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Produced by the NASA Center for Aerospace Information (CASI)
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1. R £PORT NO. 

CR-174878 

4.TIT[E AND SUBTITLE 

PRELIMINARY DESIGN OF A SUPERSONIC CRUISE 

AIRCRAFT HIGH-PRESSURE TURBINE 

L. D. Aceto, J. C. Ca1derbank 

9.PERFORMING ORG. NAME AND ADDRESS 

UNITED TECHNOLOGIES CORPORATION 

Pratt & Whi tney Engineering Division 

12. SPONSORING AGENCY NPlt1E AND 

National Aeronautics and Space Administration 

Lewis Research Center 

21000 Brookpark Road, Cleveland, Ohio 44135 

1 5. sUpp[ EMENTARY NOT ES 

. kEPORT DATE 

November 2, 1983 

PERFORMING ORG. CODE 

PWA-5923-14 

11. CONTRACT OR GRANT NO. 

NAS3-20357 

Final Report 

14. SPONSORING AGENCY CODE 

533-04-12 

Final Report, Project Manager: Frederick C. Yeh, Inter~al Fluid Mechanics 

Division, NASA Lewis kesearch Center, Cleveland, Ohio 44135 

1 6. ABSTRACT 

Development of the supersonic cruise aircraft engine continued in this National Aeronauti cs and Space Admi nistrati on (NASA) sponsored Pratt & Whitney program for the Preliminary Design of an Advanced High-Pressure Turbine. Airfoil cooling concepts and the 

technology required to implement these concepts received particular emphasis. Previous 

supersonic cruise aircraft mission stUdies were reviewed and the Variable Stream Control 

(ngine (VSCE) was chosen as the candidate for the preliminary turbine deSign. The design 

was evaluated for the supersonic cruise mission. The advanced technology to be generated 

from these designs showed benefits in the supersonic cruise application and subsonic 

cruise application. The preliminary design incorporates advanced single crystal materials, thermal barrier coatings, and oxidation resistant coatings for both the vane and 

blade. The 1990 technology vane ~nd blade designs have cooled turbine efficiency of 92.3 

percent, 8.05 percent Wae cooling and a 10,000 hour life. An alternate design with 1986 

technology has 91.9 percent efficiency and 12.43 percent Wae cooling at the same life. 

To achieve these performance and life results, technology programs must be pursued to 

provide the 1990's technology assumed for this study • 

. Supersonic Cruise Aircraft 

High-Pressure Turbine 

General 

19. SECURITY cLASs THIS (REPT)120. SECURITY cLASs THIS (PAGE)1 21 . NO. PGS 122. PRICE * 

UNCLASSIFI ED UNCLASSIFI ED 120 

* For sale by the National Technical Information Service, Springfield, vA 22161 

e 

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--- ---
TABLE OF CONTENTS 

SECTION TITLE PAGE 

1.0 SUMMARY 1 

2.0 INTRODUCTION 3 

3.0 ENGINE AND MISSION SELECTION 5 3.1 Engine Selection 5 3.2 Mission Selectiun 7 

4.0 BASIC TURBINE PARAMETER DEFINITION AND AERODYNAMIC DESIGN 12 4.1 Aerodynamic Design 12 4.2 Aerodynamic ronc1usions 26 4. 3 Mechanic~ 1 [,; ,:,".1 gn Consi derati ons 26 

5.0 C'URABILITY DESiGN 27 5.1 Configurationdl Studies 27 5. 1 . 1 Coolant Precooling 28 5.1.2 Trailing Edge Internal Cooling Geometry 39 5.1.3 Blade Leading Edge Study 41 5.1.4 Thermal Bat'rier Coating with Film Cooling 45 5.1.5 Low Heat Load Design 50 5.~ Preliminary Design 54 5.2.1 Technology Projection and Desi gn Features 57 

I. 5.2.2 Design Criteria 58 5.2.3 Mission Life Analysis 59 

r 5.2.4 Vane Durability Design 64 5.2.5 Bl ade Durabil i ty Desi gn 84 

F 

, 

~ 6.0 BENEFITS EVALUATION 105 , . 6.1 Supersonic Benefit 105 6.2 Subsonic Benefit 105 

7.0 CONCLUSION 110 

8.0 RECOMMENDATIONS 111 

APPENDIX A - HIGH-PRESSURE TURBIh~ AIRFOIL CCORDINATES 112 

SYMBOLS and ABBREVIATIONS 119 

RErERENCES 120 

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Number 

3-1 

3-2 

4-1 

4-2 

4-3 

4-4 

4-5 

4-6 

4-7 

4-8 

4-9 

4-10 

5-1 

5-2 

5-3 

5-4 

5-5 

5-6 

5-7 

LIST OF ILLUSTR".itOH~ 

Title 

Variable Stream Control Engine Flowpath 

Advanced Supersonic Transport Flight Profile 

and Reserves 

Primary Elements of the Preliminary Design Process 

Original ~SCE Turbine Flowpath High Mach Numbers 

Required Turbine Redesign 

High-Pressure Turbine Parametric Study Identified 

Improved Design 

Comparison of Original and Final Flowpath 

Turbine Geometry 

Airfoil Geometry 

Vane (Mean) Pressure Distribution 

Blade (Mean) Pressure Distribution 

Blade Mach Numbers Contour Map 

Blade Losses 

Convective1y Cooled Vane Leading Edge 

Vane Convective Leading Edge Cooling 

Convective1y Cooled Vane Leading Edge Using Precooled 

Air; (Suction Surface) 

Convective1y Cooled Vane Leading Edge Itsing Precooled 

Air; (Suction Surface) 

Vane Leading ~dge Convective Cooling; Effect of Precooling 

on Coolant Flow 

Vane Convective Leading Edge; Momentum Mixing vs Exit PreS~iJre Ratio 

Vane Leading Edge Convective Cooling; Cooling Flow 

vs TemperatJre 

iii 

~ - --------------...,..---""'----- 

Page 

8 

9 

13 

15 

16 

17 

18 

19 

22 

23 

24 

25 

29 

30 

31 

32 

33 

34 

36 

~~ 

'=
, 

.. 

5-8 

5-9 

5-10 

5-11 

5-12 

5-13 

5-14 

5-15 

5-16 

5-17 

5-18 

5-19 

5-20 

5-21 

5-22 

5-23 

5-24 

5-25 

5-26 

LIST OF ILLUSTRATIONS (continued) 

Title 

Vane Leading Edge Convective Cooling; Cooling ~, 

vs Precooling 

Influence on Thrust Specific Fuel Cunsumpt10n Change of Vane Precooled Airflow from Ledding Edge to Blade 

Cooling With No Weight P~nalty 

Vane Trai11ng Edge Concepts 

Vane Interface Temperature Distr1butions on Pressure 

Surface Distance, (1990 Technology) 

Blade Leading Edge COl1figurations 

A1rfoil Cooling Effectiveness One-Dimensional Analysis 

Airfoil Coolir.g Effectiveness Benefits with Thermal 

Barrier Coating 

Heatload Parameter vs Film Effectiveness; Blade e 0.713, 

T'J c • 0.34 

Heat Exchanger Efficiency vs HC AC/M Cp 

Blade Geometry for Low Heat Load Design 

Blade Load1ng for Low Heat Load Design 

Heat Transfer Reduction on Pressure Surface for Low 

Heat Load Design 

Blade Interface Temperature Reduced Pressure Wall 

Design Gas Temperature Profile (1990 Technology) 

Design Gas Temperature Profile (1986 Technolo~v) 

Mission Performance Increments 

Ambient Temper~ture Probability Distributi~n at 

16,154 m (53,000 ft) 

Variation of Combustor Exit Temperature as a Function 

of Variation of the Ambient Temperature 

Turbine Vane Thermal Barrier Coating Spalling Life vs 

Design Metal Temperature 

iv 

Page 

37 

38 

40 

42 

43 

46 

47 

49 

51 

52 

53 

55 

56 

60 

61 

62 

65 

66 

67 
LIST OF ILLUSTRATIONS (continued) 

FilJure 

N:.mber Title Page 

5-27 Turbine Blade Mission Creep Life ~s Average Design 68 Temperature 

5-28 Vane Cooling Configuration 70 

5-29 Vane Typical Trailing Edge Geometry 71 

5-30 Vane Coo'11ng Air Flows (1990 Technology) 73 

5-31 Vane Interface Temperature Distribution on Pressure 

Surface (1990 Technology) 74 

5-32 Vane Interface Temperature Distribution on Suction Surface Distance (1990 Technology) 75 

