Energy efficient engine high-pressure turbine detailed design report

TABLE OF CONTENTS Page SECTION 1.0 SUMMARY SECTION 2.0 INTRODUCTION SECTION 3.0 DESIGN OVERVIEW 3.1 Design Goals and Challenges 3.2 High-Pressure Turbine General Description 3.3 Design Performance Data SECTION 4.0 HIGH-PRESSURE TURBINE AERODYNAMIC DESIGN 4.1 Overview lO 4.2 Component Aerodynamic Design lO 4.2.1Flowpath Definition II 4.2.2 Airfoil Definition 4.2.3 High-Pressure and Low-Pressure Turbine Matching 4.2.4 Aerodynamic Efficiency Status 4.2.5 Supporting Technology Programs 4.2.5.1Uncooled Rig Technology Program 4.2.5.2 Supersonic Cascade Program 4.2.5.3 Leakage Program SECTION 5.0 AIRFOIL DURABILITY 30 5.1 Overview 3O 5.2 Turbine Vanes 3O 5.2.1 Vane Cooling Management System 5.2.2 Vane Materials 34 5.2.3 Turbine Vane Durability Assessment 5.3 Turbine Blades 38 5.3.1 Blade Cooling Management System 5.3.2 Blade Materials 42 5.3.3 Turbine Blade Durability Assessment SECTION 6.0 SECONDARY AIRFLOW SYSTEM 48 6.1 Overview 48 6.2 Secondary Flow Summary 6.3 Secondary Flow System Design Features 6.4 Thermal Analysis SECTION 7.0 COMPONENT MECHANICAL DESIGN 7.1 Overview 59 7.2 Turbine Rotor Assembly 7.2.1 General Description 7.2.2 Blades 63 7.2.2.1 Mechanical Design Features 7.2.2.2 Airfoil Vibration Analysis 7.2.3 Blade Attachment 7.2.3.1 Mechanical Design Features 7.2.3.2 Structural Analysis iii PREC-E_ii'_G PAG_ BLA_K NOT FIL,_ED

TABLEOF CONTENTS (Continued)

Page

7.2.4 Disk 78

7.2.4.1 Mechanical Design Features 78

7.2.4.2 Structural Analysis 78

7.2.5 Sideplate and Vortex Plate 80

7.2.5.1 Mechanical Design Features 80

7.2.5.2 Structural Analysis 85

7.2.6 Air Seals 85

7.2.6.1 Mechanical Design Features 85

7.2.6.2 Structural Analysis 85

7.3 Vane and Inner Case 93

7.3.1 General Description 93

7.3.2 Vanes 93

7.3.2. l Mechanical Design Features 93

7.3.2.2 Structural Analysis 96

7.3.3 Inner Support, Tangential on-Board

Injection and High-Pressure CompressorSeal 96

7.3.3.1 Mechanical Design Features 96

7.3.3.2 Structural Analysis 97

7.4 Outer Caseand Outer Airseal 100

7.4.1 Mechanical Design Features lO0

7.4.2 Structural Analysis I04

7.5 Active Clearance Control System 104

7.5.1 General Description IO6

7.5.2 Blade Tip Clearance Definition I07

7.6 Number4 & 5 Bearing Compartmentand Lubrication

System If2

7.6.1 General Description I]2

7.6.2 Bearing Mechanical Design Features If2

7.6.3 Seals ll5

7.6.4 Lubrication System ll5

7.7 Turbine System Weight Summary ll7

SECTION 8.0 HIGH-PRESSURE TURBINE COMPONENT TESTRIG DESIGN

I18

8.1 Introduction

If8

8.2 General Description and Major Features If8

8.3 Mechanical Design 120

8.3.1 Rotating Hardware

8.3.2 Bearing and seals

8.3.3 Air Seals

8.3.4 Static Hardware

8.4 Secondary Flow System and Thrust Balance

8.5 Rig Instrumentation

8.5.1 Performance Instrumentation

8.5.2 Structural Integrity Instrumentation 135

8.6 Facility/Rig Adaptation

SECTION 9.0 CONCLUDING REMARKS

APPENDIX A

VANE ANDBLADE AIRFOIL COORDINATES

APPENDIX B

LIST OF SYMBOLS

REFERENCES

iv LIST OF ILLUSTRATIONS Title Number 2-I Logic Diagram of High-Pressure Turbine Effort 3.2-I Energy Efficient Engine High-Pressure Turbine Component I0 4.2.1-I High-Pressure Turbine Flowpath Definition 4.2.2-I Vane Assumed Profiles For Inlet Temperature and Pressure Loss 4.2.2-2 Vane Root Section Aerodynamic Contour and Pressure Distribution 4.2.2-3 Vane Mean Section Aerodynamic Contour and Pressure Distribution 4.2.2-4 Vane Tip Section Aerodynamic Contour and Pressure Distribution 4.2.2-5 Turbine Vane Stacking 4.2.2-6 Effect of Inlet Temperature and Vane Loss Profile on Blade Inlet Angle 4.2.2-7 Blade Root Section Aerodynamic Contour and Pressure Distribution 4.2.2-8 Blade One-Quarter Root Section Aerodynamic Contour and Pressure Distribution 4.2.2-9 Blade Mean Section Aerodynamic Contour and Pressure Distribution 4.2.2-I0 Blade One-Quarter Tip Section Aerodynamic Contour and Pressure Distribution 4.2.2-II Blade Tip Section Aerodynamic Contour and Pressure Distribution 4.2.2-12 Turbine Blade Stacking of 4.2.5-I Turbine Uncooled Rig Performance Trends Showing Benefits Decreasing the Cx/U and Increasing the AN z Parameter 4.2.5-2 Efficiency Gains Associated with High Blade Reaction Levels 4.2.5-3 Spanwise Distribution of Total Pressure Loss for the Profiled Wall Cascade LIST OF ILLUSTRATIONS (Continued) Number Title

Page

Predicted Vane Cooling Losses Compared to Test Results Effect of Trailing Edge Ejection Flow on Blade Pressure Coefficient 4.2.5-6 Plane Cascade Base Blade, Mean Section Trailing Eage Cooling Air Ejection 4.2.5-7 Results of Blade-Disk Model Testing Showing Leakage in the Attachment Area Is Less Than Predicted Sealing Effectiveness of Rear Sideplate Design Results With and Without W-Seals in the Blade Platform Area 4.2.5-I0 Promising Feather Seal Configurations Evaluated 27 4.2.5-II Energy Efficient Engine High-Pressure Turbine Vane Inner Attachment Leakage Rig 4.2.5-12 Comparison of 0.025 to 0.050 cm (0.010 to 0.020 in) Thick Two-Piece Overlapping Feather Seals 4.2.5-13 Two-piece Overlapped Seals in Ground Versus Electro- discharge Machined Slots 5.2-I Combustor Exit Profile Used for Turbine Vane Durability Assessment for Flight Propulsion System at Hot Day Sea Level Takeoff Operating Conditions 30 5.2.1-1 Turbine Vane Cooling Design 31 5.2.1-2 Turbine Vane Inner Platform Cooling Scheme 33 5.2.1-3 Turbine Vane Outer Platform Cooling Scheme 33 5.2.1-4 Turbine Vane Inner Platform Heat Transfer Coefficients 5.2.1-5 Turbine Vane Outer Platform Heat Transfer Coefficients 33 5.2.1-6 Suction Surface Film Effectiveness 5.2.1-7 Vane Thermal Analysis Results 35 5.2.1-8 Pressure Wall Film Temperatures 36 vi LIST OF ILLUSTRATIONS (Continued) Title Number 5.2.1-9 36 Suction Wall Film Temperatures 5.2.1-I0 37 Vane Surface Temperature Profile 5.2.1-II Vane Surface Heat Transfer Coefficients Used to Determine Surface Temperature Profiles 5.2.3-I Vane Limiting Strain Cycle 5.2.3-2 Predicted Strain 39 5.3-] Turbine Vane Exit Profile for Flight Propulsion System at Hot Day Sea Level Takeoff Operating Conditions 5.3.1-I 41 Turbine Blade Cooling System 5.3.1-2 43 Blade Thermal Analysis Results (Midspan Location) 5.3.1-3 44 Surface Temperature Profile 5.3.1-4 Blade Surface Heat Transfer Coefficients Used to Determine Surface Temperature Profiles 5.3.3-I Blade Limiting Strain Cycle 45 5.3.3-2 Predicted Strain 5.3.3-3 Blade Temperature Comparison with Creep Limits for Flight Propulsion System at Hot Day Sea Level Takeoff Operating Conditions 47 5.3.3-4 Effects of Wall Thickness Tolerance on Blade Life 47 6.2-I 49 High-Pressure Turbine Secondary Flow System Map 6.3-I 52 Secondary Flow System Design Features 6.3-2 52 Mini Tangential On-Board Injection System 6.3-3 53 Flow Characteristics of Blade Coolant Supply System 6.4-I 55 Detailed Turbine Model Used for Thermal Analysis 6.4-2 Typical Transient Response of Turbine Disk During Snap Acceleration/Deceleration 56 vii LIST OF ILLUSTRATIONS (Continued) Number Title

Page

6.4-3 Detailed Turbine Case and Outer Airseal Model Used in Thermal Analysis 57 6.4-4 Results of Temperature Analysis 58 7.1-I High-Pressure Turbine Mechanical Configuration 60 7.2.1-I High-Pressure Turbine Rotor Assembly 60 7.2.1-2 Energy Efficient Engine High Rotor Imbalance Response 61 7.2.1-3 Energy Efficient Engine High Rotor Critical Speeds and Mode Shapes (Percent Rotor Strain Energy) 62 7.2.2-I Turbine Blade Mechanical Configuration 63 7.2.2-2 Radial Taper in Turbine Blade To Minimize Stress 65 7.2.2-3 Degree of Tilt in the High-Pressure Turbine Blade to Achieve the Desired Balance Between Gas Bending Load Stresses and Stresses Resulting from Centrifugal Loads 65 7.2.2-4 Uncoated Blade Wall Thickness and Internal Rib Design 66 7.2.2-5 Core Support Method 67 7.2.2-6 Blade Tip Squealer Geometry 67 7.2.2-7 Graphical Display of Turbine Blade Defined by the NASTRAN Analytical Technique 68 7.2.2-8 Single Crystal Major Axes Orientation 69 7.2.2-9 High-Pressure Turbine Resonance Diagram 70 7.2.2-I0 Blade Vibration Characteristics in First Mode 70 7.2.2-I l Blade Vibration Characteristics in Second Mode 71 7.2.2-12 Blade Vibration Characteristics in Third Mode 71 7.2.2-13 Blade Vibration Characteristics in Fourth Mode 72 7.2.2-14 Blade Vibration Characteristics in Fifth Mode 72 f

i

viii LIST OF ILLUSTRATIONS (Continued)

Title Page

Number 7.2.2-15 Crystallographic Orientation Secondary Orientation 7.2.2-16 Predicted Frequencies Versus Crystal Angle 7.2.2-17 Axial Shift in Crystal Orientation Configuration 7.2.2-18 Predicted Frequency with Modified Blade Attachment Stress Results 7.2.3-I Features 7.2.4-I High-Pressure Turbine Disk Design 7.2.4-2 Elliptical Cooling Air Supply Hole 7.2.4-3 Schematic Showing Disk and Blade Deflections Caused by First Coupled Mode Vibration Boundary Conditions Used in Disk Structural Analysis Flight Propulsion System Disk and Sideplate Calculated Minimum Low Cycle Fatigue Life Levels 7.2.5-I Disk Sideplate and Vortex Plate Design 7.2.5-2 Typical Sideplate Loads 7.2.5-3 Blade and Sideplate Assembly Sequence 7.2.5-4 Anti-Torque Pin Details 7.2.6-I Turbine Seal Designs Rear Thrust Balance Seal 7.2.6-2 Thrust Balance Seal Rotor and Stator Resonance Diagram 7.2.6-3 Thrust Balance Seal Rotor and Stator Coincidence Diagram 7.2.6-4 8g 7.2.6-5 Number 4 Bearing Buffer Seal Assembly 7.2.6-6 Number 4 Bearing Buffer Seal Rotor and Stator Resonance Diagram ix LIST OF ILLUSTRATIONS (Continued) Number Title 7.2.6-7 Number 4 Bearing Buffer Seal Rotor and Stator Coincidence 9O Diagram 7.2.6-8 High-Pressure Compressor Discharge Seal Assembly 91 7.2.6-9 Seal Rotor and Stator High-Pressure Compressor Discharge Resonance Diagram 92 7.2.6-I0 Seal Rotor and Stator High-Pressure Compressor Discharge Coincidence Diagram 92 7.3.1-I Inner Case Mechanical Hi_jh-Pressure Turbine Vane and De sign 93 7.3.2-I Turbine Vane Assembly 94 7.3.2-2 Design Approach to Vane Leakage Control 94 7.3.2-3 Uncoated Vane Minimum Wall Thickness Distribution 7.3.3-I Turbine Inner Support, Tangential On-Board Injection System, and High-Pressure Compressor Seal Land Support System 7.3.3-2 Nozzle Configuration for the Primary Tangential On-Board Injection System 7.3.3-3 Nozzle Configuration for the Secondary or Mini Tangential On-Board Injection System 7.3.3-4 High-Pressure Compressor Discharge Seal Clearance Summary 7.3.3-5 Inner Vane Case Stress Summary 7.3.3-6 lO0 Case Structure Deflection Summary 7.4-1 High-Pressure Turbine Outer Case and Outer Airseal Assembly lOl 7.4.1-I lOl Mechanical Design Features 7.4.1-2 Details of the Outer Airseal I03 Shoe Design 7.4.1-3 Outer Case and Outer Airseal I03 Materials Map X LIST OF ILLUSTRATIONS (Continued) Number Title Page 7.4.2-I Shell Analysis Model 104 7.4.2-2 Temperature Map 105 7.4.2-3 Stress Map I05 7.5.1-I High-Pressure Turbine Active Clearance Control System I06 7.5.2-I Typical Rotor and Case Growth History ]07 7.5.2-2 Rotor and Case Response with Two Cooling Bleed Schedules 108 7.5.2-3 Effect of Mixed Bleed I09 7.5.2-4 Rotor and Case Response with Optimum Mixed Cooling Bleed Schedule llO 7.5.2-5 Resultant Blade Tip Clearances Ill 7.6.1-1 Integrated Core/Low Spool Number 4 and 5 Bearing Compartment - Temperature and Pressure Distribution at Sea Level Takeoff -3oc (+25OF) Operating Conditions ll2 7.6.1-2 Number 4 and 5 Bearing Compartment Bearing and Seal Arrangement I]3 7.6.2-I Number 4 and 5 Bearing Compartment - Locations of Maximum Stress and Deflection 114 7.6.3-I Dry Face Seal Pressure, Temperature and Speed Experience I16 7.6.4.1 Oil Scupper Line for the Number 4 and 5 Bearing Compartment 1]7 8.2-I High-Pressure Turbine Component Test Rig l]9 8.3.1-I High-Pressure Turbine Component Test Rig Critical Speed and Mode Shapes 120 8.3.3-I High-Pressure Turbine Rig Thrust Piston Seal Resonance 123 8.3.3-2 High-Pressure Turbine Rig Thrust Piston Seal Coincidence 123 8.3.3-3 Front Compartment Front Airseal (Thrust Piston Seal) 124 8.3.3-4 High-Pressure Turbine Rig Front Bearing Rear Seal Resonance 125 8.3.3-5 High-Pressure Turbine Rig Front Bearing Rear Seal Coincidence 125 xi LIST OF ILLUSTRATIONS (Continued) Number Title

Page

8.3.3-6 Front Compartment Rear Airseal 8.3.3-7 High-Pressure Turbine Rig High-Pressure Compressor Discharge Seal Resonance 8.3.3-8 High-Pressure Turbine Rig High-Pressure Compressor Discharge Seal Resonance 127 8.3.3-9 Hi gh-Pressure Compressor Discharge Seal 8.3.3-10 Hi gh-Pressure Turbine Rig Thrust Balance Seal Resonance 128 8.3.3-11 Hi gh-Pressure Turbine Rig Thrust Balance Seal Coincidence 129 8.3.3-12 Thrust Balance Seal and Damper 129 8.3.3-]3 Buffer Seal High-Pressure Turbine Rig Number 4 Bearing Resonance 8.3.3-14 High-Pressure Turbine Rig Number 4 Bearing Buffer Seal Coincidence 8.3.3-15 Number 4 Bearing Buffer Seal - Rig 8.3.4-I Active Clearance Control Growth Summary 132 8.4-I High-Pressure Turbine Component Test Rig 133 Secondary F1 ows xii

LIST OF TABLES

Table No. Title

Page

3.2-I Advanced Technology Design Concepts 3.3-I High-Pressure Turbine Predicted Performance at Aerodynamic Design Point 4.l-I General Aerodynamic Parameters 4.2.2-I Gas Triangles 4.2.3-I High-Pressure Turbine Aerodynamics After Restaggering 4.2.4-I High-Pressure Turbine Efficiency Estimate Based on Uncooled Rig Test Results Aerodynamic Design Point 4.2.5-I Turbine Uncooled Rig Efficiency 5.2.2-I Vane Materials and Coatings 5.2.3-I Vane Life 5.3.2-I Blade Materials 5.3.3-I Blade Life 6.2-I Secondary Flow Status Summary 7.2.2-I High-Pressure Turbine Blade Stress Summary 7.2.3-I Blade Attachment Stress Summary 7.2.4-I Disk Life and Stress Summary 7.3.2-I Vane Materials and Coatings I08 7.5.2-I Flight Cycle For Clearance Analysis I09 7.5.2-II High-Pressure Turbine Active Clearance Control System Ill 7.5.2-III High-Pressure Turbine Gapping Requirements Ill 7..5.2-IV High-Pressure Turbine Tip Clearance Results ll6 7.6.3-I Integrated Core/Low Spool Seal Operating Conditions xiii LIST OF TABLES (Cont'd) Tabl e No. Title

Pa e

7.7-I Preliminary Weight summary For Integrated Core/Low Spool High-Pressure Turbine Component ll7 8.3-I High-Pressure Turbine Component Rig Forced Response Results 121 8.3.2-I Front Compartment Operating Conditions 121 8.3.2-11 Operating Conditions of the Rear Compartment Seals xiv

SECTION l.O SUMMARY The high-pressure turbine designed for the Energy Efficient Engine is a single stage configuration that utilizes technology advancements in the areas of aerodynamics, structures and materials to enhance efficiency, durability and performance retention. In addition, the single stage design offers a large savings in initial engine cost, weight, and maintenance cost because of the significant reduction in the number of components, especially expensive air- cooled airfoils.

On the basis of aerothermal-mechanical analyses as well as results from sup- porting technology programs, the high-pressure turbine meets the performance and durability goals for both the integrated core/low spool and the flight propulsion system. The predicted efficiency for the flight propulsion system is 88.8 percent, which exceeds the goal of 88.2 percent. For the integrated core/low spool, the expected test efficiency is 87.1 percent, which exceeds the goal of 86.7 percent. The turbine aerodynamic design utilizes low loss performance features with low through flow velocity to achieve this efficiency.

The turbine airfoils exceed the established durability/life requirements by the use of advanced materials and efficient cooling management. Both the vanes and blades are constructed with advanced high-temperature, high-strength sin- gle crystal alloys and coated with an improved oxidation resistant coating.

For added thermal protection, vane platforms are coated with an advanced ther- mal barrier coating. Acceptable metal temperatures are maintained through the use of impingement, convection and film cooling techniques to minimize cooling without compromising durability. Total cooling requirements are 2.75 and 6.41 percent of core engine airflow for the blade and vane, respectively.

Turbine performance is enhanced by the advancements in sealing technology.

Cooling air leakage is effectively reduced by the use of full ring, boltless sideplates as well as the extensive use of feather seals and W-seals. An ac- tive clearance control system maintains close blade tip clearances throughout the operating range. In this system, the blade tip seal is constructed of an abradable, ceramic material and the blade has an abrasive tip treatment. At design conditions, the calculated clearance of 0.032 cm (0.0126 in) surpasses the goal of 0.047 cm (0.0186 in).

The high-pressure turbine in the integrated core/low spool and envisioned for the future flight propulsion system is also the same design for the High- Pressure Turbine Component Test Rig. This rig will be used to assess turbine aerodynamic behavior and performance before the component is evaluated in the integrated core/l ow spool.

Overall, the high-pressure turbine design for the Energy Efficient Engine re- presents a considerable extension in the state-of-the-art of turbine tech- nology. Much of this technology, especially the high temperature capability materials, will have wide application in derivative and future gas-turbine engines.

SECTION 2.0

INTRODUCTION

The Energy Efficient Engine Component and Development Program, sponsored by the National Aeronautics and Space Administration, is directed towards devel- oping and demonstrating the technology to achieve greater fuel efficiency for future comercial gas-turbine engines. The overall program goals outlined for the program 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 the base Pratt & Whitney Aircraft JTgD-7A base engine. To demonstrate the tech- nology to accomplish these goals, the Energy Efficient Engine Program is organized into four tasks which involve: Task l Flight Propulsion System Analysis, Design and Integration Task 2 Component Analysis, Design and Development Task 3 Core Design, Fabrication and Test Task 4 Integrated Core/Low Spool Design, Fabrication and Test A major accomplishment under the Task 2 effort has been the design of an ad- vanced high-pressure turbine system. Figure 2-I presents a logic diagram of the high-pressure turbine design effort within the overall Energy Efficient Engine Program.

The high-pressure turbine component has been designed to meet the requirements for the flight propulsion system and the integrated core/low spool. In addi- tion, the turbine design is the same for the High-Pressure Turbine Component Test Rig, which will be used to assess turbine aerodynamic performance before the component is tested in the integrated core/low spool. The design empha- sizes the utilization of advancements in the areas of aerodynamics, materials/ cooling management and structures to achieve aggressive performance and dur- ablllty design goals.

This report presents a comprehensive description of the aerodynamic and thermal-mechanical design of the Energy Efficient Engine high-pressure turbine.

A description of the high pressure turbine rig design is also presented.

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SECT ION 3.0 DESIGN OVERVIEW 3.1 DESIGN GOALS AND CHALLENGES The Energy Efficient Engine high-pressure turbine concept is a single stage configuration capable of high work extraction and high system efficiency.

Performance and durability design goals established for the turbine component address the overall program goals as well as the requirements for a future commercial flight propulsion system.

The turbine efficiency goal for the flight propulsion system is 88.2 percent.

This represents a considerable increase in efficiency compared to a current technology single-stage design. However, the expected test efficiency goal for the integrated core/low spool is lower -- 86.7 percent. This is based on the anticipation that leakage rates and part quality in the integrated core/low spool vehicle will be worse than for a fully developed flight propulsion system. In addition, restagger for compatibility with the increased annulus low-pressure turbine and rematching associated with other component perfor- mance losses further reduce efficiency.

Other key design goals include a specific work output of 448,000 J/kgm (192.96 Btu/Ibm), an expansion ratio of 4.0, a combined turbine cooling/leakage flow rate of If.2 percent of core engine flow, and a rim speed of 527 m/sec (1730 ft/sec).

In terms of component durability, the flight propulsion system design goals consist of a vane and blade life of lO,O00 hours, and disk life of 20,000 hours. In addition, an airfoil coating life goal of 6000 hours was established.

The achievement of these performance and durability goals with a single stage turbine introduces certain design challenges. The two most prominent challenges center around maintaining acceptable blade stress with the high rim speed operation and minimizing system cooling flows and leakage losses. The follow- ing paragraphs outline how these challenges have been approached, and the specific design features that contribute to meeting the turbine performance and durability goals are described in the next section.

To achieve the efficiency goal, the single stage turbine must operate at a high ratio of wheel speed to specific work (high velocity ratio) and a low ratio of through flow to wheel speed (Cx/U). The aerodynamic parameters of velocity ratio and Cx/U can be translated into blade stress and blade attachment stress, which results in a high AN 2 (the product of annulus area and wheel speed squared). It is this structural concern that constrains attain- ing a high level of aerodynamic performance with a highly loaded, single stage turbine. In the design of the Energy Efficient Engine turbine, this challenge has been addressed by the application of high strength blade and disk alloys, along with increased blade taper.

The efficient management of cooling flow is essential since significant penalties in efficiency as well as increases in fuel burned result from the large percentage airflow required to cool the turbine components. From a per- formance standpoint, the goal is to minimize the a_.lountof coolant, while maintaining acceptable metal temperatures to meet durability requirements. The turbine design addresses this challenge by utilizing advanced high temperature capability materials for the airfoils along with thermal barrier coatings and an efficient cooling management system. This combination reduces the cooling requirement by approximately 30 percent in comparison to current-technology conL1ercial engines.

3.2 HIGH-PRESSURE TURBINE GENERAL DESCRIPTION A cross sectional view of the high-pressure turbine for the Energy Efficient Engine is presented in Figure 3.2-I.

The parts of the engine comprising the high-pressure turbine are those discus- sed in this report and are shown in Figure 3.2-I. The design is based on modularity to facilitate accessibility and maintainability. Along with the high rim speed and large annulus area, the single stage design has numerous features to reduce cost, increase aerodynamic efficiency, improve cooling effectiveness, reduce cooling air leakage, and minimize performance deteriora- tion. These various technology features are listed in Table 3.2-I.

TABLE 3.2-I ADVANCED TECHNOLOGY DESIGN CONCEPTS REDUCED COOLANT FLOW CONCEPTS: REDUCED COST CONCEPTS: Si ngl e-Stage Turbine Si ngle-Stage Turbine Reduced Number of Airfoils Improved Airfoil Cooling Effec ti veness Single Crystal Airfoil Materials Thermal Barrier Platform INCREASED AERODYNAMIC EFFICIENCY CONCEPTS: Coatings Efficient Coolant Supply System High AN2/High Rim Speed Contoured Vane End Walls Low Windage Low Loss Airfoils Reduced Tip Loss Configuration and Active Clearance Control High Airfoil Loadings PERFORMANCE RETENTION CONCEPTS: REDUCED LEAKAGE CONCEPTS: Ceramic Outer Air Seal/ Reduced Leakage Length Abrasive Blade Tip Improved Gap Sealing Improved Rim Sealing W-Seals OF POOR QUALITY c/) ,,.J l.U {/) n z e,.

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..I The single stage configuration offers a potentially large savings in engine cost, weight and maintenance costs because of a significant reduction in the number of components, especially expensive air-cooled turbine airfoils. In comparison to the base engine two stage version, the single stage turbine has 43 percent fewer airfoils. Also, the overall number of turbine parts has been reduced by elimination of the second stage disk, seals, outer airseals and related components.

The capability to operate at higher blade loadings because of a reduction in airfoil number results in substantial performance gains. Other performance enhancement features include contoured vane endwalls and an active clearance control system to minimize operating clearances at cruise conditions.

A reduction in cooling flow with no compromise in durability is obtained pri- marily through the use of advanced high-temperature materials and efficient airfoil cooling methods. Both the vane and blade materials are single crystal alloys which offer superior creep strength properties, along with good resist- ance to thermal fatigue. In addition, the vane platforms are coated with a thermal barrier coating for additional temperature capability. The internal cooling system of the airfoils is efficiently designed to promote a high heat transfer rate with a minimum of coolant airflow.

A substantial reduction in leakage has been achieved by the use of advanced sealing techniques and the relatively fewer number of airfoils, which reduces leakage sources. Besides the ceramic seals in the active clearance control system, leakage is reduced in the blade attachment by the use of full-ring boltless sideplates and W-seals. Also, feather seals are employed in the vane inner and outer platforms to reduce leakage.

Turbine performance retention in the flight propulsion system will be enhanced by an advanced blade tip sealing system. The turbine blade incorporates an abrasive material tip treatment that is used in conjunction with an abradable ceramic outer airseal.

