33184 I.I0687 0.86410 0.275 0.36507 1.12593 0.87145 0.300 0.39830 1.14178 0.87676 0.325 0.43153 1.15450 0.88011 0.350 0.46476 l .1641 l 0.881 52 0.375 0.49799 l .17067 0.881 Ol 0.400 0.531 22 l .1741 9 0.87858 0.425 0.56446 1.17464 0.87420 0.450 0.59769 1.17203 0.86782 0.475 0.63092 1.16630 0.85938 0.500 0.66415 1.15739 0.84878 0.525 0.69738 1.14522 0.83589 0.550 0.73061 1.12967 0.82053 0.575 0.76384 I.II062 0.80246 0.600 0.79707 1.08790 0.78132 0.625 0.83030 1.06129 0.75680 0.650 0.86353 1.03054 0.72845 0.675 0.89676 0.99533 0.69535 0.700 0.92999 0.95525 0.65587 0.725 0.96322 0.90979 0.60928 0.750 0.99645 0.85832 0.55753 0.775 1.02968 0.80024 0.50254 0.800 1.06291 0.73519 0.44533 0.825 1.09614 0.66307 0.38642 0.850 1.12937 0.58408 0.32619 0.875 1.16260 0.49870 0.26482 0.900 1.19583 0.40754 0.20247 0.910 1.20913 0.36963 0.17728 0.920 1.22242 0.33095 0.15198 0.930 1.23571 0.29157 0.12653 0.940 1.24900 0.25150 0.I0097 0.950 1.26229 0.21082 0.07530 0.960 1.27559 0.16955 0.04953 0.970 1.28888 0.12774 0.02364 0.980 1.30217 0.08542 -0.00236 0.990 1.31546 0.04264 -0.01494 l .000 1.32876 -0.00247 -0.00247 THIRD-STAGE BLADE COORDINATES MEAN SECTION (HOT RADIUS : 20.15401) PERCENT X X Y TOP Y BOT 0.0 -0.00023 0.73695 0.73695 O.Ol 0 O. O1306 0.78223 0.72504 O. 020 0.02635 0.81 000 0.7 3230 0.030 0.03964 0.83638 0.7 4137 0.040 0.05293 0.86145 0.7 5749 0,050 0,06622 0.88529 0,77233 0.060 0.07951 0.90798 0.78620 0.070 0.09280 0.92959 0.79915 0.080 O. 10609 0.9501 9 0.81126 0.090 O.l 1 938 0.96980 0.82259 O.lO0 O.13267 0.98849 0.83322 0.125 O.16589 l .03143 0.85701 O, 150 O. 1991 l l .06936 0.87728 0.175 0.23234 l .10272 0.89447 0,200 0.26556 l .131 86 O. 90889 0.225 0.29879 l .l 5704 0.92076 0.2 50 O. 33201 1. 17848 O. 93026 0.275 0.36524 l .19634 0.93752 O. 300 O. 39846 l .21 078 0.94262 0.325 0,43169 l .22190 0.94562 O. 350 0.46491 l .22978 0.94656 0.375 0.49814 l .23447 O. 94545 0.400 O. 531 36 l. 23601 O. 94229 0,425 0.56459 l .23441 0.93701 0.450 0.59781 l .22965 0.92956 0.475 0.63104 l .22172 0.91987 0,500 0,66426 l .21055 O, 90777 0.525 0.69749 1.19606 0.89308 O. 550 0.73071 l .17814 0.87563 0.575 0.76394 l. 15665 0.85523 0.600 0.7971 6 l .13142 0.831 55 0.625 0.83039 l .I0221 0.80348 0.650 0.86361 l .06873 O. 76915 0.675 0.89683 l .03061 0.72744 O. 700 O. 93006 O. 98738 0.67944 0.725 0.96328 0.93842 0.62626 0.750 O. 99651 0.8831 l O. 56977 0.775 l .02973 0.82096 O. 51133 0.800 l .06296 0.7 5178 0.451 52 0.825 l .0961 8 0.67570 0.39064 0.850 l .12941 0.59312 0.32892 0.87 5 l , 16263 0.50466 0.26645 0.900 l. 19586 0.41 I07 0.20330 O. 910 l .2091 5 0.37237 O. 17789 O. 920 l .22244 0.33300 O.l 5240 0.930 l .23573 0.29303 0.12678 0.940 l .24902 0,25249 O.l Ol l l 0.950 l .26231 0.21142 0.07535 0.960 l .27560 O. 16986 0.04953 0,970 l ,28889 0.12785 0.02361 O. 980 l .3021 8 0.08542 -0.00238 0.990 l .31547 0,04260 -0.01494 l .000 l .32876 -0.00246 -0.00246 THIRD-STAGE BLADE COORDINATES TIP SECTION QTR (HOT RADIUS = 21.46500) PERCENT X X Y TOP Y BOT 0.0 O.O001 5 0.83054 0.83054 O.OlO 0.01343 0.87454 0.81895 0.020 0.02672 O. 90126 0.82592 0.030 0.04001 0.92657 0.83677 0.040 0.05329 0.95056 0.8501 l 0.050 0.06658 O. 97334 0.86438 0.060 0.07986 0.99498 0.87770 0.0931 5 l .01554 0.8901 l 0.070 0.080 O. I0644 l .0351 l O. 901 69 0.090 O.11972 l .05372 0.91248 O.lO0 0.13301 1.07142 0.92257 0.125 0.16622 l .11200 0.94502 0.150 O.19944 l .14772 0.96395 0.175 0.23265 l .17902 0.97979 0.200 O. 26587 l .20623 O. 99284 0.29908 l .22960 l .00333 0.225 0.250 0.33230 l .24933 l .O1142 0.275 0.36552 l .26560 l .O1721 0.300 0.39873 l .27851 l .02080 0.325 0.43195 l .2881 7 l .02221 0.350 0.46516 1.29465 1.02148 0.375 0.49838 l .29796 l .01859 0.400 0.531 59 l .2981 6 l .Ol 352 0.425 0.56481 l .29522 l .00618 0.450 0.59802 l .2891 l 0.99647 0.475 0.63124 l .27978 0.98428 0.500 0.66445 l .2671 8 0.96937 0.525 0.69767 l .251 l 7 0.95149 0.550 0.73088 l .231 65 0.93024 0.575 0.7641 0 l .20841 0.90497 0.600 0.79731 l .l 81 25 0.87477 0.625 0.83053 l .l 4990 0.83877 0.650 0.86374 l .l 1396 0.7 9569 0.675 0.89696 l .07301 0.7451 5 0.700 0.93017 l .02642 0.69025 0.725 0.96339 0.97349 0.63312 O.750 0.99660 0.91 365 O. 57452 0.775 1.02982 0.84662 0.51476 0.800 l .06303 O. 77247 0.45403 0.825 l .09625 0.69163 0.39247 0.850 l. 12946 O. 60477 0.33020 0.875 l .16268 O. 51269 0.26731 0.900 l .19590 0.41 61 9 0.20386 0.910 1.20918 0.37652 0.17835 0.920 l .22247 0.33631 O.l 5276 0.930 l .2357 5 0.29560 O. 12706 0.940 1.24904 0.25442 0.I0132 0.950 l .26233 0.21281 0.07 549 0.960 l .27561 O. 17081 0.04961 0.970 l .28890 0.12844 0.02365 0.980 l .30218 0.08574 -0.00239 0.990 l .31547 0.04271 -O.Ol 494 l .000 l .32876 -0.00247 -0.00247
THIRD-STAGE BLADE COORDINATES
TIP SECTION (HOT RADIUS = 22.77499)
PERCENTX X Y TOP Y BOT
0.0 0.00063 0.94790 0.94790
O.Ol 0 0.01391 0.98978 0.93692
0.020 0.0271 9 l .01483 O. 94350
0.030 0.04047 l .03835 0.95372
0.040 0.05375 l .06054 O. 96625
0.050 0.06703 l .08150 0.97972
0.060 0.08031 l .lO134 O. 99230
0.070 0.09360 l. 1201 l l .00398
0.080 O. 10688 l . 13790 l .Ol 485
0.090 O. 1201 6 l .l 5475 l .02495
O.lO0 O. 13344 l . 17071 l .03435
0.125 O.16664 l .20705 I .05510
O. 150 O. 19985 l .23872 l .07235
0.175 0.23305 l .2661 6 I .08650
0.200 O. 26625 l .28969 I .09785
0.225 0.29946 l .30958 I. I0659
0.250 0.33266 l .32603 I .11286
0.275 0.36586 l .33921 l .l 1677
0.300 0.39907 1.34921 I.I1836
0.325 0.43227 l .3561 3 1 .I 1767
0.350 0.46547 l .36005 1. 11468
0.375 0.49868 1.36098 1.10935
0.400 O. 53188 l .35893 I . lOl 60
0.425 0.56508 l .35390 l .09132
0.450 O. 59829 1.34585 1.07831
0.47 5 0.63149 l .33470 l .06233
O. 500 O. 66469 l .32038 l .04306
0.525 0.69790 1.30276 1.01999
O. 550 0.731 lO l .28167 0.99227
0.575 0.76430 l .25691 0.95848
0.600 0.79750 l .22821 0.91 674
0.625 0.83071 l .l 9523 0.8681 5
0.650 0.86391 l .15750 0.81 525
0.675 0.8971 l l .11444 0.75933
0.700 0.93032 1.06526 0.70147
0.725 0.96352 l .0091 5 0.64231
0.750 O. 99672 0.94556 O. 5821 5
0.775 l .02993 0.87439 0.521 lO
0.800 l .0631 3 0.7 9592 0.45929
0.825 l .09633 0.71082 0.39680
0.850 l. 12954 0.61 996 0.33370
0.87 5 l .16274 0.52428 O. 27006
O. 900 l . 19594 0.42463 0.20592
O. 910 l .20922 0.38384 O. 18012
O. 920 l .22251 O. 34256 O. l 5426
0.930 l .23579 0.30085 O. 12831
O. 940 l .24907 0.25874 0.10232
0.950 l .26235 0.21627 0.07625
O. 960 l .27563 0.17346 0.05009
0.970 l .28891 O.13035 0.02392
O. 980 l .3021 9 0.08695 -0.00235
0.990 l . 31547 0.04329 -0. 01495
l .000 l .32876 -0.00247 -0.00247
FOURTH-STAGE BLADE COORDINATES ROOT SECTION (HOT RADIUS = 17.65700) PERCEN T X X Y TOP Y BOT 0.0 -0.00048 0.67553 0.67553 O.Ol 0 0.01281 0.70826 0.66642 O. 020 0.0261 0 0.72740 O. 67095 0.030 0.03939 0.74509 0.6781 9 0.040 O. 05269 0.7 6240 0. 68709 0.050 0.06598 0.77935 0.69692 O. 060 0.07927 O. 7 9591 O. 70634 0.070 0.09256 0.81 21 l O. 71527 0.080 0.I0585 0.82791 0.72375 0.090 O.1191 5 0.84334 0.73178 O. lO0 O.13244 0.85838 0.73940 0.125 0.16567 0.89433 0.75670 O. 150 O. 19890 O. 92787 0.771 66 0.175 0.2321 3 0.95897 0.78448 O. 200 O. 26536 0. 98761 0.79528 0.225 0.29859 1.01375 0.80416 0.250 0.331 83 l .03735 0.81120 0.275 0.36506 l .05837 O. 81646 0.300 0.39829 l .07677 0.81 999 0.325 0.43152 l .09248 0.82182 0.350 0.46475 l .10543 0.821 94 0.37 5 0.49798 l .l 1557 0.82038 0.400 O. 531 21 l. 12281 0.8171 l 0.425 0.56444 l. 12707 0.81 21 l 0.450 O. 59767 l. 12825 0.80535 0.475 0.63090 l. 12623 0.79674 O. 500 0.6641 4 l .12091 0.78624 0.525 0.69737 l .l l 213 0.77374 O. 550 0.73060 l .09975 0.7 591 l 0.575 0.76383 l .08360 0.7421 9 0.600 0.79706 1.06348 0.72275 0.625 0.83029 l .0391 6 0.70056 0.650 0.86352 l .Ol 039 0.67521 0.675 0.89675 0.97689 0.64624 0.700 O. 92998 0.93829 O. 61 302 0.725 0.96322 0.89421 O. 57482 0.750 0.99645 0.8441 8 0.531 31 0.775 l .02968 0.78770 0.48308 0.800 l .06291 0.72442 0.431 l 9 0.825 l .0961 4 0.65422 0.37647 0.850 l. 12937 O. 57720 O. 31 945 0.875 l .16260 0.49371 0.26055 0.900 l. 19583 0.40425 O. 20005 0.910 l. 2091 2 0.36694 O. 17 544 O. 920 I .2 2242 O. 32882 O. l 5058 0.930 l .23571 0.28991 O. 12558 0.940 1.24900 0.25029 0.10035 0.950 l .26229 0.20997 0.07495 0.960 l .27559 O. 16901 0.04934 0.970 l .28888 O.l 2743 0.02358 O. 980 l .3021 7 0.08529 -0.00236 0.990 l. 31546 0.04260 -0.01495 l .000 l .32876 -0.00247 -0.00247