5-33 Geometry of Suction Surface Shaped Holes 76 

5-34 Film Effectiveness vs Nondimensional Distance Cylindrical 77 and Shaped Holes 

5-35 Geometry of Conical Showerhead Holes 77 

5-36 Conical Holes Cooling Air Flow Rates 78 I 5-37 Vane Hole Geometry (1990 Technolcgy) 79 

r 5-38 Vane Cooling Flows (1986 Technology) f 80 

r· 5-39 Vane Interface Temperature Distribution on Pressure ~ 

Surface Distance (1986 Technology) 81 

5-40 Vanp. Interface Temperature Distribution on Suction Surface 

(1986 Technology) 82 

5-41 Vane Hole Geometry (1986 Technology) 83 

5-42 Blade Cooling Configuration 85 

5-43 Blade Trailing Edge Cooling Configuration Geometry) (Representative 87 

5-44 Blade Trailing Edge Cooling Losses vs Blockage 88 

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Number 

5-45 

5-46 

5-47 

5-48 

5-49 

5-50 

5-51 

5-52 

5-53 

5-54 

5-55 

5-52 

5-57 

5-58 

5~59 

6-1 

LIST OF ILLUSTRATIONS (continued) 

Title Page 

Blade Average Pull Stress vs ~pan for N2-10.146 RPM 89 

(Average Deterioration) 

First Blade Trailing Edge Thickness Efficiency Sensitivity 90 

1990 Cooling Flow Distribution 91 

Blade Creep at Critical Locations (1990 Technology) 92 

Blade Interface TPI1Iperatl;re Distribution on Suction 

Surface Distance (1990 Technology) 94 

Blade Interface Temperature Distribution on Pressure 

Surface Distanr.e (1990 Technology) 95 

Blade Mean Section Isothermals Based on 1990 Technology 96 

Bl ade Hole Geometry (1990 Technology) 97 

Blade Trailing Edge and Midchord Cooling Air 

Optimization (1990 Technology) 98 

Blade Mean Section Isothermals Based on 1990 Technology 99 

Blade Surface Temperature Distribution on Pressure 

Surface Distance (1986 Technology) 100 

Blade Surface Temperature Distribution on Suction 

Surface Distance (1986 Technology) 101 

Blade Creep at Critical Locations (1986 Technology) 102 

1986 Cooling Flow Distribution 

Blade Film Hole Geometry (1986 Technology) 

ATSFC vs AN2 

vi 

108 

104 

109 

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Table 

Number 

3-1 

3-11 

3-1rl 

3-IV 

4-1 

4-II 

4-111 

4-IV 

5-1 

5-11 

5-III 

5-IV 

5-V 

5-VI 

6-1 

6-11 

6-III 

7-1 

LIST OF TABLES 

Title 

Engine Weight Summary 34~ k~/sec (750 1b/sec) Ai rf1 ow S 11e 

Emissions Comparison 

High-Pressure Turbi::e Comparison 

Definition of Supersonic and Typical Subsonic Flight Missions 

Turbine Parameters for 1986 and 1990 Designs 

Turbine P~rameters 

Comparison of Losses With 50 and 60 Blades 

Efficlency Breakdown 

Precoo1er Configuration 

Results of Augmented Channel and Impingement Cooled 

Showerhead Leading Edge Comparison 

VSCE Materials Projections 

Superscnic Cruise Aircraft En~ine Mission Life 

Design Parameters 

VSCE Design Point Performance Supersonic Cruise 

Breakdown of Blade and Vane Damage at Various 

Operatin9 Conditions 

Preliminary Design Conclusion~ 

1990 Technology Features Compared to Current Subsonic Engine 

1990 Technology - Subsonic Application Benefits 

Conclusions 

Vii t" 

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7 

10 

11 

14 

20 

21 

26 

39 

44 

58 

63 

63 

64 

105 

106 

107 

110 

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SECTION 1.0 

SU~tMARY 

Development of the supersonic cruise aircraft engine continued in this 

National Aeronautics and Space Administration (NASA) sponsored Pratt & Whitn~y 

contract for the preliminary design of an advanced high-pressure turbine. This 

program was specifically directed toward executing a preliminary design of the 

high-pressure turbine, with particular emphasis upon airfoil cooling concepts. Technology was assumed for 1986 and 1990 time frames to implement the,e 

concepts. The preliminary design demonstrated that the high-pressure turbine 

required for the 1990's supersonic cruise engine is feasible, assuming the 

availability of the advanced technologies that are required. In order for 

these technologies to be available in a timely manner, specific high 

temperature turbine programs should be initiated. In addition. a detailed 

turbine design should be initiated to confirm and identify more fully t!e 

high-pressure turbine requirements. The ~echnologies that need to be developed would also benefit both commercial subsonic and military engines. 

Previous supersonic cruise aircraft studies (ref's. 1 through 4) were reviewed 

as part of the current contract. and the Variable Stream Control Engine (VSCE) 

was chosen as a candidate for the preliminary turbine design. Other engines considered in previous studies (Low Bypass Engine. Inverted Flow Engine. and 

Variable Cycle Engine) differ in concept and performance. but their turbines 

have the same technology requirements as the VSCE. Hence. ti'e results of the 

current study apply in general to the other engine ~,urbines. The supersonic crufse mission defined in the ref. 2 was chosen for this study. but the 

resulting airfoil design was also evaluated in a comparable lIlilitary mission 

and a commercial subsonic application. The advanced technologies required in 

these airfoils were so fundamental that benefits were shown in all 

applications. 

The VSCE turbine defined the geometric size. temperatures. pressures. \peed. cooling air requirements. and component life for the preliminary design criteria. The resulting single-stage turbine has a nominal pressure ratio of 2.0~ Its design extends the ~irfoi1 stress levels to an AN2 value of 38.7 x 

101u cm2/min2 (6.0 x 1010 in2/min2) while setting aerodynamics at 

technology levels assumed to be available for the 1990's applications. 

The vane design has two separate internal chambers. and relies on efficient 

internal cooling and external film cooling. A leading edge insert provides an 

effective distribut~on of cooling air which exits via showerhead and film 

holes with sufficient cooling margin to maintain airfoil life over a large 

range of adverse conditions. The trailing edge cavity is tapered to provide uniform tr~i1ing edge flow into the radial pr~ssure field of the controlled 

vortex design. The design also allows for future utilization of 

three-dimensional aerodynamic concepts. The spanwise rib ties ~he suction and 

pressure surfaces while separating the trailing and leading edge cooling flows. Skewed trip strips are used on both the pressure and suction side inner 

walls to augment the heat transfer with the available (supply) pressure • 

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.... ~ _______________ ...IL.--~-- -."-.~--
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The blade has a six-cavity design with three separate cooling air feeds to the 

leading edge. midchord. and trailing edge regions. Tip cooling is provided 

from the leading edge cavity. This concept distributes the coolest air to the 

highest heat load areas. As in the vane. skewed trip strips are used. However. 

the trailing edge utilizes hiq~ blockage pedestal cooling with a cast trailing edge. The blade centour is designed to reduce the external heat load while 

mai ntai ni ng aerodynamic performanc;,~. 

Advanced single crystal materiuls. thermal barrier coatings. and oxidation 

resistant cnatlngs are used fot' both the vane and blade. This material 

technology is expected to be ,; .. ail able for a 1990 design start. New component techr~logy is also assumed to oe available and will result in improved combustion exit temperature distributions and in increased airfoil cooling capabi ,tty. 

These 1990 vane and blade designs meet the performance and durability goals to 

provide a cooled turbine efficiency of 92.3 p!rcent. 8.05 percent Wae cooling and a 10.000 hour life. An alternate design with 1986 technology will achieve 

91.9 percent efficiency and 12.43 percent Wae cooling at the same life. It is 

crucial that the advanced aerodynamic. coeling. material. and fabrication 

technologies identified in this program be developed in the next four or five 

years if t.hese designs are to become part of a 1990 engine. 

The supersonic cruise engine technologies when applied to a typical subsonic 

engine (266.892 N (60.000 lb) thrust. two-spool. two stage high-pressur~ 

turbine) demonstrate substantial potential savings in thrust specific fuel 

consumption (T5FC). The increased AN2 and blade material provide a 0.58 

percent benefit in T5FC. Airfoil aerodynamics. thermal barrier system. and 

burner technology advances equally contribute a total of approximately 1.0 

percent in T5FC benefit. 