3.3 DESIGN PERFORMANCE DATA The engine performance data used to design the high-pressure turbine are a combination of flight propulsion system predictions and integrated core/low spool expectations. This design approach was taken to maximize the performance of the high-pressure turbine during integrated core/low spool testing, while still demonstrating a component that is representative of the flight propul- sion system. As a result, the design uses power and flow levels expected for the integrated core/low spool, while assuming predicted flight propulsion system levels for the aerodynamic losses. This approach results in a high- pressure turbine exit flow parameter that is incompatible with the design inlet flow parameter of the low-pressure turbine. This flow parameter discrep- ancy requires that the high-pressure turbine blade be restaggered for the in- tegrated core/low spool test.

The design performance data for the high-pressure turbine is summarized in Table 3.3-I. This table also compares this design performance with similar data for the flight propulsion system and integrated core/low spool from which it was derived.

TABLE 3.3-I HIGH-PRESSURE TURBINE PREDICTED PERFORMANCE AT AERODYNAMIC DESIGN POINT (10,668 m (35000 ft), 0.8 Mach Number, Standard Day) Component Design Flight Propulsion Integrated Core/Low (Before Restagger) System Prediction Spool Expectation N (RPM) 13233 13178 13233 Total Wc/a (%Wae) 16.84 15.95 16.84 HPT Wc/a (%Wae) 14.10 13.19 14.10 FPIn (W_P) 16.98 16.82 16.98 FPou t (W_P) 66.74 65.98 68.16 Pr 3.998 4.032 4.084 Ah (Btu/sec) 13384 13086 13409 Efficiency (%) 87.9* 88.8 87.1 Efficiency Goal (%) -- 88.2 86.7 *Based on flight propulsion system aerodynamic losses and integrated core/low spool cooling flows

SECTION 4.0

HIGH-PRESSURE TURBINE AERODYNAMIC DESIGN

4.1 OVERVIEW

The aerodynamic definition of the high-pressure turbine was based on a series

of analyses to establish the flow_ath, airfoil contours, solidity and matching

characteristics that achieve the highest level of performance within the con-

straints of the basic mechanical definition. Moreover, the final design was

influenced by results obtained from different supporting technology programs

that were in progress concurrent with the design process. These programs

served as design and diagnostic tools to address unique requirements of the

component design. The different programs consisted of the Uncooled Rig Program,

the Supersonic Cascade Rig Program and the Leakage Program. The contribution

of these programs to the turbine aerodynamic design is summarized in Section

4.2.5.

As a result of these design analyses and supporting technology programs, the

low throughflow velocity and high reaction aerodynamic approaches were

selected as the basis for the component design. Also, the airfoil design

philosophy of low uncovered turning, low exit wedge angle was maintained. The

general parameters governing the aerodynamic design of the turbine component

are listed in Table 4.l-I.

TABLE 4.l-I

GENERAL AERODYNAMIC PARAMETERS

F1 ight Propulsion

System

Sea Level Takeoff

Design Point

(Mn 0.8; I0,668 m (35000 ft)

(Hot Da_,)

STAGES

],324,49] (]92.])

PTIN Pa (psia)

CET K(OR) 1633 (2940) 1708 (3076)

RIT K (OR) 1561 (2811) 1641 (2955)

N (RPM) 13232. 13866.

AH (Btu/sec) 13384.

16.984

FPin (W _T/PT)

14.10

_IA (%Wae)

4.0

Reac tio n

43.0 percent

0.556

Velocity Ratio

NASA Work Factor (Ah/u 2) l.62

0.351

Cx U

4.06 x lO lO 4.46 x lO lO

AN ( (IN2 RPM 2)

481 (l 580)

5O4 (1655)

URIM m/see (ft/sec)

Clearance cm (fn_ 0.0469 (0.0185J

Efficiency (Design) 87.9 percent

I

_r _ _

ORIG!!:_L F:--,._-'_':. :_'

OF POOR QUALITY

As described previously in Section 3.3, a mixture of cycle performance data

from flight propulsion system predictions and the integrated core/low spool

expectations was used in the design of this turbine. The design intent was to

satisfy flight propulsion system requirements through application of the

results gained from the supporting technology programs. As work progressed it

became apparent that the high-pressure turbine in the integrated core/low

spool design should be matched to the expected integrated core/low spool per-

formace. The turbine design, therefore, uses the power and flow requirements

of the integrated core/low spool while assuming the level of aerodynamic

losses in the flight propulsion system. The high-pressure turbine exit flow

parameter resulting from this approach does not match the integrated core/low

spool low-pressure turbine flow parameter (Section 4.2.3). To achieve the

desired low-pressure turbine inlet aerodynamic conditions in the integrated

core/low spool, the high-pressure turbine blade was restaggered open 0.25

degree from its aerodynamic definition.

4.2 COMPONENT AERODYNAMIC DESIGN

4.2.1 Flowpath Definition

The flowpath for the high-pressure turbine component is presented in Figure

4.2.1-I, showing the turbine elevation and axial length. In the single stage

configuration, there is a total of 24 vanes and 54 blades.

TIP ROOT MEAN 43 (17) -- 4.356 CM VANE BX (1.715 IN) A 3.424 CM 2.946 CM 2.540 CM Z '_ _ BLADE BX (1.348 IN) (1.16 IN) (1.00 IN) n v 42.118 (16.582) _._ 40.949 41.018 (16.1221 (16.1491 40 (161 Z ...I n.- uJ I-- VANE Z IJJ {J 38 (151 ill Z (.9 Z iii

o

rr 35 (14) =b_ 35.255 (13.880) 35.135 t u.

(/) :::) ( 13.8331 34.742 (3 (13.678) ,< rr

1 I I I

33 (131 2.5 5.0 7.5 10.0 (0) (1) (2) (3) (4)

AXIAL LENGTH, CM (IN)

Figure 4.2.1-I

Hi gh-Pressure Turbi ne Fl owpa th Defi ni tion

]0

The vanes are characterized by aerodynamic sections having a blunt leading

edge and a long chord with the maximum airfoil thickness near the leading

edge. The inner vane endwall is cylindrical, while the outer is contoured in

an "S" shape (Figure 4.2.1-1). The large aerodynamic section thickness reduces

the losses associated with the introduction of cooling flow and the S-wall

design reduces endwall pressure losses. In addition, the number of vanes mini-

mizes blockage as well as optimizes the positioning of the vanes with the

combustor fuel injector nozzles.

The blades are highly tapered with a conical inner wall. Blade axial chord

taper was selected to accommodate stress requirements. Also, the stage is

defined with a high reaction level to enhance efficiency. On the basis of

design analyses, it was shown that higher reaction levels produced higher

efficiencies. The maximum allowable reaction level for this turbine design was

limited to 43 percent because increasing the axial pressure load on the rotor

resulted in excessive bearing loadings and consequent durability concerns.

Experimental testing in the supporting Uncooled Rig Program demonstrated the

benefit of the 43 percent reaction level in comparison to a lower level of 35

percent.

4.2.2 Airfoil Definition

Several analytical techniques were used to establish the turbine airfoil

definitions. A streamline computer design simulation generated the radial

aerodynamic environment. This information serves as input to the interactive

airfoil design system for a definition of the external contour of the airfoil.

Analyses are then performed to ascertain pressure distribution and boundary

layer characteristics. On the basis of these results, iterations of the air-

foil shape are made to optimize pressure distribution, boundary layer and low

loss characteristics.

Using these methods, vane sections were analytically defined. The sections

were designed so that the flow accelerated past the throat area with low,

smooth backend diffusion. The uncovered turning and exit wedge angle were

optimized to minimize pressure loss. Blade sections were designed to the same

pressure distribution criteria as the vane. Uncovered turning and exit angle

were optimized to reduce blade profile, trailing edge and shock losses. Results

from the Supersonic Cascade Program verified the final airfoil shapes. Signi-

ficant test results relating to the airfoil aerodynamic design are presented

in Section 4.2.5.

Velocity triangle data pertaining to the final airfoil definitions are presen-

ted in Table 4.2.2-I. These triangles are similar to those used in the testing

of turbine blades with a 43 percent reaction level (second build of the Un-

cooled Rig Program).

The assumed temperature profile for the combustor exit and the vane pressure

loss profile used in the design is shown in Figure 4.2.2-I.

Figures 4.2.2-2 through 4.2.2-4 present the airfoil contours of the vane root,

mean and tip sections, respectively, along with a summary of the section

aerodynamics and corresponding static pressure distribution. Figure 4.2.2-5

shows a two-dimensional schematic of the vane section stacking.

TABLE 4.2.2-1

GAS TRIANGLES*

Root

Mean

zlp

VANE

IN (deg) 90 ° 90 ° 90 °

OUT (deg) 11.6 ° 10.3 ° 9.1 °

M IN 0.09 0.08 0.07

M OUT 1.0 0.92 0.85

e GAS (deg) 78.4 ° 79.7 ° 80.4 o

BLADE

IN (deg) 33.5 ° 42.7 ° 63.6 °

OUT (deg) 15.9 ° 16.9 ° 17.7 o

Mr IN 0.36 0.25 0.14

Mr OUT 1.22 1.24 1.28

e GAS (deg) 130.6 ° 120.4 ° 98.7 o

OUT (deg) 38.0 ° 43.8 o 48.4 °

M OUT (deg) 0.54 0.52 0.52

*Based on flat Inlet temperature and flat loss profiles.

• CET PROFILE IS GOAL -10.14

"vAF"_° _OSS':LT'_ ASC A 0E / 701_

91 < ,.o

_lO.O 6 3

0.8

., I I I I I I

0.6

0 2.0 4.0 6.0 8.0 10.0

% SPAN

Figure 4.2.2-1 Vane Assumed Profiles For Inlet Temperature and Pressure Loss

r OF POOR _iO;¢=_d (.

o.o (o.o_ _- VANE ROOT SECTION 0.5 (0.21 ,-- PRESSURE DISTRIBUTION 1.0 10.4) -- 1.O 1.5 (0.6) -- 2.0 (0.8) -- O,9 2.5 (1.01 -- 3.0 (1.2) -- I-- O- 0.7 3.5 (1.41 -- 0.6 4.0 (1.6} -- 4.5 (1.8) 0,5: 5.0 (2.0) --

o.4i I I 7- I

00( 020 040 060 080 _,00

Z 5.5 12.2) X/B 6.O _t_ (2.4) 6.5 12.6i RADIUS 35.255 crn 113.880 in) 7.0 (2.8) # FOILS 24 AXIAL CHORD 4.356 crn 11.715 in) 7.5 (3.0 L.E. DIAMETER 1.334 cm (0.525 _n) T.E. DIAMETER 0.163 cm ((0.064 inl 8.0 13.2 UNCOVERED TURNING 10.0 ° 8.5 13.4] EXIT WEDGE ANGLE 4.0 ° MMA X 1.012 9.0 (3.61 INLET FOIL ANGLE 90.0 ° EXIT FOIL ANGLE 11.562 ° 9.5 CHORD 10.256 cm 14.038 in) 10.0 (4.0 10,5 (4.2 0,0 1.0 2.0 3,0 4.0 (0.0) (0.4) (0.8) (1.2) (1.6) CM (IN)

Figure 4.2.2-2

Vane Root Section Aerodynamic

Contour and Pressure

Di stri buti on

0.0 10.0) -- 0.5 (0.2) -- 1.0 10.4) -- VANE MEAN SECTION 1.5 10.6) PRESSURE DISTRIBUTION 2.0 (0,8) 2.5 I1.0) -- 1,0 %"T_'2_"-- _ __ 3.0 11.2) 3.5 11.4) 4.0 11.6) F- O.

4,5 (1.8 5.0 i i.

5.5 (2.21 6.0 i2.4

0.4- I I I I i

6.5 (2.6 0.0 0.2 0.4 0.8 0.9 1.0 7.0 (2.8 X/B 7.5 (3.(; 8.0 (3,2 RADIUS 38.103 CM (15.001 IN) 8,5 13,4) # FOILS 24 AXIAL CHORD 4.356 CM (1.715 IN) 9,0 (3.6 L.E, DIAMETER 1,334 CM (0.525 IN) 9,5 {3.8 T.E. DIAMETER 0,163 CM (0.064 IN) UNCOVERED TURNING 9.0 ° 10.0 14.0 EXIT WEDGE ANGLE 4.0 ° 10.5 (4.2 MMA X 0,927 INLET FOIL ANGLE 90.0 ° 0,0 1.0 2.0 3.0 4.0 EXIT FOIL ANGLE 10.312 = (0.01 (0.41 (0.8) I1.21 11,6) CHORD 10.899 CM (4,291 IN) CM (IN)

Figure 4.2.2-3 Vane Mean Section

Pressure Di s tri bution

Aerodynamic Contour and

13 I

OF .r:t)_R QUaLiTY

0.0 (OO) VANE LEADING EDGE TIP SECTION O.5 I0.2) PRESSURE DISTRIBUTION 1.0 (0,4} 1.5 (0.61 1.1 2.0 (0.81 1.0 2.5 (1 .OI 0.9 3.0 (1.2) P/PT 0.8 3.5 (1.4) 0.7 A 4.0 (1.6) 0.6 z 4.5 SIB_ 0,5 5,0 12.01

0. I I I I I

_J 5,5 12.2) 0.00 0.20 0.40 0.60 0.80 1,00 6.0 (2,4) X/BX 6,5 (2,6) 7.0 12.8l RADIUS 42.11B CM (16.582 IN) # FOI LS 24 7.5 (3.OI AXIAL CHORD 4.356 CM (1.715 IN) 8.0 (3.2} L.E. DIAMETER 1.333 CM 10.525 INi 8.5 (3,41 T.E. DIAMETER 0.162o CM (0.064 IN) UNCOVERED TURNING 10._) 9.0 (3,6) EXIT WEDGE ANGLE 4.0 9.5 (3.8) MMAX 0.85_) 10.0 (40) INLET FOIL ANGLE 90,0 e EXIT FOIL ANGLE 8.509 10.5 (4.2) 1 I I I I III II CHORD 11,829 CM (4,657 IN} 0.0 1,0 2.0 3.0 4.0 (0.0) IO4) (0.8) 11.21 (1.6) CM (IN)

Figure 4.2.2-4 Vane Tip Section Aerodynamic Contour and Pressure Distribution

Figure 4.2.2-5 Turbine Vane Stacking

]4

Similar design information is presented for the blade in Figures 4.2.2-6

through 4.2.2-12. As shown in Figure 4.2.2-6, the blade inlet angle has been

adjusted to account for differences between a rig and engine environment

(temperature, cooling and secondary airflows). The result is a blade design

with the root and tip sections that are undercambered 5 and I0 degrees,

respectively, and the mean section is overcambered 8 degrees in comparison to

the rig contours.

Figures 4.2.2-7 through 4.2.2-II present blade section contours for the root,

one-quarter root, mean, one-quarter tip and tip section, respectively. These

figures also provide a sumary of the aerodynamic properties as well as the

pressure distributions. Figure 4.2.2-12 shows the blade stacking.

Airfoil coordinates are given in Appendix A.

14(] + DESIGN INLET = GAS ANGLE METAL ANGLE WITH PROFILES 1 O0 GAS ANGLE INTO BLADE 8O FLAT TEMP/PRESS PROFILES 6o 4O

+ I I I I

0 20 40 60 80 1O0

% SPAN, BLADE INLET SIDE

Figure 4.2.2-6 Effect of Inlet Temperature and Vane Loss Profile on Blade

Inlet Angle

]5

OF POOR QUALITY

4.5 (1,8) BLADE ROOT SECTION PRESSURE DISTRIBUTtON 4.0 11,6) 1.0 0.8 3,5 11.4: 3.0 (1.2) 0.7 P/PT 0.6 2.5 (1.0) 0.5 0.4 0.3 U 2.0 0.50 0.60 0.70 0,.80 0.90 1.00 X/BX RADIUS 34.742 CM (13.678 IN) ii t FOILS 54 AXIAL CHORD 3.424 CM (1.348 IN) L.E. DtAMETER 0.393 CM t0.156 IN) 0.5 (0.; T.E. DIAMETER 0.1_2 CM (0.060 IN) UNCOVERED TURNING 8.0- 1,0 (0,41)_ EXIT WEDGE ANGLE 2.0 ° SNLET WEDGE ANGLE 30.0 ° 0.0 I0.0) M N iNLET 0.349 0,0 0,5 1.0 1,5 2.0 2.5 30 3.5 M N EXIT 1.189 (0.0) (0,2) (0.4l (0.6) 10.8) (1.01 (1.2t (1.4) INLET FOIL ANGLE 38.5 ° EXIT FOIL ANGLE 1592 ° CM (IN) MMA X 1,559 CHORD 4.242 CM (1.67 IN)

Figure 4.2.2-7 Blade Root Section Aerodynamic Contour and Pressure

Distribution

.. B LADE_ 1/4- ROOT SECTION 4.5 (1.8) PRESSURE DISTRIBUTION 1.0 _ --- , 4.0 ci .6i 3.5 (1.4) E7 P/PT 0:6 3,0 (1.21 o:s 0:4' 2.5 (1 .O) i 0,2 O.O0 0.20 0.41) 0.60 0.80 1.1_ (J 2.0 (0.8} X/BX RADIUS 36,311 CM (14.296 IN) //FOILS 54 lO (o4i AXIAL CHORD 3.185 CM (1.254 IN) L.E. DIAMETER. IN. 0.393 CM (0.155 IN) T.E. DIAMETER, IN. 0.152 CM (0,060 IN) 0.5 10.21 UNCOVERED TURNING 7.0 ° EXIT WEDGE ANGLE 2.0 ° IN LET WEDGE ANGLE 30.0 ° 0,0 10_0) M N INLET 0.371 0.0 1.0 2,0 3.0 M N EXIT 1.202 (0.01 (0.4) 10.8) (1 2( INLET FOIL ANGLE 32.7 = EXIT FOIL ANGLE 16.39 CM (IN) MMA X 1.603 CHORD 4.267 CM (1.68 IN)

Fi gure 4.2.2-8 Blade One-Quarter Root Section Aerodynamic Contour and

Pressure Distribution

16 i

BLADE MEAN SECTION PRESSURE DISTRIBUTION 4.0 (1.6) 3.5 (1.41 0.9 DING EDGE 0.8 • 3.0 (1.2) 0.7 P/PT 0.E 2.5 (1.0) 0.5 Z 0.4 _ _E tj 2.0 IO.8} 0.3' 0.2 0.00 0.20 0.40 0.60 0.80 1.00 X/B" RADIUS 37.879 CM (14.913 IN} 1.0 (0.4) # FOILS 54 AXIAL CHORD 2.946 CM ll .160 INI L.E. DIAMETER 0.394 CM (0.15_5 IN) T.E. DIAMETER 0.152 CM (0.060 IN) UNCOVERED TURN)NG 6.0 EXIT WEDGE ANGLE 2.0 = INLET WEDGE ANGLE 30.0 ° M N INLET 0.320 M N EXIT 1.217 0.0 1 ,O 2.0 3.0 (O.O} (0.4) (0.8) (1.21 INLET FOIL ANGLE 34.0 ° EXIT FOIL ANGLE 16.80 ° CM (IN) MMA X 't .702 CHORD 4.956 CM ( I. 77 IN)

Figure 4.2.2-9 Blade Hean Section Aerodynamic Contour and Pressure

Di s tribu tion

BLADE TRAILING EDGE 1/4 TIP SECTION 5.0 (2.0) F PRESSURE DISTRIBUTION 4.5 (1.8)F_ 0.9 1'0 '_'EADING EDGE 0.8 0.7 0.6 P/PT 0.5 0.4

i',,2

0.3

\\

0.2

, I I I I I

0.1 0._ 0.20 0.40 0.60 0.80 1 .IXl X/BX 2,0 10.81 -- _ RADIUS 39.448 CM (16.531 IN) 1.6 (0. 6) -- _ f/ FOILS 54 AXIAL CHORD 2.746 CM (1.081 IN) L.E. DIAMETER. IN. 0.393 CM (0.155 IN) 1..0 IO:,, -- _ T.E. DIAMETER. IN. 0.152 CM (0.060 IN) UNCOVERED TURNING 5.00 EXIT WEDGE ANGLE 2.0°0 INLET WEDGE ANGLE 30.0

I

M N INLET 0.232 3.0 M N EXIT 1.238 O.O 1 .O 2.0 I1.21 INLET FOIL ANGLE 42.7 ° (0.0) IDA} (O.B) EXIT FOIL ANGLE 17.29 ° CM (IN) MMA X 1.76E CHORD 4.877 CM (1.92 IN)

Fi gure 4.2.2-10 Blade One-Quarter Tip Section Aerodynamic Contour and

Pressure Distribution

]?

I

OF POOR C";* "_-'_'

RADIUS 41.018 CM (16.149 IN.)

# FOILS 54 AXIAL CHORD 2.545 CM I1.002 IN) BLADE TRAILING EDGE TIP SECTION LE. DIAMETER 0.394 CM (0.155 IN) PRESSURE DISTRIBUTION T,E. DIAMETER 0.152 CM '(0.060 IN) UNCOVERED TURNING 4.0 ° EXIT WEDGE ANGLE 2.0 ° 1.0 INLET WEDGE ANGLE 30,0' 0.9 M N MAX INLET 0.137 M N EXIT 1.269 INLET FOIL ANGLE 74 ° 0.8 0,0 (0,0} -- EXIT FOIL ANGLE 17.72 ° MMAX 1.702 I 0.7 0.5 (0.2 0.6 1.0 (0.4 0,4 (_ 1.5 (0.61 0.3 0_2 0.0 0.2 0.4 0.6 0.8 1.0 2.0 (0.8) X/BX _'_ I I I I I 1 I I I I 2.5 ll.O) 0.0 0,5 1.0 1.5 2.0 2.5 3.0 35 40 45 5.0 (0,0) 10.2) i0.4) (0.61 {0.81 (1.01 (1.21 14) 1.6) (1,B) 12.0l CM (IN)

Figure 4.2.2-11 Blade Tip Section Aerodynamic Contour and Pressure

Di s tribu tio n

Figure 4.2.2-12 Turbine Blade Stacking

4.2.3 High-Pressure and Low-Pressure Turbine Matching

To achieve the desired low-pressure turbine inlet aerodynamic conditions in

the integrated core/low spool, the high-pressure turbine blade stagger angle

was opened 0.25 degree from its aerodynamic definition. As indicated in Table

4.2.3-I, the net result of restaggering is a slight penalty in high-pressure

turbi ne efficiency.

TABLE 4.2.3-I

HIGH-PRESSURE TURBINE AERODYNAMICS AFTER RESTAGGERING

HPT

Restaggered HPT

Designed HPT Run

Run at LPT FP

(Initial IC/LS) At LPT FP

(Final IC/LS)

FPHp T IN 16.983 16.983

17,023

FPHp T OUT 66. 562 68.165

68.165

PR HPT 3.98 4.093

4.084

HPT Reaction 43. percent 43.8 percent

42.4 percent

_THPT BASE 0 to -0.3 percent

0 to -0.15 percent

Mn HPT OUT 0.523 0. 554

0.539

HPT OUT 43.8 degrees 43.0 degrees

44.0 degrees

LPT Converg

VI ROOT 1.4 1.35

1.4

BI ROOT 1.3 1.25

1.3

- Conclusion:

Restagger + 0.25 Degree To Get Back A_HPT, Bearing

Load (i.e., Reaction), and LPT Aerodynamics

4.2.4 Aerodynamic Efficiency Status

The high-pressure turbine efficiency estimates, based on results acquired from

the Uncooled Rig Program and the aerodynamic design, are summarized in Table

4.2.4-I. On the basis of these estimates, the efficiency goals for both the

integrated core/low spool and flight propulsion system are attainable at the

goal tip clearance 0.046 cm (0.0185 in) and exceeded at the status tip clear-

ance 0.032 cm (0.0126 in). The component design efficiency level is also

confirmed.

4.2.5 Supporting Technology Programs

In support of the high-pressure turbine aerodynamic design, several technology

programs were conducted to experimentally assess the critical advanced design

concepts. These programs included the Uncooled Rig Program, Supersonic Cascade

Program and Leakage Program. Results acquired from these efforts provided the

necessary technical guidance and insight to ensure a viable aerodynamic design.

]9

J

I

TABLE 4.2.4-I

HIGH-PRESSURE TURBINE EFFICIENCY ESTIMATE BASED ON

UNCOOLED RIG TEST RESULTS

Aerodynamic Design Point

(10,668 m (35000 ft), 0.8 Mach No., Standard Day)

Flight

Integrated

Propuls ion Component Core/Low

System (%) Design (%)

Spool (%)

Uncooled Rig (Build 2) 91.1 91.1 91.1

Coating -0.2 -0.2 -0.2

Cooling -3.7 -3.8 -3.8

Trailing Edge Blowing +i. 1 +i. 1 +1.1

Blade Rest agger - -0.i -0.i

Leakage

- -0.1

Clearance

- -0.2

Wind age

- -0.2

Part Quality - - -0.8

Engine Rematch

- -0.2

EFFICIENCY ESTIMATE AT:

Goal Clearance 0.0472 cm (0.0186 in) 88.3

88.1 86.6

Status Clearance 0.0320 cm (0.0126 in) 88.8

- 87.1

GOAL EFFICIENCY (%)

88.2 - 86.7

The Uncooled Rig Program was basically directed towards establishing the

uncooled aerodynamic efficiency base and verifying the principal aerodynamic

design assumptions, specifically the benefits of increased stage reaction

level and low ratio of throughflow to wheel speed (Cx/U). The Supersonic

Cascade Program focused on determining the performance characteristics of

different vane endwa11 geometries and blade configurations, including loss

characteristics with cooling flow. The Leakage Program was structured to

investigate potential leakage sources in the turbine design and define

approaches to reduce leakage within the constraints of the component design

and sealing concepts evaluated.

Salient results from these programs, which influenced the turbine aerodynamic

design, are summarized in the following sections. A complete description of

the results of each program is contained in the following NASA Technical

Reports: Energy Efficient Engine High-Pressure Turbine Uncooled Rig Technology

Program (CR-165149, Reference l), Energy Efficient Engine High-Pressure

Turbine Supersonic Cascade Technology Report (CR-165567, Reference 2), and

Energy Efficient Engine High-Pressure Turbine Leakage Technology Report

(CR-165202, Reference 3).

4.2.5.1 Uncooled Rig Technology Program

Test results obtained from the Uncooled Rig Program demonstrated that sub-

stantial efficiency gains could be achieved by designing a single stage

turbine to operate at a low ratio of throughflow to wheel speed (Cx/U)

with an attendant high blade attachment stress and at higher turbine reaction

Ievel s.

The benefits of decreasing the Cx/U value and increasing the AN2 para-

meter, as established by rig testing, are summarized in Figure 4.2.5-1. As

shown, the measured perfomance of the Energy Effi£ient Engine turbine con-

figuration is clearly superior to turbines with AN L parameters at state-of-

the-art levels. The benefit of operating at a low Cx/u translates into

1.15 percent increase in turbine efficiency over current performance levels.

AN 2(IN 2_ RPM 2) E3 4.6 x 101° STATE-OF-ART 3.4 x 101° 92 1_ A_TPREDICTED = 1.1% / _. A_ MEASURED = 1'1°//° MASSAVGU/''_'` E3 BLD#1

7(%)

SINGLE STAGE HPT _

84 I I I I I

3.5 4.0 4.5 5.0 5.5 6.0 PR

Figure 4.2.5-I Turbine Uncooled Rig Performance Trends Showing Benefits of

Decreasing the Cx/U and Increasing the AN L Parameter

Figure 4.2.5-2 shows the positive effect achieved by increasing the turbine

stage reaction level. At the design point pressure ratio, testing with a

higher reaction level -- 43 percent as opposed to the lower reaction level of

35 percent-- produced a 0.8 percent improvement in performance.

Table 4.2.5-I presents the predicted uncooled rig efficiency, along with the

measured values. As indicated, the goals were surpassed. Overall, the results

from this program have established the uncooled aerodynamic efficiency of the

high-pressure turbine at gl.l percent and have verified the feasibility of the

advanced turbine aerodynamic design concepts.