FOURTH-STAGE BLADE COORDINATES
ROOTSECTION (HOT RADIUS = 19.09200)
QTR
PERCEN T X X Y TOP Y BOT
0.0 -0.00044 0.68756 0.68756
O.Ol 0 O. 01286 0.731 l l 0.6761 l
0.020 0.0261 5 0.75754 0.68301
0.030 0.03944 0.7827 2 0.69374
0.040 0.05273 0.80674 0.7 0691
0.050 0.06602 0.82968 0.72105
0.060 0.07932 0.851 59 0.73437
0.070 0.09261 0.87253 0.74683
0.080 O. 10590 0.89256 0.7 5853
0.090 O.l l 919 O. 91169 0.76951
O.lO0 O. 13248 0.92999 0.77984
O.125 0.16571 0.97227 0.80306
O.l 50 O. 19894 l .00995 0.82298
O.175 0.23217 l .04337 0.83997
0.200 0.26540 l .07284 0.85433
0.225 0.29863 1.09859 0.86625
0.250 O. 33186 l .12078 0.87591
0.275 0.36509 l. 13955 0.88341
0.300 0.39832 l. 15505 0.88885
0.325 0.43155 l .16735 0.89228
0.350 0.46478 l. 17650 0.89375
0.37 5 0.49801 l. 18256 0.89326
0.400 O. 53124 l .18555 0.89081
0.425 0.56447 l .18546 0.88639
0.450 O. 59770 l .18230 0.87992
0.475 0.63093 l. 17601 0.87135
0.500 0.6641 6 l .16655 0.86058
0.525 0.69739 l. 15384 0.84745
0.550 0.73062 l .13775 0.831 81
0.575 0.76385 l .l 1818 0.81340
0.600 0.79708 l .09496 0.791 88
0.625 0.83031 l .06785 0.76682
0.650 0.86354 l .03662 0.73754
0.67 5 0.89677 l .00092 0.70309
0.700 O. 93000 O. 96035 O. 66208
0.725 0.96323 O. 91436 O. 61403
0.7 50 O. 99646 0.86232 O. 56106
0.775 l .02969 0.80364 0.50512
0.800 l .06292 O. 73797 0.4471 8
0.825 l .0961 5 0.66523 0.38770
O. 850 l .12938 0.58566 0.32704
0.87 5 l. 16261 0.49977 0.26534
O. 900 l .19584 0.4081 9 0.20276
O. 910 l .2091 3 0.3701 4 O. 17750
O. 920 l .22242 0.331 35 O.l 5213
0.930 l .23571 0.29186 0.12663
O. 940 l .24900 0.25170 O.l Ol 03
0.950 l .26230 0.21095 0.07533
O. 960 l .27559 O. 16962 0.04954
0.970 l .28888 O. l 2777 0.02364
O. 980 l .3021 7 O. 08543 -0.00236
0.990 l .31546 0.04264 -O.O1494
I .000 l .32876 -0.00247 -0.00247
FOURTH-STAGE BLADE COORDINATES HEAN SECTION (HOT RADIUS = 21.24500) PERCEN T X X Y TOP Y BOT 0.0 0.00008 0.81316 0.81316 O.OlO 0.01336 0.85744 0.80149 0.020 0.02665 0.88438 0.80852 0.030 0.03994 0.90993 0.81946 0.040 0.05322 0.93416 0.83290 0.050 0.06651 0.95718 0.84728 0.060 0.07980 0.97906 0.86071 0.070 0.09308 0.99987 0.87323 0.080 0.I0637 1.01967 0.88490 0.090 0.I1966 1.03851 0.89580 O.lO0 0.13294 1.05645 0.90599 0.125 0.16616 1.09758 0.92869 0.150 0.19938 1.13383 0.94786 0.175 0.23260 l .l 6561 0.96395 0.200 0.26581 l .l 9329 O. 97726 0.225 0.29903 l .21709 0.98801 0.250 0.33225 1.23723 0.99636 0.275 0.36546 1.25387 1.00243 0.300 0.39868 1.26714 1.00630 0.325 0.431 90 l .2771 3 l .00802 0.350 0.4651 l l .28392 l .00761 0.375 0.49833 1.28753 1.00506 0.400 0.53155 1.28800 1.00036 0.425 0.56477 1.28533 0.99342 0.450 0.59798 1.27946 0.98415 0.475 0.63120 1.27038 0.97245 0.500 0.66442 1.25802 0.95809 0.525 0.69763 1.24225 0.94085 0.550 0.73085 1.22297 0.92038 0.575 0.76407 1.19999 0.89619 0.600 0.79728 1.17310 0.86754 0.625 0.83050 1.14204 0.83337 0.650 0.86372 I.I0645 0.79194 0.675 0.89694 1.06590 0.74267 0.700 0.93015 1.01981 0.68861 0.725 0.96337 0.96749 0.63198 0.750 0.99659 0.90836 0.5?366 0.775 1.02980 0.84210 0.51409 0.800 1.06302 0.76876 0.45350 0.825 1.09624 0.68871 0.39206 0.850 1.12945 0.60257 0.32989 0.875 1.16267 O.SlllO 0.26708 0.900 1.19589 0.41513 0.20369 0.910 1.20918 0.37563 0.17820 0.920 1.22246 0.33557 0.15264 0.930 1.23575 0.29501 0.12696 0.940 1.24903 0.25396 0.I0125 0.950 1.26232 0.21246 0.07544 0.960 1.27561 0.17055 0.04957 0.970 1.28890 0.12826 0.02363 0.980 1.30218 0.08563 -0.00239 0.990 1.31547 0.04267 -0.01494 l .000 1.32876 -0.00247 -0.00247 FOURTH-STAGE BLADE COORDINATES QTR TIP SECTION (HOT RADIUS : 23.03799) PERCEN T X X Y TOP Y BOT 0.0 0.00074 O. 97496 O. 97496 O.Ol 0 0.01402 l . Ol 616 O. 96415 0.020 0.02730 l .04072 0.97060 0.030 0.04058 l .06371 0.98061 0.040 0.05386 l .08540 0.99286 0.050 0.0671 4 l. 10587 l .00606 0.060 0.08042 1 .12523 l .Ol 841 0.070 0.09370 l .14354 l .02986 0.080 O.l 0698 l. 16087 l .04051 0.090 O. 12026 l .17728 l .05040 O.l O0 O. 13354 l .l 9282 l .05960 O.125 0.1667 4 l .2281 3 l .07987 O. 150 O. 19994 l .25883 l .09665 O. 175 0.2331 4 l .28537 l .l I036 0.200 0.26634 l .30804 l .12125 0.225 0.29954 1.32712 1.12954 0.250 0.33274 l .34281 l .13534 0.275 0.36594 l .35526 l. 13878 0.300 0.3991 4 l .36458 l. 13987 0.325 0.43234 1.37087 1.13865 0.350 0.46554 1.37418 1.13509 0.375 0.49874 l .37454 l .12916 0.400 O. 531 94 l .371 95 l .l 2073 0.425 0.5651 5 l .36641 l .I0970 0.450 0.59835 l .35788 l .09583 0.475 0.63155 l .34628 l .07886 0.500 0.66475 1.33154 1.05839 0.525 0.69795 1.31351 1.03384 0.550 0.731 l 5 l .29202 l .00432 0.575 0.76435 1.26688 0.96847 O. 600 0.79755 l .23779 O. 92442 0.625 0.8307 5 l .20441 0.87361 0.650 0.86395 l. 16626 0.81 91 l 0.675 0.8971 5 l .12273 0.76242 0.700 0.93035 l .07300 0.70418 0.725 0.96355 l. 01626 0.64469 0.7 50 O. 99675 O. 951 98 O. 58421 0.77 5 l .02995 0.88006 O. 52288 0.800 l .0631 5 0.80082 0.46080 0.825 l .09635 O. 71497 0.39808 0.850 l .12955 O. 62340 0.33475 0.875 l .l 627 5 0.52704 0.27091 O. 900 l .19595 0.42676 0.20656 0.91 0 l .20923 0.38572 O.18068 O. 920 l .22251 0.34422 O. 15474 0.930 l .23579 0.30227 O.12871 0.940 l .24907 0.25995 O.l 0264 0.950 l .26235 O. 21726 0.07649 0.960 l .27563 O. 17424 0.05026 0.970 l .28892 O.l 3093 0.02401 O. 980 l .30220 0.08732 -0.00234 0.990 l .31548 0.04347 -O.O1495 l .000 l .32876 -0.00247 -0.00247 FOURTH-STAGE BLADE COORDINATES TIP SECTION (HOT RADIUS : 24.83200) PERCENT X X Y TOP Y BOT 0.0 O.OO164 l .l 9628 l .19628 0.01 0 0.01491 l .22908 l .18716 0.020 0.02818 l .24824 l .19170 0.030 0.04145 l .26550 l .19896 0.040 0.05472 l .281 86 l .20788 0.050 0.06800 1.29739 1.21778 0.060 0.08127 l .3121 2 l .22713 0.070 0.09454 l .32609 l .23578 0.080 O. I0781 l .33933 l .24378 0.090 0.12108 1.35188 1.25119 O.lO0 O.13435 l .36375 l .2 5803 0.125 0.16753 l .39065 l .27280 0.150 0.20071 l .41 382 l .28452 0.175 0.23388 l .43350 l .29345 0.200 0.26706 l .44989 l .29973 0.225 0.30024 l .4631 4 l .30349 0.250 0.33342 l .47339 l .30479 0.27 5 0.36659 l .48071 l .30364 0.300 0.39977 l .4851 6 l .30004 0.325 0.43295 l .48680 l .29391 0.350 0.4661 3 l .48565 l .2851 5 0.375 0.49931 l .48169 l .27361 0.400 0.53248 l .47492 I .25902 0.425 0.56566 1.46529 1.24108 0.450 0.59884 l .45274 l .21 925 0.475 0.63202 l .4371 8 l .l 9282 0.500 0.66519 1.41850 l .16060 0.525 0.69837 1.39654 1.12084 0.550 0.731 55 l .371 l l I .07328 0.575 0.76473 l .34196 1.02083 0.600 0.79790 l .30879 O. 96547 0.625 0.83108 1.27119 0.90813 0.650 0.86426 l .22863 0.84928 0.67 5 0.89744 l .l 8042 0.78920 0.700 0.93061 l .12580 0.72807 0.725 0.96379 l .0641 5 0.66605 0.750 0. 99697 0.99512 O. 60321 0.775 l .0301 5 O. 91 870 0.53966 0.800 l .06332 0.83526 0.47546 0.825 l .09650 0.74539 0.41 066 0.850 1.12968 0.64988 0.34528 0.875 l .16286 0.54954 0.27946 0.900 l .19604 0.4451 5 0.21 31 l 0.910 1.20931 0.40241 0.18642 O. 920 l .22258 O. 3591 8 O. 15968 0.930 l .23585 O. 31549 O.13286 0.940 l .24912 0.271 37 O.10600 0.950 l .26239 O. 22686 0.07907 .