Throughout the entire process. the manufacturing feasibility was continually assessed. Trailing edge thickness. wall thickness. airfoil taper. and 4"ternal 

complexity wpre based upon expected developments in fabrication technology. 

The performance and life results indicate that the original goals can ~e met 

provided the technology programs are pursued. 

2 

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SECTION 2.0 

INTRODUCTION 

Pratt & Whitney participated in a series of pr09r~ms under the NASA sponsored Supersonic Cruise Aircraft Research (SCAR) and Variable Cycle Engine (VCE) 

programs with the common objectives of identifying and evaluating key elements 

in propulsion system technology for the second generation advanced supersonic aircraft. These programs defined four overall engine confi9urations: the Low 

Bypass Engine (LBI'), the Inverted Flow Engine (IFE), the VClrfable Stream 

Contro' Eng~ne (V~:E) and the Variable Cycle Engine (VCE). Tht aircraft/engine 

syster.l requirements for supersonic transport applicl:.ions and engine performance requirements for the 1990's were defined in these studies. 

Critical components and the related general technology requirements were 

identified in these programs. 

In the turbines, the previous studies concluded the most critical advanced 

technology features to be those associated with achieving the required durability without using large quantities of cooling air which would impose a 

penalty on the engine thennodynamic cycle and turbine efficiency. The reasoils 

that supersonic cruise engines are very sensitive to the hot section 

durability technology are the very high projected turbine inlet temperatures and cooling air temperatures during long periods of supersonic operation, and 

the cycle characteristics whir~ cau:e high stresses in the turbine blades at 

these hi gh temperature level... Consequently, the need for very advanced 

materials, design techniques and cooling schemes was indicated. However, these 

previous studies did not address the problem of heat transfer analysis and 

design of the high-pressure turbine, which must withstand the stringent requirements of the supersoni~ cruise aircraft. 

The engine cross sections generatld in the previous studies defined the 

overal; turbine requirements of pressure ratio, airflow, speed and physical size. The four selected engine configurations defined a series of highly efficient but similar turbines which basically differed only in size. Similar 

advanced aerodynamic design goals and effective cooling air use were assumed 

for each turbine so that high levels of turbine efficiency could be achieved. 

A high-pr~ssure turbine pressure ratio of approximately 2.3, characteristic of 

current commercial transport engines, was chosen as a result of optimization 

studies. 

The preliminary design of the Supersonic Cruise Aircraft high-pressure turbine, conducted under the current contrar.t, further investigated the 

details of advanced cycle engine design requirements. The high-pressure turbine flowpath, airfoils and cooling design features were defined in detail assumin~ the availability and use of the emerging technologies identified 

under current research and developmer,\o work. The cooling configurations and 

materials selected r':present advanced state-of-the-art concepts. Pratt and 

Whitney's Standard DeSign System was applied to perform detailed aerodynamic design anG heat transfer analyses leading to the selected designs. 

3 

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paw's Standard Design System includes a series of computer codes which IMthematf~ally describe the physics of aerodynamic and heat transfer phenomena known to apply to turbines. Details currently too difficult to model 

analytically, or t.hose not yet fully understood, are simulatel1 by 

experimenta lly tietermi ned empi rical correlations. ~hese corr-elat ions are 

updated as ne~ ~alibration data are developed from rig and engine tests, and 

as the physical models are expanded to include new findings. 

Work under the current cont-act was performed under three technical tasks: 

4 

First, an advanced variable cycle engine and a supersonic flight misSion 

iPPTTcable to commercial aircraft were selected on the basis of the 

previous NASA sponsored program (Section 3). The Variable Stream Control 

Engine (VSCf) was chosen mainly because of its favorable thrust-to-weight ratio. Other differences between the candidate engines were found to be small. 

Second, tne aerodynamic and durability designs of the VSCE high-pressure turbine were defined using a series of trade studies (Sections 4 and 5). The aerodynamic design evolved from the original VSCE flowpath, after a 

blade loading optimization study using mean line analysis was completed. Airfoil cooiing schemes were selected with assistance from configurational 

s tudf es of advanced coo 11 ng concepts. 'fhe blade a i rfofl shape was a ho 

modified as a result of the configurational studies to reduce heat loads. 

Feasibility of the selected configurations was demonstrated by completing 

a preliminary design and detailed heat transfer analysis of the critical 

span cross section. 

Third, technology feat~res which are critical to the deve10pment of 

advanced cooling concepts in this program were identified. Benefits of 

this technology to advanced subsonic engines were also evaluated (Section 6) to show the overall generic value. 

-- ... 
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SECTION 3.0 

£NGINE AND MISSION SELECTION 

The init14l t.ask in the high-pressure turbfne del)fgn was the definftfon of the 

engfne and overall turbfne configuratior.. and the selection of an operating 

rrissfOIl for \~h~ch the high-pressure turbine afrfofls were desfgned. 

A St:r ie!) of previou!l supersonfc crufse engfne studfes idpntiffea several candf~ate engine conffguratfons. These studies identified operating missions 

for a supersonic cruise aircraft fn commercial use. Results of these studies. 

which also generated noise. cost. and gross weight trade factors to assess 

economical commercial usage, \~ere used as the basis for the engine selection 

for this high-pressure turbine prelfminary design. A supersonic operating 

mission was estimated for military applfcatiJns based on criteria consistent 

with the commercial studies. 

3.1 Engine Selection 

A review was made of the four candidate engine configurations I/hich were 

studied in the previous NASA sponsored Supersonic Cruise Aircraft Research 

programs. The candidate configurations are: the Variable Stream Control 

Engine. the Low Bypass Engine. the Inverted Flow Engine and the Variable Cycl~ 

Engir.e. These engines are briefly summarized in the following paragraphs. 

Variable Stream Control Engine (VSCE) Concept 

The VSCE is an advanced duct burning turbofan concept that makes extensive use 

of variable geometry components. A flexible throttle schedule allowed the ~ndependent variation of the two coannular exhaust stream~. This unique ~~heduling capability provided the inverse velocity profile neerled to take 

advantage of the inherent jet noise re~uction benefit at takeoff. while at 

subsonic and supersonic flight conditions the exhaust velocities can ~~ 

matched to provide a flat profile for high propulsive efficiency. The VSC~ 

turbine design incorporates a simple stage high-pressure turbine and a two 

stage low-pressure turbine. The design definition, performance and operatin~ 

features of this propulsion concept were discussed in detail in NASA CR-1597~0 

( re f. 2). 

Lm'i BypHS Engine (LBE) COricept 

The LBE ;s an advanced. nonaugmented. twin spool turbofan in which the design 

bypass ratios were in the 0.2 to 0.5 range. In the parametric version of this 

engine the bypass ratio was chosen to be 0.4. the primary and the bypass 

streams were partially mixed. and the exhaust gases discharged through a 

common variable area ej~ctof' nozzle. The refined LBE-4S0A incorporated a high effe~tiveness forced mixer to provide improved supersonic thrust specific fuel 

consumption through better mixing of the primary and fan streams. The 

high-p"essure spool of the engine consisted of a variable geometry compressor 

and an advanced si ng1 e stage turbi np wi ",h hi gh temperature ca.;:abil ity. 

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Inverted Flow Engine (IFE) Concept 

The IFE is an advanced, r.Jnaugmented, twin spool turbofan in which the design 

bypass ratios are in the 0.4 to 0.6 range. The engine configuration 1s similar 

to the Low Bypass Engine (LBE) except that the fan and core streams were 

inverted through a flow inverter downstream of the low-pressure turbine and 

discharged through independent nozzles rather than mixing the streams and 

exhausting through a c~on variable area ejector nozzle. Inversion of the two 

exhaust streams enables the IFE to achieve an inverted velocity profile which 

provides the potential to attain reduced engine noise levels, similar to that 

obtained with the VSCE. In the IFE the high velocity outer stream is formed by 

the turbine exhaust flow, and the inner stream in the nozzle by the fali 

discharge flow. Maintaining a constant jet velocity ratio and constant 

corrected total airflow while the engine is being throttled at takeoff 

requires that the IFE have a variable core stream (outer stream) nozzle area 

and a variable fan stream nozzle area. 

Variable Cycle Engine (VCE) Concept 

The VCE has a single spool, two stage turbine design and a turbine bypass 

system. This concept bypasses part of the compressor air around the 

burner/turbine and injects it into the nozzle system. The bypass air is 

modulated with power setting to produce inflight matching oenefits. The 

turbine bypass systems will be inoperative during the subsonic and supersonic cruise segments of the flight cycle. 