OR_GI;_L PAGE t_

OF POOR QUALITy

92B

OF PO0 QUt LITY

_ E3 #2 MASS AVG

(

9O =43, DESIGN POINTS _ E3 #1 MASS AVG -- R EACTI ON = 35% "_,_ E3 #1 E3 #2 RPS DESIGN 0.35 0.43 A//PRED 0.5% RPS MEAS 0.33 0.37 AI,/MEAS 0.8%

I I I I I

3.5 4.0 4.5 5.0 5.5 6.0 PR

Figure 4.2.5-2 Efficiency Gains Associated with High Blade Reaction Levels

TABLE 4.2.5-1 TURBINE UNCOOLED RIG EFFICIENCY Predicted (%) Measured (%) Build 1 90.3 90.4 Build 2 90.8 91.1

4.2.5.2 Supersonic Cascade Program

In the Supersonic Cascade Program, two vane endwall geometries and three blade

sections were evaluated to ascertain their influence on turbine performance.

The vane endwall configurations included a contoured (S-wall) cascade and a

straight wall cascade. Test results are presented in Figure 4.2.5-3, showing

the spanwise distribution of total pressure loss for the two configurations.

The data trends show that the S-wall design demonstrated substantially less

pressure loss, a total of 17 percent, resulting primarily from the lower

secondary loss in the S-wall cascade.

The ability to assess the losses of cooling and leakage flows is essential in

order to properly predict the performance characteristics of a cooled turbine.

To demonstrate the ability to predict these losses in a transonic environment,

a cooled vane cascade test was conducted. In Figure 4.2.5-4, test data are

compared to the analytical predictions. The good correlation of results veri-

fies the ability of the design system to predict these effects.

0.12 L_P/P MASS AVG _> PLANE CASCADE RESULTS PLANE 2,34% S WALL 1.91 % 0.10 NOTE: THE "S" WALL WAS RETAINED IN THE COMPONENT DESIGN 0.08 0.06 _._t i l

o. R

I _jr_/s WALL * I '. _:_'- " _PLANE WALL /(_/_ "_1_ 0.02 I_ --.\ ,,_- _#,

01---- LA.A OF,L'DW'LL-- "

0 20 40 60 80 100 PERCENT SPAN

Figure 4.2.5-3 Spanwise Distribution of Total Pressure Loss for the Profiled

Wall Cascade

O PLANE CASCADE DATA .... PREDICTION PRESSURE SIDE SUCTION SIDE 3,0 2.0

0 0

Ap/p 1.0

,,f(:T_-"

/kP/P 2.0-

(%)

(%)

O

I 1 I I

0.

1.0 2.0 1.0 0 1.0 2.0 WC/W_ , PERCENT

WC/W °°, PERCENT

to Test Results

4.2.5-4

Predicted Vane Cooling Losses Compared

Fi gure

OF FOCR _U:._Lri'Y

Blade cascades tested included overcambered and straightback designs which

were evaluated against the candidate design (base blade configuration) for the

Energy Efficient Engine high-pressure turbine. Cascade results demonstrated

that no additional performance benefits were achieved by using the straight-

back design. Figure 4.2.5-5 shows the response of blade base pressure coeffi-

cient with trailing edge coolant injection and exit Mach numbers, and Figure

4.2.5-6 shows the effect of trailing edge ejected cooling flow on mean section

total pressure loss. These results show that a significant performance im-

provement can be obtained with proper use of ejected cooling air.

0.1 DESIGN POINT_ t%_) WC/A - 2.0% M °O- 1.24 f v -0.1 CPBAv G -O.2 -0.3 • SOLID AIRFOIL _ mJ O HOLLOW AIRFOIL 0% WC/WM 1 _ A t % WC/WM -0.4 O 2% WC/WM n 3% WC/WM FLAGGED SYMBOLS DENOTE HEATED RETEST POINTS -0.S I I I I I I I I

I

0.6 0.7 0.8 0.9 1,0 1.1 1.2 1.3 1.4 M E _ MEAN, MASS AVERAGED

Figure 4.2.5-5 Effect of Trailing Edge Ejection Flow on Blade Pressure

Coefficient

8.0 EXIT MACH NO.

ISENTROPIC (M2i) = 1.3 [_ AIRFOIL #4 0 AIRFOIL #5 6.0 Loss _PTPT

D

(%) 4.(

0 0

O

2.0

o.o I I I

o.o _.o 2.0 3.0

COOLING FLOW Wc W c = cooling air flow rate _% RATIO W m W m = mainstream air flow rate

Fi gure 4.2.5-6

Plane Cascade Base Blade, Mean Section Trailing Edge Cooling

Air Ejection

OF POOR QUALITf

4.2.5.3 Leakage Program _ l _ d _ _ l _ . _ _ b P _ . . . .

The Leakage Program was conducted to evaluate techniques for leakage reduction

in the high-pressure turbine. Based on this effort, the low leakage technology

in the Energy Efficient Engine high-pressure turbine has been successfully

substantiated. Test models were used to simulate component leak paths as well

as to assess leakage reduction concepts. These nodels simulated the blade-disk

attachment and the vane inner and outer platfom attachment seals.

The results of blade-disk attachment testing disclosed that leakage in this

area could be significantly reduced by paying careful attention to tolerances

along the contact surfaces between the vibration damper and platform contact

surface. As shown in Figure 4.2.5-7, attachment leakage is less than predicted,

thereby demonstrating the effectiveness of the blade dampers in sealing the

platforms.

Other tests were conducted to verify the full ring sideplate, W-seal design.

The data presented in Figure 4.2.5-8 show that a flat rear sideplate against

the disk face results in low leakage levels. Also, W-seals are effective in

controlling leakage, as indicated by the results in Figure 4.2.5-9.

80.0 - - (4 DAMPERS) TION _-BLADE PLATFORM- 7 10.0 - \ /, BLADE GAP

- __/_t/__

- OA.PER-

1 ,, - I FRONT LCONTACT POINT

Z_ VARIATIONS CAUSED - ¢ BY DAMPER O0 INTERCHANGE 1.0 l l I l I I I I I I I I l l I I I I 0.01 0.10 0.6 PDAMP - PAMB PDAMP

Figure 4.2.5-7 Results of Blade-Disk Model Testing Showing Leakage in the

Attachment Area is Less Than Predicted

8O O CURVED SI DEPLATE - 0.0025 CM (0.001 IN.) SHIM - (3 BLADES) 0 CURVED SIDEPLATE- 0.0051 CM (0.002 IN.)SHIM O CURVED SIDEPLATE -0.0051 CM (0.002 IN.) SHIM + 0.0051 CM (0.002 IN.) SHIM AT DAMPER RADIUS -- /PREDICTION _" // (EQUIVALENT TO A ,,¢ !

/// CURVED-SIDEPLATE) o ,I-- v // PAMB r'_/

,'LOCATION 7 /-/"

OF SHIM FOR "_ _ I 'El// '_ F_ >;, UNIFORM ._ _ " _ 0" 7- DEFLECTION -"_J _---"_.0229 CM (0.009 IN.) / \ _CRITICAL J DEFLECTED -'_,,_ _ _____ THROTTLING /o REAR SIDEPLATE "__ ._ AREA (NOT TO SCALE) "_ -_ PSUPPLY k SHIM 0.(}025 CM (0.001 IN.)

1.0 0.10 0.9 0.01 PSUPPLY - PAMB PSUPPLY

Figure 4.2.5-8 Sealtng Effectiveness of Rear Sideplate Design

70. -- F (3 BLADE LENGTH) ICTION

10.-

WITHOUT SEAL BLADE PLATFORM PLATFORM GAP O

--h/2

L_'-____ REAR _i _DpAL_PTE 1.

I 1 I I 1 t I I 11 I I I I I 1 I I I I , I

0.01 0.10 1.0

PDAMP - PAMB PDAMP

Figure 4.2.5-9

Results With and Without W-Seals in the Blade Platfom Area

Leakage tests were also performed to assess the effectiveness of feather seal

perturbations for leakage control in the vane inner and outer platform attach-

ments. Figure 4.2.5-I0 shows the various feather seal configurations tested,

and Figure 4.2.5-II presents the test rig arrangement, along with the major

components. All of the test rigs incorporated gaps similar to those expected

in the turbine component assembly. Testing demonstrated that significant

reductions in leakage were achieved by eliminating the feather seal inter-

sections and plugging the seal gaps. These results are shown in Figure 4.2.5-12

by the comparison of 0.025 to 0.050 cm (O.OlO to 0.020 in) thick two piece

overlapping seals in electrical discharge machined slots. In addition, the

results indicated that minimizing surface waviness was essential for obtaining

a good sealing surface. These results are presented in Figure 4.2.5-13.

Two-piece feather seal with overlap _

(e)

(a)

Two-piece feather seal Two-plece feather seal_

(b)

Two-piece feather seal wT_thPuitCev'_rlOng" feather seal without overlap and

(c)

F°u r-_ ie/_efla t h/_ Two-piece feather seal without overlap, with _'_ front and re___ I I I I I I I I I Fj'M_ ORIG!NAL ;:AC,._ _.":;;

(h)

OF POOR QUALiT't

(d)

Figure 4.2.5-I0 Promising Feather Seal Configurations Evaluated

.J IE Z _J

L

tl.JJ

0.01!

NO PLUG (2 PLATFORM_ FRONT PLUG 0.01C

j %_-

J BOTH PLUGS 0.005 - OVERLAPPED _ 0.025 CM (0.010 IN.) THK 0.050 CM (0.020 IN.) THK

I I

0 I

1.0 1.5 2.0 2.5

PRESSURE RATIO

Figure 4.2.5-12 Conparison of 0.025 to 0.050 cm (0.010 to 0.020 in) Thick

Two-Piece Overlapping Feather Seals

0.015 N PLUGS m 0.010 -- _TOTALLY PLUGGED

u O u O

GROUND • ELECTRO-DISCHARGE MACHINED

o I I I

1.0 1.5 2.0 2.5 PRESSURE RATIO

Figure 4.2.5-13 Two-piece Overlapped Seals in Ground Versus Electrodischarge

Machi ned S1 ots

29 [

ORIGINAL PAGZ _

OF POOR QUALITY

SECTION 5.0

AIRFOIL DURABILITY

5.1 OVERVIEW

The Energy Efficient Engine high-pressure turbine design emphasizes operation

at a moderately high combustor exit temperature with a minimum of cooling to

maximize fuel efficiency. This must be achieved, however, with no compromise

in component durability.

Turbine airfoil durability goals for both vanes and blades are lO,O00 hours of

service life for the flight propulsion system and 50 hours hot section life at

28oc (84OF) day sea level takeoff conditions for the integrated core/low

spool. The lO,O00 hour goal is established in terms of international missions,

and reflects an equivalent of 2200 missions. These goals are achieved through

the combination of improved cooling effectiveness and advanced high-temperature

capability materials.

5.2 TURBINE VANES

The themal design, including durability assessment, of the turbine vanes is

based on the combustor exit temperature profile shown in Figure 5.2-I. This

profile represents an atypical worst case situation at hot spot locations,

showing a nearly flat radial profile with a maximum temperature of 1888oc

(3431OF) which reflects an appreciably high design pattern factor of 0.42

and a deteriorated engine condition.

(3500) _J o o:: 164B :::) I.- (300O) TT4 = 1500=C (2733"F) n.."

,,, TT3 = 577oC (1071OF) a..

PATTERN FACTOR = 0.42 LLI I- (2500)

i I l l I i I J

0 20 40 60 80 1 O0 % SPAN

Figure 5.2-I

Combustor Exit Profile Used for Turbine Vane Durability

Assessment for Flight Propulsion System at Hot Day Sea Level

Takeoff Operating Conditions

30 I

5.2.1 Vane Cooling Management System

Vanes in the Energy Efficient Engine high-pressure turbine, as aerodynamically

defined, are physically larger than turbine vanes in current gas-turbine

engines because of the selection of a low number of airfoils. This increases

vane surface area and causes inherently longer surfaces that must be cooled.

In turn, the longer surfaces necessitate a greater number of film cooling rows

to maintain acceptable metal surface temperature levels.

A diagram of the vane cooling system is presented in Figure 5.2.1-I. As indi-

cated, the vane requires only a total of 6.41 percent of core engine inlet

flow for cooling, excluding inner and outer platform surfaces. Internal sur-

faces are cooled by convection, while external surfaces are film cooled. In

comparison to current commercial engines operating at similar turbine stator

inlet temperatures, this represents a 1.2 percent reduction in cooling flow

requirements.

0.73% .91% 2.31% 2.72% 0.325% 1.52% 0.325% 1.52% TOTAL COOLING FLOW 6.41% WAE

Turbine Vane Cool ing Design

Figure 5.2.1-I

O_IG_..',_, F.;_ _3

OF POOR QUt,_LiTY.

Cooling air enters the vane from the tip and the root at a pressure of

2,840,657 Pa (412 psia) and temperature of 577oC (1071OF) at the

durability design condition of sea level takeoff, hot day. Exact percentages

of flow are depicted in Figure 5.2.1-I. The coolant is distributed within the

internal structure of the vane, which is designed with three cavities. These

cavities are convectively cooled through the use of sheet metal impinge_nt

tubes that fit into the three cavities. To provide maximum strength against

bulging deformation, two ribs tie the pressure wall and suction wall to-

gether. In addition, the vane trailing edge is convectively cooled by cooling

air flowing through a series of pedestals or braces between the vane walls.

Cooling flow passes around the pedestals and is discharged through a slot in

the trai I i ng edge.

The front cavity impingement tube is supplied cooling air to convectively cool

the walls of the cavity. After the internal surface is cooled, the coolant is

discharged through an array of showerhead holes in the leading edge as well as

a set of holes downstream of the leading edge on the suction wall to provide a

cooling film over the external surface. The showerhead holes are angled

radially, as opposed to the axial angular orientation of the film cooling

holes on the suction and pressure walls, for more effective heat transfer in

the thick leading edge region.

The middle and rear cavities also contain impingement tubes from which

numerous cooling air jets are impinged against the vane inside surface.

Cooling air flows in a chordwise direction and is discharged through axially-

angled film holes in order to provide film cooling. A portion of the cooling

air in the rear cavity is channeled through the trailing edge pedestal and

discharged at the vane trailing edge. The size and spacing of the pedestals

have been selected to provide the desired convective cooling and cooling flow

level s.

Because of the somewhat unusual aerodynamic contour of the vane, the stag-

nation point on the airfoil appears on the pressure surface. This surface is

film cooled to offset this heat load by two sets of two rows of holes, approxi-

mately 0.058 cm (0.023 in) in diameter. The suction surface incorporates three

rows of cooling holes, approximately 0.050 cm (0.020 in) in diameter.

The cooling scheme, including cooling hole arrangement and flow distribution

for the inner and outer vane platforms, is shown in Figures 5.2.1-2 and

5.2.1-3, respectively. The outer platform requires nearly 0.5 percent flow,

while the inner platform requires 0.32 percent flow. The platform cooling

scheme is based on Pratt & Whitney Aircraft experience and utilizes impinge-

ment cooling from under the platform and convection cooling from drilled

holes. The large size of fhe airfoil makes exclusive use of convection cooling

infeasible because of the increased friction loss resulting from the longer

holes. Also, the requirement for more vane material to accommodate these

longer holes would result in a significant increase in weight.

Impingement cooling is employed for the portion of the platform adjacent to

the pressure side. Cooling holes are incorporated on the side and aft rails.

If required, additional impingement cooling can be provided on the platform

suction side. The heat transfer coefficients used for inner and outer platform

analysis are shown in Figures 5.2.1-4 and 5.2.1-5.

VANE I.D. PLATFORM COOLING FLOW 0.32% VANE O.D. PLATFORM COOLING FLOW 0.49% THERMAL BARRIER COATING 0.10% WAE ON GAS PATH 0.05% WAE SIDE OF PLATFORM THERMAL BARRIER COATING ON GAS PATH SIDE OF PLATFORM 0.32% WAE PLATE IMPINGEMENT PLATE 0,03% WAE

Figure 5.2.1-2 Turbine Vane Inner Figure 5.2.1-3 Turbine Vane Outer

Platfom Cooling Scheme

Platfom Cooling Scheme

F_ REGION 1, h - 4,50 _'l REGION 1. h - 450 ] REGION 2, h - 840 []REGION 2, h _ 840 L_ REGION 3. h ~ 1040 _ REGION 3, h - 1040

OF poOR QUAL|TIf

Turbine Vane Outer

5.2.1-5

Turbine Vane Inner Figure

Figure 5.2.1-4

Platform Heat Transfer

Platforg Heat Transfer

Coefficients

Coefficients

An assessment of the effectiveness of the vane cooling management system is

summarized in Figures 5.2.1-6 through 5.2.1-10. The effectiveness of the film

over the long suction wall is compared to two-dimensional flow on a flat plate

in Figure 5.2.1-6. As shown, the Energy Efficient Engine design is conserva-

tive relative to the flow on a flat plate. Results of a themal analysis,

using the combustor exit profile shown in Figure 5.2-1, are presented in

Figure 5.2.1-7. As shown by this isothem plot, the highest calculated metal

temperature is 1226oc (2239OF) on the suction side wall adjacent to the

third cavity.

The effectiveness of the film cooling technique for the pressure and suction

surfaces is corroborated by the film temperature distributions. A profile of

the pressure surface film temperature is presented in Figure 5.2.]-8 and a

similar profile of the suction surface film temperature distribution is shown

in Figure 5.2.1-9. Of particular importance in Figure 5.2.1-9 is the fact that

the leading edge film holes provide effective film protection for the entire

suction surface with 2.31 percent of the total engine flow. A resulting pro-

file for vane surface temperature is presented in Figure 5.2.1-10. The heat

transfer coefficients used in the profile analysis are shown in Figure

5.2.1-II.

5.2.2 Vane Materials

A summary of the turbine vane materials and coating is presented in Table

5.2.2-I. The base alloy is SC 2000 nickel base single crystal material, which

affords both high strength and high temperature capability. For the integrated

core/low spool, vanes will be fabricated from PWA 1480 single crystal material.

This material provides an approximate lOOC (50OF) metal temperature advan-

tage over directionally solidified material (PWA 1422) used for turbine air-

foils in modern Pratt & Whitney Aircraft commercial engines. The impingement

tubes are of Inconel 625 sheet metal stock.

The external surface of the vane is coated with an oxidation-erosion resistant

coating, PWA 270 (NiCoCrAly). A themal barrier coating (PWA 264) is applied

to the hot gas surface of the platform.

The materials and coatings planned for the flight propulsion system are

advanced derivatives of those selected for the integrated core/low spool. The

base alloy is a second-generation nickel base single crystal material (SC

2000). This material will provide an additional lOOC (50OF)metal tempera-

ture improvement over the PWA 1480 material.

For the flight propulsion system vane, both the internal and external surfaces

are coated. The external surface is coated with an advanced overlay coating

(PWA 286) that provides improved resistance to oxidation and erosion. An

aluminide coating (PWA 275) is also applied for the internal cavity to enhance

oxidation resistance. The platforms are coated with an advanced thermal

barrier coating (TBC lO0) for added temperature capability.

0°8 m CURRENT DESIGNS

OF F oo'_

0.6 -- 0.5 %" '_' % %%'%" 2-D FLOW ON PLATE ,% 0.4 --

. /

% TG-T F 0.3 -- %, %, TG-T C 0.2 -- EEE EXTRAPOLATION / S = SURFACE DISTANCE FROM STAGNATION POINT D = COOLING HOLE DIAMETER

I I I I I 1 1 I I I I I I

0.1 200 400 600 800 60 80 100 4O S/D FROM STAGNATION PT.

T G = TEMPERATURE, GAS STREAM T F = TEMPERATURE, FILM T C = TEMPERATURE, COOLANT SUPPLY

Figure 5.2.1-6 Suction Surface FilD Effectiveness

1010 (1850) 1093 (2000) 954 (1750) 1226(2239) 1148 (2100) 1148 (2100) 176 (2150) 982 (1800)

°C(°F)

1176 (2150) 1176 (2150) 1176 (2150)

Figure 5.2.1-7 Vane Themal Analysis Results

OF poOR QU/_LFIY

m (3500) (..)

o u..." 1648 _ o (3000) UJ nr I.-- < n- uJ Q,.

LU I-- 1371 -- (25OO)

I

V

!

I I I

(200O) I

5 7 10 (1) (2) (3) (4) (5) DISTANCE FROM STAGNATION PT, CM (IN.)

Figure 5.2.1-8 Pressure Wall Film Temperatures

m (350O)

"---I

A u. 1648 o (3000)

d

o n-- :Z) I-- <( /,: LU O. 1371 (2500) uJ I- 12000)

I

i I I I I I

0 2 5 7 10 12 15 (7) (1) (2) (3) (4) (5) (6) DISTANCE FROM STAGNATION PT, CM (IN.)

Figure 5.2.1-9 Suction Wall Film Temperatures

(23O0) (2200) o_ 1148 (21001 °_ 1093

F

_ 120001 _ 1037 119001 uJ a.

_; 982 ,,I 118001 F- 117001

k

871 PRESSURE SIDE SUCTION SIDE (1600) I I I = I I I I I I I I I -4.0 0 4.0 8.1 17.1 -12.1 -8.1 (-1.6) (1.6) (3.2) (4.8) (-4.8) (-3.2) DISTANCE FROM STAGNATION PT, CM (IN.)

Figure 5.2.1-I0 Vane Surface Temperature Profile

SUCTION SURFACE PRESSURE SURFACE I I I I I I I 2 5 7 10 12 15 17 (1) (2) (3) (4) (5) (6) (7) DISTANCE FROM STAGNATION POINT, CM (IN.)

Figure 5.2.1-II Vane Surface Heat Transfer Coefficients Used to Determine

-Surface Temperature Profiles

TABLE 5.2.2-1 VANE MATERIALS AND COATINGS Flight Propulsion System Integrated Core/Low Spool SC 2000 PWA 1480 Base Alloy (Advanced Single Crystal) (Single Crystal) PWA 286 PWA 270 External Coating (Advanced NiCoCrAIY Overlay) (NiCoCrAIY) PWA 275 None Internal Coating (Aluminlde) TBC i00 PWA 264 Platform Coating (Advanced Ceramic Thermal (Ceramic Thermal Barrier Barrier Coating) Coating)

5.2.3 Turbine Vane Durability Assessment

To evaluate vane durability characteristics, estimates of strain were made

during transient engine operation. Transient strains essentially cause two

strain cycles per flight mission, and the effects of both were inc_.uded in

estimating cyclic life. Results show that the vane leading edge experiences

the greatest total strain range (0.7 percent). The relative strain at the

leading edge during the two strain excursions -- takeoff and reverse thrust

-- as opposed to other flight modes, is indicated in Figure 5.2.3-1. Strain

ranges for the other areas of the vane are identified in Figure 5.2.3-2.

The calculated vane life for the flight propulsion system and the integrated

core/low spool is presented in Table 5.2.3-I. Oxidation life and cracking life

exceed the goal values by lO00 hours or 300 cycles.

5.3 TURBINE BLADES

The thermal design of the turbine blades was based on the exit temperature

average profile produced by the vanes. This profile is shown in Figure 5.3-I,

and as indicated, the peak of the profile (gas path temperature of 1425oc

(2598OF)) occurs at the 65 percent span location. Consequently, the design

of the blade cooling system was tailored to match the spanwise temperature

ranges reflected by this profile in combination with the stress profile.

5.3.1 Blade Cooling Management System

The turbine blade cooling system design for the Energy Efficient Engine, like

the vane, relies on the efficient management of coolant to maintain acceptable

metal temperatures with minimum cooling air. The cooling system design was

verified by the results acquired from a supporting flow visualization model

test program.

0 ¸ IDLE I I TWO STRAIN EXCURSIONS PER FLIGHT MISSION -0.1 OF T,3L:T, _;_,-,.,;.

-0.2 Z -0.3 FLIGHT IDLE < CE I-- CRUISE o_ APPROACF --0.4 THRUST

\

REVERSE --0.5 -0.6 ;LTO V 537 648 760 871 982 1093 1204 (1000) (1200) (1400) (1600) (1800) (2000) (2200) TEMPERATURE, °C (o F)

Figure 5.2.3-I Vane Limiting Strain Cycle

TOTAL LOCATION STRAIN RANGE 0.7% LE 0.54% RIB PRESSURE SIDE 0.6% 0.37% SUCTION SIDE 0.25% TE

Figure 5.2.3-2 Predicted Strain

P,AGE I$

TABLE 5.2.3-1

II_ ._UALI_ "_

VANE LIFE FLIGHT PROPULSION SYSTEM Calculated Required Oxidation 6000 Hours* 7000 Hours* Cracking I0000 Hours ii000 Hours (2200 Flight Missions) (2500 Flight Missions) INTEGRATED CORE/LOW SPOOL Oxidation 50 Hours (Hot Time) i00 Hours (Hot Time) *i0,000 hours achieved with one recoating (2600) TT4.1 = 1425°C (2598°F) (2400) TT3 = 577°C (1071°F) _" 1204 o (2200) _j" uJ n,- I-'- (2000) I-- (18OO)

I I I I I

20 40 60 80 1O0 % SPAN.

Fi gure 5.3-I

Turbine Vane Exit Profile for Flight Propulsion System at Hot

Day Sea Level Takeoff Operating Conditions

A schematic of the blade cooling design is presented in Figure 5.3.1-1, show-

ing the internal passages and distribution of cooling air. The blade internal

geometry is cooled through convection, while external surfaces are locally

film cooled from the leading edge showerhead holes and tip pressure side

holes. There are no film cooling holes on either the pressure or suction

surfaces of the airfoil.

The total cooling air requirement is 2.75 percent of the core engine inlet

flow. As indicated in Figure 5.3.1-I, cooling air enters the blade through

three root passages. The cooling flow is supplied to the blade root at a pres-

sure of 1,661,646 Pa (241 psia) and a temperature of 556°C (I033°F) at the

cooling design condition of hot day sea level takeoff.

BLADE COOLING FLOWS (TOTAL 2.75%)

0.26%

///#

0.04% 0.40% 2.05% ORIG!NA= PAC 7.

OF

0.70% 1.72% 0.33%

Turbine Blade Cooling System

Figure 5.3.1-I

4]

The front passage supplies approximately 25 percent of the total blade flow to

cool the leading edge and tip. This flow is discharged through a series of

0.038 to 0.050 cm (0.015 to 0.020 in) diameter showerhead holes at the leading

edge and an array of 0.050 to 0.063 cm (0.020 to 0.025 in) diameter holes at

the tip. Since the predicted gas temperatures are substantially lower at the

airfoil span extremes, as shown earlier in Figure 5.3-I, the leading edge

cooling hole pattern has been tailored to accor._odate these temperature re-

quirements. In addition, trip strips are integral with the internal passage

design to promote a higher heat transfer rate.

The middle passage utilizes approximately two-thirds of the total blade cool-

ing flow. Flow circulates through the internal cavity, making two spanwise

excursions, and enters the trailing edge passage where it passes through an

array of pedestals before being discharged into the gas path. Trip strips are

also used in the passages to raise the heat transfer coefficient. At the tip

and root sections, turning vanes are used to reduce aerodynamic flow separa-

tion and the attendant pressure loss penalty. Sizing of the flow areas was

carefully selected to avoid areas of diffusion that could lead to flow separa-

tion and high pressure loss.

The third blade passage directs a small percentage of cooling air to cool a

portion of the root area and supplements trailing edge cooling. Flow injection

improves the flow distribution in the root turn area with essentially no

penalty on the cooling supply pressure.

Verification of the coolant passage design was accomplished by the use of a

five times size flo_ model (Reference 4), which duplicated the conplete

internal shape of the blade.

The capability of the cooling system design to maintain acceptable metal tem-

peratures was confirmed by a thermal analysis. Figure 5.3.1-2 presents an

isotherm plot of the blade laid-span section. As indicated, the average metal

temperature is 954oC (1750OF). The highest predicted temperature is

slightly above I093oc (2000OF), and occurs on the suction wall surface

near the first rib. A profile of pressure and suction surface temperatures is

shown in Figure 5.3.1-3. The heat transfer coefficients used in the profile

analysis are shown in Figure 5.3.1-4.