O. 960 l .27566 O. 181 98 0.05205 0.970 l .28893 O. l3677 0.02497 0.980 l .30220 0.091 24 -0.0021 6 0.990 l .31548 0.04542 -0.01491 l.O00 1.32875 -0.00244 -0.00244 FOURTH-STAGE VANE COORDINATES ROOT SECTION (HOT RADIUS : 17.60899) PERCENT X X Y TOP Y BOT 0.0 -0.001 53 0.62329 0.62329 O.Ol 0 O.O1315 0.67021 O. 6121 9 0.020 0.02782 0.701 65 0.62070 0.030 0.04250 0.7 3134 0.63345 0,040 0.05718 0.75944 0.64829 0.050 0.07186 0.78606 0.66294 0,060 0.08654 0.81131 0.6771 8 0.070 O. lO122 0.83529 0.69099 0.080 O. 11590 0.85807 0.70437 0.090 O. 13058 0.87971 O. 71732 O.lO0 O. 14526 0.90029 0.7 2982 0.125 O.181 96 0.94739 0.75910 O.150 O. 21 866 0.98885 0.78548 0,175 0,25536 l .0251 7 0.80886 0.200 0.29205 l .05681 0.8291 8 0,225 0.32875 l .08407 0.84634 0.250 0.36545 l .10722 0.86029 0.275 0.40215 l .l 2646 0.87099 0.300 0.43885 l .141 97 0.87839 0.325 0.47555 l. 15388 0.88249 0,350 0,51225 l .16227 0,88326 0.375 0.54894 l. 16722 0.88071 0.400 O. 58564 l. 16879 0.87486 0.425 0.62234 l, 16699 0.86575 0.450 0.65904 l .16181 0.85343 0.475 0.69574 l .l 5324 0.83794 O. 500 0.73244 1.14123 0.81 937 0.525 0.7691 4 l .l 2572 0.79780 O. 550 0.80583 l .10661 0.77330 0.575 0.84253 l .08376 0.74599 0.600 0.87923 l ,05704 0.71 595 0.625 O. 91593 l .02620 0.68329 O. 650 O. 95263 O. 991 O0 0.6481 2 0.67 5 0.98933 0.9511 l 0.61054 0.700 l .02602 0.90613 0.57068 0,725 l .06272 0.85559 0.52862 0.7 50 l .09942 0.79926 0.48449 0.775 l. 1361 2 0.7371 l 0.43838 0.800 l. 17282 0.66946 0.39039 0.825 l .20952 0.59677 0.34062 0.850 l .24622 O. 51 967 0.2891 8 0.87 5 l .28291 0.43879 0.2361 2 0.900 l .31 961 0.35477 O.18157 0.91 0 l ,33429 0.32041 O. 15935 O. 920 l .34897 O. 28564 O. 13690 0.930 l .36365 0.25052 O.11424 O. 940 l .37833 0.21507 0.09136 0.950 1.39301 0.17932 0.06828 0.960 l .40769 0.14327 0.04498 0.970 l .42237 0.10697 0.02150 0.980 l .43705 0,07042 -0.0021 9 0,990 l .45173 0.03366 -0.01803 I .000 l .46641 -0.00572 -0.00572 FOURTH-STAGE VANE COORDINATES QTR ROOT SECTION (HOT RADIUS = 19.178500 PERCENT X X Y TOP Y BOT 0.0 -0.00069 0.7 2757 0.72757 O.Ol 0 O.O1398 0.77184 0.71714 0.020 0.02865 0.801 78 0.72508 0.030 0.04332 0.83055 0.73693 0.040 0.05800 0.85821 0.75058 0.050 0.07267 0.88480 0.76412 0.060 0.08734 O. 91 038 0.77747 0.070 O. 10201 O. 93498 0.79059 O. 080 O. 11668 O. 95865 0.80348 0.090 0.I 3135 0.98140 0.81 614 O.lO0 O. 14602 l .00328 0.82855 O. 125 O. 18270 l .05434 0.85846 O. 150 O. 21 938 l .10050 0.88669 0.17 5 0.25605 l .l 4205 0.91306 0.200 0.29273 l .17925 0.93744 0.225 0.32941 l .21230 0.95965 0.250 0.36609 l .241 37 0.97954 0.27 5 0.40276 l .26658 0.99688 0.300 0.43944 l .28802 l .Ol 153 0.325 O .4761 2 l .30579 l .02325 0.350 O. 51 280 l .31 995 l .031 86 0.37 5 0.54947 l .33050 l .03716 0.400 0.58615 1.33748 1.03894 0.425 0.62283 l .34087 l .03702 0.450 0.65950 l .34062 l .031 24 0.475 0.69618 1.33672 1.02143 0.500 0.73286 1.32908 1.00749 0.525 0.76954 1.31761 0.98930 0.550 0.80621 1.30218 0.96681 0.575 0.84289 1.28263 0.94000 0.600 0.87957 l .25877 0.90888 0.625 0.91625 l .23036 0.87351 0.650 0.95292 l .l 971 2 0.83398 0.675 0.98960 l .15866 0.79039 0.700 l .02628 l .11454 0.74292 0.725 l .06296 l .06417 0.69170 0.750 l .09963 l .00693 0.63695 0.775 l. 13631 0.9421 5 0.57883 0.800 l. 17299 O. 86937 O. 517 57 0.825 l .20967 0.7881 9 0.45335 0.850 l .24634 0.69847 0.38638 0.875 l .28302 0.60024 O. 31685 0. 900 l .31 970 0.49379 O. 24496 0.910 l .33437 0.44898 0.21557 0.920 l .34904 0.40297 O.l 8584 0.930 l .36371 0.35578 0.1 5580 0.940 l .37838 O. 30745 O.l 2542 0.950 1.39305 0.25802 . 0.09474 0.960 l .40772 O. 20754 0.06376 0.970 l .42239 O.15605 0.03248 0.980 l .43707 O. I0359 0.00094 0.990 l .45174 0.05020 -0.01802 l .000 l .46641 -0.00572 -0.00572
FOURTH-STAGE VANE COORDINATES
MEAN SECTION (HOT RADIUS : 20.74800)
PERCENTX X Y TOP Y BOT
0.0 0.00055 0.88029 0.88029
O.Ol 0 O.O1521 0.92137 0.87052
0.020 0.02987 0.94894 0.87772
0.030 0.04453 0.97547 0.88848
0.040 0.05918 l.OOlOl 0.90071
0.050 0.07384 1.02561 0.91282
0.060 0.08850 ] .04930 O. 92475
0.070 0.I0316 1.07210 0.93651
0.080 O. 11782 1 .09405 0.94806
0.090 O. 13248 l .l l 518 0.95943
O.lO0 0.14714 1.13550 0.97060
0.125 0.18378 1.18296 0.99754
O. 150 0.22043 l .22587 l .02302
0.175 0.25708 I .26448 l .04687
0.200 0.29372 I .29900 1.06895
0.225 0.33037 I .32957 l .08904
O. 250 O. 36701 l .35633 l .10696
0.275 0.40366 1.37939 1.12250
0.300 0.44031 1.39882 1.13540
0.325 0.47695 1.41468 1.14542
0.350 0.51360 l .42700 l .15232
0.375 0.55025 l .43579 l . l 5580
0.400 O.58689 l .44105 l. 15561
0.425 0.62354 l .44276 l. l 5149
0.450 0.66019 1.44085 1.14318
0.475 0.69683 l .43527 l. 13047
O. 500 0.73348 l .42593 l. If31 8
0.525 0.77013 l .41270 l .09118
0.550 0.80677 1.39544 1.06440
0.575 0.84342 l .37395 l .03282
O. 600 0.88006 l .34802 O. 99648
0.625 0.91671 l . 31734 0.95548
0.650 0.95336 l .28157 0.90999
0.675 0.99000 l .24028 0.86020
O. 700 l .02665 l .19290 0.80634
0.725 l .06330 l. l 3875 0.74866
O. 750 l .09994 1.07714 0.68745
0.775 l .13659 l .00737 0.62294
0.800 l .17324 0.92899 O. 55543
0.825 l .20988 0.841 61 0.48515
O. 850 l .24653 0.74517 0.41 238
0.875 l .2831 8 0.63974 0.33730
O. 900 l .31982 0.52574 0.26016
O. 910 l .33448 0.47785 0.22877
O. 920 l .34914 0.42869 O. 19708
0.930 l .36380 0.37834 O. 1651 l
O. 940 l .37846 0.32681 O.l 3289
0.950 1.39311 0.27418 0.10040
O. 960 l .40777 0.22046 0.06766
0.970 l .42243 O.16573 0.03469
O. 980 l .43709 O. l I002 O.OOl 48
0.990 l .45175 0.05338 -0.01803
l. 000 l .46641 -0. 00572 -0.00572
FOURTH-STAGE VANE COORDINATES QTR TIP SECTION (HOT RADIUS = 22.31750) PERCENT X X Y TOP Y BOT 0.0 0.00201 l .06069 l .06069 O.OlO 0.01666 l .10085 l .05108 0.020 0.03130 l .12797 l .05804 0.030 0.04594 l .15384 l .06846 0.040 0.06059 1.17855 1.08004 0.050 0.07523 l .20214 l .09149 0.060 0.08988 l .22470 l. I0272 0.070 0.I0452 l .24626 l .11375 0.080 O. l 1916 l .26688 l .12457 0.090 0.13381 l .28659 l. 13517 O.lO0 0.14845 l .30543 l .14553 0.125 0.18506 l .34895 l .17036 0.150 0.22167 l .38770 l . 19353 0.175 0.25828 1.42201 1.21486 0.200 0.29489 l .4521 6 l .23416 0.225 0.33150 1.47837 1.25122 0.250 0.3681 l l .50080 l .26583 0.275 0.40472 l .51 960 l .27774 0.300 0.44133 l .53486 l .28668 0.325 0.47794 l .54668 l .29238 0.350 O. 51 455 l .55507 l .29459 0.375 0.551 l 6 l .56009 l .29302 0.400 0.58777 l .56171 l .28740 0.425 0.62438 l .55993 l .27747 0.450 0.66099 1.55468 1.26304 0.475 0.69760 l .54592 l .24392 0.500 0.73421 l .53353 l .21999 0.525 0.77082 l .51738 l .l 9116 0.550 0.80743 l .49731 l. 15747 0.575 0.84404 l .47 310 l .11893 0.600 0.88065 l .44449 l .07567 0.625 0.91726 l .41 l 13 l .02786 0.650 0.95387 l .37259 O. 97574 0.675 0.99048 l .328 31 O. 91949 0.700 l .02709 l .27757 0.85945 0.725 l .06370 l .21948 0.79582 0.750 l .lO031 l. l 5320 0.72893 0.775 l .13692 l .07794 0.65902 0.800 l .17353 0.99314 0.58638 0.825 l .21014 0.89852 0.51 120 0.850 l .24675 0.7941 4 0.43379 0.875 l .28336 0.68030 0.35429 0.900 l .31 997 O. 55768 0.27293 0.910 l .33461 0.50633 0.23990 0.920 l .34926 0.45375 O. 20661 0.930 l .36390 0.40001 0.17307 0.940 l .37854 O. 34514 O.13930 0.950 l .39319 0.28924 0.10529 0.960 l .40783 0.23232 0.071 05 0.970 l .42248 O. 17447 0.03661 0.980 l .43712 O.l 1572 O.OOl 94 0.990 l .45176 0.0561 5 -0.01804 l .000 l .46641 -0. 00572 -0.00572