These engine studies were performed for each engine over a range of thrust 

size, weight, noise, emissions and fuel burned, and cost trade-off studies 

were conducted. These' 'udies chose the ';SCE concept as being the most 

promising. Table 3-1 shuws that the VSCE is the lightest engine for the sar.,e 

flow sileo The two stage turbine, single spool VCE, based upon a limited 

amount of study, proved to be the heaviest by approximately 50 percent. 

TABLE 3-1 

ENGINE WEIGHT SUMMARY 

340 kg/sec (750 lb/sec) Airflow Size 

VSCE 3915 kg (8630 lb) 

IFE 4491 kg (9900 lb) 

LBE 4273 kg (9420 lb) 

VCE 5481 kg (12280 lb) estimated 

A comparison of emissions (TaDle 3-11) to the Environmental Protection Agency (EPA) rules indicates that no engine concept in this group strictly meets the 

code. However, the VSCE engine has the most potential to be developed to meet 

these goals. 

The VSCE also had the lowest thrust specific fuel co~sumption at a 

representative subsonic cruise thrust (ref. 3). 

6 

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EPA CLASS T5 RULE 

VCE 

IFE 

LBE 

VSCE 

TABLE 3-11 

EMISSIONS COMPARISON 

EPA Parameter (EPAP) 

1000g emission/1000 16 fn-hr/cyc1e 

NOx CO 

5.0 7.8 

8.9 12.2 

6.9 11. 2 

6.9 11.2 5. 1 14.4 

A military engine (STJ-562) was considered for the present study but was 

rejected since it has a substantially lower AN2 level. In addition, its size 

is small and therefore a poor base to extrapolate benefits in the larger supersonic cruise engines. 

High-Pressure Turbine Comparison 

The hi gh-pressure turbi nes Qf the VSCE, LBE, IFE and VCE concepts wi th the 

sa'lle level of technology are compared in Table 3-III. The three turbines are 

single stage designs, drive similar compressors, and result in similar 

expansion ratios. The mean velocity ratio and axial velocity ratio o. these 

turbines are essentially identical due to their cOIl1~arab~e te~hnology. The 

stress levels in terms of AN2 were set. at 36.7 I. 10 cm /min (5.7 x 1010 in2/min2). 

The review of the candidate Supersooic Cruise Ai'rcraft Research (SCAR) engines concluded that the VSCE (Figure 3-1) was the most Attractive engine concept in 

terms of di rect operati ng cost and noise. These rl'su1 ts wer:~ ;.>resented to and approved by the NASA Project Manager. 

3.2 Mission Selection 

In establishing advanced supersonic engine hot section mission life and dura~ility design requirements for these studies, an assessment was made of 

the various advanced supersonic transport missions in terms of operating time 

spent at the most sev~re temperatui"es and maximum rotor speeds. Figure 3-2 

shows a typical flight profile for an advanced supersonic transport, including 

reserve operating requirements for cruise to alternate landing sites and hold. 

7 

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~ ~l __________ ..,_~...o..-. ______ ... __ _ 

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-'10 

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i5 100 125 150 In 

AXIAL LENGTH 

Fi gure 3-1 Variable Stream Control Engine Flowpath 

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'OCA - BEST CRUISE ALTITUDE 

Fi gure 3-2 Advanced Supersor,i~ Transport Flight Profile and Reserves 

r£ .. b___ ..-..-_"~. __ ... __________ '"""--~ __ 

9 
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. 

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Reference 1 summarizes the time for an entirely supersonic mission and for two 

mixed missions, which include subsonic legs added to satisfy overland cruise 

requirements. However, the subsonic elements reduced the compressor 

tempt:atures, the combustor temperature, and the high spool speeds. As a 

result, the severity of the mission is reduced in direct propol,tion to the 

length of the subsonic cruise element. It was decided, and approved by NASA, 

to include a subsonic cruise leg, but to retain the more solid aggressive mission. The mission with the 556 km (300 naut mile) subsonic leg and a block 

cycle time of 190.6 minutes was selected as being most representative of 

expected commercia~ service and used as primary mission for durability design. As indicated in Table 3-IV, more than 50 percent of the operating time is 

spent at the most severe temperature and stress conditions. Table 3-IV also 

compares the supersonic mission with a typical subsonic mission and indicates 

the increased severity of a supersonic cruise application. Based upon a 10,000 

hour life, a subsonic engine will accumulate 150 hours at maximum conditions 

in contrast to 5,000 hours for the supersonic cruise aircraft engine. 

TABLE 3-111 

HIGH-PRESSURE TURBINE COMPARISON 

VSCE LBE IFE VCE 

Number of Stages 1 1 1 2 Pressure Ratio 2.40 2.32 2.25 4.02 Ah, BTU/LBM 142.00 146.00 142.60 

Mean Ve10citR Ratio UMI~2g Jli G 

0.58 0.578 0.581 0.58 

CX/U Axial Velocity Ratio 0.60 0.60 0.60 0.60 Ex~ t ~1ach No. 0.49 0.53 0.49 AN x 10-9 IN2/MIN2 57.70 57.90 57.70 58.0 N1. RPM Rotor Speed 11 ,190 9451 9305 6144 Inlet Flow Parameter 72.70 103.40 108.70 150.2 W'If/PT 

Exit Swirl Angle 16.60 26.40 17.80 

10 

~ _ ~~-r ... ~.,.:.a' or 

-- --------
Flight Segment 

Tax i 

Takeoff 

Subs on icC limb 

Subs on i c Cru ise 

Supersonic Climb 

Supersonic Cruise 

Descent 

Tax i 

Total Time 

DeSigned for 

10,000 Hour Life 

TABLE J-IV 

DEFINITION OF SUPERSONIC AND 

TYPICAL SUBSONIC FLIGHT MISSIONS 

Altitude 

m (ft) 

o (0) 

335 (1,100) 

11,000 (36,0139) 

11,000 (3f,,089) 

11 , 1~:; ( 36,500) 

16, :54 (53,000) 

o (0) 

Supersonic Cruise 

Tirr~ 

Mn i!.:!!l 

,~.O 

~.3 0.7 

0.3-0.9 17.4 

0.9 18.5 

1.3 14.7 

2.3 99.7 

24.6 

5.0 

m:b 

Conlpre5Sor Ex i t 

Temperature 

K (OF) 

1483 (2210) 

1376 (2017) 

1253 (1796) 

15130 (2385) 

1755* (2700*) 

*5000 Hours at 

Max Cond it i on 

TypIcal 

Subsonic r~iHion 

Time 

~!!!l 

7.5 

1.5 

,9.2 

39.3 

15.0 

7.5 

~ 

Compres sor Ex it 

Temperature 

K (OF) 

1670+ (2546+) 

1550 (2330) 

1478 (2200) 

+250 Hours at 

l'1ax Cond it ion 

Critical mission operating points were selected for consideration in design of 

the VSCE high-pressure turbine. These operating points and the most 

significant operating environment for each point were: 

1) t~keoff. sea level static - high specific thrust. low noise; 

2) takeoff. cutback power. 335 m (1.100 ft). 0.3 Mn - low noise; 

3) subsonic cruise. 1.100 m (36.089 ft). 0.9 Mn - low thrust. low fuel 

consumption. inlet flow-matching; 

4) subsonic climb. 11.125 m (36.500 ft). 1.3 Mn - high thrust; and 

5) su~ersonic cruise. 16.154 m (53.000 ft). 2.32 Mn - high thrust. low 

fuel consumption. 

A supersonic flight mission for military t~ansport applications also was 

required for evaluating benefits under this program. On the basis of expected flight routes and supersonic operation restrictions. the military mission was 

selected to be ide'ltica1 to the ("onmerci~1 mission. Mi~ itary flights most 

likely to be made by the supersrnic dlrcraft will be over very long distances 

on intercontinental routes. sim',lar' to the commercial flights. Additionally. for all peace time operation an(i controlled military conflicts. the noisE' 

regulations dictating takeoff and climb conditions and operating routes of 

commercial aircraft were also applied to military operation. 

- ,

..-------------.'"----- ... .. 