5.3.2 Blade Materials

A summary of the materials and coatings selected for the turbine blade is

presented in Table 5.3.2-I for the integrated core/low spool and flight pro-

pulsion system. For both applications, the base material and coatings are the

same as for the turbine vane. The base material is an advanced nickel base

single crystal alloy and the blade is coated with an advanced oxidation

resistant coating. The major difference is that a thermal barrier coating on

the platfom is not required.

OF pOOR Q UAL_'_r'_

(1800°F) 982°C 898oc (1800°FI 11650OF) 1982°C 982°( 1800OFI (1800°F 926°C ,(1700°F} 982°C (1800°F 982°C (1800°F) 926°C 926°C i(1700°F) (1700°F)

Blade Thermal Analysis Results (Midspan Location)

Fi gure 5,3,1-2

43 l

ORIGINAL F,_,C.7_ !_

OF POOR QUALi"_: (2050) F (2OOO) SUCTION SIDE PRESSURE SIDE 1065 ; (1950) _ (1900) =v o . 1010 (18501 D I,- _ 982 LU 118001 LU I'- 954 (1750l (1700) (165Ol I I I I I I I (1600)_ -7.11 -6.09 -5.08 -4.06 -3.04 -2.03 -1.01 1,01 2,03 3,04 4.06 5.08 6.09 7.11 (-2.80) (-2.40) (-2.00) (-1.60) (-1.20) (-0.80) (-0.40) (0.40) (0.80) (1.20) (1.60) (2.00) (2.401 (2.80) DISTANCE FROM STAGNATION PT. CM (IN.)

Figure 5.3.1-3 Surface Temperature Profile

Z -- 1200 -- LuU" 10OO-- / I PRESSURE SURFACE L) ' 800 --

°21C C I

u_ 600 -- _ (/3 Z _ _ _ SUCTION SURFACE I-- <_I-- 200 UJ

-r o I I I l I l I I I I I I I I I I

O 0.5 1.0 1.5 2.0 2.5 3.0 3.5 4.0 4.5 5.0 5.5 6.0 6.6 7.1 7.6 {0.2) (0.4) (0.6) (0.8) (1.0) (1.2) (1.4) (1.6) (1.8) (2.0) (2.2) (2.4) (2.6) (2.8) (3.0) DISTANCE FROM STAGNATION POINT, CM (IN.)

Figure 5.3.1-4 Blade Surface Heat Transfer Coefficients Used to Deten,line

Surface Temperature Profiles

TABLE 5.3.2-1 BLADE MATERIALS Flisht Propulslon System Intesrated Core/Low Spool SC 2000 Base Alloy PWA 1480 (Advanced Single Crystal) (Single Crystal) PWA 286 External Coating PWA 270 (Advanced NiCoCrAIY) (NiCoCrAIY) PWA 275 None Internal Coating (Aluminide)

5.3.3 Turbine Blade Durability Assessment

An analysis of transient strains encountered during the flight cycle resulted

in predicted lives that meet the durability goals. Characterization of blade

strain properties showed that the first rib is subjected to the highest strain

(0.55 percent range). Rib strain as a function of flight mode and corresponding

operating temperature is shown in Figure 5.3.3-I. Strain ranges for other

parts of the blade are indicated in Figure 5.3.3-2.

0.6, (COLD RIB) -- _ SLTO

_ f/ i]

CRUISE

z"

0.3 < _r TWO STRAIN EXCURSIONS PER FLIGHT MISSION 0.2 -- GROUND I _- FLIGHT _ 0.1 -- I IDLE !

!

!

o --J ,vl I I I I I I

0 426 537 648 760 871 982 1093 (800) (1000) (1200) (1400) (1600) (1800) (2000) TEMPERATURE, °C (OF)

Figure 5.3.3-I Blade Limiting Strain Cycle

TOTAL LOC. STRAIN RANGE RIB 0.55% L.E. 0.28% PRESSURE SIDE 0.25% SUCTION SIDE 0.37% T.E. 0.27%

Figure 5.3.3-2 Predicted Strain

The calculated blade life for oxidation and cracking failure modes is tabula-

ted in Table 5.3.3-I. Durability goals for both the integrated core/low spool

and flight propulsion system are surpassed by an appreciable margin, as shown

by these results. Cracking as a result of creep is defined as the life limit-

ing failure mode. Figure 5.3.3-3 presents the calculated blade temperatures.

TABLE 5.3.3-I BLADE LIFE FLIGHT PROPULSION SYSTEM Required Calculated Oxidation 6,000 Hours 16,000 Hours Cracking* I0,000 Hours 16,000 Hours (2200 Flight Missions) (3500 Flight Missions) INTEGRATED CORE/LOW SPOOL Oxidation 50 Hours (Hot Time) 400 Hours (Hot Time) Creep 50 Hours (Hot Time) 80 Hours (Hot Time) *Cracking due to interacting creep and low cycle fatigue 12O4

OF POOR QL_AL|TI

(22OO1 B 12000) CALCULATED f TEMPERATURES o_ (1800) _ 871 (16001 (1400) e4e I I I I I 11200) 0 20 40 60 80 100 % SPAN

Figure 5.3.3-3 Calculated Blade Temperature for Flight Propulsion System at

Hot Day Sea Level Takeoff Operating Conditions

As part of the durability assessment, blade life sensitivity to wall thickness

tolerances was analyzed to evaluate the possible impact of core shifts. The

predicted effects of wall thickness on blade life are summarized in Figure

5.3.3-4. Creep life is shown to be relatively insensitive to small tolerance

variations with a single crystal material.

1.5 m / ' \\\\_\\\ ' .J uJ _> 1.o B i- T_ON SIDE_ _ _r_ _ _ "" THIN SUCTION SIDE

•

, , , O,_ _ 0.038 0.025 0.012 0 0.012 0.025 0.038 10.015) (0.01 O) (0.005) (0.005l (0.010) i0.015l THICKNESS, CM (IN.)

Figure 5.3.3-4 Effects of Wall Thickness Tolerance on Blade Life

I

SECTION 6.0 SECONDARY AIRFLOW SYSTEM 6.1 OVERVIEW The high-pressure turbine secondary flow system is designed to maximize the use of secondary air for cooling and thrust balance as well as minimize para- sitic leakage and the attendant performance penalty. The primary design fea- tures that enhance leakage control and contribute to a higher overall system perforuance incl ude: A tangential on-board injection (TOBI) system for positive blade coolant flow supply Front rim cavity mini tangential on-board injection (TOBI) system Boltless and full ring rotor sideplates A multi knife-edge, stepped high-pressure compressor discharge seal.

The benefits derived from these features, in conjunction with improved sealing concepts, result in the utilization of only 14.10 percent of the core engine inlet flow for cooling the high-pressure turbine disk, blades, vanes and case in the flight propulsion system.

Much of the sealing technology incorporated in the secondary flow system design has evolved from a leakage supporting technology program. This effort was instrumental in providing design guidance for the vane feather seals, blade damper seals and outer air seal. Significant results from the leakage program are described in Section 4.2.5.3.

6.2 SECONDARY FLOW SUMMARY Figure 6.2-I shows a secondary flow map of the high-pressure turbine, identi- fying cooling air flows, secondary cooling flows, leakage flow rates, and the static pressure at various points. Table 6.2-I presents a status sugary of the flow rates of the major turbine components in both the flight propulsion system and integrated core/low spool. Secondary flow totals for the flight propulsion system and integrated core/low spool are 14.10 and 14.56 percent of the total inlet flow, respectively. The slight difference in total flow is based on the assumption that the flight propulsion system will incorporate a higher level of feather seal technology in the vane and outer seals. These numbers differ from those shown for Wc/a in Table 3.3-I since they are status numbers and represent flows calbulated after the parts were designed.

Table 3.3-I lists the flows estimated before completion of the design.

The design features that contribute to these low flow rates are described in the following section.

OF POOR QUALITY DISK AND BLADE COOLIN(

I

I

[]

I

\ ACTIVE CLEARANCE CONTROL AIR O AIR COOLING %Wae AT ADP O ECONDARY COOLING %WaeAT ADP LEAKAGE FLOW % Wae AT ADP F"] STATIC PRESSURE % PT3 AT ADP Turbine Secondary Flow System Map Figure 6.2-I High-Pressure TABLE 6.2-1 SECONDARY FLOW STATUS SUMMARY PERCENT Wae A_RODYNAMIC DESIGN POINT Integrated Core/ Flight Propulsion Low Spool System DISK 0.60 0.60 Front Rim Cavity 0.40 0.40 Rear Rim Cavity i. O0 1.00 Sub Total BLADE 2.75 2.75 Foil Cooling Flow 0.19 0.19 Sldeplate Cooling 0.23 0.23 Leakage 3.17 3.17 Sub Total VANE 6.41 6.41 Foll Cooling Flow 0.81 0.81 Platform Cooling 1.83 1.40 Leakage 9.05 8.62 Sub Total CASE 1.03 0.99 Outer Air Seal Cooling 0.25 0.25 Active Clearance 0.06 0.06 Flange Leakage 1.34 1.30 Sub Total 14.56 14 .i0 TOTAL 5O 6.3 SECONDARY FLOW SYSTEM DESIGN FEATURES The main design features in the secondary flow system are shown in Figure 6.3-I. These include a blade coolant supply tangential on-board injection (TOBI) system, a front rim cavity mini tangential on-board injection (TOBI) system, boltless rotor sideplates, and a multi knife-edge, stepped high- pressure compressor discharge seal.

The blade tangential on-board injection system is a high efficiency cascade design to ensure positive supply of blade cooling air, which is furnished from the high-pressure compressor inner bleed. Since the system is pressure bal- anced, inner and outer seals are not necessary. Also, only a small percentage of air flow (O.l percent of the total engine inlet flow) is required around the flow guides at the nozzle discharge plane. Furthermore, because the system is balanced to accomodate the gas path inner diameter pressure, cooling flow is insensitive to rim seal clearances.

The mini tangential on-board injection system shown in Figure 6.3-2 preswirls the front rim cavity and thereby reduces windage heat up that would increase front side plate temperature. The swirl field also provides a radial pressure gradient between the blade supply tangential on-board injection system and the gas path static pressure, thereby effectively linking the blade supply pres- sure to the gas path leading edge inner diameter pressure. This keeps the ratio of blade supply pressure to gas path inner static pressure fixed and independent of seal leakage, attachment leakage and blade flow area.

The full ring, boltless sideplate configuration adopted for the Energy Efficient Engine high-pressure turbine considerably reduces the leakage in- herent in Conventional segmented sideplates. A diagram of the blade coolant supply system, showing pressure and flow characteristics, is presented in Figure 6.3-3. The design blade cooling flow is 2.75 percent of core engine air flow and is supplied to the blade at an inlet pressure of 53.2 percent of the total pressure at the high-pressure compressor exit. The tangential on-board injection dump pressure of 43.1 percent of the total high-pressure compressor exit pressure is increased to 48.2 percent by the free vortex pressure re- covery, and up to 53.2 precent by solid body rotation in the disk feed passages. As indicated, the full ring sideplate arrangement is very effective in reducing leakage. In addition, the use of W-seals at the interface of the rear plate and blade platfom successfully controls leakage flow.

Other pertinent rotor and turbine case secondary flow features include the fol lowl ng: o Rim Seals -- These seals are positioned at the leading and trailing edge of the blade platfom to isolate the disk cavity from the gas path. In essence, these serve as flow guides that function like fish- mouth seals to prevent hot gas ingestion and the resulting heat up of the disk rim.

High-Pressure Compressor Discharge Seal -- The compressor discharge seal is a nine knife-edge structure with an abradable seal land. This seal is designed to operate at a very close clearance 0.031 cm (0.0125 in), permitting minimum leakage flow. The calculated leakage flow rate is 0.4 percent of the total core engine flow.

51 I

OF POOR QUALITV

J

/ MINI TOBI BOLTLESS, FULL RING SIDEPLATE Figure 6.3-I Secondary Flow System Design Features Figure 6.3-2 Mini Tangential 0n-Board Injection System OF PCOR QUP, LITY 598°C (1110°F) J3 43.1 512°C LOW PRESSURE LOSS BLADE COOLANT SUPPLY SYSTEM " 0.1 (955°F) • BOLTLESS TO MINIMIZE PRESSURE DROP AND LEAKAGE

/ • FULL RING SlDEPLATES WITH W-SEAL TO MINIMIZE

LEAKAGE " 48.2 [ 1990 ° F ) 0.002 Figure 6.3-3 Flow Characterfstfcs of Blade Coolant Supply System Rear Seal -- A seal at the rear of the high-pressure turbine is used to provide thrust balance. The net thrust of the high-pressure rotor is presently calculated at 22,241 N (5000 lb) at sea level takeoff condi ti ons.

Buffer Seal -- This seal, which is located in the proximity of the rear seal, is used to separate cool low-pressure compressor discharge air from high-pressure turbine bore cooling air for the rear bearing compartment.

Active Clearance Control System -- Turbine case active clearance control is achieved through an internal system supplied with a mix- ture of tenth and fifteenth-stage high-pressure compressor bleed flow, which totals 0.25 percent of the total engine inlet flow. The impingement of this air cools the full ring rails to control the clearance of the outer air seals. The thermal environment of the front and rear rails is precisely matched to prevent adverse themal graidents.

Outer Air Seals -- These seals are ceramic coated and impingement cooled by combustor secondary cooling air. Since the pressure of this cooling air is relatively high, it is channeled to the intersegment gap to cool the exposed intermediate seal layer as well as prevent gas path ingestion.

W-Seals and Feather Seals -- Both sealing techniques have proved effective in minimizing leakage during experimental model_ tests.

These seals are used extensively in the high-pressure turbine design.

6.4 Thermal Analysis The effectiveness and adequacy of the secondary flow system has been verified on the basis of results acquired from thermal analyses. Figure 6.4-I shows the deteiled thermal model element breakup of the rotor and portions of the static structure required to establish boundary conditions. Figure 6.4.-2 shows an example of time transient responses for the disk during a snap acceleration/ deceleration. As expected, the rim thermal response is the fastest, while the bore response is slowest.

Turbine case and outer seal temperatures were generated with the thermal model shown in Figure 6.4-3. Results of the analysis are presented in Figure 6.4-4, which shows a time transient response of elements at the ceramic interface in the front and rear outer air seal rails for a snap acceleration/deceleration.

The ceramic interface temperature is a key parameter for outer air seal structural Integrity.

(1100) F °EL 482 -- " 426 ],..,. /I (_ _ _,_ _

,_oo, / ._.o_ _ "_T_o_

371 i,-,,, 1_,-4 r 315 _ _ ' "_,_

,_ ,6oo,1.._ /Z-"

260 _ (500) 204 ' (400) 6700 6850 7000 71 50 7300 7450 7600 7750 TIME, SECONDS Typical Transient Response of Turbine Disk During Snap Fi gure 6.4-2 Accelerati on/Decel erati on :5 o o g, QO L.' L_ _ _1_ :._ _ _ _,__..._-_._" I, ,I (1400) iI I I I l J _L I L' 1- | l F 1 I i I (1300) CERAMIC INTERFACE £ (1200) LL (1100) O rj 537 o - (lOOO) LU REAR RAIL rr 482 I-- (900) ,< FRONT RAIL rr 426 Lu (800) ,,,., _; 371 iii t- (700) (600) 26O (500: (400) 6 700 6850 7000 7150 7300 7450 7600 TIME, SECONDS Figure 6.4-4 Results of Temperature Analysis

SECTION 7.0 COMPONENT MECHANICAL DESIGN 7.1 OVERVIEW The mechanical definition of the high-pressure turbine component evolved through an iterative process based on the results of the aerodynamic effort and various supporting technology programs. Where necessary, the turbine design for the flight propulsion system has been modified to meet specific requirements associated with the integrated core/low spool. The resultant con- figuration is illustrated in Figure 7.l-l. To facilitate design and analysis, the total effort was divided into the analysis of the following subassemblies: the rotor, vane and inner case, outer case and outer airseal, and number 4 and 5 bearing compartment.

7.2 TURBINE ROTOR ASSEMBLY 7.2.1 General Description The high-pressure turbine rotor assembly is illustrated in Figure 7.2.1-I. The primary elements in this assembly are the blade, disk, disk sideplates, vortex plate and high-pressure turbine to high-pressure compressor bolted joint.

Design details pertaining to these components are discussed in the following sec ti ons.

The Energy Efficient Engine high-pressure rotor construction is different from most previous Pratt & Whitney Aircraft designs in that the rotor is straddle mounted. This arrangement eliminates the bearing compartment forward of the high-pressure turbine disk and places it after the disk. The advantage of straddle mounting the rotor is shown by the results of a dynamics analysis. A stiff high rotor equipped with a soft effective front spring mount (No. 3 bearing) eliminates any high critical speed response in the engine operating range, as indicated in Figure 7.2.1-2.

Figure 7.2.1-3 shows that the most serious critical speed mode is a free-free mode with lO0 percent strain energy. This mode occurs well above the rotor red line speed at 27,600 rpm. A pitch mode with 4.2 percent rotor strain energy occurs at 7450 and a bounce mode occurs at 4950 rpm, respectively, which are both below idle speed.

Another possible critical mode is during startup when a stationary rotor be- comes bowed as a result of thermals from residual heat in a nonoperating engine. This mode occurs below idle. Acceptable bowed rotor start character- istics are achieved with the aid of viscous film dampers on the bearing outer races.

A significant concern in the design of the joint between the turbine front hub and compressor rear hub was to provide adequate joint strength to prevent separation in the event of blade loss and ensuing high rotor imbalance loads.

High strength, cobalt alloy MP 159 was selected as the bolt material for this joint. Bolt tensile stress at steady state conditions was calculated to be 863,918,440 Pa (125,300 psi), and assembly tensile stress was I,I18,336,560 Pa (162,200 psi). Steady state thread shear stress was 364,045,440 Pa (52,800 psi) and assembly principal shear was 847,370,920 Pa (122,900 psi). All stresses were within allowable limits for this material.

OR|G_N_L P_G" |_

OF POOR QUALITY _1 I_ 53.0 CM (20.9 IN) --!

12.9 (5.1 IN) R DISCHARGE SEAL NO. 4-5 BEARING COMPARTMENT BLADE NOTE: DIMENSIONS ARE NOMINAL, RADII ARE MEASURED FROM VANE INNER 40.6 CM ENGINE CENTER LINE (16.0 IN) R VA_ 48.0 CM OUTER CASE AND (18.9 IN) R OUTER AIR SEAL Figure 7.l-I High-Pressure Turbine Mechanical Configuration .9 IN) R 39.3 CM (15.5 IN) _I I 11.9 CM _['_"-'_1_--_ 12.9 CM (5.1 IN) R (4.7 IN) 12.9 CM (5.1 IN) R HIGH PRESSURE TURBINE NOTE: DIMENSIONS ARE NOMINAL.

TO HIGH PRESSURE COMPRESSOR RADII ARE MEASURED FROM BOLTED JOINT ENGINE CENTER LINE 22.3 CM (8.8 IN) (8.0 IN) R VORTEX PLATE SK SIDEPLATES 40.6 CM (16.0 IN) R

f

Figure 7.2.1-I High-Pressure Turbine Rotor Assembly OF POOR QUAL_I'Y

HIGH ROTORI_ITH STIFFEFFECTIVE FRONT SPRINGMOUNT

L_ W Q_ ._J Q_ r_ i,, 0 ._J

5 10 15 2"0 25 30

N2 ROTOR SPEED-KRPM

HIGH ROTORWITH SOFT EFFECTIVE FRONTSPRINGMOUNT

uJ uJ 0_ n- O uJ l_u n- Z X c_

3O

5 i0 15 2O 25

N 2 ROTOR ,SPEED-KRPM Figure 7.2.1-2 Energy Efficient Engine High Rotor Imbalance Response

RPM = 4950

_ SHAFT I

_ CASES w-, ......

HPT OF poOR QUAL_"I'Y HPT BOUNCE MODE (2,6_) RPM- 7450 ILl I'-- --I FAN DUCTS HPT

/ -

LU LOW ROTO I'-- . G ..I ILl n,."

(4,2%)

_HPC HPC PITCH MODE RPM =" 27, 600 LM C3 F- L :s uJ > ROTOR l=- ..J LIJ n.* Figure 7.2.1-3 Energy Efficient Engine High Rotor Critical Speeds and Mode Shapes (% Rotor Strain Energy)

7.2.2 Blades

OF. POOR QUAL;IY 7.2.2.1 Mechanical Design Features The high-pressure turbine blade mechanical design is shown in Figure 7.2.2-I.

The mechanical design features of the blade are tailored to accommodate the increased AN 2 parameter, the benefits of which were demonstrated in the high-pressure turbine Uncooled Rig Program, as discussed in Section 4.1.2.1.

The corresponding increased wheel speed associated with the higher AN2, however, results in a higher stress load. To compensate for these higher stresses, the blade is constructed from advanced high strength and high temperature capability materials. It is also tapered, incorporates a slight degree of tilt and has minimum wall thickness. The calculated average stress loads are shown in Table 7.2.2-I.

_COATING PWA 270 SQUEALER MATERIAL: FPS' SC 2000 IC/LS)PWA 1480 1 PC BLADE CASTING HINED DAMPER-SEAL SURFACE 5 TOOTH ATTACHMENT Figure 7.2.2-I Turbine Blade Mechanical Configuration TABLE 7.2.2-I HIGH-PRESSURE TURBINE BLADE STRESS SUMMARY Calculated Force/Area Stress MPa (kpsi)* 325,434 (47.2) Root 287,513 (41.7) One-Quarter Root 230,286 (33.4) Mean Extended Neck 31 0,266 (45.0) *Sea level takeoff condition, 28°C (84°F), high-pressure rotor speed = 14,045 revolutions per minute (rpm) The blade is constructed from a one-piece casting of a nickel base single crystal alloy. An advanced overlay coating is also used on the airfoil surface to provide additional resistance from oxidation and erosion. The materials selected for the blade are SC 2000 for the flight propulsion system and PWA 1480 for the integrated core/low spool. SC 2000 is a second-generation stngle crystal alloy still under development. It has improved stress rupture, creep, thermal fatigue, notch fatigue strength, and coated oxidation resistance over currently used directionally-solidified alloys. PWA 1480 is a currently avail- able first-generation single crystal alloy with a 10oc (50OF) lower tem- perature capability than SC 2000. The blades for the integrated core/low spool were designed based on SC 2000 properties and cooling requirements, then checked to ensure that they met the integrated core/low spool life goals.

Although PWA 270 coating is planned for use in the flight propulsion system to enhance blade life, it is not required for the experimental hardware.

Taper is used to control the radial distribution of blade mass so centrifugal forces due to wheel speed can be held to an acceptable level. The degree of taper in the blade design is illustrated in Figures 7.2.2-2, and was deter- mined as part of an analysis that assessed the combined effects of taper, wall thickness, and blade tilt on airfoil stresses.

The degree of tilt in the airfoil design was established through an analysis aimed at balancing the stresses resulting from gas bending loads as well as those associated with centrifugal loads. Gas bending loads dominate at low engine speed whereas centrifugal loads dominate at high engine speeds. The degree of tilt in the blade design to accomplish the desired stress balance is shown in Figure 7.2.2-3.

Blade wall thickness and cooling passage ribs were also tailored to provide the desired distribution of radial mass. Figure 7.2.2-4 shows the nominal thicknesses established for the root, mean and tip sections. The thin trailing edge wall thickness of 0.0508 cm (0.020 in) including coating thickness was dictated by the aerodynamic concern to maintain a trailing edge wedge angle of 2 degrees and a cooling air discharge slot width of 0.050 cm (0.020 in).

The internal cooling configuration of the blade required provisions to support the ceramic core through the blade tip, while the blade is being cast (Figure 7.2.2-5). Following casting, these holes are plugged by welding in closure plugs.

A blade tip squealer and abrasive tip cap have been selected in order to de- crease the amount of performance deterioration that would normally result in the event of blade tip rubs (see Section 7.6 for further discussion on blade tip clearances). The blade tip squealer was designed to be compatible with the tip cap configuration by incorporating a 0.127 cm (0.050 in) nominal wall thickness, which will provide adequate surface area for the application of an abrasive grit. The predicted performance gain expected by using only the squealer with a 0.088 cm (0.035 in) wall thickness is approximately 0.4 percent. Increasing the wall thickness to 0.127 cm (0.050 in) decreases this performance gain by O.l percent. However, using an abrasive tip cap will more than offset this slight performance penalty. The blade tip squealer design is shown in Figure 7.2.2-6.

--LT

- LR_ AXIAL Figure 7.2.2-2 Radial Taper in Turbine Blade To Minimize Stress I U TIP TRUE 3/4 TIP RADIAL MEAN 1/4 ROOT I-TIP )OT

\ i

LINE _i \ - 314 TIP STACKING _ [ - MEAN I - 1/4 ROOT U - ROOT Degree of Tilt in the High-Pressure Turbine Blade To Achieve Figure 7.2.2-3 the Desired Balance Between Gas Bending Load Stresses and Stresses Resulting from Centrifugal Loads OF pOOR _ U'-r>-L:'i''y Z e.- ,p a-- L 4_ t-- t-- e- ,p.

e- k-- _J _.- .._ I OJ (D ofm b- 0_"'_'_ FACE _S OF POOR QUALITY CORE SUPPORT ACCESS HOLES

%

Figure 7.2.2-5 Core Support Method Figure 7.2.2-6 Blade Tip Squealer Geometry

7.2.2.2 Airfoil Vibration Analysis

A s;;_uctural analysis of the single crystal blade was perfomed using the

NASTRAN analytical technique. This technique uses a three-dimensional finite element analysis capable of showing airfoil vibratory response in terms of mode shape and natural frequency.

A graphical display of the high-pressure turbine blade defined by the NASTRAN technique is presented in Figure 7.2.2-7. The model utilizes plane stress ele_nts for the airfoil skin and ribs, while the blade neck is simulated with beam elements. Attachment flexibilities are included as springrates to ground.

L_ Figure 7.2.2-7 Graphical Display of Turbine Blade Defined by the NASTRAN Analytical Technique In addition, this model included the unique stiffness properties associated with the single crystal atomic orientation whose characteristics could not be evaluated by using the classical beam analytical models for blade vibration.

This enabled optimizing the crystallographic orientation of the blade to ensure sufficient margin from critical vibratory excitations throughout the engine operating range.

The initial NASTRAN blade analysis was based on an engine configuration which had 14 struts in the transition duct downstream of the high-pressure turbine.

Also, the single crystal axes orientation was arbitrarily aligned with the engine axis system (primary axis in the radial direction, secondary in the axial direction (see Figure 7.2.2-8)), recognizing that the secondary axis _OlO> and {I00} could be re-oriented spacially, if required. Results of this initial analysis indicated unacceptable frequency margins in the engine operating range.

OR,_G{NAL P_ _S ¢.,/ _. ,,, OF POOR _LI,_Lt1_ ROTATI ON (010) AXIAL REARWARD ( 100} I = PRIMARY GRAIN I GROWTH (RADIAL) CRYSTAL (0 °) (001) I ORIENTATION I MAXlMUM IN-PLANE ELASTIC MODULUS

N

rRD LINE ALONG TE TE Figure 7.2.2-8 Single Crystal Major Axes Orientation Figure 7.2.2-9 presents the estimated vibratory frequencies for the first five modes of blade vibration. The critical engine order lines of 14E and 24E, which correspond to turbine transition duct strut and first-stage turbine vane passing frequencies, respectively, are shown with the estimated frequency margins in the cruise-to-redline high-pressure rotor speed range. As indicated, there is insufficient margin between the 14E order line and the third vibra- tory mode (7.3 percent) and the 24E order line and fifth vibratory mode (2.6 percent).

Analysis of blade vibrations in these five resonant modes indicated that the trailing edge region from mid-span to tip was the area of greatest vibrational magnitude. This can be seen in the airfoil isodeflection characteristics illustrated in Figures 7.2.2-I0 through 7.2.2-14.

To achieve acceptable margins, two methods of corrective action were taken.

The first was an engine configurational change to reduce the number of transi- tion duct struts from 14 to II. This resulted in an lie order line signifi- cantly below the 14E line, and permitted tuning of both first and second modes so that adequate margin was provided for the first mode below minimum cruise engine speed and for the second mode above redline speed.