FOURTH-STAGE VANE COORDINATES
TIP SECTION (HOT RADIUS = 23.88699)
PERCENTX X Y TOP Y BOT
0.0 0.00370 l .26887 l .26887
O.Ol 0 0.01833 l .30908 l .25925
0.020 0.03296 l .33632 l .2661 9
0.030 0.04758 l .36182 l .27653
0.040 0.06221 l .38575 l .28780
0.050 0.07684 l .40825 l .29879
0.060 0.09146 l .42944 l .30949
0.070 0.10609 1.44943 1.31990
0.080 O. 12072 l .46829 l .33001
0.090 O. 13534 l .48609 l .33980
O. l O0 O. 14997 l .50291 l .34927
O. 125 O. 18654 l .54098 l .37143
0.150 0.22311 1.57392 1.39132
0.17 5 0.25967 l .60223 l .40872
0.200 0.29624 1.62633 1.42346
0.225 0.33281 1.64651 1.43531
O. 250 0.36938 l .66301 l .44406
0.275 0.40595 l .67601 l .44948
0.300 0.44251 1.68564 1.45136
0.325 0.47908 l .69200 1.44948
0.350 0.51 565 l .6951 8 l .44366
0.37 5 0.55222 l .69522 l .43368
0.400 O. 58878 l .6921 l l .41 940
0.425 0.62535 l .68584 l .40069
0.450 0.661 92 l .67638 l .37747
0.475 0.69849 l .66366 l .34965
0.500 0.73505 l .64757 l .31 726
0.525 0.77162 l .62799 l .28030 0.550 0.80819 1.60472 1.23887 0.575 0.84476 l .57753 l .19306 0.600 0.881 33 l .54612 l .14303 0.625 0.91789 l. 51008 l .08892 O. 650 0.95446 l .46889 l .03097 0.675 0.99103 l .42185 0.96935 0.700 1.02759 l .36800 0.90431 0.725 l .0641 6 l .30622 0.83602 0.750 l .I0073 l .23544 0.76477 0.775 I. 13730 l .15464 0.69071 0.800 l. 17387 l .06320 O. 61409 0.825 1.21043 0.96087 0.53507 0.850 l .24700 0.84788 0.45387 0.87 5 l .28357 0.72485 0.37063 0.900 l .32014 0. 59279 0.28555 0.91 0 l .33476 0.53767 0.25103 0.920 l .34939 0.48134 0.21 624 0.930 l .36402 0.42388 O. 18120 0.940 l .37864 O. 36536 O.l 4591 0.950 l .39327 0.30587 O.l I038 0.960 l .40790 0.24543 O. 07462 0.97 0 l .42253 0.18415 0.03863 0.980 l .43715 0.12206 0.00243 0.990 l .45178 0.05922 -0.01805 l .000 l .46641 -0.00573 -0.00573 FIFTH-STAGE BLADE COORDINATES ROOT SECTION (HOT RADIUS = 17.74400) PERCEN T X X Y TOP Y BOT O.O001 5 0.47744 0.47744 0.0 O.OlO O.O1372 0.5041 5 0.46878 0.020 0.02758 0.51 946 0.471 54 0.030 0.04144 0.53334 0.47653 0.05531 0.54674 0.48284 0.040 0.050 0.06917 0.55966 0.48921 0.08304 0.5721 l 0.49518 0.060 0.09690 0.5841 l 0.50078 0.070 O.l I076 O. 59565 O. 50602 0.080 0.090 0.12463 0.60674 0.51089 O. 13849 0.61739 0.51 543 O.lO0 0.125 O. 17315 0.6421 3 0.52534 0.150 0.20781 O. 66421 O. 53332 0.175 0.24247 0.68372 0.53947 0.27713 0.70072 0.54390 0.200 0.225 O. 31179 0.71524 0.54670 O. 34644 0.7 2736 O. 54793 0.250 0.275 0.381 l 0 0.7371 0 0.54766 0.300 0.41 576 0.74450 O. 54593 0.45042 0.74958 0.54279 0.325 0.350 0.48508 O. 75238 O. 53830 0.375 0.51974 0.75291 O. 53247 0.400 0.55440 0.75118 O. 52534 0.425 0.58906 0.74721 O. 51 693 0.450 0.62372 0.7 4098 O. 50726 0.475 0.65838 0.73251 0.49634 0.500 0.69303 0.7 21 79 0.4841 9 0.525 0.72769 0.70881 0.47081 0.550 O. 76235 0.69355 0.45620 0.575 0.79701 0.67600 0.44038 0.600 0.831 67 0.65613 0.42332 0.625 0.86633 0.63391 0.40503 0.650 0.90099 0.60930 0.38549 0.93565 0. 58227 0.36469 0.675 0.700 0.97031 O. 55276 0.34261 0.725 l .00496 0.52073 0.31923 0.750 l .03962 0.4861 0 O. 29453 0.775 l .07428 0.44883 0.26845 0.800 I. 10894 0.40889 0.24097 0.825 I. 14360 0.36631 O. 21 204 0.850 l .17826 0.321 13 0.18162 0.875 l .21292 0.27343 0.14962 0.900 I .24758 O. 22330 O. l 1600 0.910 l .26144 0.20259 0.10206 0.920 l .27530 O. 1 81 53 0.08786 0.930 l .28917 O.1601 l 0.07 336 0.940 l .30303 O. 13835 0.05857 0.950 l .31690 O.l 1626 0.04348 0.960 l .33076 0.09384 0.02808 0.970 l .34462 0.071 l l O.Ol 236 0.980 l .35849 0.04806 -0.00369 0.990 l. 37235 O. 02473 -0. 01471 l .000 l .38621 -0.00228 -0.00228 FIFTH-STAGE BLADE COORDINATES (HOT RADIUS = 19,79700) ROOT SECTION QTR PERCENT X X Y TOP Y BOT 0.0 O.OOl 88 0.44382 0.44382 O.O1 0 O.O1532 0.47595 0.43483 0.020 0.02877 0.49497 0.43923 0.030 0.04221 O. 51299 0.44630 0.040 0.05565 O. 53038 0.45494 0.050 0,06910 0.5471 8 0.46389 0.060 O. 08254 O. 56338 0.47252 0.070 0.09598 0.57899 0,48083 0.080 O. 10943 0.59404 0.48883 0.090 O. 12287 0.60851 0.49652 O. lO0 0.13631 0.62244 0.50389 O.l 25 O,16992 0.65488 0.52091 O.l 50 O. 20353 0.68404 O. 53591 0.175 0.23714 0.71002 0.54883 0.200 0.27075 0.73291 0.55967 0.225 0.30436 0.75278 0,56839 0,250 O. 33797 0.76969 0.57498 0.275 0.371 57 0.78371 0.57940 O. 300 0.40518 0.79486 0. 58166 0,325 0.43879 0.80318 0.58172 0.350 0.47240 0.80871 0.57959 0.37 5 0.50601 0.81143 0.57527 O. 400 O. 53962 0.81136 O. 56874 0.425 0.57323 0.80850 0.56002 0.450 0.60683 0.80283 O. 54910 0.475 0.64044 0.79433 0.53601 0.500 0.67405 0.78298 O. 52074 0.525 0.70766 0,76872 0,50333 0.550 0.741 27 0.7 51 52 0.48377 0.57 5 0.77488 0.73130 0.4621 l 0.600 0.80849 0.70800 0.43837 0.625 0.84209 0.68153 0.41257 0.650 0.87570 O. 65177 O. 38475 0.67 5 0.90931 O. 61862 0.35494 0.700 0.94292 0.581 92 0.32317 0.725 0.97653 0.54153 0.28949 0,7 50 l .OlOl 4 0.49726 0.25393 0.77 5 l .04374 0.44896 O, 21653 0.800 l .07735 0.39661 O. 17733 0.825 l .11096 0.34024 O.l 3639 0.850 l .14457 0.27999 0.09373 0.87 5 l .1781 8 O. 21606 0.04941 O. 900 l .21179 0.14869 0.00348 O. 91 0 l .22523 0.12085 -0.01535 0,920 l .23867 0.09253 -0.03441 0.930 l .2521 2 0.06374 -0.05373 0.940 l .26556 0.03450 -0.07329 0.00484 -0.09308 1.27900 0.950 -0.02523 -0.I1312 1.29245 0.960 -0 5568 -0.13339 1.30589 0.970 -0 8652 -0.15390 1.31933 0.980 -0 1770 -0.16470 1.33278 0.990 -0. 5184 -0,15184 1.34622 1.000 FIFTH-STAGE BLADE COORDINATES MEAN SECTION (HOT RADIUS = 21.85100) PERCENT X X Y TOP Y BOT 0.0 0.00882 0.46390 0.46390 0.02180 0.50030 0.45401 O.OlO 0.020 0.03479 O. 52243 0.45930 0.030 0.04777 0.54355 0.46764 0.040 0.06075 0.56374 0.47783 0.050 0.07374 O. 58306 0.48853 0.08672 0.601 51 0.49903 0.060 0.09971 0.61 91 5 0.50931 0.070 O. 11269 0.63600 0.51 936 0.080 O.12567 0.65208 0.52916 0.090 O.lO0 0.13866 O. 66743 O. 53871 O.171 l 2 0.70269 0.56145 0.125 O. 20358 O. 7 3376 O. 58240 0.150 0.175 0.23604 0.76090 0.60139 0.200 0.26850 0.78433 O. 61 825 0.30096 0.8041 9 0.63276 0.225 0.250 0.33342 0.82064 0.64474 0.275 0.36588 0.83379 0.65399 0.300 0.39834 0.84375 O. 66032 0.325 0.43080 0.85058 0.66355 0.350 0.46326 0.85435 0.66354 0.375 0.49572 0.85510 0.66013 0.400 O. 52818 0.85286 O. 65328 0.425 0.56064 0.84766 O. 64290 0.450 0. 5931 0 0.83947 0.62903 0.475 0.62556 0.82831 0.61168 0.500 O. 65802 0.81414 O. 59096 0.525 0.69048 0.79691 O. 56698 0.550 0.72294 O. 77660 0.53991 0.575 0.75540 0.7531 2 0.50990 0.600 0.78786 0.72637 0.4771 6 0.625 0.82032 0.69626 0.44188 0.650 0.85278 0.66264 0.40426 0.675 0.88524 0.62537 0.36449 0.91770 O. 58424 O. 32275 0.700 0.725 0.9501 6 0.53902 0.27924 0.750 O. 98262 0.48952 O. 2341 0 0.775 l .O1508 0.43560 0.18748 0.800 l .04754 0.37725 0 .l 3953 0.825 l .08000 O. 31457 0.09037 0.850 l .11246 O. 24774 0.0401 3 0.875 l .14492 0.17702 -O.Oll 12 0.900 l. 17738 0.I0274 -0.06328 0.910 l. 19037 0.07 210 -0.08437 0.920 l .20335 0.04097 -0. I0559 0.930 l .21633 0.00936 -0. 