11 
, 

SECTION 4.0 

BASIC TURBINE PARAMETER AND AERODYNAMIC DESIGN DEFINITION 

The deSign of high performance, high-pressure turbines is a complex series of 

compromises that result from the need to meet both aerodynamic performance and 

durability goals. Typically, prior to the start of the high-pressure turbine 

deSign, turbine goals are set for engine and mission studies to allvw early overall evaluation of engine weight, fuel burn and operating cost. Having 

demonstrated overall engine benefits, preliminary deSign studies of the 

turbine are conducted to evaluate the feasibility of mepting assumed goals. Gen~rally, several design iteratio~s are necessary in order to achieve 

agreement between design goals and the actual deSign. A natural outcome of the 

preliminary design iterations is the refining of the goals, the technology required to meet these goals, and the associated trades of performance, weight and cost. 

The major elements of the preliminary turbfne design process used in the 

preliminary design of a Supersonic Cruise High-Pressure Turbine Program 'is 

shown in Figure 4-1. Input to this process consisted of certain key elements: 

initial component aerodynamic and mechanical parameters, technology and 

material prOjections for the 1990s, engine deSign tables and the mission 

definition. Current design systems were used to: verify the performance and 

life, define the number of airfoils and airfoil contours, flowpath and 

mechanical speeds, and configure the cooling design of the blade and vane. 

Results of the previous Supersonic Cruise Aircraft Research (SCAR) engine studies provided the starting point for the high-pressure turbine preliminary 

design. The chosen mission cycle and engine cycle established the turbine 

parameters (iable 4-1) and the initial flowpath. Studies were performed with 

meanline and streamline analysis to refine the flowpath and to obtain velocity triangles, reaction level, annulus area and wheel speed. Airfoil cross 

sections for the vane and blade were deSigned to provide minimum aerodynamic 

losses and to establish the heat loads on the vane and blade. Internal design of the airfoil was performed and cooling air distributed to obtain the design goals fc:" metal temperatures and life. T~e cooled aerodynamic loss was then 

determined for both the vane B.nd blade and the overall turbine efficiency established. 

4.1 Aerodynamic Design 

Aerodynamic design goals set for the VSCE high-pressure turbine design were 

maintained for the supersonic transport engine. The goals for the VSCE design, relative to the Energy Efficient Engine design, are: airfoil 10dding, +17 

percent; profile loss, -10 percent; and endwa11 loss, -15 percent. 

12 

The aerodynamic design was started with the selection of the mission profile, the performance requireme~ts, the VSCE turbine, and the above aerodynamic goals. AN2 equal to 38.7 x 1010 cm2/min2 (6.0 X 1010 in2/min2), 

used for both the 1990 and 1986 engines, is an increase from the 36.7 x 1010 

cm2/min2 (5.7 X 1010 in2/min2) used In previous VSCE studies. This 

increase was decided on before the start of this program. 

t 

! ~t _______________ --"----- _ .. ,_.-. - '"lIIj; .... -... --
. 

~. r 

Inputs 

Component 

Aerodyn6mi cs / 

Updated 

TechnologJ 

Project ions 

Pre 1 iminary 

De si<Jn 

Analysis 

Mated a 1 s / Component 

Updated Mechanical 

Technology Parameters/ 

Updated Projections Technology 

Projections 

-

I High Spool I 

I comT~ 

In i 

Engine 

Tab 

I 

tial 

Design 

les 

~----------------~-------~--------------------~ 

Outputs L 

Support ] Tec!,no10gy 

Requ ired 

Fi gure 4-1 

Materi a 1 

Requirements 

Ai rfo i 1 

eros s 

Sections 

HPT 

Per fa rmance I 

I 

U~da1:e I L..-____ ~ 

Primary Elements of the Preliminary Design Process 

13 

! .. L _______ .. _____________ ...II.-. _____ ... _.-__ 
• 

~ . 

TABLE 4-1 

TURBINE ~ARAMETERS FOR 1986 AND 1990 DESIGNS 

o Single Stage 

o Maximum Turbine Inlet Temperature - 1755 K to 1811 K (2700·F to 2800·F) 

o Compressor Discharge Temperature - 922 ~ (1200·F) 

o High-Pressure Turbine Design Life - 10,000 Hours, 5,000 Hours at 

~laximum Speed, ~laximum Turbine 

Inlet Temperature, Maximum Com- p;~essor Di scharge Temperature 

o Velocity Ratio - 0.58 

o Axial Velocity Ratio (Cx/U) - 0.60 

o AN2 - 38.7 x 1010 cm2/min2 (6.0 X 1010 in2/min2) 

The initial design studies centered on the original VSCt study engine flowpath 

at conditions modified for the current design program. The predicted pressure distribution on the blade surface is shown in Figure 4-2. The initial study identified an undesirable blade loading distribution which created a 1.35 

maximum Mach number on the blade suction surface. The high Mach number lowered 

turbine efficiency, which in turn increased operating temperatures and 

prevented meeting durability goals. 

SUbsequebnt1 y

i

, a Plarametrihc stubdY was conductehd

l 

tO

d 

re

h

fi

A

n

N

e

2

the 6tuorbxine10design and to 0 ta n a ower Mac num er design but 0 t ~ to. 10. 

Mean1ine turbine designs were performed over a speed range of +10 percent and 

velocity ratios from 0.4 to 0.8. The mean radius and annulus area of these 

turbines were varied. This study, graphically summarized in Figure 4-3, 

provided a 0.5 percent net incre~Ge in turbine efficiency using initial 

cooling loss estimates. The chosen design reduced blade maximum Mach number 

from 1.35 to 1.1, trailing edge Mach number from 0.93 to 0.86, the axial 

velocity ratio (Cx/U) from 0.60 to 0.43, and the rotor speed from 11,194 RPM 

to 10,075 RPM. The reduced speed was acceptable to the VSCE high-pressure spool. The original flowpath and the selected low Mach number flowpath are 

compared in Figure 4-4. Turbine parameters for the selected engine configuration are presented in Table 4-11. The turbine airfoil geometry is 

shown in Figures 4-5 and 4-6 • 

14 

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Figure 4-2 

0 0.20 040 0.60 0.80 100 

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Original VSCE Turbine Flowpath High Mach Numbers Required Turbine Redesign 

- ----------------------""'----- 

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LOW MACH NUMBER DESIGN 

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11 

High-Pr€ssure Turbine Parametric Study Identified Improved Design 

.- ... 

13 
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52 BOO 

12075191 

RADIUS, 

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VANE 

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11356561 

LOW MACH NUMBER DESIGN 

Figure 4-4 Comparison of Original and Final Flowpath 

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AYIAL. eM (IN) 

Figure 4 -5 Turb In!"? GeOIl1C' try 

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Figure 4-6 Airfoil Geometry 

BLADE (MEAN) 

19 

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20 

TABLE 4- I I 

TURBIN[ PARAMETERS 

Radius B Min r'ollt Airfoil Section cm (in) x ~in ·)out cm(in) (<reg) (deg) 

Vane R 39.85 (15.69) 4.36 (1.72) 0.152 0.965 90.0 15.3 

M 46.28 (18.22) 5.33 (2.10) 0.147 0.832 90.0 15.3 

T 52.71 (20.75) 6.27 (2.47) 0.146 0.737 90.0 15.3 

Blade R 39.73 (15.64 ) 5.81 (2.29) 0.459 0.811 33.5 22.7 

11 46.28 (18.22 ) 4.62 (1.82 ) 0.283 0.862 53.8 22.7 

T 52.83 (20.80) 3.45 ( 1 .36 ) 0.195 1. 01 93.5 22.7 

Vane design philosophy has be~n to select a minimum number of vanes to reduce 

aerodynamic loss, cooling air requirements and parts cost. However, 

compensating concerns of controlling p'Jatform alignment, and hence leakage, 

tends to drive the selection toward reduced platform length and increased nu~ber of vanes or at least clusters of vanes. The preliminary aerodynamic and 

structural considerations set the number of vanes at 30. 

The trend in advanced turbine rotor designs is to use fewer blades whic~ 

results in lower parts and maintenance costs, and increased flexibility of the 

rotor system structural design. Since the number of airfoils along with the 

chord length also affect aerodynamic performance and blade pull, a separate 

evaluation was conducted to determinl the number of blades that ~ould be uJed 

in the VSCE design, consistent with tt.e assumed design goals. 

P.erodynami closs, a sum of the bl ade profil e and secondary (end\'/all) losses, 

was evaluated for the range of 50 to 60 blades. The aerodynamic loss and the 

total loss, including estimated cooling effects, are shown to be relatively insensitive to the number of airfoils. This results from the compensating 

trends of the profile and endwall losses as illustrated in Table 4-III. 