2%

5TH MODE .t._:-_ lSTVANE. \

v 2,,E/D,F_USE RCASE_

MIN. CRUISE _ t STRUTS / 't PRIMARY NOZ I 5_ A 4TH MODE (./3 r.J 4-- O _7.3% 3RD MODE ,_..,.,," 14E (HOT STRUTS) X v >.

3 -- I Z uJ Cl J IJJ 1ST MODE 2-- n.* u..

1 -- I I I I I I I I 9 10 11 12 13 14 15 16 N 2 SPEED ( X 10 3 RPM) Figure 7.2.2-9 High-Pressure Turbine Resonance Diagram E3 HPT BLADE LE TE 1ST MODE / FREQUENCY = 1920 CPS Figure 7.2.2-10 Blade Vibration Characteristics in First Mode • 70 Y E 3 HPT BLADE LE TE I I I ,17 .15 .13 ,11 \2 \ 2ND MODE / FREQUENCY = 2825 CPS Fi gure 7.2.2-II Blade Vibration Characteristics in Second Mode E 3 HPT BLADE ,16

\ ,C >8

4\_ 7 3RD M_UENCY = 3560 C_PS Figure 7.2.2-12 Blade Vibration Characteristics in Third Mode ?!

ORIG!I','_.L F_,-:< ' E3 HPT BLADE 12 13 TE LE 4TH MODE / FREQUENCY = 4375 CPS Figure 7.2.2-13 Blade Vibration Characteristics in Fourth Mode E3 HPT BLADE LE TE 5TH MODE / FREQUENCY = 5675 CPS Figure 7.2.2-14 Blade Vibration Characteristics in Fifth Mode

The second area of corrective action, which was required to provide adequate

frequency margin at redline speed for both the liE second mode and 24E fifth mode resonances, involved investigation of the impact of single crystal secon- dary axis orientation on blade vibration characteristics. Figure 7.2.2-15 shows how the _odulus of elasticity varies along the secondary axis of the crystalline structure, thereby indicating that re-orientation of the secondary crystallographic axis could be expected to affect the frequencies of the various vibratory modes. To determine this variation, the crystal was rotated about the radial axis {001} in a clockwise manner. The resulting mode frequen- cies were plotted as a function of this rotation angle. The results are shown in Figures 7.2.2-16. The final requirement for crystal secondary axis control was predicated on obtaining the highest second and fifth mode frequencies.

This was obtained by rotating the crystal secondary axis 25 degrees. Because of the significant trailing edge clockwise motion in these modes, it was necessary to align the maximum in-plane modulus of elasticity with chord line "A" running through the blade trailing edge, as shown in Figure 7.2.2-17.

(001) E = 18 x 106 <i00) E = 18 x 106 J'ARBITRARY I (ii0) E = 33 x 106 Figure 7.2.2-15 Crystallographic Orientation

73 I

OF PCOR QUALI'I_f 1.08 MODE MODE 1.(> 2ND 1.02 uJ u3 ,m 3RD MODE 1.00 1ST MODE 0.98 0.96 I I I I I 0.94 0 10 20 30 40 50 60 70 80 90 ROTATION ANGLE ABOUT 001 AXIS , DEGREES Figure 7.2.2-16 Predicted Frequencies Versus Crystal Secondary Orientation Angle ROTATION (010) AXIAL REARWARD (100) PRIMARY GRAIN "_. 25 ° _ REQUIRED ROTATION FROM INITIAL ORIENTATION GROWTH RADIAL (001) "_ ,_' / / / / MAXIMUM IN-PLANE _" / ELASTIC MODULUS \ / ._ / ',,. / ,% LINE "A" CHORD LINE ALONG TRAILING EDGE T°E.

Figure 7.2.2-17 Axial Shift in Crystal Orientation An analysis of this r,lodified configuration demonstrated acceptable frequency margins for both liE and 24E engine order lines. The results are presented in the resonance diagram in Figure 7.2.2-18.

Blade platfom frequencies were calculated to ensure that no vane passing resonances would occur in the engine operating range. Both lie and 24E reso- nances were predicted to be well above redline speed.

The reduced velocity flutter parameter for the high-pressure turbine blades is 2.1, which is well below the established design limit for shroudless turbine blades. The long chord and high frequency characteristics of the blade contri- bute to this stability. Blade damper load was set at 4448 N (1000 Ib) to pro- vide effective control of both buffet stress and resonant stress. The buffet stress analysis assumed that the combustor liner pressure drop was 2.5 percent of total pressure at the high-pressure compressor exit location.

Axial gapping between the blade, first vane, and hot strut was also analyzed.

The spacing along a streamline at the outer diameter between the blade trail- ing edge and the hot strut leading edge was established at three times the maximum thickness of the strut. The axial meanline gap between the blade lead- ing edge and the first vane trailing edge was established at approximately 0.33 times the vane meanline axial chord.

25 _ CRYSTAL ANGLE 11 HOT STRUTS i 5TH MODE I / 10.7% r- _24 E =_ JI"7 IST VANE \ == /(_u i Of FFUSER CASE I / I z _ STRUTS f / I P3 \PRIMARY NOZ./ 4TH MODE x 3RD MODE 2NO MODE z 12.2% =o _r 1 tE (HOT STRUTS) U_ 1ST MODE 2_ / I 1 I I I [ o I 9 10 11 12 13 14 15 16 N2 SPEED (× lO' RPM) Figure 7.2.2-18 Predicted Frequency with Modified Configuration ?5

7.2.3 Blade Attachment

7.2.3.1 Mechanical Design Features

The attachment for the high-pressure turbine blade is a five-tooth firtree

design that evolved from a four-tooth design during the preliminary design effort for the high-pressure turbine component. Primary contributing factors to this change consisted of: (1) a flowpath revision that increased the plat- form pull by 20 percent, and (2) a blade neck extension of 0.584 cm (0.23 in), which increased neck pull by 50 percent. With the original four-tooth config- uration, the combination of these factors increased the concentrated stress in the attachment above the allowable level. Consequently, the five-tooth geo- metry was adopted.

7.2.3.2 Structural Analysis A stress analysis of the blade attachment geometry was conducted by using a finite element technique. Results are su_m_arized in Table 7.2.3-I for the nominal stresses of shear, bending, bearing, and tension. These results are based on the most stress-limiting tooth in the firtree geometry.

As indicated by these results, the predicted shear and bearing stresses exceed the maximum allowable limit by 8 and II percent, respectively. The remaining values are substantially below limit, with the exception of tension, which is only marginally higher than the acceptable limit. However, rig testing of single crystal root attachment specimens has been successfully conducted both at and above these stress levels with no evidence of material distress. On the basis of these experimental data, the blade attachment configuration is structurally adequate and the predictions are somewhat conservative.

Additional stress analyses were perfomed using two-dimensional finite element technology to ensure concentrated attachment stresses provided adequate life in the integrated core/low spool and the flight propulsion system. Life goals are lO00 and 12,000 cycles, respectively, assuming the appropriate materials.

The results of this analysis are presented in Figure 7.2.3-I. Estimates for the integrated core/low spool show that all stresses are below the maximum level, thereby attaining the lO00 cycle goal. However, for the higher cyclic life requirements of the flight propulsion system, blade stresses are lower than the allowable limit but stresses are also higher on the disk by approxi- mately the same percentage. This result suggests that the disk broach geometry can be optimized to achieve the desired stress balance between the blade and disk to meet the 12,000 cycle life goal.

OF PO0_-,_QUALITY TABLE 7.2.3-I BLADE ATTACHMENT STRESS SUMMARY Stress Mode Location % Above (Or Below) Design System Shear Blade Disk Blade -48 Bending Disk -63 Blade Beam ng Disk Tension Blade +2 Disk +l conducted at these levels without material di stress.

*Rig testing has been * = GREATER THAN ALLOWABLE = LESSTHAN ALLOWABLE FPS IC/LS PERCENT OF PERCENT OF PERCENT OF PERCENT OF BLADE DISK BLADE DISK _IIT][ ALLOWABLE ALLOWABLE ALLOWARLE !lit I_ ..... ALLOWABLE FOR 12.000CYCLES FOR 1,000 CYCLES FOR 1,000 CYCLES

,OR .OOOCYCLES

-32% -28% --5% +B% -20% -16% -5% *B% -15% -10% -2% +_2% -16% --11% -15% -4% -9% -9% -39% -31% Figure 7.2.3-I Blade Attachment Stress Results 7?

7.2.4 Disk 7.2.4.1 Mechanical Design Features The high-pressure turbine disk that evolved from the detailed design effort is illustrated in Figure 7.2.4-I. Design requirements for this disk were defined by the high rim speed that resulted from the high AN 2 parameter selected for the basic turbine aerodynamics. Characteristics of the disk required by this parameter are a thick bore region and use of advanced high strength materials.

The disk rim features firtree attachments to hold the blades, shelves to sup- port the boltless front and rear sideplates, a flange to support the vortex plate, and curved elliptical cooling air supply holes which transport the cooling air from the tangential on-board injection nozzle to the blade roots.

The design was based on flight propulsion system requirements and the use of advanced MERL 80 material. This same disk configuration will be used in the integrated core/low spool. However, the material will be PWA I099, which has adequate mechanical properties to meet the life requirements of the demonstra- tor vehicle and is currently available.

Details of the elliptical cooling air supply hole are illustrated in Figure 7.2.4-2. This design was predicated on the desire to keep rim width to a minimum (i.e., minimize material mass in the rim area), control rim breakout stress concentrations, and provide the required coolant air flow to the blade root. Again, because of the high rim speeds, the primary concern was stress concentration in the rim area.

7.2.4.2 Structural Analysis Results of a stress analysis for the areas of the rim, blade attachment, and sideplates indicated that average tangential stresses are within allowable limits for the flight propulsion system disk and slightly less than allowable for the integrated core/low spool to meet its lOO0 cycle life requirement.

Further refinement will bring these stresses within a11owable limits. Burst margin is adequate. Disk web thickness was controlled to avoid possible low order (2E through 4E) first coupled mode resonances occurring at high speed.

First coupled mode vibration is defined as that mode shape of the disk and blade where disk deflection is in a fore and aft direction and accompanies (coupled) bending of the blade (see Figure 7.2.4-3). Principal life and stress characteristics are shown in Table 7.2.4-I.

TABLE 7.2.4-I DISK LIFE AND STRESS SUMMARY Flight Propulsion Integrated Core/ • _stem Low Spool 35 8.5 Bore Life (x 107, 3 cycles) Rim Life (x I0 -° cycles) I00 80 Burst Margin l .22 l .244 Average Tangential Stress MPa (kpsi) 796,349 (I15.5) 759,462 (llO.15) OF POOR QU_LITy 1_-11.9 CM (4.7,N)----_._ (I.6 IN} 20.3 CM 4,0 CM WEB (8.0 IN) R VORTEX SHELF ATE 25.4 (10.0 IN) R

/ 1

' !

CURVED, COOLING AIR PASSAGE 32.5 CM (12.8 IN) Figure 7.2.4-I High-Pressure Turbine Disk Design Features Figure 7.2.4-2 Elliptical Cooling Air Supply Hole ?9 ORIGINAL PA_ -c- [S OF POOR QUAL!TY 1ST COUPLED VIBRATION MODE 2 NODAL DIAMETERS # . i S

,/

'\ (_) k b "_ + I !

I # _.4 L--- i w

1_2

+ = MOTION TOWARD OBSERVER - MOTION AWAY FROM OBSERVER Figure 7.2.4-3 Schematic Showing Disk and Blade Deflections Caused by First Coupled Mode Vibration Low cycle fatigue (LCF) requirements for the disk, sideplates and vortex plate were set at 12,DO0 cycles. In order to assess disk low cycle fatigue character- istics, a finite element analysis was undertaken. Figure 7.2.4-4 shows the finite element break-up used, along with the boundary conditions assumed for the analysis. To account for stresses generated along the curved elliptical cooling air passage, a three-dimensional boundary integral equation analytical model was used. The combined results of these analyses are summarized in Figure 7.2.4-5, which shows the low cycle fatigue lives for the critical areas of the turbine rim. As shown, calculated lives in three critical areas exceed requirements. Since the low cycle fatigue properties of notched MERL 80 and PWA I099 materials are identical, the same life is predicted for both the flight propulsion system and the integrated core/low spool.

7.2.5 Sideplates and Vortex Plate 7.2.5.1 Mechanical Design Features The disk sideplates and vortex plate are illustrated in Figure 7.2.5-I. The full ring sideplates perform a dual function -- blade retention and sealing in the rim areas. Axial loads at the critical sealing areas are generated by centrifugal loads, acting on the sideplates that are canted away from the disk rim. Undesireable retention holes in the sideplates were eliminated by design- ing the sideplates to be trapped by bayonet connectors (engaged lugs and slots) instead of conventional bolting arrangements. Stress analysis showed that conventional bolting designs resulted in unacceptably high bolt and bolt hole stresses. Leakage is controlled by damper seals sealing axially along the blade platform and by a W-seal trapped by the blade and rear sideplate. Typical sideplate loads are shown in Figure 7.2.5-2.

ORIGIHAL P_SGE |_ OF POOR QUALITY (29,862 Ibs AXIAL AXIAL PRESSURE PRESSURE =333,501 Pa =1,176,873 Pa (48.37 PSI) (170.69 PSI) PRESSURE _ = 206,844 Pa -" (30 PSI) = 0.1424 CM (0.0561 IN) I 6RADIAL 6AXIA L= 0.0 cm (0.0 in) SPEED = 13,921 RPM Figure 7.2.4-4 Boundary Conditions Used in Disk Structural Analysis

OF pOOR QU_L_

NOTE: IC/LS LIVES ALL > 1000 CYCLES 25,000 FRONT REAR SiDEPLATE S I D EP LA T E ._.._._..._ _,_, 60,000 20,000 )_ 40,000 25,000 \ 40,000 CYCLES VORTEX PLATE 30 0OO > 1OO,0OO DISK Figure 7.2.4-5 Flight Propulsion System Disk and Sideplate Calculated Minimm,1 Low Cycle Fatigue Life Levels PLATE DISK VORTEX BAYONET SNAP FEATURES DIA. RETENTION _-_._

\\

L

RIM CAVITY SEALING PRESSURE LOAD

/

I OUTER DIAMETER CONTACT LOAD SIDEPLATES W/SEAL BLADE 82 Figure 7.2.5-1 Disk Sideplate and Vortex Plate Design ORIG!NAL _"'_ [3 OF POOR QUALI,_y LOADS /SIDEPLATE 339,397 N (76,300 Ib) 320,715 N (72,100 Ib)_ _230,416 N (51,8OO Ib) 301,587 N (67,800 Ib) 12,O10 N (2,700 Ib) 22,241 N (5,000 Ib) Figure 7.2.5-2 Typical Sideplate Loads There are several other advantages with the boltless sideplate configuration.

These include elimination of broach angle/bolt compatibility problems, reduc- tion in rim cavity size requirements and elimination of flow restrictions caused by the proximity of the bolt to the rim cooling air source.

Anti-torque features are provided by the blade metering plates that engage the front sideplate and by torque pins for the rear sideplate. The assembly se- quence, depicted in Figure 7.2.5-3, consists of the following sSeps. The front sideplate is assembled first by heating the disk, then placing the plate into position through the bayonet arrangement and rotating the plate 20 degrees to engage the disk slots and sideplate lugs. The turbine blades, which are in- stalled next, lock the front sideplate in place by extension of the blade meter plate located at the bottom of the blade root attachment. The rear side- plate assembly is similar to the front sideplate. After it has been rotated into position, anti-torque pins are positioned to lock the plate. Figure 7.2.5-4 illustrates the anti-torque pin details.

The function of the vortex plate is to contain the blade cooling air and provide a passage for free-vortex pressure rise to augment the pressure of cooling flow exiting the tangential on-board injection nozzle. Pumping action through the curved elliptical hole further raises the pressure before entering the blade root cavity. The vortex plate is fastened to the disk in the same manner as the front sideplate. The retaining feature is boltless, thereby eliminates the need for windage covers and provides a clean external design.

Anti-torque for the vortex plate is provided through use of a blind rivet.

i STEP 1 FORWARD SIDEPLATE I !

I J STEP 3 STEP 2 INSTALL REAR SIDEPLATE INSTALL BLADE & DAMPER Assembly Sequence Figure 7.2.5-3 Blade and Sideplate • CM (0.025 IN.) + 0.008 CM (0.003 IN.)

0.813CM (0.320 IN.) DIA 0.838 CM (0.330 IN.)

oe eoOi QY gl

Figure 7.2.5-4 Anti-Torque Pin Details 7.2.5.2 Structural Analysis A detailed finite element analysis of the sideplates and vortex plate was conducted concurrently with that for the disk to determine stress rupture and low cycle fatigue lives. The analysis identified critical stress locations at the front sideplate and rear sideplate snap diameter fillets where centrifugal loading of the plates are transraitted to the disk. The limiting low cycle fatigue lives at these corresponding highest stress locations are shown in Figure 7.2.4-5. Vortex plate stresses are relatively low resulting in high low cycle fatigue lives.

The limiting stress rupture points occur on the front sideplate outer diameter where the temperature is at 732oc (1350OF) and also at the front sideplate snap diameter fillet where the temperature is down to 593oc (llO0OF), but the stress is higher than at the outer diameter. Stress rupture lives at both of these locations is greater than 1000 hours at hot day sea level takeoff conditions. Stresses at all other sideplate and vortex plate locations are lower than the above two and therefore have stress rupture lives greater than lO00 hours.

A buckling analysis of the front sideplate was also conducted to assess the possible effects of thermal gradients encountered during the flight cycle.

This analysis indicated that no buckling occurs at any point in time during the flight cycle.

7.2.6 Air Seals 7.2.6.1 Mechanical Design Features The high-pressure turbine rotor includes three knife-edge air seals. These seals are illustrated in Figure 7.2.6-I and include the high-pressure com- pressor discharge seal, the No. 4 bearing compartment buffer seal and the thrust balance seal located on the rear side of the rotor. To avoid adding a bore to support the thrust balance seal, it has been designed within its self- sustaining radius, and thrust balance tuning is obtained by varying the pressure within the piston area cavity.

7.2.6.2 Structural Analysis Vibratory characteristics of both the stationary and rotating parts of the seals were analyzed to ensure adequate resonance margin, concidence margin and flutter stability in the engine operating range.

The resonance characteristics of rotating and stationary members were analyzed to ensure that a frequency margin to resonance exists at all speeds through redline to prevent vibratory excitation from the rotor.

Coincidence studies were also completed to confirm adequate coincidence margins. Coincidence occurs when the mode shape of the rotating member matches the mode shape of the stationary member at the same operating condition. At such a coincidence, energy transfer can occur from one member to the other, leading to excessive deflection and self destruction.

Analyses were also conducted on the seals, concerning aerodynamic flutter.

Stability criteria expressed as stability energy must be met for both rotating and stationary members in order for the seals to be acceptable.

OF pOOP. Q _i_L__'_ BUFFER SEAL NO. 4 BEARING HPC DISCHARGE SEAL THRUST BALANCE SEAL Figure 7.2.6-I Turbine Seal Designs The rear thrust balance seal assembly is shown in Figure 7.2.6-2, and loca- tions for axial and radial restraint are depicted. A frequency response analy- sis of the assembly indicated that margins for resonance and coincidence met or exceeded requirements for commercial applications. The predicted resonance margins are 46 percent for the stationary seal land and 20 percent for the rotating knife-edge seal. The estimated coincidence margin for this seal assembly is 69 percent. The resonance and coincidence curves are shown in Figures 7.2.6-3 and 7.2.6-4. Frequencies divided by nodal diameter are plotted as a function of nodal diameter with the point of limiting margin indicated.

Maximum stability energy was calculated to be less than 0.0005 cm-kg/cm (O.OOl in-lb/in), predicting flutter free operation. The dimensions shown in Figure 7.2.6-2 are minimum requirements for vibration-free operation.

As an added precaution to vibratory excitation, dampers were designed for both the rotating and non-rotating members. Their configurations are shown in Figure 7.2.6-2.

Details of the No. 4 bearing buffer seal assembly are shown in Figure 7.2.6-5.

Predicted resonance margins for this seal are 37 percent for the stationary seal land and greater than lO0 percent for the rotating knife-edge seal, both well in excess of commercial requirements. Coincidence margin for this seal is 48 percent, as shown in Figures 7.2.6-6 and 7.2.6-7. Like the thrust balance seal, the Number 4 bearing buffer seal has a calculated maximum stability energy less than 0.0005 cm-kg/cm (O.OOl in-lb/in) and is predicted to be free from flutter. Again, the dimensions shown in Figure 7.2.6-5 are minimum re- quirements to ensure vibration-free operation.

As with the thrust balance seal, damper rings were designed for the buffer air seal as a precautionary measure, to prevent vibratory excitation. The dampers were provided in the form of split rings fitting in grooves in the forward and aft sides of the seal, as shown in Figure 7.2.6-5.

ORIGINf_L P?;,_Z ['3 OF POOR QL.:iLL)'YY AXIAL 15.811 CM (6.225 IN) - R ROTATING PART RADIAL 0.25 CM 10.10 INI (NOM,) 19.189 CM (7.555 IN) -R ) 718 0.30 CM INOMJ NARYPART (0.13 IN) (MINI (0,13 INI (MIN} RADIAL AINT R = 30.543 CM (12.025 IN) AXIAL RESTRAINT /_ ROTATING DAMPER CUT INTO 12 SEGMENTS ,1_--II- (AMS 5596) ,ILL" .069cm - .0719/N.)THICK NON-ROTATING DAMPER SLOTTED INTO 36 FINGERS (AMS 5596) .142 cm -- .173cmTHICK (.056 -- .068 IN.)

Figure 7.2.6-2 Rear Thrust Balance Seal

87 I

- OF POOR ..... _'_-_" 1000 m • FREQUENCY W/RESPECT TO PART ANALYZED oO n O 600 -- Z -,...

,oo UM,T NG

/ | LLIMITING MARGIN 200 I RESONANCE o I I I I I I I I i ! i 2 3 4 5 6 7 8 9 10 11 12 NODAL DIAMETER Figure 7.2.6-3 Thrust Balance Seal Rotor and Stator Resonance Diagram 1200 -- "FREQUENCY W/RESPECT 8OO TO GROUND n U Z LIMITING f STATOR/ND _"" I MARGIN 2 3 4 5 6 7 8 9 10 11 12 NODAL DIAMETER Figure 7.2.6-4 Thrust Balance Seal Rotor and Stator Coincidence Diagram OR_GINF, L P;,CZ ;3 OF POOR QUALITY RADIAL R = 11.32 CM (4.46 IN) 0.48 CM (0.19 IN) MIN) :rOTATING KNIFE-EDGE 0.317 CM (0.125 IN) MIN INCO 718 STATIONARY PART AXIAL RE_: DIAL RESTRAINT 0.241 CM (0.095 IN) NOM.)

R = 19.50 CM (7.68 IN) 0.241 CM (0.095 IN) NOM.

RESTRAINT BUFFER AIR SEAL (PWA 1003) SPLIT RING DAMPERS (AMS 5699) .229cm-.234cm DIA (.090 IN.-.092 IN.)

Figure 7.2.6-5 Number 4 Bearing Buffer Seal Assembly *FREQUENCY W/RESPECT 1600 F TO PART ANALYZED ,.-.r_,_ _;._iL _ ":" OF r_'_. "'' 1400 F i LIMITING MARGIN _ (..)

d 6oo Z _ '_'_ f ROTOR/ND ATOR/ND I -- " M--ARG N ......

2 3 4 5 6 7 8 9 10 11 12 NODAL DIAMETER Figure 7.2.6-6 Number 4 Bearing Buffer Seal Rotor and Stator Resonance Diagram 1600 -- 1400-- 1200 m 1000 -- d 800 f ROTOR/ND LIMITING MARGIN STATOR/ND 600 F 400 F *FREQUENCY W/RESPECT TO GROUND 200 F I I I I I I 1 I I ! t l O-- 2 3 4 5 6 7 8 9 10 11 12 NODAL DIAMETER Figure 7.2.6-7 Number 4 Bearing Buffer Seal Rotor and Stator Coincidence Diagram 9O ORIGINAE ,PAGE lg OF POOR QUAL1TY The analysis of resonance and coincidence characteristics of the high-pressure compressor discharge seal, shown in Figure 7.2.6-8, indicated more than ade- quate frequency margins throughout the engine operating range. Predicted mar- gins are 35 percent for the stationary seal land and greater than lO0 percent for the rotating knife-edge seal. The coincidence margin for the seal assembly is greater than 100 percent, as shown in Figures 7.2.6-9 and 7.2.6-I0. Minimum dimensions to ensure vibration-free operation are shown in Figure 7.2.6-8.

Again as a precautionary measure, sheet metal finger dampers were provided for the non-rotating portion of the seal, as shown in Figure 7.2.6-8.

MATERIALS: MERL 80 EXCEPT AS NOTED HIGH ROTOR SPEEDS: REDLINE; 14.270 RPM GROUND IDLE; 9.526 RPM HPC BORE BASKET AMS 4928 (PWA 1007}

\ ,, .. ,,T BORE BASKET

FL / I I I I 22 FLOW HOLES I / STATIONARY _ DAMPER SLOTTED INTO 40 FINGERS (AMS 5596) .097cm -- .109cm THICK (.038 IN.) -- (.043 IN.)

Figure 7.2.6-8 High-Pressure Compressor Discharge Seal Assembly 9] ORIGINAL " _ OF POOR _u_,_-,, "FREQUENCY W/RESPECT TO PART ANALYZED f ROTOR/ND d 800 Z LIMITING B MARGIN f STATOR/ND RESONANCE I i i i I I I I i i i 2 3 4 5 6 7 8 9 10 11 12 NODAL DIAMETER Figure 7.2.6-9 High-Pressure Compressor Discharge Seal Rotor and Stator Resonance Di agram 1400 *FREQUENCY W/RESPECT TO GROUND 1600 f .

1,oo[

L 600 B { LIMITING ) f STATOR/ND I I I I I I I I I I I 2 3 4 5 6 7 8 9 10 11 12 NODAL DIAMETER Figure 7.2.6-I0 High-Pressure Compressor Discharge Seal Rotor and Stator Coi ncidence Diagram OF POOR QU_U'I'V" 7.3 VANE AND INNER CASE 7.3.1 General Description The vane and inner case assembly is illustrated in Figure 7.3.1-I. The primary elements are the vanes and the vane support structure, the tangential on-board injection (TOBI) system, and the high-pressure compressor discharge seal sup- port structure. Design details relating to these elements are described in the following sections.

6.0 CM (2.4 IN) 19.5 CM (7.7 IN)R HIGH PRESSURE (4.0 IN) COMPRESSOR DISCHARGE SEAL 3.0CM (1.2 IN) SYSTEM 0,5 cry1 (0.2 IN) 32.5 CM (12.8 IN) R

i

VANE INNER SUPPORT NOTE: DIMENSIONS ARE NOMINAL. 6.8 CM 40,8 CM RADII ARE MEASURED FROM (2.7 IN} 16.1 IN)R ENGINE CENTERLINE Figure 7.3.1-I High-Pressure Turbine Vane and Inner Case Mechanical Design 7.3.2 Vanes 7.3.2.1 Mechanical Design Features The turbine vane assembly, shown in Figure 7.3.2-I, comprises 24 vanes sup- ported at the outer surface by bolts through the two holes in the outer flange.

These bolts absorb the radial loads as well as provide circumferential re- straint. This avoids excessive torque loads at the inner support structure and improper loading of the diffuser case struts. To minimize leakage caused by vane twisting, both the inner and outer surfaces are clamped along a chordal cut. By having a chordal cut, axial tilting of the vane, introduced by differ- ential axial growth between the inner case and outer combustor case, is allow- ed to occur without binding up or opening a leak path.

Vane platforms are sealed by feather seals inserted into various slots on the platform surfaces, as shown in Figure 7.3.2-2, to prevent leakage of compressor discharge air into the turbine flowpath. Results from the leakage supporting technology program were incorporated in the design by minimizing feather seal slot intersections, plugging end gaps, and controlling slot straightness.