12693 0.940 l .22932 -0. 02269 -0. 14838 0.950 l. 24230 -0.05517 -0.16995 0.960 l .25529 -0.08806 -0. 191 64 0.970 l .26827 -0.12135 -0. 21342 0.980 l .281 25 -0. 15499 -0.23503 0.990 l .29424 -0.18899 -0. 24045 l .000 l .30722 -0.22597 -0.22597 FIFTH-STAGE BLADE COORDINATES TIP SECTION (HOT RADIUS = 23.90401) QTR PERCENT X X Y TOP Y BOT 0.0 -O.Ol 089 O. 55933 0.55933 O.Ol 0 0.00229 0.59048 0.55036 0.020 O. O1548 O. 60895 O. 55441 0.030 0.02866 0.62642 0.56104 0.040 0.041 84 0.6431 2 0.56914 0.050 0.05502 0.6591 0 0.57748 0,060 0.06820 O. 67437 O. 58563 0.070 0.081 38 0.68895 0. 59357 0.080 0.09456 0.70286 0.60133 0.090 0.I0774 0.71 61 4 0.60886 O. I00 O.12092 0.72878 0.61617 0.I 25 O.15388 0.75774 0.63343 O. 1 50 O. 18683 0.7 8309 O. 64908 0.175 O. 21978 0.80499 0.66296 0.200 0.25273 0.82360 0.67489 0.225 0.28569 0.83903 0.68467 O. 250 O. 31864 O. 85140 O. 69208 0,275 0,35159 0.86077 0.69694 O, 300 O, 38455 0.86722 0.69904 0.325 0.41750 0.87078 0.69816 0.350 0.45045 0.87148 0.6941 5 0.37 5 0.48340 0.86935 0.68683 0.400 O. 51 636 0.86438 O. 67610 0.425 0.54931 0.85658 0.66190 0.450 0.58226 0.84591 0.6441 9 0.475 0.61522 0.83233 0.62300 O. 500 O. 6481 7 0.81 580 O. 59840 0.525 0,681 l 2 0.79625 0.57052 0.550 0,71407 0.77358 0.53949 0.57 5 0.74703 0.74769 0.50552 0.600 0.77998 0.71 845 0.46879 0.625 0.81293 0.68571 0.42951 O. 650 0.84588 O. 64925 O. 38789 0.675 0.87884 0.60886 0.34413 0.700 0.91179 0.56425 0.29842 0.725 0.94474 O. 51 517 0.25095 0.7 50 O. 97770 0.46143 0.201 90 0.775 l .01065 0,40295 O. l 5140 0.800 l .04360 0.33972 0.09961 0.825 l .07655 0.27185 0.04665 O. 850 l .10951 O.l 9951 -0.00737 0.875 l .l 4246 O.l 2298 -0.06232 O. 900 l .17541 0.04255 -0. II 81 4 0. 910 l .I 8859 0.00937 -0.14069 O. 920 l .20177 -0.02436 -0.16336 0,930 l ,21495 -0.05861 -0. 18614 O. 940 l ,22814 -0.09338 -0.20903 0.950 l .24132 -0.12861 -0.23202 0.960 l .25450 -O.16432 -0.2551 3 0.970 l .26768 -0.20047 -0.27832 O. 980 l .28086 -0. 23703 -0. 301 62 O. 990 l .29404 -0.27399 -0.321 31 I .000 l .30722 -0.31 281 -0.31 281 FIFTH-STAGE BLADE COORDINATES TIP SECTION (HOT RADIUS = 25.95799) PERCEN T X X Y TOP Y BOT 0.0 -O.O1771 0.66940 0.66940 O.OlO -0.00446 0.69252 0.66053 0.020 0.00879 0.70548 0.661 88 0.030 0.02204 0.71746 0.66520 0.040 0.03529 0.72900 0.66941 0.050 0.04854 0.7401 l 0.67360 0.060 0.06178 0.75081 0.67771 0.070 0.07503 0.76109 0.68172 0.080 0.08828 0.77095 0.68563 0.090 O.l O153 0.78040 0.68945 O.lO0 O.11478 0.78944 0.6931 5 0.125 0.14791 0.81023 0.70190 O.150 O. 18103 0.82847 0.70980 0.175 0.21 41 5 0.8441 8 0.71671 0.200 0.24728 0.85737 0.7 2243 0.225 0.28040 0.86804 0.72676 0.250 0.31 352 0.8761 7 0.72941 0.27 5 0.34665 0.88176 0.73010 0.300 0.37977 0.88479 0.72845 0.325 0.41289 0.88523 0.7241 2 0.350 0.44602 0.88303 0.71 668 0.375 0.4791 4 0.87816 0.70578 0.400 O. 51226 0.87055 0.691 08 0.425 0.54539 0.8601 5 0. 67241 0.450 O. 57851 0.84688 0.64972 0.475 0.61163 0.83064 0.62309 O. 500 0.64476 0.81133 O. 59278 0.525 0.67788 0.78884 O. 55910 0.550 0.71 lO0 0.76302 0.52245 0.575 0.7441 3 0.73372 0.48322 0.600 0.77725 0.70073 0.44178 0.625 0.81037 0.66384 O. 39846 0.650 0.84350 0.62280 0.35357 0.675 0.87662 0.57737 0.30735 0.700 0.90974 0.52748 0.26001 0.725 0.94287 0.47317 0.21172 0.750 0.97599 0.41465 O.16262 0.775 l .0091 l 0.35221 O.l 1285 0.800 l .04224 0.28624 0.06249 0.825 l .07536 O. 2171 4 O.O1162 0.850 l. I0848 0.14531 -0.03966 0.875 l. 141 61 0.0711 3 -0.09134 0.900 l. 17473 -0.00506 -0. 14334 0.910 l. 18798 -0.03604 -0. 16422 0.920 l .201 23 -0.06727 -0. 1851 5 0.930 l .21448 -0.09875 -0.20612 0.940 l .22773 -0. 13045 -0. 2271 3 0.950 l .24098 -0. 16236 -0.24817 0.960 1.25422 -0.19448 -0.26926 0.970 1.26747 -0.22677 -0.29037 0.980 l .28072 -0.25924 -0.31 152 0.990 l .29397 -0.29188 -0.33208 l .000 l .30722 -0.3261 0 -0.3261 0
FIFTH-STAGE VANE COORDINATES(WANGEA2)
ROOTSECTION (HOT RADIUS = 17.70000)
PERCENTX X Y TOP Y BOT
0.0 -0.00044 O. 38687 0.38687
O.Ol 0 0.01400 0.42832 0.37688
0.020 0.02845 0.45635 O. 38404
0.030 0.04289 0.48317 0.39480
0.040 0.05733 0.50884 0.40691
0.050 0.07177 0.53342 0.41870
0.060 0.08622 0.55696 0.4301 l
0.070 O. 10066 0.57950 0.44113
0.080 O. 11510 0.60109 0.45179
0.090 0.12955 0.62176 0.46206
O.lO0 O.14399 0.641 57 0.471 95
0.125 0.18009 0.687 41 0.49504
O.l 50 0.21 620 0.72840 0.51 582
0.175 0.25231 0.76487 0.53429
0.200 0.28841 0.79710 0.55047
0.225 0.32452 0.82533 0.56439
0.250 0.36063 0.84976 0.57607
0.27 5 0.39673 0.87056 0.58552
0.300 0.43284 0.88786 0.59276
•0.325 0.46895 0.90178 0.59781
0.350 0. 50505 O. 91240 0.60069
0.375 0.54116 0.91980 0.60141
0.400 O. 57727 0.92404 0.59999
0.425 O. 61337 0.92515 0.59645
0.450 O. 64948 0.9231 8 O. 59080
0.47 5 0.68559 0.91 81 2 0.58306
O. 500 O. 721 69 O. 90999 0.57324
0.525 0.75780 0.89876 0.561 35
O. 550 0.79391 0.88442 0.54742
0.575 0.83001 0.86692 0.53145
0.600 0.86612 O. 84621 O. 51347
0.625 0.90223 0.82221 0.49346
0.r650 0.93833 0.79484 0.471 48
0.675 0.97444 0.76397 0.44749
0.700 1.01054 0.72951 0.421 56
0.725 1.04665 0.69126 0.39365
0.750 l .08276 0.64904 0.36379
0.77 5 1 .11886 0.60266 0.33201
0.800 1 .15497 O. 55201 0.29829
0.825 l .l 9108 0.49700 0.26266
0.850 1.22718 0.43772 0.22513
0.87 5 1.26329 0.37422 0.18571
0.900 l .29940 0.30671 0.14441
0.91 0 l. 31384 0.27861 O.l 2737
0.920 l .32828 O. 24993 O. 11002
0.930 1.34272 0.22068 0.09238
O. 940 l .35717 O.l 9085 0.07444
0.950 1.371 61 0.16050 0.05619
0.960 l .38605 O.12962 0.03766
0.970 I .40049 0.09825 0.01884
0.980 l .41494 0.06638 -0.00030
0.990 l .42938 0.03405 -O.Ol 370
l .000 l .44382 -0.001 35 -O.OOl 35
FIFTH-STAGE VANE COORDINATES ROOT SECTION (HOT RADIUS = 19.66000) QTR PERCENT X X Y TOP Y BOT -0.00006 0.58849 0.58849 0.0 O.OlO 0.01438 0.62893 0.57871 0.02882 O. 65625 O. 58563 0.020 0.030 0.04326 0.68240 O. 59604 0.040 0.05770 0.70745 0.60769 0.050 0.07 21 4 0.7 3145 0.61 902 0.060 0.08658 0.7 5444 0.62998 O.10102 0.77648 0.64057 0.070 0.080 O. l 1546 0.79759 0.65080 0.090 O. 12989 0.81782 0.66067 O.lO0 O. 14433 0.8371 9 0.67015 0.125 0.18043 0.88204 0.69224 0.150 0.21 653 0.9221 2 0.71198 0.175 0.25262 0.95773 0.72937 0.200 0.28872 0.98913 0.74434 0.225 0.32482 l .Ol 654 0.75692 0.250 0.36091 l .04014 0.76706 0.275 0.39701 l .06005 0.77477 0.300 0.43311 l .07642 0.78002 0.325 0.46921 l .08931 0.78282 0.350 O. 50530 l .09881 0.78314 0.375 0.54140 l .10498 0.78100 0.400 O. 57750 l. 10785 0.77637 0.425 0.61359 l .10742 0.76926 0.450 O. 64969 l .I0371 0.7 5970 0.475 0.68579 1 .09671 0.74764 0.500 0.72188 1.08637 0.73314 0.525 0.75798 1.07267 0.71 618 0.550 0.79408 1.05553 0.69677 0.575 0.83018 1.03486 0.67493 0.600 0.86627 1 .01 057 0.65067 0.625 0.90237 0.98250 O. 62402 0.650 0.93847 0.95051 O. 59500 0.675 0.97456 O. 91438 O. 56361 0.700 1.01 066 0.87388 0.52990 0.725 1.04676 0.82870 0.49387 0.750 1.08285 0.77852 0.45556 0.775 1.11895 0.72299 0.41500 0.8OO 1 • 15505 0.661 99 0.37223 0.825 1.19114 O. 59546 0.32724 0.850 1.22724 O. 52359 O. 28011 0.875 1.26334 0.44664 O. 23083 0.900 1.29944 0.36502 0.17948 