... ~ _____ • ____ ~ __ """-___ ....... n ___ ~: ... _ 

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TABLE 4-II I 

COMPARISON OF LOSSES WITH 50 AND 60 BLADES 

Number of Airfoils 

ProffleAP/P 

Endwall 

Coo11ng 

Total 

50 

.0114 

.0120 

.0118 

.0352 

60 

.0122 

.0104 

.0118 

.0344 

This loss difference between 50 and 60 blades is considered to be 

insignificant (2 percent of the loss). Consequently, the number of blades was 

selected to be 50 in order to take advantage of the potential cost benefits of 

the smaller number. 

The meanline study defined the inlet and exit aerodynamics. The Pratt & 

Whitney airfoil design system was used to define the airfoil contour. This 

system defines the inlet and exit aerodynamics, 1ead';ng edge and trailing edge 

contours, wedge angle, and overall loading and modifies the chordwise loading distribution to obtain the minimum aerodynamic loss. In configuration studies 

(Section 5.1.5) this process was extended to include minimizing the external 

heat load. 

Design results for the vane and blade pressure distributions are presentfd in 

Figures 4-7 and 4-8. The blade pressure distribution reflects the effects of a 

redllced exit ~Iach number. Intra-blade ~lach number for mean section is shown in 

Figure 4-9. 

Boundary layer analyses were performed to quantify aerodynamic losses. The 

airfoil contours and their res~lting aerodynamics provided the input for the 

STAN-5 modeling of the viscid loss calculations. The Pratt & Whitney design 

system also models the shock losses. Figure 4-10 shows the radial distribution 

of suction sirte boundary layer transition from laminar to turbulent and the 

predicted losses. The aerodynamic losses are ihown to be between 1.0 to 1.5 

percent, including the shock loss. Based on these results, a preliminary efficiency of 92.3 percent for the high-pressure turbine was obtained as shown 

on Table 4-IV • 

21 

t! I.t _________________ '""'----. _ ...... -

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,t 07 Ci: 

06 

0,5 

04~----____ ~ __________ ~ __________ ~ __________ ~ ________ --J 

a 0.2 0.4 0,6 0.8 1.0 

X/BX 

Figure 4-7 Vane (Mean) Pressure Distribution 

..•

fie
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t ' 

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0.4 L..-___ ........ ____ --L. _ ____ '--____ L.-___ ~ 

a 0,8 10 

Fi gure 4-8 

0,2 0.4 0,6 

X/BX 

Blade (~lean) Pressure Distribution Redesigned Blade Achieved 

Reduced Mach Numbers 

23 

L 

-----... -------------.:&..---- .... --

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X 

aJ 

>= 

15 

1.3 

11 

09 

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Figure 4-9 

.• i. 

-06 -0.4 -02 o 02 0.4 

X/BX 

81 ade ~lach Numbers Contour Map 

./ 

? 

CURVE CURVE 

LABEL MACH NUMBER 

1 01 

2 02 

3 03 , 04 

5 05 

6 0.6 

7 07 

8 08 

9 09 

11 1.0 

0.6 0.8 1.0 12 14 16 
I 

INOID3t11. N31nat1n1. : 

0 

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1/1 

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1/1 

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-- ---------
o Current Technology 

TABLE 4-IV 

EFFICIENCY BREAKDOWN 

Including AN2, Velocity Ratio, Pressure Ratio Effect 

o Advanced Technology 

Profile and Secondary Loss Reductions 

Increased Annulus 

Increased Loading 

Reduced Cooling Loss 

4.2 Aerodynamic Conclusions 

91.0% 

+0.6% 

+r 5 

-0.2 

+0.4 

92.3% 

The high-pressure turbine efficiency estimate based upon the preliminary aerodynamic design is summarized in Table 4-IV. This study identified an 

increased efficiency of 0.7 percent above that predicted in the original SCAR 

programs. This 92.3 percent efficiency in the high-pressure turbine was 

achieved with 10 percent fewer airfoils. It is further concluded that the 

benefits obtained in this design study are transferable to other advanced 

aircraft engine designs. 

4.3 Mechanical Design Considerations 

Blade pulls and stresses were evaluated for the 50 blade design. Included were 

studies to demonstrate the benefit of axial chord tapering between the airfoil 

root and tip on blade centrifugal stress. The chord taper ratio (tip chord/root chord) was determined by setting the airfoil tip and root load 

coefficients at the design goal level (17 percent higher than that in the 

Energy Efficient Engine). Sections between the tip and root were then assumed 

to be linear with the radius. The airfoil defined in this manner resulted in 

blade root centrifugal pull and stress increases of 7 and 5 percent 

respectively, relative to the Energy Efficient Engine blade. Further reduction 

of the pull and stresses is possible by more complex tapering of the chord and 

wall thickness redistribution in the radial direction, which should be 

considered during a final design program. 

26 

~i 

l 1..1:. ___________________ .... _ .... __ _ 

.~ 

\;;
SECTION 5.0 

DURABILITY DESIGN 

The operating requirements of ~;upersonic cruise aircraft engines are unusual 

in that the maximum turbine tem~eratures and rotor spe~ds occur during the 

long time cruise portien of the mi~sion. At the selected mission and a 10,000 

hour total design life, the hot section parts experience the near maxiMum 

combustor exit temperature, rotor speed and cooling air temperature for over 

5000 hours. In contrast, the hot section parts ot a typical subsonic engine would require to withstand only about 150 hours of exposure at these 

conditions over the same 10,000 hours total life. This difference shifts the 

major design criteria of the high-pressure turbine life in supersonic applications from takeoff and transient cyclic operation, to creep or 

oxidation related modes. 

Advanced cooling techniques coupled with material improvements are necessary 

to achieve durability consistent with the engine application gOJls and the 

target efficiency. To accomplish this, a three step design approach was used: 

first, configurational studies for screening durability technologies considered available for t~e supersonic cruise aircraft engine; second, 

mission life analysis to establish the design metal temperatur~ limits; and 

third, preliminary cooling designs of the vane and bl~de were completed, utilizing the configurational study results and coolant flows dete~ined in 

earlier studies. Detailed heat transfer analysis was conducted to assure that 

the combined cooling features satisfy overall durability requirements. 

5.1 Configurational Studies 

Advanced turbine airfoil cooling designs generally comprise a number of 

individual local cooling schemes tailored to meet cooling and geometry requirements of each part of the airfoil (leading edge, midchord, and trailing edge). Such individual cooling schemes for the supersonic cruise aircraft 

engine were selected using configurational studies described in this section. 

The criteria for inclusion in the subsequent preliminary design was the 

potential for improving engine performance either by reducing cooling airflows 

or improving cooling air management. 

The following schemes were evaluated: 

o 

o 

Vane coolant precooling to reduce vane coolant flow or reuse the vane 

coolant in the blade; 

Vane trailing edge cooling configurations and its compatibility with 

aerodynamic design; 

o Blade leading edge configurations; 

o Thermal barrier coating with advanced film cooling; and 

o Airfoil design for reduced heat load. 

27 

.. ~_. _________________ ...Ia-____ M ... _.-••• _ .~:; """"<." ._. _______ . __ _ 
, 

,; 

~" 

t· 

r 

a, 

t: 

5.1.1 Vane Coolant Precooling 

The use of a precooler was investigated as a n~ans of reducing the cooling airflow and associated momentum mixing losses for the first stage vane. Two 

distinct cooling schemes were investigated. The first scheme utilized 

precooled air to cool the vane leading edge cavity. The precooled air passes 

between the vane walls and the insert. exiting as film on the suction surface 

(Figure 5-1). A suction surface exit is required because of the increased 

pressure drop requirements of internal convection and the precooler relative 

to a film cooled design. Using a pressure surface (low Mach number) exit would 

not allow sufficient pressure drop to drive the precooler and additionally provide the required irternal con'Jection. As can be seen from Figure 4-7. the 

pressure surface has surface pressures very nearly equal to the leading edge 

total pressure in the leading edge cavity region. Utilizing this region a~ an 

exit would provide an available pressure ratio for cooling of only 1.09 

((PT4/PT3) / (~S/PT3)) not accounting for b~ckflow margin. This 

pressure ratio is nsufficient to drive the precooler as well as to provide sufficient internal cooling. 