CHORDAL'CU_ I.D. CLAMP I O°D, UPPORT LOAD REACTIONS Figure 7.3.2-1 Turbine Vane Assembly M_CLE3ANCE I'D" SUPPORT 'W' SEAL _ CHORDAL CONTACT F CONTACT CHORDAL O,D. CLAMP COMBUSTOR BAND / Fi gure 7.3.2-2 Design Approach to Vane Leakage Control OF POOR QUALITY The vane materials and coatings selected for the flight propulsion system and integrated core/low spool are presented in Table 7.3.2-I. The advanced materials and coatings noted for the flight propulsion system vanes are re- quired to meet the aggressive durability and temperature goals. Durability and life goals for the integrated core/low spool pemit the use of state-of-the-art materials. As shown, an internal coating is not required for the integrated core/low spool vanes. Airfoil minimum wall thickness data are shown in Figure 7.3.2-3. Nominal wall thickness at trailing edge including coating is 0.055 cm (0.022 in).

TABLE 7.3.2-I VANE _TERIALS AND COATINGS Flight Propulsion Integrated Core/ STstem Low Spool SC 2000 PWA I¢B0 Base Alloy (advanced single crystal) (single crystal) PWA 286 PWA 270 External Coating (advanced NiCoCrAly) (NiCoCrAIY) PWA 275 None Internal Coating (aluminide) TBC 100 PWA 264 Platform Coating (advanced ceramic (ceramic thermal barrier themal barrier coating) coating) O.O38 CM 0.038 CM (0.015 (0.015 IN.)

THICKNESS UNIFORMLY :KNESS TAPERED UNIFORMLY TAPERED MIN WALL THICKNESS - 0.132 CM ROOT, MEAN AND TIP SECTIONS (0.052 IN.l: COATI NGS FPS -- ADVANCED VAPOR DEPOSITION 0.122 CM 203 CM OVERLAY (NiCoCrAIY) (0.080 IN.)

0.005-0.013 CM (0.002-0.005 IN.) THICK ICILS-- PWA 270 OVERLAY 0.198 CM (0.078 IN.) r (NiCoCrAIY) 0.122 CM (0.048 IN.) CM 0.005-0.013 CM FOR FULL SPAN (0.050 IN,] ) (0.002--0.005 IN.) THICK OF AIRFOIL Figure 7.3.2-3 Uncoated Vane Minimum Wall Thickness Distribution

ORi=,r..__L F_,G'_ [_ OF POOR QUALITY' 7.3.2.2 Structural Analysis The structural analysis of the high-pressure turbine vanes is addressed as part of the durability studies. See Section 5.2.3 of this report.

7.3.3 Inner Support, Tangential On-Board Injection, and High-Pressure Compressor Seal Support 7.3.3.1 Mechanical Design Features The high-pressure turbine inner support, tangential on-board injection system, and high-pressure conpressor seal land support assembly are shown in Figure 7.3.3-I.

The inner vane support provides structural support to both the inner portion of the vane and the compressor discharge seal land. The inner portion of the vane is connected to the support through bolts that tie both the vane and inner combustor liner together into the support. The compressor discharge seal is supported from the inner region of the case through a bolt that also traps the primary tangential on-board injection nozzle. Air is brought into the center of this support case froB the combustor area to supply air to the pri- mary and secondary tangential on-board injection nozzle.

Although parts for the integrated core/low spool are designed using forgings and Inconel 718 welded construction, it is anticipated that cast materials would be used for a flight propulsion system.

DISCHARGE COMPR ESSOR_ SEAL LAND I]l ,jPR,MARY TOB, COMBUSTOR LINER TURBINE VANE Turbine Inner Support, Tangential On-Board Injection System, Figure 7.3.3-I and High-Pressure Compressor Seal Land Support System

The primary tangential on-board injection nozzle, shown in Figure 7.3.3-2, is

designed to be fabricated from Hastelloy X material. The 36 vanes are brazed into slots in both the inner and outer wall s. By keeping the primary tangential on-board injection nozzle a separate bolted-on unit, flexibility is provided for easy modification to provide proper flow. One way of modifying the flow is to cut back the trailing edge of the vanes and, thereby adjust the nozzle area.

The secondary (or mini) tangential on-board injection nozzle is shown in Figure 7.3.3-3. The primary purpose of this nozzle is to swirl the coolant flow to the front side of the turbine disk, thereby reducing windage heat up of the front side plate. The vortex action of this nozzle affects the pressure level of the pressure balanced tangential on-board injection system described in Section 6.4 of this report.

The high-pressure compressor discharge seal is designed to maintain minimum clearance at all operating conditions. By using a high expansion Tinidur material rotating member, the seal permits component assembly using a rela- tively large clearance that will close as parts heat up, providing a running clearance between 0.027 and 0.035 cm (O.Oll and 0.014 in) depending on axial location. Clearances are set to ensure no rubs during maneuvers and to mini- mize the amount of rub during startup operation.

Figure 7.3.3-4 summarizes the seal clearances at two typical engine operating conditions. A more detailed discussion of clearances is presented in Section 7.6 of this report.

7.3.3.2 Structural Analysis Vane durability analysis is discussed in Section 5.0 of this report. The major concern in the inner case design was to provide a lightweight design with suitable load-carrying ability for the first vane blow off loads, while at the same time minimizing deflections. Total blow off load of the first vanes, acting in an aft direction on the vane inner and outer attachment points, is 222,410 N (50,000 Ib) at sea level takeoff. Half of this amount is taken by the inner case.

A detailed shell analysis was conducted at hot day sea level takeoff. At this condition, there is a large pressure load as well as thermal gradient. The resulting stresses are sumarized in Figure 7.3.3-5, and as indicated, the values are quite low relative to the high strength of the Inconel 718 material.

Thus, the impact from axial and radial deflections was the primary concern.

The same shell analysis was used to determine the magnitude of these deflec- tions, and the resulting axial deflections are summarized in Figure 7.3.3-6.

Calculated radial clearances were not considered critical except at the com- pressor discharge seal, where maintenance of seal clearance tolerances is important. The critical axial deflection was at the vane inner support, as noted. Excessive axial motion at this location could create undesirable leak- age flow paths in the vane platfor_ and at the interface of the inner vane and blade platform.

OF POOR QUALITY I _AREA ADJUSTABLE 0n-Board Figure 7.3.3-2 Nozzle Configuration for the Primary Tangential Injection System

\\

SECTION A-A Figure 7.3.3-3 Nozzle Configuration for the Secondary or Mini Tangential 0n-Board Injection System OF POOR QUALITY • MIN. ASSEMBLY CLEARANCE

®

OF 0.038 CM (0.015 IN) • DESIRED SLTO CLEARANCE OF 0.0317 CM (0.0125 IN) • NO RUB AT MANEUVER 0.0193 CM (0.0076 IN) • RESULTS

(9 @ @

COLD GAP 0.043 CM (0.017 IN) 0.073CM (0.029 IN) 0.040 CM (0.016 IN) PINCH PT _60 SEC. DECEL. _ 10-12 SEC. ACCEL, _" 80 SEC. DECEL, AVG GAP @: • ADP 0.030 CM (0.012 IN) 0.035 CM (0.014 IN) 0,027 CM (0.011 IN) • SLTO 0.0317 CM (0.0125 IN) _- Figure 7.3.3-4 High-Pressure Compressor Discharge Seal Clearance Sugary • STRESS LEVELS RELATIVELY LOW • CONTROL AND LIMIT AXIAL FELT METAL AND RADIAL DEFLECTIONS 96.527 MPa (14 KSI) (INCONEL X) 344.740 MPa 206.844 MPa (30 KSl) 27.579 MPa HASTELLOY X) 296.476 MPa (43 KSI) <206.844 MPa (30 KSI) 461.952 MPa (67 KSI) 48.264 MPa (INCONEL 718) (7 KSI) 358.530 MPa (52 KSI) @ SLTO HOT DAY i 193.054 MPa (28 KSI) HOOP (WASPALOY) ® Figure 7.3.3-5 Inner Vane Case Stress Summary ORIGINAL PAGE |_1 OF POOR QUAL1TY ACTUA L A L LOWANCE CM (IN) CM (IN I (_ ---0.304 (0.120) 0.381 (0.150) SHIM (_) 0.190 (0,075) 0.50 (0.20) (_) 0.190 (0.075) 0.381 (0.150) (_ 0.190 (0.075) 0.635 (0.250) (_) ---0.127 (0.050) 0.635 (0.250) (E) +0.127 (0.050) 0.381 (0.150) RELATIVE TO ---- _1)" VANE O.D.

Figure 7.3.3-6 Case Structure Deflection Sur.lmary 7.4 OUTER CASE AND OUTER AIRSEAL The high-pressure turbine outer case and outer airseal assembly is illustrated in Figure 7.4-I. The primary components of this assembly are the outer case, the front and rear outer airseal support rails, the ceramic outer airseal shoe, the cooling air impingement ring, and the active clearance control and cooling air nanifolds. Design details relating to this assembly are discussed in the following sections.

7.4.1 Mechanical Design Features The mechanical design features of the outer case and outer airseal are shown in Figure 7.4.1-I.

The outer case provides support for the outer diameter attachment of the inlet guide vane as well as support for the outer airseal shoes. The vane outer dia- meter support is carried through a flange and cone out to the flange connect- ing the turbine outer case to the rear combustor case. A W-seal is provided between this cone and the front surface of the outer airseal support rail.

This W-seal prevents leakage of high-pressure compressor discharge air that is brought through the flange to provide cooling for the outer airseal shoes.

I00

OF F3CR

+,+

5.0 CM (2.0 IN) 2.2 CM (0.9 IN) 6.6 CM (2.6 IN) 48.0 CM (18.9 IN) R NOTE: ACTIVE CLEARANCE DIMENSIONS ARE NOMINAL. RADII I_ 6.8 CM CONTROL AIR_ I_ (2.7 IN) .. ARE MEASURED FROM ENGINE CENTERLINE.

Figure 7.4-1 High-Pressure Turbine Outer Case and Outer Airseal Assembly COOLING AIR IMPINGEMENT RINE HPT BLADE TIP GAP OUTER AIRSEAL SHOE "W" SEAL COOLING AIR MANIFOLD HPT INNER VANE SUPPORT REAR OUTER AIRSEAL SUPPORT RAI L HPT CASE FRONT OUTER AIR ACTIVE CLEARANCE CONTROL SEAL SUPPORT RAI L AIR MANIFOLD Figure 7.4.1-I Mechanical Design Features lOl An internal active clearance control system is used for optimizing blade tip clearance during all engine operating conditions. This is accomplished by impinging controlled temperature air on the outer airseal support rails (Figure 7.4.1-1) to move the blade outer air seal shoes radially. Temperatures of the rear and front rails are kept stmilar for uniform deflection of both rails. The controlled temperature air enters the active clearance control manifold through eight bosses in the high-pressure turbine case, flows through holes in the active clearance control manifold to impinge on the outer airseal support ratls, and axtally discharges through holes in the rear outer airseal support rail to an annulus between the hot strut outer diameter fairing and the high-pressure turbine case. Further discussion of the active clearance control system is contained in Section 7.5 of this report.

The blade outer airseal shoe, shown in Figure 7.4.1-2, features a ceramic coating 0.327 cm (0.129 in) thick over a PWA 655 (cast Inconel 713) shoe nominally 0.254 cm (0.100 in) thick. Slots in the shoe reduce its spring rate.

This prevents the ceramic material from being overstressed as it cools follow- ing engine transient operation.

To minimize cooling air leakage, W-seals are also used on the front and rear hook areas of the shoe and feather seals ape used at the circumferential ship lap joint between shoes. The ship lap protects the feather seals from pressure pulses caused by the passing blades and serves as a backup (but less effective) seal. It also prevents a large step between shoes in the event of a blade rub.

If one end of a shoe deflects due to a rub, the ship lap will force the end of the next adjacent shoe to deflect equally and simultaneously.

The shoes are held radially inward by the pressure differential across the shoes. Vibrations are damped by a combination of the axial spring load applied by the W-seals, pressure loadings and metal-to-metal contact in the ship lap and feather seal areas.

High-pressure compressor discharge air cools the outer airseal shoes (Figure 7.4.1-I). Air is metered through holes in the inlet guide vane support, passes through holes in the front outer airseal support rail to the cooling air mani- fold, then flows through radial holes in a circumferential impingement ring to cool the outer diameter of the outer airseal shoes.

After the outer alrseal shoes are cooled, portions of the air flow to three different areas. The majority of the air flows through holes at one circumfer- ential end of the shoes into the gap between shoes to prevent intrusion of gas path air into the gap and to lower the metal and ceramic temperatures adjacent to the gap. A small amount of the air leaks past the W-seals located between the outer airseal shoes and the support rails. Some of the air passes through axial holes in the hook area of the rear outer airseal support rail to main- tain the temperature of the rear rail the same as that of the front rail.

The materials used in the outer case and outer airseal assembly are shown in Figure 7.4.1-3. Waspaloy and Inconel 718 materials were selected for the case wall and the common wall between the cooling air and active clearance control manifolds to provide the required dynamic resistance to satisfy the blade containment criteria, while maintaining an acceptable material thickness.

The ceramic is an abradable yttrium stabilized zirconia. The development goal is for lO:l abradable volume ratio with grit imbedded blade tips.

FOOT FEATHER SEAL SLOT METAL ).327 CM (0.129 IN) Y203 STABILIZED CERAMIC ) ZrO 2 Figure 7.4.1-2 Details of the Outer Airseal Shoe Design INCO 718 INCO 713 WA WASI WASPALOY TINIDUR INCO 718 WASPALOY WASPA LOY INCO 718 Figure 7.4.1-3 Outer Case and Outer Airseal Materials Map OF POOR QUALr_I 7.4.2 Structural Analysis Structural analysis of this assembly was accomplished through the use of the shell analysis model shown in Figure 7.4.2-I. This model includes pressure loadings, vane and shoe support reactions, and thermal gradients based on a steady-state sea level takeoff, hot day engine operating condition. Static pressures are shown in Figure 7.4.2-I and the metal temperature distribul;ion is shown in Figure 7.4.2-2.

Typical stresses resulting from this analysis are shown in Figure 7.4.2-3. All are well below allowables at the assumed steady-state condition, thereby satisfying both flight propulsion system and integrated core/low spool design requirements.

.... AMBIENT PRESS.

LOAD VANE LOAD _ OAS LOAD OAS LOAD E_ STATIC PRESS. - % PT3 Figure 7.4.2-1 Shell Analysis Model 7.5 ACTIVE CLEARANCE CONTROL SYSTEM Large turbine blade operating clearances have a deliterious affect on component efficiency, and consequently a negative impact on overall system fuel consump- tion. The goal, therefore, is to maintain tip clearances as close as possible, while avoiding rubs during normal operating conditions.

Since the high-pressure turbine in the Energy Efficient Engine is a high per- formance system, blade tip clearances are designed for minimum leakage. The oal blade tip clearances are 0.068 cm (0.027 in) at takeoff and 0.047 cm 0.0186 in) at cruise (the aerodynamic design point). To achieve optimum clearances during all flight conditions, the high-pressure turbine is designed with an active clearance control system.

o C (OF) FOR STEADY STATE - SLTO - 28°C {84°F) DAY WITH 15TH STAGE ACC AIR I I (1170°F) 1105°F) 783°C (1442 (1119°F) (1110°F) (1147%) (1114%) • 535°C (995°F) 530 ° C 515°C (959%) 552°C (1027°F) (997°F) 498 °(3 (929 °F _(965°F) 501 °C (935°F °C (1257 °F) 499°C (931°F 518 °C (966 °F) 539°C (1003 ° F) Figure 7.4.2-2 Ter, lperature Map STEADY STATE - SLTO - 28°C (84°F) DAY WITH 15TH STAGE ACC AIR G H = .,t19.632 MPa (-13 KSI O H = -179.265 MPa ('26 O H =.-227.528 MPa (-33 KSI) i o B = -+ 586.058 MPa (-+185 KSI) --_ O H = -193.054 MPa (-28 KSI) o H =-351.635 MPa (-51 KS!

= -151.686 MPa (-22 KS1} (7 H = 48.264 MPa (7 KSI) O H = 158.580 MPa (23 KSI) o H = 165.475 MPa (24 KSI) (7 8 = + 448.162 MPa (+ 65 KS! ) H = 131.001 MPa (19 KSI) O H = 158.580 MPa (23 KSI) o H = 193.054 MPa (28 KSI) (7 H = 220.634 MPa (32 = -+ 206.844 MPa (+ 30 KSI) O H = 275,792 MPa (40 o H = 275.792 MPa (40 KSI) o H = 289.582 MPa (42 KSI O H = 213.739 MPa (31 KSI) o"6 _ + 186.160 MPa (+27 KSI) Figure 7.4.2-3 Stress Map OF POOR V.-,'--' _" 7.5.1 General Description Clearance between the blade tip and outer air seal varies with engine opera- ting conditions as a result in the changes in temperature and speed. With active clearance control, close turbine operating clearances are maintained by controlling the relative growth between the rotor and external cases. This is accomplished by changing the thermal expansion rate of the case through the introduction of controlled temperature cool ing air.

The active clearance control system in the high-pressure turbine is shown schematically in Figure 7.5.I-I. In this system, the primary components are the outer case, front and rear outer airseal support rails, ceramic outer air seal shoe, cooling air impingement ring, and active clearance control and cooling air manifolds. Details pertaining to the mechanical design and structural analysis of these components are contained in Section 7.4.

Optimum blade tip clearances are achieved during all operating conditions by impinging controlled temperature air on the outer air seal support rails. The introduction of this cooler temperature air lowers the turbine case metal temperature and corresponding themal expansion, thereby controlling the radial movement of the seal shoes towards the blade tip. Air for the active clearance control system is supplied from two sources in the high-pressure compressor: the tenth stage and the fifteenth stage. The mix of air from these sources is contingent on the engine operating condition and the desired gap closure.

COOLING AIR IMPINGEMENT RIN, VANE BLADE OAS COOLING DLING AIR AIR MANIFOLD HPT IGV REAR OAS SUPPOF SUPPORT RAI L ACC AIR MANIFOLD FRONT OAS ACC AIR SUPPORT RAIL FROM 10TH OR 1STH STAGE Figure 7.5.1-I High-Pressure Turbine Active Clearance Control System OF pOOR (_'J_"' The high-pressure compressor bleed air enters the active clearance control manifold through eight bosses on the high-pressure turbine case. It then flows through holes in the active clearance control manifold to impinge on the outer air seal support rails. At this point, the air is discharged through holes in the rear outer air seal support rail into an annulus between the turbine intermediate case support strut outer diameter fairing and the high-pressure turbine case.

7.5.2 Blade Tip Clearance Definition In general, the rotor structure grows at a _ore rapid rate than the turbine case during engine acceleration from idle to takeoff power because of the centrifugal forces acting on the rotor. This results in the smallest clearance, referred to as the pinch point gap, occurring shortly after acceleration to takeoff power. Figure 7.5.2-I shows a typical radial growth time history for the rotor and case, indicating the occurance of the pinch point.

An analysis of turbine rotor and case growth compatibility suggested that the pinch point occuring at takeoff could be eliminated by using fifteenth stage high-pressure co_pressor bleed air. However, it was not apparent that utiliza- tion of only fifteenth stage air would optimize clearances throughout the air- craft flight cycle.

CASE PINCH

/ L .o,o. b

RADIAL GROWTH 'q-START, _ _ ACCEL _ _'_ DECEL T-- COLD IDLE TAKEOFF IDLE TIME Figure 7.5.2-I Typical Rotor and Case Growth History As a result, a more comprehensive fltght cycle analysis of rotor and case response growth was conducted. The analysis was based on the fltght profile information presented in Table 7.5.2-I. In addition, the analysis assumed the use of two turbine case cooltng schedules for active clearance control. Trends for rotor and case response with the two cooltng schedules are presented in Flgure 7.5.2-2.

The first schedule involved utilizing fifteenth stage high-compressor bleed air from idle through climb power until attaining an altitude of 6096 m (20,000 it). At this point, cooling flow was switched to the tenth-stage bleed.

As shown in Figure 7.5.2-2, this case cooling schedule provided the rapid case growth necessary to eliminate the pinch point at takeoff, but produced less than optimum clearance at both climb and cruise flight conditions.

TABLE 7.5.2-I FLIGHT CYCLE FOR CLEARANCE ANALYSIS DOMESTIC MISSION 1296 km (700 nm) Power Altltude/speed Time Segment Settlr_L M/Mn (ft/Haeh No.) (mln) 'Start and Idle-Taxl Gr. idle 010 7.5 Takeoff T.O.

0-457/0-0.39 (0-1500/0-0.39) 2.0 Climb H. Cl.

457-10,668/0.39--0.80 (1500-35,000/0.39-0.80) 17.4 Cruise 0.85-0.82 M. Cr.

10,668/0.80 (35,000/0.80) 56.6 Descent FI. idle 10,668-457/0.80-O.39 (35,000-1500/0.80-0.39) 20.4 Approach/Landlng 0.30 T.O. - FI. idle 457-O/0.39-0.17 (1500-0/0.39-0.17) 2.0 Reverse M. rev.

0/0.15 0.2 Idle-Taxl Gr. idle 0/O 7.5 113.6 OAS (15TH STAGE BLEED BELOW 6,096M [20,000 FT] OAS (15TH STAGE BLEED IOTH ABOVE 6,096M [20,000 FT]) AT IDLE, IOTH ABOVE IDLE) "___: f _ PINCH POINT . ,,_ O.381 // x lA (O.150 RADIAL 0.254 (0.100) DEFLECTION, ,/ii CM/I N I / !

0.127 !

(O.050) !

!

100 SEC !

I,

pt-----SNAP ACCEL _-_L'_ SNAP DECEL -_ I I

COLD IDLE $LTO IDLE ADP MAX N 2 TIME Figure 7.5.2-2 Rotor and Case Response with Two Cooling Bleed Schedules ]08 The second cooling schedule assumed fifteenth-stage bleed air would be used through idle to eliminate the take off pinch point and tenth stage air would be used for the remainder of the flight cycle. This resulted in a second pinch point occurring at climb power (Figure 7.5.2-2). However, by mixing fifteenth and tenth stage air, this pinch point was eliminated. This effect is shown in Figure 7.5.2-3, along with the initial two cooling schedules for comparison.

The final active clearance control bleed schedule is presented in Table 7.5.2-II. The resulting rotor and outer air seal response with this schedule is shown for the flight cycle in Figure 7.5.2-4.

Results from this analysis were used in establishing the high-pressure blade tip and seal gapping requirements as well as final blade tip operating clearances. Factors considered in establishing tip clearances include: (1) themal and centrifugal gradients; (2) tolerances, eccentricities and rotor whirl; and (3) maneuver and cowl loads.

15TH STAGE BLEED BELOW 6,096M (20,000 FT) 0.381 10TH ABOVE 6,096M (20,000 FT) (0.150) OPTIMUM RESPONSE / WITH MIXED BLEED _f f ....... 15TH STAGE BLEED AT IDLE, ...\ 10TH ABOVE IDLE 0.254

"\/

(0.100) RADIAL DEFLECTION CM/IN.

0.127 (0.050) 't, lr-

,- SNAPDECE --'t I

• SNAP ACCEL

I F" "

SLTO IDLE ADP COLD IDLE TIME Figure 7.5.2-3 Effect of Mixed Bleed TABLE 7 • 5.2-II HIGH-PRESSURE TURBINE ACTIVE CLEARANCE CONTROL SYSTEM Bleed System For ACC All 15th Idle Mixed 10th and 15th ACCEL-SLTO All 10th ADP ORIGINAL F,_C:_ ;3 OF POOR QUALITY 0.381 (0.150 0.254 (0.100) RADIAL DEFLECTION CM/I N.

0.127 (0.050) Figure 7.5.2-4 Rotor and Case Response with Optimum Mixed Cooling Bleed Schedule A summary of the high-pressure turbine gapping requirements, the minimums needed to prevent a rub, is presented in Table 7.5.2-III for the flight condi- tions of idle, sea level takeoff and cruise. In arriving at these values, a number of considerations were used to establish a turbine build clearance.

First, the clearance must be large enough to accommodate machining tolerances and eccentricities. Also, nominal unbalances in the rotor will produce addi- tional rotor whirl motion for which more clearance must be provided. Further- more, during the startup process, it is possible that the rotor can be bowed because of unsymmetric themal gradients. Although this resultant r, mtion is not very well defined, it is preferred to have the additional clearance to acconw_odate such motion. Finally, during flight conditions, additional deflec- tion of the cases occur under nomal maneuver and cowl aerodynamic loads. All these clearances are added up to provide the final gapping requirements.

The final blade tip clearances, using the active clearance control schedule of Table 7.5.2-II and recognizing the gap requirements of Table 7.5.2-III, are tabulated in Table 7.5.2-IV. The comparison of the clearance goal and predic- ted tip clearance status shows that the goals have been exceeded, thereby resulting in an improvement in turbine efficiency.

The rotor and seal response throughout the flight cycle is shown in Figure 7.5.2-5. The critical design point is the pinch point, which occurs approxi- mately six seconds into the snap acceleration. The clearance of 0.034 cm (0.O134 in) is maintained throughout the acceleration segment. This results in the clearance during start to idle being greater than required and also results in the clearance at the aerodynamic design point being greater than the requied 0.020 cm (0.0079 in). Although clearances at these two conditions are greater than required, they are still significantly less that the design goal s.

OF POOR r,l, _. _ "'" TABLE 7.5.2-III HIGH-PRESSURE TURBINE GAPPING REQUIREMENTS ADP Start-ldle ACCEL-SLTO cm (in) cm (in) cm (in) 0.0058 (0.0023) 0.0058 (0.0023) 0.0058 (0.0023) Tolerances 0.0086 (0.0034) 0.0086 (0.0034) 0.0086 (0.0034) Eccentricity 0.0025 (0.0010) 0.0025 (0.0010) 0.0025 (0.0010) Rotor Whirl 0.0025 (0.0010) 0.0165 (0.0065) Normal Maneuvers 0.0005 (0.0002) 0.0005 (0.0002) Cowl Loads 0.0457 (0.0180) Bowed Rotor Whirl 0.0199 (0.0079) 0.0339 (0.0134) 0.0626 (0.0247) TOTAL TABLE 7.5.2-IV HIGH-PRESSURE TURBINE TIP CLEARANCE RESULTS Goal cm (in) Status cm (in) 0.1739 (0.0685) Cold 0.1244 (0.049) Idle 0.0685 (0.027) 0.0340 (0.0134) SLTO 0.0472 (0.0186) 0.0320 (0.0126) ADP OAS .. ---- __"'_ _ _L_Lo.0340 CM (0.0134 IN) R ESU LT _ I \- (0.0340CM [0.01342N] NECESSARY) E TIP ,/ /

/ / \ / 0.0320 cM

I / _ t (0.0126 IN) RESULT "--[---'t':l_i-- _ / (0 0200 CM [0 0079 IN]

I//-J- O.O ,OCM,O.O,,,N) RESOLT " NECESSARY'

RADIAL DEFLECTION

'

_2NJ _,ES_L 7 y a (0.0627 CM i0.0247 IN] NECESSAR ) I / !

/ 100 SEC

/F-I

'-" t.,------S.A,,A_CEL

SLTO IDLE ADP COLD IDLE TIME Resultant Blade Tip Clearances Figure 7.5.2-5 III

7.6 NUMBER 4 AND 5 BEARING COMPARTMENT AND LUBRICATION SYSTEM

7.6.1 General Description The number 4 and 5 bearing compartment for the Energy Efficient Engine high- pressure turbine incorporates existing service-proven technology and parts.

The use of parts from current production engines minimizes expensive develop- merit and fabrication costs, and also reduces the degree of risk normally associated with the development of advanced technology hardware.

All parts in the number 4 and 5 bearing compartment meet Energy Efficient Engine design requirements, including mechanical stiffness, load carrying, imbalance, and high speed capability. Salient details of the number 4 and 5 bearing compartment for the integrated core/low spool are shown in Figures 7.6.l-I and 7.6.1-2.