0.910 1 .31388 0.33116 O. 1 5836 0.920 1.32831 0.29665 O.1 3691 0.930 1 .34275 0.26153 O.11 514 0.940 1 .3571 9 0.22583 0.09304 0.950 1 .37163 O. 18956 0.07062 0.960 1.38607 O. 15277 0.04789 0.970 1 .40051 O. 11 549 0.02483 0.980 1.41 495 0.07773 O.OO1 47 0.990 1 .42939 0.03952 -O.O1 37 0 l .000 1 ,44382 -O.OO135 -O.OO1 35 FIFTH-STAGE VANE COORDINATES MEAN SECTION (HOT RADIUS : 21.62000) PERCENT X X Y TOP Y BOT 0.0 0.00032 0.78462 0.78462 O. Ol 0 0.01475 0.82168 0.77540 0.020 0.0291 9 0.84611 0.78149 0.030 0.04362 0.86963 0.79073 0.040 0.05806 0.89226 O. 801 06 0.050 0.07249 0.91404 O.811 l 5 0.060 O. 08693 O. 93500 0.82095 0.070 O. I0136 0.9551 6 0.83046 0.080 O. 11580 0.97455 0.83969 0.090 O. 13023 0.9931 8 0.84861 O. lO0 O. 14467 l .Ol I09 0.85723 0.125 O.l 807 6 1.05278 0.87739 O.150 O. 21 684 1 .09030 0.89554 0.17 5 0.25293 1 .I 2382 0.91158 0.200 O. 28902 1 .l 5355 O. 92541 0.225 0.3251 l 1 .17962 0.93696 0.2 50 O. 36119 1.2021 2 O. 94614 0.275 0.39728 1.22118 0.95285 0.300 0.43337 1.23686 0.95703 0.325 0.46946 1 .24920 0.95860 0.350 O. 50554 l .25824 0.95749 0.375 0.54163 1.26399 0.95362 0.400 O. 57772 l .26647 O. 94696 0.425 0.61381 l .26564 0.93743 0.450 0.64990 l .261 48 0.92499 0.475 0.68598 1 .25395 0.90962 0.500 0.72207 1.24295 0.89129 0.525 0.7581 6 l .22843 0.86999 0.550 0.79425 l .21 024 0.84570 0.575 0.83033 l. 18826 0.81845 0.600 0.86642 l .16232 O. 78824 0.625 0.90251 l. 1321 9 0.75509 0.650 O. 93860 l .09765 0.71908 0.67 5 0.97468 l .05834 0.6801 9 0.700 l .01 077 l .Ol 394 0.63853 0.725 l .04686 0.96390 0.5941 2 0.750 l .08295 0.90770 O. 54706 0.775 l .l 1903 0.84479 0.49740 0.800 l. 15512 0.77491 0.44523 0.825 l. 191 21 0.69796 0.39062 0.850 l .22730 O. 61 422 0.33366 0.875 l .26338 0.52409 0.27444 0.900 l .29947 0.42822 0.21 305 O. 910 l. 31391 0.38841 O. 18790 0. 920 l .32834 O. 34783 O. 16242 0.930 l .34278 0.30653 O. 13663 O. 940 l .35721 O. 26455 O. 1 I050 0.950 l .37165 0.22193 0.08407 0.960 l .38608 O.17871 0.05734 0.970 l .40052 O.13493 0.03029 O. 980 l .41495 0.09063 0.00296 0.990 l .42939 0.04585 -O.O1369 l .000 l .44382 -O.OO134 -O.OOl 34 FIFTH-STAGE VANE COORDINATES TIP SECTION (HOT RADIUS = 23.58000) QTR PERCENT X X Y TOP Y BOT 0.0 0.00067 0.97251 0.97251 0.01 0 0.01 511 1.00440 0.96384 0.020 0.02954 1.02469 0.96844 0.030 0.04397 1.04427 0.97568 0.040 0.05840 1 .06316 O. 98359 0.050 0.07283 1.08137 0.99133 0.060 0.08726 1.09892 O. 99886 0.070 O. 10169 1.11 580 1 .00620 0.080 O. 11613 1. 13205 1.01 334 0.090 0.I 3056 1. 14767 1.02027 0.I00 0.14499 1 .16269 1.02698 0.125 0.18107 1.19761 1 .04276 O. 150 O. 2171 5 1 .22893 1 .05702 O. 17 5 0.25323 1 .25678 1 .06965 0.200 0.28930 1.28124 1.08047 0.225 0.32538 1.30241 1.08938 0.250 0.36146 1.32034 1.09620 0.27 5 0.39754 1 .33512 1.10078 0.300 0.43362 1.34673 1.10295 0.325 0.46970 1.35524 1.10255 0.350 0.50578 1 .36062 1 .09941 0.375 0,54186 1 .36288 1 .09339 0.400 0.57793 1.36201 1.08430 0.425 0.61401 1 .35794 1.07204 0.450 0.65009 1.35065 1.05648 0.475 0.68617 1 .34005 1 .03751 0.500 0.72225 1.32608 1 .01508 0.525 0.75833 1.30860 O. 98913 O. 550 0.79441 1 .28750 O. 95965 0.575 0.83048 1.26263 0.92667 0.600 0.86656 1.23380 0.89024 0.625 0.90264 1.20076 0.85044 0.650 0.93872 1 .16327 0.80738 0.675 0.97480 1.12096 0.76116 0,700 1.01 088 1 .07 345 0.71196 0.725 1.04696 1.02021 0.65990 0.750 1.08304 0.96073 0.60517 0.775 1 .I 1 911 0.89453 0.54791 0.800 1.15519 0.821 32 0.48832 0.825 1.191 27 0.74092 0.42651 0.850 1 .22735 0.65342 0.36269 0.875 1.26343 0.55902 0.29698 O. 9(30 1 .29951 0.45815 0.22955 0.91 0 1 .31394 O. 41609 0.2021 l 0.920 1 .32837 0.3731 0 O.17443 0.930 1.34280 0.32923 0.14651 0.940 1.35723 O. 28451 O. l 1835 0.950 1 .371 67 0.23897 0.08996 0.960 1 .38610 O. 1 9268 0.061 36 0.970 1 .40053 O. 14563 0.03256 0.980 1.41496 0.09789 0.00353 0.990 1.42939 0.04949 -0.01370 1 .000 1 .44382 0.001 35 -0.001 35 26O FIFTH-STAGE VANE COORDINATES TIP SECTION (HOT RADIUS = 25.53999) PERCENT X X Y TOP Y BOT 0.0 0.00094 l . l l 157 l .l l 157 O. Ol 0 0.01537 l. 13977 l .10297 0.020 0.02980 l .15724 l .I0632 0,030 0.04423 l. 17407 l. II 196 0.040 0.05865 l .19025 l . 11797 0.050 0.07308 l .20583 l .12388 0.060 0.08751 l .2207"9 l .12968 0.070 O.lO194 l .2351 8 l .13537 0.080 O. 11637 I .24899 l .14093 0.090 O.13080 l .26222 l .14638 O.lO0 0.14523 1.27491 1.15169 0.125 0.18130 l .30427 l .16435 O. 150 0.21737 l .33038 l .17604 O.175 0.25344 l .35335 l .18664 0.200 O. 28952 l .37326 l .19597 0.225 0.32559 1.39019 1.20389 0,250 0.36166 1.40419 1.21019 0,275 0.39773 l .41532 l .21467 0.300 0.43380 1.42358 1.21706 0.325 0.46988 l .42901 l .21709 0,350 O. 50595 l .431 56 l ,21445 0.375 0.54202 l .43125 l .20885 0.400 O. 57809 l .42803 l .19992 0.425 O. 61 416 l .42187 l .18736 0,450 0.65024 1.41269 1.17086 0.475 0.68631 l .40041 l. 15016 0.500 O. 72238 l .38494 l .12506 0.525 0.75845 l .3661 3 l .09546 0.550 0.79452 1.34386 1.06136 0,575 0.83060 l .31791 l .02281 0,600 0.86667 l .28808 0.98001 0.625 0.90274 l .25406 0.93320 0.650 0.93881 l .21 554 0.88267 0.675 0.97489 l. 17206 0,82873 0.700 l .O1096 l .1231 2 0.771 75 0.725 l .04703 l .06803 0.71201 0.7 50 l .08310 l .00624 O. 64985 0.77 5 l .l 1917 0.93720 O. 58553 0.800 l . 15525 0.86060 0.51933 0.825 l. 19132 0.77628 0.45144 0.850 l .22739 0.68440 0.3821 l 0.87 5 l .26346 O. 58523 0.31146 0.900 l .29953 0.47930 0.23970 0.91 0 l .31396 0.43517 O. 21070 0.920 l .32839 O. 39008 O.181 55 0.930 l .34282 0.3441 0 O. 15226 0,940 l .35725 0.29724 O. 12283 0,950 1.37168 0.24957 0.09326 0.960 l .38611 0.20112 0.06357 0.970 l .40054 O.l 5195 0.03376 0.980 l .41496 O. 10207 0.00383 0.990 1.42939 0.05154 -0.01369 l .000 l .44382 -O.OO134 -O.OO134
APPENDIX E
APPENDIX E LOW-PRESSURE TURBINE EXIT GUIDE VANE AIRFOIL COORDINATES ROOT LEADING EDGE SECTION (HOT RADIUS = 17.82300) PERCENT X X Y BOT Y TOP 0.0 0.0 -l. 58466 -l. 58466 0.010 0.06332 -I .50484 -l .62788 0.020 0.12664 -1. 44639 -i .59315 0.030 -I .55836 0. 18996 -I .38943 0.040 0.25328 -i. 33392 -i .52401 0.050 0.31660 -1.27980 -1.49010 0.060 0.37992 -I .22703 -1.45662 0.070 - I . 17558 -1.42357 0.44324 O. 080 0.50656 -I. 12540 - i. 39095 0.090 0.56988 -i .07646 - 1. 35877 0.100 0.63320 -I. 02872 -I. 32699 0.125 -i .24940 0.79150 -0. 91445 O. 150 O. 94980 -0.80707 -i. 17439 O. 175 i. 10810 -0. 70621 -I.i0193 0.200 I .26640 -0.61151 -I.03198 0.225 -0.96451 1.42470 -0.52269 0.250 I .58300 -0. 43948 -0.89948 O. 275 1. 74130 -0.36166 -0.83688 0.300 1.89960 -0. 77668 -0.28903 0.325 2.05790 -0.22139 -0.71885 0.350 2.21620 -0.66338 -0. 15860 -0.61024 O. 375 2. 37450 -0. 10051 0.400 -0.55942 2.53280 -0.04699 0.425 -0.51091 2.69110 0.00207 0.450 2.84940 0. 04678 -0.46469 0.475 3.00770 -0.42075 0.08722 0.500 -0.37908 3.166OO 0. 12347 0.525 3.32430 0. 15560 -0.33967 0.550 3.48260 0. 18366 -0.30253 0.575 3.64090 -0.26763 0.20771 0.600 -0.23499 3.79920 O. 22780 -0.20460 0.625 3. 95750 O. 24395 0.650 4.11580 0.25620 lO. 17645 0.675 4. 27410 0.26456 -0.15056 0.700 -0. 12693 4.43240 0.26905 0.725 4.59070 0.26968 -0. 