Flow turbu1ators were place~ on both the insert and vane walls to achleve 

maximum internal heat transfer augmentation. Internal heat transfer wa~ 

tailored to maintain a constant interface temperature by varying the insert to 

vane gap. This gap decreased chordwise from the leading edge towards the rear 

of the leading edge cavity as the coolant temperature increased. Heat transfer 

coefficients were limited by the available pressure drop and friction 

generated by the cooling passages. Maximum available heat transfer 

coefficients at the critical rear cavity location (maximum Tco~l~nt) are 

shown as a function of cooling flow in Figures 5-3 and 5-4 for the suction 

wall and pressure wall respectively. Wall thicknesses on the vane were thinned 

to 0.127 cm (0.050 in) to maximize the cooling effectiveness. 

The pressure and suction walls were analyzed separately using an idealized 

one-dimensional heat transfer model to determine the minimum flow required to 

cool the vane walls. 

The idealized model (Figure 5-2) assumed a flat surface at a uniform interface 

temperature. With these assumptions the required internal heat transfer 

coefficient could be determined and plotted as a function of coolant flow and 

the extent of precooling (Figures 5-3 and 5-4). Figures 5-3 and 5-4 also show 

the attainable heat transfer coefficients using state-of-the-art internal 

turbul ators. 

The minimum cooling airflow required to convectively cool the vane leading edge can be less than that required for the film cooled design as shown in 

Figure 5-5. This convective cooling scheme. although reducing the leading edge cooling airflow. did not reduce the overall momentum mixing losses. Momentum 

mixing losses are a function of the ~ooling flow level as well as the ratio of 

coolant exit velocity to local gas velocity. Increasing the coolant velocity 

to reduce momentum mixing losses reqires an increasing of the upstream driving static pressure. As can be seen from Figure 5-6. fur the leading edge convective cooling this would require increasing the internal static to 

28 

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~~

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, 

f 

T AllO"ED INTEPNAl 

COOUNG - MINIMUM 

COOUNG FLOW 

o 127CM 

10050lNI 

• COOLING TAILORED TO MAINTAIN CONSTANT TINTERFACE TO ATTAIN 

MINIMUM FLOW REQUIREMENT 

• WALL MINIMUM THICKNESS A 1 0.' 27 CM (0.050 IN) TO ATTAIN 

MAXIMUM INTERFACi: COOLING 

Figure 5-1 Convective1y Cooled Vane Leading Edge 

29 

.~ 

~
r r' 

~' , 

r 

m__._ 

30 

"IDEALIZED" MINIMUM FLOW LEADING EDGE 

9, 

o ,. h. Ag [ T g - TI J 

6 .. m cp [ T c - T c J 2 , 

0L • h. A [T - T J L g g I 

SOLVING FOR h2 

h :& 

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Figure 5-2 Vane Convective Leading Edge Cooling 

.... ---------------""".---""'---- ... ... 

h2 - hREQUIRHl 

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S' A Tf Uf T HE ART 

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COOLANT 

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~~----~--~--~~-L~~~--------~ ____ ~ ____ L-~ __ ~~~_J 02 OJ 06 10 20 3,0 !i.0 100 

Figure 5-3 Convectively Cooled Vane Leading Edge Using Precooled Air; (Suction Surface) 

31 

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r... - ~;; ~""'."'~' .... 

...... -----------"'"'----. ...... - -- -- 

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() 2 

Fi gure 5-4 

05 2 0 ~ 0 I() 0 

~onvectively Cooled Vane Leading Edge Using Precooled Air; (Pressure Surface) 

... 

.t.

\:
, 

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QJ 

1) 

;:. 

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We A FILM COOLfO LfAOING EOGf INO PAfCOOLINGI 

" 

l)~ ________ ~ __________ ~ __________ ~ __________ ~I __________ ~ 

o 56 111 167 -222 

101 1001 I .. 2001 I 3001 I - 4001 

DELTA COOLANT PRECOOLING. K (OF) 

. 278 

I 500) 

Figure 5-5 Vane Leading Edge Convective Cooling; Effect of Precooling on 

Coolant Flow 

r£> 

~ ~~~"'~...r-; •. .. ----------------------""'---

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20 

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Figure 5-6 

, 

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[ . 

r 

34 

2"<> I/"de I 196 u e I 320 u FI PRfCOOLINGI 

1 75% - 251"e I - 420°F) 

~ ... ------ 6PIP LEADING EDGE 

FILM COOLED 

1,6 1.8 20 

P IP 

SINNER SEX IT 

Vane Convective Leading Edge; Momentum Mixing vs Exit Pressure 

Ratio 

r£> 

~ --"'"." -'-.' . i ... _______ 

c __________ "". _____ , _____ •• u •••• ______ '" ___________ . __ _ 
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r 

external static pressure ratio to 1.7 in order to reduce the loss levels to 

the base film cooled levels of the leading edge. This increase in internal 

static pressure would reduce the assumed internal convective effectiveness due 

to the reduced pressure drop available for cooling and require more cooling flow. To maintain momentum mixing losses equal to the base film cooled case, 

supply pressures in excess of compressor discharge ~ressure are required. The 

film cooled base case flow levels, shown for the leading edge in Figures 5-3, 

5-4, and 5-5, reflect initial flow level estimates. Total leading edg~ flow 

levels were reduced from this initial estimate of approximately 3.8 percent Wae to 3.35 percent Wae when detailed analyses were conducted for the film 

c(io1ed design. 

The vane trailing edge could not utilize the precooler to reduce coolant flow 

levels and momentum mixing losses. Utilization of a precoo1er requires elimination of the pressure side film cooling due to the increased pressure 

drop requirements of the precooler and internal convection. Removal of the 

pressure side film significantly increases the film temperature at the 

trailing edge of the vane. This increase in film temperature raises the 

cooling flow requirement of the trailing edge slot beyond that of the total 

flow required for the film cooling case even with precooling of 422 K to 477 K 

(300·r to 400·F). Cooling lostes would increase dramatically since the 

trailing edge is a much higher loss region relative to the pressure wall. 

To elim~nate the momentum mixi:fg loss of the leading edge cooling air, a 

second sc..;'eme was reviewed in which the va.,e front cavity was cooled by passing precooled air from the outer diameter to the inner diameter of the 

vane and using the same air to cool the fir.;t stage blade. An opti.1um or 

lIidea111 cooling scheme was analyzed to detemine the maximum potential for 

reducing the cooling airflow and increase engine performance with this design. Vane cooli~g configurations were tailored so that the vane cooling airflow and 

exit temperature matched the blade cooling requirements (Figure 5-7). 

Precooler pressure, temperature and flow requirements were established over a 

range of tangential on-board injection (TOSI) pressure ratios (Figure 5-8), 

and based on the flow characteristics and requirements of the vane and blade. 

A range of TOSI pressure ratios was presented due tJ the significant impact of 

the TOSI on turbine pump work and blade cooling air temperature. The lIideal ll 

system considered does offer an improvement in I~ngine thrust specific fuel 

consumption (Figure 5-9). 

A heat exchar.ger and its associated plumbing, however, would weigh approximately 54 kg (120 1b) to achieve 477 K (400·F) temperature drop with 

the allowable pressure drop. This added weight would result in an aircraft 

thrust specific fuel consumpt';un (TSFC) penalty of 0.23 percent (Table 5-1) bringing the net TSFC benefit down to 0.13 percent at most. This small gain is 

based on an "ideal ll system and the actual benefit, if any, would be 

significantly less. Because of the extremely small potential of this system coupled with the inherent risk of damage to the blade from ingestion of loose 

parts, it was not considered as a viable technology option for the supersonic cruise application. 

e 

.. t ___________________ .A-~. ___ .......... _ .~.~ ..• -' "- - ---------------

35 
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111 

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36 

• THRU FLOW TO BLADE 

• 100% TOBI EFFICIENCY 

• rutl PR • 1,0 

ATCOOLER 233K 

756 

(9001 

289K I· 420 0 FI 

811 

110001 

867 

(11001 

922 

(12001 

T CA AT BLADE SUPPLY, K (OF) 

BLADE COOLING REQUIREMfN'r 

978 

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1033 

(14001 

Vane Leading Edge Convective Cooling; Coolant Flow vs Temperature 

l

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PR ~ TANGENTIAL ON BOARI) INJECTOR nOBl1 

PRESSURE RA 110 

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70~--______ ~ ________ ~. ________ ~ __________ ~ ______ ~ 

4.0 

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20 

10 ~--------~--------~----____ -L ___ o 56 

101 1 1001 

- 111 

1 2001 

- 167 

1 - 3001 

,n COOLANT PRECOOLING, K (OF) 