LPT SHAFT HPT SHAFT 232°C (450 °F) 176% 176 ° C 432°C (350 ° F) (810%) (840 ° F) 6% PT3 239(, PT3 490°C (915 ° F) 468°C(875°F) 12% PT3 476°C (890°F) OPERATING CONDITION: SLTOSTD-3 °C(25°F) TURBINEINTERMEDIATE CASE STRUT Figure 7.6.1-I Integrated Core/Low Spool Number 4 and 5 Bearing Compartment - Temperature and Pressure Distribution at Sea Level Takeoff -3 C (+25OF) Operating Conditions 7.6.2 Bearing Mechanical Design Features The number 4 and 5 bearings for the integrated core/low spool are the same as those used in a current Pratt & Whitney Aircraft engine. These bearings, in addition to meeting all established integrated core/low spool llfe require- merits, have logged a significant amount of service experience, and, therefore provide a low risk base for application in the integrated core/low spool. In addition, the cost of designing, manufacturing and testing new bearings is eliminated. The bearings for the number 4 and 5 bearing compartment are discussed in the following paragraphs.

OR_GL_,cZ_ F_C;Z __i 112 OF POOR QUALITY OF _OOR QUALITY e- c; z e- e- i.

¢0

i

Z o e- r_" I.a- E l.

w c; z I-- o- e- rv, i r- La.I .c; z e" Z t_ o \ I.a_ z c; z I i _'Z z L.LI I--_ or..,,") ._l r,," oe,-- ...J ,_ r_,,_ OUJJ The number 4 roller bearing for the integrated core/low spool measures 165 mm at the inner diameter and 222 mm at the outer diameter; has 16 na_ diameter rollers and operates at 2.3 X 106 DN. The beartng is oil damped, spring centered, preloaded and under race cooled. The viscous oil damper and soft centering spring are incorporated to satisfy high rotor dynamic response criteria. The preloadtng benefit minimizes roller skidding. The under-race cooling feature provides positive cooling to manage themal expansion through- out all operating conditions. 0il flow is 7 kg/min (16 1b/rain) and heat gene- ration is rated at 475,141 J/rain (450 Btu/min). The expected B1 life of the number 4 bearing is greater than 500 hours, and the expected BIO life is greater than 2500 hours. This bearing may have to be optimized to meet flight propulsion system durability and life requirements.

The integrated core/low spool number 5 bearing is also a 165 mm inner dia- meter, 222 mm outer diameter size bearing, with 16 m diameter rollers, but operates at much lower speed than the number 5 bearing; 0.64 X 106 DN. It has zero preload and reduced internal radial clearance. Its oil flow is 1 kg/min (4 lb/mtn) with a heat generation of 31,676 J/rain (30 Btu/min). A viscous oil film damper has been incorporated on the outer ring to control low rotor vibratory reponse. Integrated core/low spool life rating for this bearing is B1 life greater than 10,000 hours and BIO life greater than 50,000 hours.

The results of a structural study to determine maximum stress and deflection are shown in Figure 7.6.2-I.

The number 5 bearing for the flight propulsion system is similar to the number 4 bearing, but with the inner diameter borereduced to 160 mm and the outer diameter increased to 230 ram.

.P 227,528 MPa (55 KSI) OF POOR QUALITY Figure 7.6.2-1 Number 4 and 5 Bearing Compartment - Locations of Maximura Stress and Deflection 7.6.3 Seals The seals used in the number 4 and 5 bearing compartment of the integrated core/low spool are derived from existing service-proven designs for minimized risk and decreased cost. The operating environment of the integrated core/low spool is less severe than that of the existing design. Therefore, the seals more than adequately meet the life and durability requirements of the integrated core/low spool.

Carbon seals were selected in place of knife edge seals because of the increased radial deflection of the high-pressure turbine rotor associated with the oil-damped number 4 bearing. The resulting large radial gaps would be detrimental to labyrinth seal operation. Carbon seals minimize the breather flow and limit total engine oil consumption to 0.22 I/hr (0.06 gal/hr).

The seal arrangement in the number 4 and 5 bearing compartment consists of a number 4 front carbon seal, an intershaft seal consisting of back-to-back carbon seals, and a number 5 rear carbon seal. These seals are dry-face, short carbon with cooled rotating seal plates. Maximum surface rubbing speed is 135 m/sec (445 ft/sec) for the flight propulsion system and 140 m/sec (462 ft/sec) for the integrated core/low spool, operating in a low-temperature and low- pressure environment at 68,948 Pa (10 psi) seal pressure drop. Current engine experience on similar seals has been accumulated at this rubbing speed with seal pressure drops of 379,214 Pa (55 psi). The integrated core/low spool seals are derived from existing hardware. The seals used in the flight propul- sion system feature a high quality carbon grade with excellent durability characteristics.

To reduce leakage, the low-pressure bearing compartments are buffered by cold low-pressure compressor air bleed discharge. The system features a single mainshaft deoiler in the front compartment. Air is introduced to the bearing compartment between the low and high-speed shafts. The buffering air is bled from the low-pressure compressor, routed through the center shaft, bled into the bearing compartment and then discharged through blowdown tubes to the deoiler in the main compartment. The main feature of this system is the self- regulating design, which eliminates the need for extensive valving and pumps.

Analysis of the seal pressure drop indicates that positive seal pressure can be obtained at all operating conditions. This positive seal pressure can be achieved at sea level idle conditions by either reducing the exit bleed flow or increasing the idle speed.

Integrated core/low spool seal operating conditions are summarized in Table 7.6.3-I. A summary of dry face seal pressure, temperature, and speed experi- ence is presented in Figure 7.6.3-I.

7.6.4 Lubrication System The lubrication system for the number 4 and 5 bearing compartment is part of a system that features positive oil management to provide sufficient cooling and lubrication flow. One of the main features of this system is the self- regulation arrangement, which eliminates the need for expensive pumps and extensive plumbing. As a result, the plumbing system is simplified, the size and number of scavenge pumps reduced, and the need for a pressure regulating valve eliminated.

OF POOR QUALITY TABLE 7.6.3-I INTEGRATED COP.E/LOW SPOOL SEAL OPERATING CONDITIONS Hiah Rotor Seals Low Rotor Seals ADP SLTO IDLE ADP SLTO IDLE Rubbing Speed 133 (439) plsec (ft/sec) 140 (462) 102 (335) 39 (129) 39 (129) It (37) Seal _p Pa (psi) 20,684 (3) 48,263-68948 (7-10) <6894<(1) 20,684 (3) 48,263-68,948 (7-10) <6894<( 1 ] Heat Generation (Stu/mln) 131 139 93 26 24 Air Temperature oc (OF) 232 (450) Front 232 (450) Rear 426 (800) • E 3 IC/LS OPERATING CONDITIONS LESS SEVERE THAN EXPERIENCE 689,480 (1001 454o C (850°F) - 510°C (950°F) 551,584 (80) SEAL I 3710C (700 oF )-398 °C (750 °F ) PRESSURE 413,688 i (60) _315°C (600°F)-454 °C (850 ° F) DROP 777-rrr Z Pa/PSI 275,792 (40) 137,896 E 3 FPS IC/LS (20) O x

I I 123:C (45°°FI I

91 (300)

,3501 121 I 01 137,4501 152 I5 1

MAXIMUM RUBBING SPEED - M/SEC (FT/SEC) Figure 7.6.3-1 Dry Face Seal Pressure, Temperature and Speed Experience An additional feature of the lubrication system is the blowdown system with oil and air scavenged together through the turbine transition duct strut. The integrated core/low spool uses two struts and the design ensures that the oil and air mixture flows through each of the struts without backing up through the scavenge lines. This design also eliminates the high cost of installing a complex internal plumbing arrangement within a single strut. With the plumbing divided between two struts, the fabrication cost and the associated operation- al risk are substantially reduced.

Oil is transported from the main engine oil pump to two radial oil scoops in the rear compartment, one scoop servicing the front bearing and the other servicing the rear. The scoop efficiency for the integrated core/low spool is 60 percent, which is conmensurate with scoop efficiency levels of most current comercial engines. The scoop arrangement for the flight propulsion system is being refined for an efficiency increase.

I16

OF"PO07:C::"" .-.-v

Adequate leakage provisions have been supplied to ensure that no oil touches any of the hot parts. To address this requirement, an oil scupper line has been incorporated to provide a drain through which oil can be transported away from hot parts. This arrangement is depicted in Figure 7.6.4-I.

IC/LS REAR COMPARTMENT LUBRICATION SYSTEM HIGH ROTOR LOW ROTOR NO. 4 NO. 5 SEAL SEAL SEAL SEAL G BEAL SPLIT 4:3 SPLIT 3:4

t

t

5 KG/MIN (12 LB/MIN) TO SCOOP 18 KG/MIN 140 LB/MIN) TO SCOOP • SCOOP EFFICIENCY 60% • TOTAL OIL FLOW (INCLUDING NO. 4 DAMPER) 25 KG/MIN (57 LB/MIN) • LINE SIZES ESTABLISHED; JETS NEED BENCH FLOW TESTS • HIGH CAPACITY PUMP REQUIRED Figure 7.6.4.1 Oil Scupper Line for the Number 4 and 5 Bearing Compartment 7.7 TURBINE SYSTEM WEIGHT SUMMARY Preliminary weight analyses were conducted for the high-pressure turbine as configured for the integrated core/low spool. Results of these analyses are presented in Table 7.7-I. A detailed weight assessment will not be performed until the final flight propulsion system preliminary design update.

TABLE 7.7-I PRELIMINARY WEIGHT SUMMARY FOR INTEGRATED CORE/LOW SPOOL HIGH-PRESSURE TURBINE CO_ONENT Itern Disk and Seals 150 (330) 32 (71) BIade Assembly 26 (58) Vane Assembly 45 (99) Tangential On-Board Injection System (TOBI) 106 (233) Outer Case Assembly To tal 359 (791) ll7

SECTION 8.0 HIGH-PRESSURE TURBINE COMPONENT TEST RIG DESIGN 8.1 INTRODUCTION The Energy Efficient Engine High-Pressure Turbine Component Rig was designed to establish the performance base for the turbine and verify the advanced aerodynamic/thermodynamic design concepts of the program. The rig design con- figuration incorporates structural criteria and mechanical constraints consis- tent with experimental hardware.

The following sections provide (1) a general description of the rig, including salient rig assembly and safety features, (2) information describing the mech- anical design, and (3) a description of the instrumentation used to monitor turbine performance and rig structural integrity.

8.2 GENERAL DESCRIPTION AND MAJOR FEATURES The turbine component rig is designed to test the full size high-pressure tur- bine component at the design differential pressure, but at reduced temperature and absolute pressure levels. It is intended to confirm the aerodynamic per- formance of full sized integrated core/low spool hardware prior to running the integrated core/low spool test. In addition, it will provide a check of pre- dicted cooling and leakage flows and operation of the active clearance control system. A cross section of the rig is illustrated in Figure 8.2-I.

The rig assembly includes an inlet section, rotor and vane assembly, and ex- haust section. The rotor and vane assemblies and active clearance control sec- tion are mainly component hardware (suitable for integrated core/low spool operation) while the inlet, exhaust and outer case sections are rig unique hardware. The rig features separate controls for all secondary flows as well as main flow. In addition, a separate system is provided for the active clear- ance control system, which covers an approximate 148oc (300OF) temperature range for clearance change. The main flow temperature will be 426oc (800OF) with the appropriate secondary air temperature ratio to simulate engine condi- tions. A circumferential traverse instrumentation ring has been provided to acquire a more thorough mapping behind the vanes and blades.

The rig incorporates an active clearance control system, which will be evaluated during testing. The internal hardware of the active clearance control system for the integrated core/low spool is also used in the rig.

Special consideration was given to the type of _aterial used in certain areas of the rig. For example, rig hardware exposed to main and secondary airflow was designed using stainless steel or comparable rust-resistant alloys, to prevent contamination of coolant passages. High strength materials are used in the high temperature regions of the rig, while external rig hardware is of less expensive low carbon steel material.

OF _OOR QUALi'_'y

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The rig safety systems are characterized by three primary modes. The first is an explosive system activated by the occurrence of an overspeed condition. The matn stream airflow bypasses the rig and re-enters in the exhaust duct. The second system is a pop-valve, which also allows the airflow to bypass the rig.

In addition to speed, the pop-valve can be excited by loss of oil flow, bearing compartment adverse pressure gradients, excessive vibration, and excessive bearing temperature. The third system is an alarm system, which is activated when the limits of various rig parameters are exceeded.

8.3 MECHANICAL DESIGN 8.3.1 Rotating Hardware The rotating components in the rig consist of the front and rear bearing com- partments, rotor shaft and rotor assembly.

The rig rotor design was analyzed for critical speeds and the rotor tie bolts were analyzed for blade loss capability. In addition, the rotor air seals were reviewed for resonance and coincidence. The critical speeds and mode shapes determined from the rotor dynamics study are shown in Figure 8.3.1-I. Because the mode shapes indicate a critical condition, a forced response analysis was conducted. The resultant bearing loads corresponding to 0.002 cm (0.001 in) of bearing support vibration at the speeds shown were judged to be acceptable (Table 8.3-I).

The calculated stress in the rotor tie bolts resulting from a blade loss situation is 482,636 MPa (70 kpsi), which is well under the O.2-percentyield stress level of Inconel 718 material.

Seal dampers were provided for the rotor air seals as a conservative measure to avoid any resonance or coincidence problem, as discussed in Section 8.3.3.

RPM = 4500 RPM = 9850 UJ a Q RO _ F- TAILCONE /

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:E r < LU > < CASE f UJ DAMPER n" _""'' TA I LC ONf n,- uJ "_ _I.,u AXIAL LENGTH AXIAL LENGTH Figure 8.3. I-I High-Pressure Turbine Component Test Rig Critical Speed and Mode Shapes OF POOR QUALITY

TABLE 8.3-1

HIGH-PRESSURE TURBINE COMPONENT RIG

FORCED RESPONSE RESULTS

Rotor Bearing Loads for 0.002 cm (0.001 in)

Bearin 9 Support Vibration k_ (Ib)

Speed

Front BearinB Rear Bea_rin_ 4500 34 (75) 138 (306) 9850 230 (508) 386 (851) 8.3.2 Bearings and Seals The bearing compartments were designed using existing parts with minor modi- fications. Key features of the front and rear bearing compartments are de- scribed as follows.

The front bearing has a 220 mm bore and a 320 mm flanged outer diameter. Its maximum DN is 2.1 x 106 , and its maximum load is less than 4,535 kg (lO,O00 Ib). The calculated Bl life is greater than 250 hours.

The front compartment seals are the dry face type with an oil-cooled carbon rubbing plate. Stackpole 2080 carbon grade seals are used for improved durability. Air temperature in the compartment is estimated to be 65oc (150°F). The seal operating conditions are listed in Table 8.3.Z.-I.

TABLE 8.3.2-1 FRONT COMPARTMENT OPERATING CONDITIONS Seal Differential Pressure Pa (psi) Rubbing Speed Condi ti on ( AP) m/sec (ft/sec) FORWARD REAR Aerodynamic Design Point (ADP) 227,528 (33) 220,633 (32) IZl (400) Maximum Speed 262,002 (38) 255,107 (37) 134 (440) Differential /% P Range* 186,159-351,634 (27-51) 179,264-344,740 (26-50) * Range expected to be encountered during rig operation The design of the rear compartment bearing for the rig is the same as that of the number 4 bearing for the integrated core/low spool. This oil-damped bear- ing has a 165 mm bore and a 222 mm outer diameter and is equipped with a soft centering spring. It also features preloadtng and under race oil cooling. Its maximum radial load is less than 453 kg (1000 lb), and its calculated B1 life is greater than lO00 hours. The maximum DN is 1.6 x lO 6, The design of the rear compartment forward seal is the same as that of the number 4 seal for the integrated core/low spool, and the rig rear seal is derived from the integrated core/low spool number 5 rear seal. Both are the dry-face type seal with oil cooled rubbing plates. Stackpole 2080 carbon grade seals are used for improved durability. The temperature of the surrounding air is maintained at 65°C (150°F). Operating conditions of the seals are shown in Table 8.3.2-II.

TABLE 8.3.2-II OPERATING CONDITIONS OF THE REAR COMPARTMENT SEALS Seal Z_P Pa(psi) Rubbing Speed Condition Both Seals m/sec (ft/sec 1 ADP 41,368 (6) 89 (293) Max Speed 41,368 (6) 98 (324) Operating Requirement: 68,948 Pa (lO psia) Compartment Pressures 8.3.3 Air Seals Extensive vibrational analysis was conducted on all high pressure drop, high speed air seals. Characteristics of the seals that were studied include resonance, coincidence and aerodynamic flutter.

A total of five air seals was reviewed in the study. These included the front and rear seals of the front bearing compartment and the three high-pressure turbine component design seals shown in Figure 7.2.6-I. These three seals, having been studied for their vibrational characteristics at integrated core/ low spool conditions, were studied for response at rig speeds and pressure levels. The study results are as follows: Front compartment front air seal (or thrust piston seal). Resonance and coincidence characteristics are as shown in Figures 8.3.3-I and 8.3.3-2. Resonance margin for the stationary part of the seal is 60 percent and 45 percent for the rotating part. Coincidence margin is 46 percent.

A sheet metal damper of 0.170 cm (0.067 in) thick sheet metal was provided to the rotating member as a conservative measure to avoid any resonance. Aerodynamic flutter was considered unlikely for this seal because of the influence of the damper and relatively low pressure drop across the seal. The seal and its damper are shown in Figure 8.3.3-3.

OF POOR QUALITY 800 J 7O0 600m SO0 ¢3 Z ,.= 4O0 ROTOR/ROTOR/ND 300 -- LIMITING MARGIN LIMITING MARGIN 2OO RESONANCE LINE I I I I I l 2 4 6 8 10 12 NODAL DIAMETER Figure 8.3.3-1 High-Pressure Turbine Rig Thrust Piston Seal Resonance m 7O01 _ w.

ROTOR/GND/ND lo0 I I I I I I 2 4 6 8 10 12 NODAL DIAMETER Figure 8.3.3-2 High-Pressure Turbtne Rig Thrust Ptston Seal Coincidence 123 J ORIGlt_AL P_G_ _3 OF POOR QUALITY DAMPER CUT INTO 8 SEGMENTS (AMS 5504) Figure 8.3.3-3 Front Compartment Front Airseal (Thrust Piston Seal) O Front compartment rear airseal. This seal's function is to act as a backup seal in the event of a malfunction of the carbon seal. A one- knife edge design was selected as best suited for this purpose in the unlikely event of a carbon seal failure.

Resonance and coincidence margins for this seal were found to be in- adequate, Figures 8.3.3-4 and 8.3.3-5 indicate no margin near the three nodal diameter frequency mode for both resonance and coinci- dence. A damping feature was provided for this seal to avoid reson- ance or coincidence. The damper is constructed of 0.160 cm (0.063 in) thick sheet metal and is provided for both rotating and nonrotating members. The configuration of the seal with dampers is shown in Figure 8.3.3-6.

The addition of dampers to this seal, along with its single knife edge design, makes it a low risk for aerodynamic flutter.

High-pressure compressor discharge seal. Resonance margin for this seal, as shown in Figure 8.3.3-7, is 83 percent for the rotor and 68 percent for the stationary member. Coincidence margin is II percent, as shown in Figure 8.3.3-8. Flutter stability for this seal was analyzed for the integrated core/low spool application (Section 7.2.6.2) and found to be acceptable. The seal is therefore adequate for the rig application because of the same pressure drops. Sheet metal dampers 0.I09 cm (0.043 in) thick were provided for the stationary member as a precautionary raeasure to provide additional safety margin. The seal configuration with dampers is shown in Figure 8.3.3-9.

OR,,=,,]_L P_.C :E 1400 _ OF POGR QUALITY LAND/NO 120C I00_ m ROTOR/ROTOR/ND NO C3 Z E 4OO 2OO RESONANCE LINE I I I l I I 2 4 6 B 10 12 NODAL DIAMETER Fi gure 8.3.3-4 High-Pressure Turbine Rtg Front Bearing Rear Seal Resonance 16OO 14OO O z o ,.= 1OOO COINCI OOO I I I I I I 60O 2 4 6 8 tO 12 NODAL DIAMETER Coi nci dence Figure 8.3.3-5 High-Pressure Turbine Rig Front Bearing Rear Seal ROTATING DAMPER SLOTTED INTO 8 FINGERS (AMS 5504) OF pO0_ ,Q_"_'_ : _I NON-ROTATING DAMPER SLOTTED INTO 36 FINGERS (AMS 5504) Figure 8.3.3-6 Front Compartment Rear Airseal 24OO 14OO O Z ROTOR/ROTOR/NO 8OO_ 4OO I I I I I I 2 4 6 8 tO 12 NODAL DIAMETER Figure 8.3.3-7 High-Pressure Turbtne R19 High-Pressure Compressor Discharge Seal Resonance OF POOR QU_-_.L.q"I 2000 -- I000 m Z ..= 1200 -- 8O0 LIMITING MARGIN 4OO I I I I I I 2 4 6 8 10 12 NODAL DIAMETER Figure 8.3.3-8 High-Pressure Turbine Rig High-Pressure Compressor Discharge Seal Coincidence DAMPERSLOTTED INTO 40 FINGERS (AMS 5596) • 097cm - . 109cm THICK (.038 IN.) -- (.043 IN.)

Figure 8.3.3-9 High-Pressure Compressor Discharge Seal Thrust balance seal. The margin for resonance on the thrust balance sea] is 20 percent for the rotor, as shown in Figure 8.3.3-]0, and 56 percent for the stationary member. Coincidence margin is 21 percent, as shown in Figure 8.3.3-11.

A damper was provided for the rotating member, 0.073 cm (0.029 in) thick, as shown in Figure 8.3.3-]2.

Flutter stability is adequate because of the analysis completed for integrated core/low spool and the similarity in rig and integrated core/low spool pressure drops. The addition of a damper provides additional confidence in flutter stability.

Number 4 bearing buffer air seal. Resonance margin for the buffer air seal was determined to be 53 percent for the stationary member and 75 percent for the rotating member, as shown in Figure 8.3.3-13. The coincidence margin was detemined to be 46 percent, as shown in Figure 8.3.3-14.

Aerodynamic flutter was detemined to be unlikely because of the low pressure drop across the seal. The analysis indicated no need for a damper on this seal.

The seal configuration is shown in Figure 8.3.3-15.

6OO SOO 40O Z 3MITING dARGIN E 2O0 MARGIN RESONANCE LINE I I I I I I 2 4 8 8 10 12 NODAL DIAMETER Figure 8.3.3-10 High-Pressure Turbine Rig Thrust Balance Seal Resonance ORIGh'L_.L _':_L'_ '_" OF POOR QUAL;,Y OF p£_"_.!_ _ ..............

IK)Om 400 _,, Z m m 2OO IO0 I I I I I l 2 4 6 8 10 12 NODAL DIAMETER Figure 8.3.3-11 High-Pressure Turbine Rig Thrust Balance Seal Coincidence DAMPER CUT INTO 1 2 SEGMENTS ) Ftgure 8.3.3-12 Thrust Balance Seal and Damper ,,_ _ i f,_.'_'_ _ _ I_,_, :_ . _ _'_ 1200 -- _" _,_,,_ _..-'__., , 800 -- ROTOR_ROTOR_ND Z RESONANCELINE I I I I I I 2 4 6 8 10 12 NODAL DIAMETER Figure 8.3.3-13 High-Pressure Turbine Rig Number 4 Bearing Buffer Seal Resonance 800 -- Q Z O 600 -- LIMITtNG MARGIN LANDIND I I I I I I 2 4 6 8 10 12 NODAL DIAMETER Figure 8.3.3-14 High-Pressure Turbine Rig Number 4 Bearing Buffer Seal CoY nci dence ORIQINAL PAGE 14 OF POOR QUALITY Figure 8.3.3-15 Number 4 Bearing Buffer Seal - Rig 8.3.4 Static Hardware The rig static structure consists of the inlet and exit ducts, vanes and cases, including the active clearance control system.

The active clearance control system is used in the test program to evaluate its effectiveness during rig testing. The internal hardware of the active clearance control system for the integrated core/low spool is also used in the rig. The integrated core/low spool outer turbine case could not be used be- cause of the requirement to place the exit probes in a specific axial location, which is incompatible with the case design. A rig-unique case was therefore designed to accommodate the required instrumentation. The material chosen for the rig case, Inconel 600, is a relatively low-cost nickel-based alloy that adequately matches the thermal expansion properties of PWA I007 (Waspaloy) material in the integrated core/low spool. The rig design was analyzed to ensure that the outer air seal, over the turbine blade tip, moves out parallel to the rig centerline in order to maintain essential chordwise clearance between the blade tip and outer airseal platfom during temperature excursions.

The active clearance growth summary is shown in Figure 8.3.4-I.

8.4 SECONDARY FLOW SYSTEM AND THRUST BALANCE The turbine component rig simulates engine cooling air and leakage flows and blade tip clearances, while preventing oil weepage from the bearing compart- ments and excessive thrust bearing loads. Figure 8.4-I presents the rig secondary flows.

Sufficient rig instrumentation will allow measurement of these key flows. It will also provide confirmation of leakage and swirl field flows with static and total pressure sensors, cooling air temperatures and windage heat genera- tion with thermocouples. Key thrust balance cavity pressures will be contin- uously monitored and processed to give an on-stand readout of thrust load.

[

ORIG'_,'.;L "+ .... - ="..,+: : .: POINT A _IPOINT B OF POOR " :+' ...... ++' ,;" ' __ ' __ POINTC A GWTH _ GWTH L_ GWTH COND, PT A PT B PT C CONDITION 1 CM (IN) CM (IN) CM (IN) * GROWTH AT RIG RUNNING 0,0279 0,0248 0,0218 TEMPS, °C (=F) 1 (0.0110) (0.0098) (0,0086) 0.0596 0,0614 0.0632 CONDITION 2 2 (0,0235) (0.0242) (0,0249) * GROWTH WITH 93,3"C (200°F) AT (CROSS-HATCH AREA) DIFF,- 0,0317 0,0365 0,0414 PTS 1&2 (0,0125) (0.0144) (0,0163) Figure 8.3.4-1 Active Clearance Control Growth Summary The high-pressure compressor rear seal air supply will be injected tangentially to simulate the high-pressure compressor discharge bleed swirl in the engine.

There will be separate control of the disk front rim cavity "mini" tangential on-board injection flow to change the swirl level and, hence, blade supply pressure. This will also affec.t thrust load. The temperature of the active clearance control air will also be controlled between 18oc (650 F) and 204°C (400OF). This control will change outer air seal (blade tip) clear- ances over the range available with the engine active clearance control system.

The bearing compartments will be vacuum-pumped to provide a positive pressure gradient across the seals, thus preventing oil weepage. The rotor has been thrust-balanced so that the maximum load is 4009 kg (8839 Ib).

8.5 RIG INSTRUMENTATION The rig incorporates sufficient instrumentation to detemine turbine aerody- namic performance and monitor rig structural integrity. This instrumentation was specifically selected or designed to maximize data acquisition capability without unduly perturbating the flow characteristics of the rig. This instru- mentation measures overall stage performance and provides basic aerodynamic data as well as airfoil and endwall aerodynamic loading information. In addi- tion, it monitors rig safety parameters and also records the performance of the "mini" tangential on-board injection system and active clearance control system. Both the cascade and the full-stage rigs have essentially the same instrumentation. A rig supervisory system will be employed to automatically control all of the secondary flow systems. All probes and wires are calibrated before instal Iati on.

The following sections describe rig perfomance and structural integrity instrumentation. A summary of rig instrumentation is presented in Table 8.5-I.

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TABLE 8.5-1 INSTRUMENTATION SUMMARY TT PT TAir Tmetal PS Miscellaneous 2 i0 - 16 Rig Super- visory Control Inlet 8O 40 Vane Surface & Shrouds Blade Outer Air 22 28 22 Laser Prox.