10556 0.750 -0.08645 4.74900 0.26644 -0.06963 O. 775 4.90730 0.25934 0.800 5.06560 0.24835 -0.05509 0.825 -0. 04286 5.22390 O. 23347 -0.03294 0.850 5.38220 0 21465 0.875 5.54050 0 19187 -0.02535 0.900 5.69880 0 16509 -0.02011 0.910 -0.01868 5.76212 0 15325 -0.01763 0.920 5.82544 0 14075 0.930 5.88876 -0.01696 0 12760 0.940 5.95208 -0.01667 0 11379 0.950 6.01540 -0.01676 0.09931 0. 960 6.07872 -0.01725 O. 08417 0.970 6.14204 -0.01812 0.06835 0. 980 6.20536 -0.01938 0. 05185 0.900 6.26868 -0.02104 0. 03467 i .000 6.33200 -O.O0001 -0.000 Ol ROOT TRAILIHG EDGE SECTION (HOT RADIUS : 19.46600) PERCENT X X Y TOP Y 80T 0 0 -O.O878O -I.87726 -I.87726 0 010 -0. 02360 - 1. 78131 -I. 92093 0 020 0.04060 -1.70450 ll.87492 0 030 0.10480 -1.63028 -1.82752 0 040 0.16900 -I.55849 -1.78088 0 050 0.23320 -1.48901 -i.73498 0 060 0.29740 -1.42172 -1.68980 0 070 0.36160 -1.35652 -1.64535 0 080 0.42580 -1.29332 -1.60161 0 090 0.49000 -1.23201 -I.55858 0 100 0.55420 -I.17254 -1.51624 0 125 0.71470 -1.03136 -1.41341 0.87520 -0.90014 -1.31479 0 150 1.03570 -0.77806 -1.22026 0 175 0 200 1.19620 -0.66445 -1.12975 1.35670 -0.55871 -i.04316 0 225 0 250 1.51720 -0.46037 -0.96040 1.67770 -0.36900 -0.88140 0 275 1.83820 -0.28425 -0.80608 0 300 0.325 1.99870 -0.20579 -0.73439 0.350 2.15920 -0.13335 -0.66627 2.31970 -0.06670 -0,60164 0. 375 2.48020 -0.00564 -0.54046 0.400 0.425 2.64070 0.05004 -0.48269 0.450 2.80120 0.10047 -0.42827 2.96170 0.14581 -0.37717 0. 475 0.500 3.12220 0.18616 -0.32935 0.525 3.28270 0.22164 -0.28476 3.44320 0.25234 -0.24338 0.550 3.60370 0.27832 -0.20518 0.575 3.76420 0.29965 -0.17013 0.600 0.625 3.92470 0.31638 -0.13820 4.08520 0.32855 -0.i0938 0.650 4.24570 0.33619 -0.08364 0.675 0.700 4.40620 0.33931 -0.06096 4.56670 0.33792 -0.04134 0.725 O. 750 4,72720 0.33201 -0.02476 O. 775 4.8S770 0.32159 -0.01120 5.04820 0.30660 -0.00066 0.800 0.825 5.20870 0.28702 0.00687 5.36920 0.26281 0.01138 0.850 0.875 5.52970 0.23391 0.01289 5.69020 0.20024 0.01140 0.900 5.75440 0.18542 0.00996 0.910 0.920 5.81860 0.16981 0.00804 5.88280 0.15342 0.00564 0.930 5.94700 0.13623 0.00275 O. 940 6.01120 0.11824 -0.00061 O. 950 0.960 6.07540 0.09944 -0.00446 0.970 6.13960 0.07982 -0.00879 6,20380 0.05937 -0.01361 0.980 6.26800 0.03809 -0.01891 0.990 6.33220 -O.O000l -O.O000l I .0o0 I/4 ROOT TRAILING EDGE SECTION (HOT RADIUS = 21.36050) Y BOT PERCENT X X Y TOP -l .88512 0.0 -0.18555 -l .8851 2 0.010 -0.12037 -1.78847 -l .92837 -1.88160 0.020 -0.05519 -I.,71107 -1.83386 0.030 0.00999 -1.63627 -1.78689 0.07517 -I.56392 0.040 -1.74068 -1.49390 0.050 0.14035 -1.69520 0.060 0.20553 -1.42608 -1.65048 0.070 0.27071 -1.36037 -1.60647 0.080 0.33589 -1.29667 -1.56319 0.090 0.40107 -1.23488 -1.52062 0.IO0 0.46625 -1.17492 -1.41724 0.125 0.62920 -1.03262 -1.31812 O. 150 0,79215 -0.90036 -1.22317 0.175 0.95510 -0.77731 -1.13227 0.200 1.11805 -0.66278 -i.04531 0.225 1,28100 -0.55620 -0.96223 0.250 1.44395 -0.45708 -0.88293 0. 275 1.60690 -0.36498 -0.80734 0.300 1.76985 -0.27956 -0.73540 0.325" 1.93280 -0.20049 -0.66703 0.350 2.09575 -0.12751 -0.60216 O. 375 2.25870 -0.06035 -0.54077 0.400 2.42165 0.00117 -0.48279 0.425 2.58460 0.05723 -0.42816 0.450 2.74755 0.I0801 -0.37687 O. 475 2.91050 0.15364 -0.32885 0.500 3.07345 0.19423 -0._8409 0.525 3.23640 0.22990 -0.24253 0.550 3.39935 0.26072 -0.20417 0.575 3.56230 0.28679 -0.16896 0.600 3.72525 0.30814 -0.13690 0.625 3.88820 0.32485 -0.10795 0.650 4.05115 0.33694 -0.08211 0.675 4.21410 0.34443 -0.05935 0.700 4.377O5 0.34735 -0.03967 0.725 4.54000 0.34570 -0.02304 0.750 4.70295 O. 33947 -0.00948 O. 775 4.86590 0.32865 0.00103 0.800 5.02885 0.31321 0.00850 0.825 5.19180 0.29312 0.01291 0.850 5.35475 0.26831 0.01427 0.875 5.51770 0.23874 0.01257 0.900 5.68065 0.20431 0.01103 0.910 S.74583 0.18916 0.00900 0.920 5.81101 0.17322 0.00648 0.930 5.87619 0.15648 0.00347 0.940 5.94137 0.13892 -0.00004 0 950 6.00655 0.12055 -0.00404 0 960 6.O7173 0.10135 1 0 . 0 0 8 5 4 0 970 6.13691 0.08131 -0.01353 0 980 0.06044 6.20209 -0.01901 0 990 6.26727 0.03870 I 000 6.33245 -0. 00001 -O.O0001 MEAN TRAILING EDGE SECTION (HOT RADIUS = 23.25500) PERCENT X X Y TOP Y BOT -0.28730 0.0 -l .75474 -l.7 5474 -0.22110 0.010 -I .66533 -l.7 9712 0.020 -0. 15490 - 1. 75500 -1. s.%19 0.030 -0.08870 -I .52907 -1.71295 0.040 -0.02250 -I .46387 -1.67149 0.050 0.04370 -1.40051 -1.63065 0.060 O. 10990 -I. 33893 -I.59040 0.070 0. 17610 =l. 27906 -1.55075 0.080 0.24230 -I. 22083 -I.51168 0.090 0. 30850 -I. 16419 -1.47320 0.I00 0. 37470 -i. 10908 -1.43529 0.125 0.54020 -0. 97769 -1.34301 0.150 0.70570 -0. 85489 -i.25421 O. 175 0.87120 -0. 74010 -1.16884 0.200 1.03670 -0.63282 -i.08681 0.225 I .20220 -0.53260 -I.00807 0.250 1. 36770 -0.93256 -0.43910 0. 275 1.53320 -0. 35197 -0.86023 0.'300 1.69870 -0. 27095 -0.79103 0.325 1.86420 -0. 19577 -0.72492 0.350 2.02970 -0. 12622 -0.66185 O. 375 2.19520 -0.06212 -0.60179 0.400 2.36070 -0.00329 -0.54470 0.425 2.52620 0.05042 -0.49055 0.450 3.69170 0.09914 -0.43932 0.475 2.85720 0. 14298 -0.3g098 0.500 3.02270 0.18204 -0.34550 0.525 3. 18820 0.21642 -0.30287 0.550 3.35370 0.24618 -0.26307 0.575 3.51920 0. 27138 -0.22608 0.600 3.68470 0.29208 -0.19189 0.625 3.85020 0.30833 -0. 16050 0.650 4.01570 0.32015 -0. 13189 0.675 4.18130 0. 32756 -0.10606 0.700 4.34670 0.33059 -0.08300 O. 725 4.51220 0.32922 -0.06272 0.750 4.67770 0.32347 -0.04521 O. 775 4.84320 0.31331 -0.03049 0.800 5.00870 0. 29873 -0.01854 0.825 5. 17420 0.27968 -0.00939 0 850 5.33970 0.25614 -0.00305 0 875 5.50520 0.22804 0.00048 0 900 5.67070 O. 19533 0.00118 0 910 5.73690 0. 18094 0.00066 0 920 5.80310 O. 16579 -0.00032 0 930 5.86930 0. 14988 -0 00175 O. 940 5. 93550 O. 13320 -0 00365 O. 950 6 00170 O. 11575 -0 00600 O. 960 6 06790 0. 09752 -0 00882 0.970 6 13410 O. 07850 -0 01210 0. 980 6 20030 0. 05869 -0 01585 0.990 6 2.6650 0. 03808 -0 02006 i .000 6 33270 =0.00000 -0.00000 I/4 TIP TRAILING EDGE SECTION (HOT RADIUS = 25.14900) PERCENT X X Y TOP Y 80T 0.0 -0.38605 -I.40187 -l .40187 -0.31886 -1.32629 -1.44336 0.010 0.020 -0.25167 -1.27422 -1.41314 0.030 -0.18448 -!.22332 -1.38327 0.040 -0.I1729 -I.17356 -1.35377 -0.05010 -1.12493 -I.32462 0.050 -I.07739 -1.29582 0.060 0.01709 0.070 0.08428 -1.03092 -1.26737 0.080 0.15147 -0.98551 -1.23927 -0.94111 -1.21151 0.090 0.21866 -0.89773 -i.18410 0.100 0.28585 -i.11705 0.125 0.45382 -0.79354 -I.05212 O. 150 0.62180 -0.69524 O. 175 0.78977 -0.60258 -0.96925 -0.92844 0.200 0.95775 -0.51533 -0.56966 0.225 1.12572 -0.43330 0.250 1.29370 -0.35632 -0.81288 -0.758i0 0.275 1.46167 -0.28421 0.300 1.62965 -0.21684 -0.70529 0.325 1.79762 -0.15408 -0.65444 -0.60554 0.350 1.96560 -0.09580 0.375 2.13357 -0.04192 -0.55858 0.400 2.30155 0.00767 -0.51355 -0.47045 0.425 2.46952 0.05305 0.450 2.63750 0.09429 -0.42926 0.475 2.80547 0.13146 -0.39000 -0.35265 0.500 2.97345 0.16461 0.525 3.14142 0.19380 -0.31722 0.550 3.30940 0.21907 -0.28372 -0.25214 0.575 3.47737 0.24046 -0.22249 0.600 3.64535 0.25800 0.625 3.81332 0.27171 -0.19478 -0.16902 0.650 3.98130 0.28161 0.675 4.14927 0.28773 -0.14522 0.700 0.29005 -0.12340 4.31725 -0.10357 0.725 4.48522 0.28859 0.750 4.65320 0.28334 -0.08576 0.775 -0.06998 4.82117 0.27429 0.800 4.98915 0.26144 -0.05624 0.825 0.24476 -0.04459 5.15712 0.850 -0.03504 5.32510 0.22421 0.875 5.49307 0.19978 -0.02764 0.900 -0.02240 5.66105 0 17142 0.910 5.72824 0 15897 -0,O2O92 0.920 5.79543 0 14587 -0.01980 0.930 -0.01903 5.86262 0 13214 0.940 -0.O1863 5.92981 0 11776 0.950 5.99700 0 10273 -0.01859 0.960 -0.01891 6.06419 0.08704 0.97o -0.01960 6.13138 0.07070 0.980 6.19857 0.05370 -0.02067 8.990 0.03603 -0.02212 6.26576 I .000 6.33295 -0.00000 -0.00000 TIP TRAILING EDGE SECTION (HOT RADIUS = 27.04300)