Seal (OAS) and Active Clearance Control (ACC) System Exi t 48 48 4 Air Angle 16 16 Vane Cooling 22 22 Tangential On- Board Injector (TOBI) Disk Bore 12 64 Cooling Exit 8 8 Cavities Bearing Com- i0 24 22 I0 Vibration 96rtments 3 speed Totals 102 PT TT PS TAIR Tmetal Laser Proximity Vibrat ion 10 Speed Air Angle 8.5.1 Perfomance Instrumentation Perfomance instrumentation measures inlet and exit flow, and blade tip clearance.

Inlet Instrumentation. Inlet instrumentation consists of 4 total pres- sure rakes, each with 10 sensors. In addition, there are 12 static pres- sure taps located on both the inner and outer flowpath. The inlet rakes are properly located circumferentially to prevent any rotor excitation.

Exit Instrumentation. Exit instrumentation consists of 4 total pressure rakes and 4 total temperature rakes each with 12 sensors. In addition, there are 12 static pressure taps located in both the inner and outer flowpath. The rakes are located on a ring that can be traversed circum- ferentially approximately 30 degrees. These rakes are positioned 90 degrees apart. Coupled with the pressure and temperature rakes, there are 4 air angle probes located at the exit plane. These probes are wedge shaped and have the ability to traverse radially as well as circumferen- tially. They are used to measure two static pressures and a total pressure, and are also used to calculate exit air angle.

Blade Tip Clearance Measurement. Four equally spaced laser probes will be used to measure and monitor the blade tip clearance. These probes will be attached to the blade outer air seal shoes, extend through the outer cases, and sealed by means of piston rings. Temperature variations in the flow to the active clearance control system will provide tip clearance changes. Probe measurements will then be used to determine any changes in performance resulting from variations in tip clearance.

Compartment Air Temperature. Air temperature measurements taken in various compartments, especially those conducting secondary air flow, will be useful in calculating performance parameters.

8.5.2 Structural Integrity Instrumentation This instrumentation verifies design assumptions and measures static tempera- tures, pressures, and bearing temperatures. Probe strain gages and accelero- meters are also used.

Static Pressure. Two turbine vanes will be instrumented on the airfoil surface at three spanwise locations. Seventeen airfoil static pressure taps will be positioned at each location, 10 on the suction surface and 7 on the pressure surface. Endwall static pressure taps will be placed at both the inner and outer flowpath platforms. The number of these used will depend on available platform area. Static pressure taps will also be located throughout the rig in order to determine cooling and leakage flows.

Static Temperature. One vane will be instrumented in the airfoil section to determine the effectiveness of the cooling scheme. In addition, instrumentation will be located on the active clearance control system to verify its operation. All cooling flow systems will be instrumented to monitor and maintain control of coolant flow temperatures.

135 [ Bearing Temperatures. Thermocouples will be located in the bearing Compartments to ensure proper bearing and seal operating temperatures.

This instrumentation will be monitored throughout the test program.

Probe Strain Gages. The fixed inlet rakes will be strain gaged at the root of _e rake to detect probe vibratory modes that might exist at test conditions.

Accelerometer. Both horizontal and vertical accelerometers will be located on the front and rear bearing supports. These wtll be monitored throughout the test program.

8.6 FACILITY/RIG ADAPTATION The facility for testing the high-pressure turbine component rig will provide an open-loop air supply system. Two natural gas burners heat the primary air to approximately 426oc (800OF). The secondary air, supplied by the main- stream air upstream of the combustor, is delivered to the rig at approximately 65°C (150°F), thus maintaining the proper main air temperature and cooling air temperature ratio. Power generated by the turbine is absorbed through two lO,O00 horsepower dynamometers. The rig is connected to the power absorption system through a coupling and gearbox. There are provisions for seven indepen- dent secondary cooling air systems. Flow for both the primary and secondary air systems is metered through critical flow venturies. An automated data secondary system will be used to process data, and a rig supervisory system will be employed to control all secondary flows.

SECTION 9.0

CONCLUDING REMARKS

The high-pressure turbine design in the Energy Efficient Engine represents a considerable advancement in turbine technology. The design capitalizes on the inherent economic and weight advantages offered with a single stage system by applying technology advancements in the areas of structures, ae,-odynamics and materials. On the basis of results acquired from design analyses and, more im- portantly, supporting experimental test programs, turbine performance and dur- ability goals are achieveable. The predicted efficiency of 88.8 percent for the flight propulsion system exceeds the goal of 88.2 percent. Durability ex- pectations are commensurate with commercial service operation.

Overall, the technology evolved through the turbine design processes and eventual demonstration during the integrated core/low spool test program will have wide application. Much of this technology, particularly high temperature capability materials, is applicable to any gas-turbine engine of the next gen- eration, including advanced derivative models of current commercial engines.

APPENDIX A

ENERGY EFFICIENT ENGINE HIGH PRESSURE TURBINE

COMPONENT VANE AND BLADE AIRFOIL COORDINATES

PRECEDING PAGE BLANK NOT RL_|EO

ORIGINAL P,_ [3

OF POOR QUALITY

EEE ENGINE VANE COORDINATES

ROOT, RADIUS =(13.880in.) 35.255 cm

PRESSURE SUCTION

SIDE SIDE

X/BX

Y - TOP Y - BOTTOM

3.66059 -1.67371 3.o_059 0.0 3.73262 -1.E_670 3.56876 O.CIO 5.76822 -1.63_69 3.51296 0.020 3.79_10 -1.63267 3._8706 O. 030 3.51>11 3._(595 -I.60566 O. 0"+0 3.83236 -|.58865 3._$33 G. 0_0 3.8_39 -1.57164 3._2_1 0.060 3._6315 3._0907 -I .55662 0.070 3.87901 -1.53761 3.3875_ O. 0_.,_ 3.89318 -! _2060 3 36_97 O.OqO 3.90665 3 3_151 -1 50359 0.100 3.937_8 3 27951 -1 _5106 0.125 3.96354 3 21_59 -1 61P53 0.150 3.9053_ 3 166_8 -1 37600 0.175 6.0G_63 3 07267 -1 333q6 0.200 6.01530 2 99351 -1 29003 0.225 6.02322 -l 24840 0.250 _.02625 2.8_13 -1 20557 0. 275 6.02620 2.76625 -1.16336 O. 300 6.01G_7 2.68276 -1.12081 O. 325 6.00399 2.59976 -1.07_28 O. 350 3.9_527 2.51528 -I.03575 0.375 3.95035 2._2o61 -0.99322 0._,00 3.92880 2._220 -0.95069 0.6?5 3._008 2.25_66 -0.90816 0.450 3.8_357 2.16_1 -0.86563 0.475 3.788_6 -0.8_310 2.07268 0.500 3.7_369 -0.78057 1.90025 0.5?.5 3.66796 -0.73806 1.E8653 0.550 5.55953 -0.69551 1.79167 0.575 3._55_0 -0.65Z98 1.6_506 0.600 3.333_9 -0.610_ 1.59725 0.625 3.19012 -0.56791 1._9600 0.650 3.0_F81 -0.52538 1.39720 0.675 2.8_281 1.29_02 -0._8_5 0.700 2.6_61 1.19066 -0._032 0.725 2.43150 !.03_$1 -0.39779 0. 750 2,_0936 0.97652 -0.35526 O. 775 1.97950 0.86606 -0.31273 .0.800 1.76356 0,75285 -0.27_20 '0,825 1.50299 0.63667 -0.22767 0.850 1.25882 0.51613 -0 1G51_ 0..375 1.0117_ 0.39080 -0 1,*_61 0.900 0.91227 0.33881 -0 12559 0.910 0.812_5 0.28551 -0 10_5_ 0.920 0.71236 0.23073 -0 09157 0.930 0.61190 0.17326 -0 07_56 0. :_,_0 0.51123 0.11320 -0 05755 O. 950 0.61033 0.06902 -0.0_053 O. 900 0.3091_ -0.01628 -0.02352 0.970 0.Z0783 -0.00651 -0.02676 0.980 0.10620 -0.0256_ 0.01050 0.990 0.00221 0.02752 -0.23526 I .000

ORIGINAL PAG_ [,_

EEE ENGINE VANE COORDINATES

OF POOR QUALITY

MEAN, RADIUS : _5.001in.) 38.103

cm

PRESSURE

SUCTION

SIDE

SIDE

X/BX

Y - TOP

Y - BOTTOM

-1,67335 0.0 3.91742 3.91742 -],65634 O.OIO 6.00945 3.82538 -1.63933 O.G20 4,04534 3. 78950 -].62232 0.030 4.07129 3.76354 -1.60531 0.O4O 4.09185 3. 74302 -1.59830 O.O.rO 4.10909 3.725C0 -1.57130 0.060 4,12553 3.70704 -1.55429 O. 070 4.1414B 3.68795 -1.53728 O. 080 4.15691 3.66797 -1.5£027 O. 090 4.17181 3.64717 -1.50326 O.IO0 4.18619 3.62561 -1.46074 O. 125 4.21973 3.56857 -1.418Z2 0.150 4.24969 3.50752 -1.37570 O. 175 4.27590 3.44289 -1.33317 0.200 3.37501 4.29813 -I,24065 0.225 3.30414 4,31615 -1.24813 O. _.50 4.3_969 3230_9 -1.2_561 O. 275 4.33_3 3,15421 -1.16309 0._o0 4.36205 3,07545 -1.12057 0.325 4.34016 2.99_32 -1.07604 0.350 4. 33231 2.91090 -1.03552 0.375 4.31796 2.82526 -0.9_300 0.400 4.29650 2.737_4 -0.95048 0.425 4.26742 Z.6_750 -0.90796 0.450 4.22769 2.55544 -0.86544 0.475 4.18239 2.461_8 -O 822_2 0.500 4.12_3! 2.36501 -O 76039 0.5£5 4.05396 2.76664 -0 73737 0.550 3.96931 3. 16610 -0 6_535 O. 575 3.86790 2.06337 -0 6_C33 0.600 3.74623 1.95_42 -O 61031 0.625 3.60056 1,_5113 -0 56779 0._50 3.43406 1.74144 -0 52526 0.675 3.25073 1.62920 -0 4C274 0. 700 3.05355 1,51428 -O 44022 O. 77.S 1.39651 2.84441 -O 39770 O. 750 2.62461 1.27566 -0 35516 O 775 1.15140 2.395_5 -0 3]366 0 gO0 2.15772 1.02343 -O 27014 O _25 0.89123 1.9]197 -0.2276] 0 t_,50 1.65Ca0 0.75_21 -0.18509 O g75 0.61152 1.3_75 -0.14257 O ._O0 |.13205 0.46157 -0.12556 0.910 1.02_39 0.39922 -0.10_55 0.920 0.91452 0.33554 -0.07155 0.930 0.80392 0.26953 -0.07454 0.940 0.69248 0.Z0168 -0. 05753 0.950 0.13099 0.58006 -0.04052 O. 960 0.46692 0.05762 -0.02351 0.970 0.35261 -0.01430 -0.00650 0.960 0.23737 -0.02674 0.0105l 0.990 0.12109 -0.02543 0.02752 I .000 0.00194 0.24012

!

ENGINE VANE COORDINATES

EEE

OF PCOR Q : +

cm

42. 118

= (16.582 in.)

LE. TIP, RADIUS

SUCTION

PRESSURE

SIDE

SIDE

Y - BOTTOM

Y - TOP

X/BX

_.31185 4.31186 -1.67286 _,Z1987 0.0 4.40384 -1.65584 6.18401 0,010 4.;3971 _.15807 0.020 6.66565 -1.02183 _.13756 0.030 4 ._8629 -1.60483 4.11964 0.0,'+ 0 6.50353 -1.5C782 4.101_0 O. 050 _.51997 -I.5}OSZ q.G_I56 0.0S0 4.53591 -1.55337 _.O_OB3 0.070 _.55135 -1.53681 4.03907 0,030 -1.51901 _.01636 0.090 _.5_071 -1.50251 3.95381 0.100 6.61452 -I.96030 3.G_042 O. 125 6.6_5_6 -1.41779 3.3_078 0.150 _.67_5 -1.375C3 3.7_732 q.175 _.69_00 -1._3_77 3.67033 O.C.O0 _.71617 -1.290Z6 3.5_012 0.225 _.73263 -1.Z:,775 3.50688 U._50 4.79524 -1.205_5 3.42079 O. 275 4.75379 -1. 1627_ 3.33197 0.300 4.75_06 -1.120_3 3._053 0 ._5 4.75778 -I 07772 3.14656 0.350 4.752_5 -I 03521 3.05012 0,375 6.7_Z9 -0 9_70 2.951Z5 0.400 4.7_6C5 -0 950:9 2.8_9_3 ,) .025 4.70398 -0 c07C8 2.7_635 0. L,50 4.674C0 -0 86510 _.6_033 0._75 _.63785 -0 8_C57 _.5_l_0 l). 5_ 0 _.5_ZO_ -0.7001b 2.42106 _,5t.5 6.535_5 -0,73765 Z.30773 0.550 _._6716 -0.6751_ Z.191_6 0.575 4.3_318 -9,G5263 2.07336 0.600 -0.61012 1.95211 0.625 -0.56762 1,8Z799 0,650 3.9771_ -0.5_511 1.70083 0.675 3.775_3 -0.46_60 1.57038 0.7C0 3.5q_45 -0.4_009 1.43639 O. 725 3._976_ -0.3_75B 1.2905_ O. 750 3.02703 -0.35507 1.15632 0. 775 2.73928 -0.31_56 1.00_I 0.800 2.q3612 -0.27006 0._5638 0.8Z5 2.119_3 -0.22755 0.69_65 0 .C,50 1.7_003 -0.10504 0.5_639 O. $75 1.45117 -0.14253 0.45812 0.900 1.31249 -0.1_553 0.38551 0.910 1.17Z#4 -0.I0_52 0.31065 0.9_0 1.030_ -0.09152 0.23384 0.930 0.0_12 -0.07452 0.15359 0.940 0.74354 -0.05751 0.0684_ O. 950 0.59040 -0.04051 -0.01430 0.960 0.45127 -0.02351 -0.02673 0.970 0.30305 -0.00650 -0.0Z5_3 O.9GO 0.1539Z 0.01050 -0.333_7 O. 9';'0 0.00152 0.0Z751 1.000

QL-AL_TY

EEE ENGINE

BLADE

COORDINATES

ROOT, RADIUS =(13.678in.) 34.742 cm

SUCTION PRESSURE

SIDE SIDE

X/BX

X

Y - TOP Y - BOTTOM

0.00006 I.o8397 1,05397

0.0 0.01352 0.010 1.13073 1.04116 0,02695 l.lC03_ O. 0:-0 1.02650 O.0_oq5 0.030 1.01756 0.05391 0.0q0 1.21SC6 1.01216 0.06737 0.050 1.00955 O.OGOS4 0.060 1.2_638 1.00941 0.09430 C. 070 1.2!018 1.01175 0.10776 0.080 1.31302 1.01680 0.12123 O.CgO 1.33_95 1.02512 0.13q69 0.100 1.35601 1.03475 0.16B35 0.125 1.405_3 1.055!5 0._$2CI 0.150 1.07394 0.2_067 0. 173 1.0_31 !.4%c57 0.26933 C,_00 1.5Z554 1.0_9q3 0.30_99 o .225 1.55756 1.10747 0.33655 O._SO 1.576t5 _._5_ 0.37030 0.275 1.6!003 1.11470 0,40396 0.300 1._3203 I.i1405 0.43762 0.325 1.64_75 I,!I069 0 ._7128 0. 350 1,663'_0 l.lOq_4 0.504_ 0.375 1.67_q6 1.09392 0.53950 0.400 1.68131 0.57226 O. 4._5 1.68_35 I C70¢5 0.60_92 0.450 !.6933g I 05409 0.6355B 0.475 1.67_16 1 03491 0.67323 0.5C0 1.66035 1Cl3!l 0.70689 0.5Z5 1.63331 0 _L363 0.74055 0.550 1.63303 0 9_143 0.77431 0.57_ 1.60610 0 93151 0._0707 0.500 1.5713_ 0.69573 0.64153 0.625 1.52772 0.66303 0.87519 0.650 1.47161 0.90385 i.39C37 0.575 0.7_241 0.94251 1,30322 0.700 0.737_4 0.97617 1.21310 0.725 0.65036 0.6%_0 0. 750 1.04348 1.01i40 0. 775 0.57_5 1.07714 0.90717 0.51_4 0.800 l.llC_O O.BOIOI 0.45397 0.525 l.lq4q6 0.6_305 0 ¢50 0.333q5 1.17012 0.50330 0 675 0.30721 1.2117_ 0.47106 0._2_;34 Q <;00 1.72524 0,42674 0.]89!8 0 910 1.23_70 0.35123 0.15270 0 9."0 1.25317 0.33535 O.IIA_6 0 930 1.26563 0.25949 0.07499 0 940 1.27909 0.24275 0.03371 0 950 1.29256 0.1957_ -0.00955 0 960 I._0602 0.14835 0 970 -0.0_621 1.31943 0.1002_ 0.960 -0.0_9Q4 _.332_5 0.05151 -0.02561 C.990 1.34641 -0.00199 1.000 -0.00199

EEE ENGINE BLADE COORDINATES

OF FOOR QULLITY

37.879 cm

MEAN, RADIUS =I14.913in.)

SUCTION PRESSURE

SIDE

SIDE

X Y - TOP

X/BX Y - BOTTOM

0.11588 1.37101 0.0 1.37101 0.12748 1.41_88 0.010 1.33092 0.13907 1.44625 0.020 1.31674 0.150_6 1.q7_ll 0.030 1.30765 0.16225 1.49979 O. Ot+O 1.30102 0.17724 1.52358 0.050 1.29729 0.18543 1.5q570 1.29611 0.19703 1.56630 0.070 I._9615 0.203_2 1.5_353 0.000 1.29501 1.60151 0.22021 0.0'?0 1.30183 0 231_0 1.62332 0.I00 1.30797 0 :5978 1.65774 0.1."5 1.32937 0 2_976 1.6_%27 O. 150 1 30718 0 31874 1.71555 0.175 1 35_5 0 3_772 1.73717 0.200 1 36_53 0 37670 1.75q29 0.;'25 I _{554 0 40568 1.76721 0.250 ! _6:'23 0 43q{,G 1.77526 O. 275 1 25399 0 4679% 1.72_'_I 0.300 1 341f6 0 49261 1.7_!77 0.3,'5 1 32603 0 52159 1.77_51 0.350 I _0652 0 55057 1.77129 O. 375 1 26_/,2 0 57955 1.75955 0.400 1 ZE_95 0 _0353 1.7q3_2 0 .(_25 1 22732 0 63751 1.72222 o. _, 50 i 19"_5 0 666_9 1.695_9 O.q7E, I. 15_;'_% 0 695:;7 1.662q3 o. 5f: 0 1.11957 0 7_/+45 1.62107 0.52.5 I .07777 0 753_3 1.57205 0.550 1.03_15 0 7_4! 1.50979 0.575 0.9S176 0 31139 1._:3295 0.6t!0 0._5_62 0 6_037 1.3502_ 0.62.5 0.B8277 0 86935 1.26463 0.6"_0 0.827_7 0 89_32 1.17735 0.675 0.76907 0 92730 1.05S67 0.7(_0 0.70328 0 95628 0.932_0 0.725 0.6q485 0 93525 0.750 _._7272 1.01424 0.81(,39 0.T;5 0.51007 1.0#322 0.72¢433 0.800 0.q_77 1.03"220 0.65081 0.225 0.36_76 1.10118 0.5_31 o.sr_,o 0.23_10 1.13016 0.4_235 0.875 0.20371 1.1591G O._q6G3 0.900 0.12659 1.17073 0.30793 0.910 0.00297 1.1_232 0.2692_ O. 9:'0 0.058C5 1.19391 0.23035 0.930 0.02427 1.20550 0.19!_7 0.940 -0.01061 1,21710 0.15220 0. 950 -0.046_6 1.22669 0.11255 0.950 -0.07201 1.24028 0.07297 0.970 -0.078_8 1._5187 0.03329 O. 930 -0.07909 1.263_6 -0.00666 0.990 -0.07_35 1.27505 -0.05190 1.000 -0.05190

OF POOR QUALITY

EEE ENGINE BLADE COORDINATES

TIP, RADIUS = 06.149 in.) 41.018 cm

SUCTION PRESSURE

SIDE SIDE

XIBX

Y - TOP Y - BOTTOM

O. 17520 0.0 1. 95437 !. 95407 0.1_22 0.010 1.99237 I. 91736 0. I g_24 0.020 2. 00598 !. 90375 0.20526 0 030 2.01502 I .89_71 0.21528 2.02,150 0 ,'Y_O I ._$036 0.22_30 2.03714 0 0_3 1.88397 0.23532 2.03233 0 060 1.88118 0.2qSE4 2,03707 O 070 1.87982 0.25537 2.04236 00CO 1.87956 0.26539 2.0q523 0 GgO 1.87915 2,0_:34 0 100 O. _. 75.11 1.87728 O. 300_6 ,?.. 05538 I .86500 0 ]25 0.3"_551 2.05_21 I .85291 0:5O 0.35057 2. G60_l 0 175 I .83251 0.37562 2.e5__3 I ._0741 0 200 0.43367 2. c_4_1 0 225 I .77C09 0,4," 57_' 2.0c,703 0.250 I. 7_',:+ _ 2 0._3077 2.03"153 0 • ?75 1. 70323 0.47:-33 2. 02274 C. 7.00 1.668_3 0.530_3 2. 0053t_ 0.3,'1.5 1.6?573 0.525?3 l. <_SqO 0 O. 350 1.5o045 0.550'_8 1.95.322 0.3;'5 1,53257 0.57604 1.92732 0._,00 I .48274 0.50109 1.6g0_9 0.425 1 43089 0.62614 I .04567 0.450 1 377I 3 0.65119 I. 7-3075 0.475 1 32164 0.67_25 I. 722_5 G.EC3 1 264_r-7 0.70130 1.6td>89 .5:5 1 20593 0.72635 1.56466 0.550 I I_{.i 3 0.751Q0 0.575 I 03CS1 0.776q5 I .Q3138 O.COO 1 02244 0.801.51 1.31932 0.b25 0 gS_B5 0.82655 0.650 0 89429 0.851GI I. 15,!47 0.675 0.B2059 0.87667 1.07i97 0,700 0. 767:.34 0.90172 O. _3941 O. 7;:5 0,69495 0.92677 0.750 0.62077 O. 9_5282 0.&2422 O. 775 O.5D7P_5 0.741(,5 0.976C0 0.6CO 0 ,_:C3Z2 1.00193 0.65093 0 .L,25 0,417_5 1.02_';8 0.57325 0.C50 0,316<;0 1.05203 0 49359 O.._75 0,275_0 0.41092 1.07733 0.';00 0.20323 O. 37704 1.0:3711 0.910 0. _ 7420 0. 34477 I .09713 0.900 0.I_510 0.31164 1.10715 0,930 0.11590 0.27_60 1.11717 0.940 O. G_668 0.24552 1.12719 0.05915 O.gSO 0.2124_, l. 13721 0,05032 0.950 0. i 7935 I • 14723 0.04 738 G .970 0. 14630 1. 15725 0.04G39 0._50 0.11315 I. 16727 0.05352 0 .<;gO 1.17729 O. 07474 0.07_7_ i.000

14.5

OF POOR QUAL|_ APPENDIX B LIST OF SYMBOLS A Annulus area ACC Active clearance control ADP Aerodynamic design point amb. Ambient Btu British thermal unit CET Combustor exit temperature CpB Base pressure coefficient Cx Axial flow velocity D Di ame ter FP Flow parameter, W_FT-/P FPS Flight propulsion system HPC Hi gh-pressure compressor HPT Hi gh-pressure turbi ne IC/LS Integrated core/low spool ID Inner diameter LE Leading edge LPT Low-pressure turbine Mach number M, M n N Mechanical speed, rpm N2 High-pressure rotor speed, r1_m OAS Outer a irseal OD Outer diameter P Pressure PF Patter ,_factor A P/P, Pr Pressure ratio LIST OF SYMBOLS (Continued) PT Total pressure R Ra nk i ne rpm Revolutions per minute S Span SLTO Sea level take-off T Temperature TE Trail i ng edge TOBI Tangential on-board injection U Tangential wheel speed Urim Rim speed Wae Engine airflow Wc/a Total cooling airflow 0 Angle Angle A Delta Deflection Efficiency Stress CT REFERENCES I. Energy Efficient Engine High-Pressure Turbine Uncooled Rig Technology Program (CR-165149) 2. Energy Efficient Engine High-Pressure Turbine Supersonic Cascade Technology Report (CR-165567) 3. Energy Efficient Engine High-Pressure Turbine Leakage Technology Report (CR-165202) 4. Energy Efficient Engine High-Pressure Turbine Cooling Model Technology Report (CR-165374) DISTRIBUTION LIST GOVERNMENT AGENCI ES NASA Headquarters 600 Independence Ave., SW Washington, D.C. 20546 Attention: RTP-6/R.S. Colladay RTP-6/C .C. Rosen RTP-6/L. Harris RRP-6/J. Facey Library NASA-Lewis Research Center 21000 Brookpark Road Cleveland, OH 44135 MS 301-2 Attention: D. L. Nored MS 301-4 (20 copies) C. C. Ciepluch MS 30i-4 J. W. Schaefer MS 301-4 P. G. Batterton G. K. Sievers MS 301-2 MS 60-3 (2 copies) Library MS 5-5 Report Control Office Tech Utilization Office MS 3-19 M. A. Beheim MS 3-5 M. J. Hartmann MS 5-3 MS 60-4 R. A. Rudey R. A. Weber MS 500-127 MS 501-10 W. C. Strack T. P. Moffitt MS 77-2 J. E. Rohde MS 77--2 R. L. Dreshfield MS 105-1 MS 500-305 A. Long W. M. Braithwaite MS 500-208 L. Reid MS 5-9 MS 501-3 AFSC Liaison Office MS 302-2 Army R&T Propulsion Lab NASA-Lewis Research Center NASA-Lewis Research Center 21000 Brookpark Road 21000 Brookpark Road Cleveland, Ohio 44135 Cleveland, Ohio 44135 Attention: D.L. Nored MS 301-2 Attention: R.J. Weber MS 500-127 NASA-Lewis Research Center NASA-Lewis Research Center 21000 Brookpark Road 21000 Brookpark Road Cleveland, Ohio 44135 Cleveland, Ohio 44135 Attention: W.C. Strack MS 501-10 Attention: C.C. Ciepluch MS 301-4 20 Copies NASA-Lewis Research Center NASA-Lewis Research Center 21000 Brookpark Road 21000 Brookpark Road Cleveland, Ohio 44135 Cleveland, Ohio 44135 Attention: J.W. Schaefer MS 301-4 Attention: T.P. Moffitt MS 77-2 NASA Lewis Research Center NASA-Lewis Research Center 21000 Brookpark Road 21000 Brookpark Road Cleveland, Ohio 44135 Cleveland, Ohio 44135 Attention: P.G. Batterton MS 301-4 Attention: J.E. Rohde MS 77-2 NASA-Lewis Research Center NASA-Lewis Research Center 21000 Brookpark Road 21000 Brookpark Road Cleveland, Ohio 44135 Cleveland, Ohio 44135 Attention: G.M. Sievers MS 301-2 Attention: R. L. Dreshfield MS 105-1 NASA-Lewis Research Center NASA-Lewis Research Center 21000 Brookpark Road 21000 Brookpark Road Cleveland, Ohio 44135 Cleveland, Ohio 44135 Attention: Library MS 60-3 MS 500-305 Attention: A. Long 2 Copies NASA-Lewis Research Center NASA-Lewis Research Center 21000 Brookpark Road 21000 Brookpark Road Cleveland, Ohio 44135 Cleveland, Ohio 44135 Attention: Report Control Office MS 5-5 Attention: W.M. Braithwaite MS 500-208 NASA-Lewis Research Center NASA-Lewis Research Center 21000 Brookpark Road 21000 Brookpark Road Cleveland, Ohio 44135 Cleveland, Ohio 44135 Attention: Tech Utilization Office Attention: L. 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1s3 [

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