PERCENT × X
Y TOP Y BOT
0.0 -0.48680
-0.63221 -0.63221
0.010 -0.41860
-0.57440 -0.67660
0.020 -0.35040
-0.54937 -0.66821
0.030 -0.28220
-0.52477 -0.65984
0.040
-0.21400 -0.50061 -0.65148
0.050 -0. 14580 -0.47687
-0.64312
0.060 -0.07760
- O. 45356 -0.63478
0.070
-0.00940 -0.43067 -0.62645 0.080 0. 05880 -0.q0820 -0.61813 0.090 0.12700 -0.38615 -0.60982 0.100 O. 19520 -0.36452 -0.60152 0.125 0.36570 -0.31224 -0.58084 O. 150 0.53620 -0.26253 -0.56023 O. 175 0.70670 -0.21534 -0.53972 0.200 0.87720 -0.17065 -0.5193O O. 225 1. 04770 -0.12844 -0.49899 0. 250 1.21620 -0.08869 -0.47880 0. 275 1.38670 -0.05136 -0.45874 0.300 i .55920 -0.01644 -0.43882 0.325 I. 72970 0.01608 -0.41905 0.350 1.90020 0.04622 -0.39945 0.375 2.0707O O,O7401 -0.38004 0.400 2.24120 O.09945 -0.36083 0.425 2.41170 0.12256 -0.34184 0.450 2.58220 0.14335 -0.32310 O. 475 2,75270 0.16182 -0.30463 0.500 2.92320 0.17799 -0.28645 0.525 3.09370 0.19186 -0.26859 0,55O 3.Z6420 0.20344 -0.25109 0.575 3.43470 0.21274 -0.23397 0.600 3.6O52O 0.21975 -0.21726 0.625 3. 77570 0.22449 -0.20099 0.650 3. 94620 0.22694 -0.16520 0.675 4.11670 0,22712 -0.16992 0. 700 4.28720 0.22501 -0.15517 0.725 4.45770 0.22062 -0.14096 0.750 4.62820 0.21394 -0.12736 0.775 4.79870 0.20498 -0.11433 0.800 4.96920 0.19371 -0.10190 0.825 5.13970 O,18O13 -0.09005 0.850 5.31020 0.16424 -0.07879 0.875 5.48070 0.14603 -0.06811 0.900 5.65120 0.12547 -0.05797 0.910 5.71940 0.11659 -0.05407 0.920 5.78760 0.10734 -0.05025 0.930 5.85580 0.09770 -0.04651 0. 940 5.92400 0.08769 -0.04285 O. 954 5.99220 0.07729 -0.039_6 O. 960 6.06040 0.06652 -0.03575 0.970 6. 12860 0.05536 -0.03232 0. 980 6.19680 0.04382 -0.02895 0.990 6.26500 0.03190 -0.02566 1 .o00 6.33320 -0.00001 -O.O0001 TIP LEADING EDGE SECTION (HOT RADIUS = 26.07001) X Y TOP Y 80T PERCENT X -l .08479 0.0 -0.43455 -I .08_79 -1.12716 0.010 -0.36687 -1.01759 0.020 -0.29919 -i. 10632 -0.97742 0.030 -i .08567 -0.23152 -0.93803 -I. 06524 0.040 -0. 16384 -0.89942 0.050 -0.09617 -0.86158 -I. 04500 0.060 -0.82451 -1. 02495 -0.02849 0.070 -0. 78818 -I .00511 0.03919 -0.98546 0.080 0. 10686 -0. 75260 0.090 0. 17454 -0. 71774 -0.96600 0.100 0.24222 -0.68361 -0. 94674 0.125 0.41141 -0.60158 -0.89943 -0.85333 0. 150 0.58060 -0.52345 O. 175 0. 74979 -0.44977 -0.80841 0.200 O. 91898 -0. 58015 -0.76468 0.225 1.08817 -0.31453 -0.72213 -0.68075 0.250 1. 25736 -0.25280 -0.64055 0. 275 1.42655 -0. 19487 0.300 I .59574 -0.14067 -0.60151 0.325 1.76493 -0.09014 -0.56364 -0.52694 0. 350 1.93412 -0.04318 0. 375 2. 10331 0.00024 -0.49141 0.400 2.27250 0.04018 -0.45706 0.425 2.44169 o. 0767o -0.42391 0.450 -0.39193 2.61088 0. 10983 -0.36117 0.475 2. 78007 O. 13962 0.500 2.94926 0. 16609 -0.33162 0.525 3.11845 0.18927 -0.50329 -0.27622 0.550 3. 28764 0.20920 0.575 3.45683 0.22591 -0.25040 0.600 3.62602 0.23939 -0.22586 -0.20260 0.625 5.79521 O. 24967 -0. 18067 0.650 3.96440 O. 25676 -0.16006 0. 675 4. 13359 0.26067 O. 700 4.30278 0.26139 -0. 14082 0.725 -0. 12296 4.47197 0.25894 -0. 10650 0. 750 4.64116 0.25330 O. 775 0.24447 -0.09147 4.81036 0.800 -0. 07786 4.97955 0.23245 -0.06572 0.825 5. 14874 0.21721 0.850 5.31793 O. 19873 -0. 05506 0. 875 -0.04590 5.48712 O. 17700 0. 900 -0.03826 5.65631 O. 15199 0 910 -0.03563 5. 72398 0.14106 0 920 5.79166 O. 12960 -0.03326 0 930 -0.03112 5.85933 0.11761 0 940 5.92701 O. 10508 -0.02924 0 950 5. 9946 9 -0.02761 0.09201 0 950 -0.02623 6.06236 0. 07840 0 970 6. 13004 0.06424 -0.02511 0 980 6.19772 -0.02424 0. 04955 0 990 6.26539 -0.02365 O. 03430 1 000 6.33307 -O.O0001 -0. O0001
LIST OF ABBREVIATIONS ANDSYMBOLS
ACC
active clearance control
ADP
aerodynamic design point
A1 aluminum
B
constant for the law of the wall
bx
axial chord
C Cel isus (centegrade) c/a cooling airflow Cf skin friction, 2 _w/pU2_ Cl climb cm centimeter pressure coefficient Cp cps cycle per second Cr cruise Cx/U ratio of throughflow velocity to wheel speed deg degree, angle EGV exit guide vane F Fahrenheit FI flight idle FPS flight propulsion system ft feet Z_h specific work H shape factor, _*/e HPC high-pressure compressor HPT high-pressure turbine Z_h/U2 work factor I idl e IC/LS integrated core/l ow spool ID inner diameter IGV inlet guide vane in inch k Von Karman's constant K spring rate LIST OF ABBREVIATIONS AND SYMBOLS (Cont'd) L Iength Ib pound LE leading edge L/H length to height ratio LPC low-pressure compressor platform gap low-pressure turbine meter m airfoil inlet Mach number Ml airfoil exit Mach number mach number Mn N Newton low-pressure rotor speed, rpm Nl OD outer diameter PDR Preliminary Design Review static pressure Ps total pressure Pt pressure at station 3
PAts
total pressure loss PT/PT turbulence energy, u2 + v2 + w2 R radius Re8 Reynolds number, Ume/u, Boundary Layer Momentum Loss Thickness revolutions per minute rpm surface distance S SLTO sea level takeoff SS steady state T tempe ra tu re TE trail ing edge TLpR Transient Liquid Phase Ti ti tan ium T/O takeoff LIST OF ABBREVIATIONS AND SYMBOLS (Cont'd) U fluctuating streamwise component of velocity Dimensionless turbulence intensity _/u-_/v* streamwi se component of turbulence intensity U+ dimensionless velocity, U/v* U wheel speed V _uctuating velocity perpendicular to test wall normal component of turbulence intensity v2 Vm Velocity Ratio, mean - Relationship between the blade wheel speed and amount of turbine work output Vm =/U 2 (No. of Stages)
V
2gJ Ah V* friction velocitY v/-_p traverse component of turbulence intensi ty Wacc active clearance control flow Wae core engine airflow platform axial width Wp normal distance from wall Y y+ dimensionless distance from wall, y v* /u angle eG platform gap angle c_T taper angle airfoil inlet angle airfoil exit angle _2 6 boundary layer thickness Aerodynamic damping 6 aero Z_ difference P density _B bending stress _Brg bearing stress _C compression stress _H hoop stress e angl e e Boundary Layer Momentum Loss Thickness p Kinematic viscosi ty TOO wall shear stress O2 frequency REFERENCES II Leach, K. and Thulin, R.: "Energy Efficient Engine Turbine Transition Duct Model Test Program Technology Report", February 1982, pending publication.
o Sharma, O. P. et al: "Energy Efficient Engine Low-Pressure Turbine Subsonic Cascade Technology Report", January 1982; NASA CR-165592 (PWA 5594-167).
o Gardner, W.B.: "Energy Efficient Engine Low-Pressure Turbine Boundary Layer Technology Report", April 1981; NASA CR-165338 (PWA 5594-141).
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