- 222 

1- 4001 

- 278 

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PA z 10 

Figure 5-8 Vane Leading Edge Convective Cooling; Coolant Flow vs Precooling 

37 

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fi

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38 

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PR 1 ANGENTIAL ON BOARD INJECTOR (TOBI) 

PRESSURE RA TIO 

- 56 

( 100) 

- 111 

(- 200) 

PR • 1.2 

- 167 

(- 3001 

- 222 

(- 400) 

- 278 

( 500) 

Influence on Thrust Specific FUi~l Consumption Change of Vane 

Precool ed Ai rf1 ow from Leadi ng I~dge to B1 ade Cool Lig With No 

Weight Penalty 

E 

Io..~ _____________________ "",-- ____ ... _. __ _ 

~ _art 

-~

'=
~' 

TABLE 5-1 

PRECOOLER CONFIGURATIO~ 

o Weight - 54 kg (120 lb)minimum 

o 

o 

Thrust Specific Fuel Consumption Benefit for Precooling 

Thrust Specific Fuel Consumption Debit Due to Weight 

Net Gain 

+ 

Blade Ingestion Risk 

o Conclusion: 

Benefits Do Not Warrant Additional Risk 

5.1.2 Vane Trailing Edge Internal Cooling Geometry 

0.36% 

0.23% 

0.13% 

In this study, advanced state-of-the-art cooling schemes fo~ the vane trailing edge were evaluated in terms of cooling effectiveness, required cooling air 

supply pressure, and compatibility with aerodynamic requirements. Four 

candidate concepts (Figure 5-10) with either pedestals or trip strips were 

considered to augment convective heat transfer. These included: 

o a baseline circular pedestal array typical of current trailing edge designs; 

o a ribbed design in which the ribs form rectangular channels for 

coolant flow passages; these ribs also incorporate trip strips on 

their surface; 

o a pear shaped pedestal array, and 

o an alternate ribbed design with skewed trip strips on the airfoil 

pressure and suction side inner walls. 

Another fifth, scheme, emp 1 oyi ng a "cut back II pressure wa 11, was also 

considered and discarded early because its cooling effectiveness is not 

currently considered to be competitive with the other candidate schemes. The 

cut back design becomes more effective in applications having aerodynamic performance more sensitive to airfoil trailing edge thickness than the 

supersonic cruise aircraft engines high-pressure turbine vane (also see 

Section 5.2). 

tr!> l - ...".....,- --';" 

' 

39 

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r 

If, 

"" ~ 

l~t '" 

1""" 

.~ 

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o 

TYPICAL PEDESTAL ARRAY 

025 

10101 

025 

(010) 

.. 

0127 

(00501 DIA 

O'oo'ol 

o __ ?~ g;?~ ~ ~?J -- '- '_. TIE 

TRIPS ON RIBS 

015 0.30 

(006) _1~2~-r---1 Ltc .~-+------- 025 

to 0101 

-j~~~ 

0,10 In- 0 25 -' =~-:=-=-~-:--::--....... =-====:: 

(004) (010) -'- --- DIMENSIONS IN eM liN) TIE 

ELONGATED PEDESTALS 

025 

'0101 

025 

(010) I. --I 

(j 127 

,0050, DIA 

AL TERNA IE RiBBED OESIGN 

TRIPS ON WAllS 

,~ ./'~ 

~-~-

Figure 5-10 Vane Trailing Edge Concepts 

." ________________ ............ , - ... -- .... ,....,.r."""_ 

... 

0064 

'0025' OIA 

T~E 

TIE 

--
Analysis was performed on the trailing edge confi9urations to determine the 

coolant flows at available supply pressure. define the corresponding heat 

transfer coefficients. and calculate the local metal temperatures. For each 

configuration. the cooling surface and the conduction areas were modeled to 

reflect the presence of pedestals. Thermal barrier coating was assumed on the 

airfoil surface for metal temperature calculations. 

The study results are presented in Figure 5-11. The results showed that the 

best trailing edge internal cooling scheme is the alternate ribbed with trips 

on walls design (skewed). The design with trips on the ribs also achieved high heat transfer coefficients on the ribs but could not generate the necessary side wall heat transfer with the available coolant supply pressure. The 

reduced flow and smaller exposed coolant surface areas of the pedestal designs 

resulted in lower cooling effectiveness. Both pedestal designs provided similar heat transfer coefficients but could not adequately cool the trailing edge with the available coolant supply pressure. 

5.1.3 Blade Leading Edge Study 

Two competitive leading edge cooling concepts (Figure 5-12) suitable for 

advanced engines were evaluated for the high-pressure turbine blade design. Both were judged to require internal convective cooling supplemented by 

showerhead film cooling. The evaluation was made on the basis of cooling effectiveness. showerhead back flow margin. and the ability to tolerate local 

film hole plugging without catastrophic failure. Manufacturability was also 

considered. Both designs used showerhead film cooling, but one design relied 

on trip strip augmented channel flow for internal convective cooling while the 

other used impingement cooling for heat transfer augmentation. 

The trip strip augmented convect·, "ely cooled design directs cooling air 

through a radial supply passage from which d portion of the flow passes 

through showerhead holes and film cools the leading edge. The remainder of the 

cooling airflow continues radially outward and is subsequently used to cool th~ blade tip. The impingement cooling design has rne additional rib which 

di\ides the radial supply passage and the leading edge passage. Coolant flow 

is ejected through holes cast in this rib onto the leading edg~ surface to 

provide impingement cooling and is then ejected through the showerhead holes 

and additional "gill" holes to provide film cooling. For the impingement cooled leading ~dge design, tip cooling air must be supplied from separate 

midchord cooling passages. 

Both designs met the operational requirements of the supersonic cruise 

aircraft engine high-r.ressure turbine. Coolant flow and heat transfer study results for both the 'full-flowing" and the "plugged" showerhead holes are 

summarized in Table 5-11. These results show that the trip strip augmented channel design supplies about 0.2 percent more coo1irg airflow to the leading edge feed. However. the impingement design must supply an additional 0.4 

percent flow from the midchord supply for tip cooling. 

41 
1322 

119201 

1300 

11880 

127a 

118401 

u. 1256 

::... 118001 

~ 

u.j 

a:: 

:::l 1233 

~ 117601 ~ 

a:: 

w 

Cl. 

~ 1211 W 

~ 117201 

w 

U 

~ 

u. 1189 a:: 

w 116801 ~ 

Z 

1167 

116401 

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t 1122 

115601 I 

0 10 2.0 

101 10 41 108) 

LE 

PEDES TAL DESIGN 

DESIGN GOAL 

30 4 1 5 1 6 1 7 1 11 2) 11 61 12 0) 12 41 12 81 

PRESSURE SURFACE DISTANCE, CM (IN) 

TRIP STRIP DESIGN 

WITH TRIPS 

ON WALLS 

8 1 9 1 

13 21 r;J 61 

Figure 5-11 

Vane Interface Temperature Distributions on Pressure Surface (1990 Technology) 

~ 

b ~t ____________ ....-. __ ~ ... u ••••• ~.~ --=< •••.. ~_,. _______ . _ ~_~_ ~ 

42 

1 t) 2 

14 (): 

It 

-fi

'=
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> 

, -

, 

, 

IMPINGEMENT COOLED LEADING EDGE TRIP STRIP AUGMENTED, 

CO,\IVECTIVEL Y COOLED LEADING EDGE 

Figure 5-12 Blade Cooling Configurations 

~ . 

e 

t:_" ________ ._~...i..-_______ ..... -__ 

43 

____ i

V
I 

r 

r 

,. 

, 

Both designs (Table 5-11) still provide more than two-thirds of the 

full-flowing convective heat transfer coefficient if the showerhead holes are 

completely plugged. Although local metal temperatures in both designs increase 

by more than 422 K (300·F) when complete showerhead plugging occurs, rapid failure$ will not follow IS indicated by experience with current showerhead 

leading edge designs. 

The trip strip augmented channel cooled leading edge design was chosen for the 

supersonic engine high-pressure turbine because of: 1) the slightly lower 

total coolant flow, 2) successful experience of this design in current engine applications, and 3) projected lower manufacturability risks. 

TABLE 5-II 

RESULTS OF AUGMENTED CHANNEL AND IMPINbEMENr COOLED SHOWERHEAD 

LEADING EDGE COMPP,RISON 

Channel Impingement 

Full Flowing Showerhead 

Coolant SupDly 80" 80" (~ of Compressor Discharge Pressure) 

L. E. Feed Flow (" Wae) 1.41't 1.23" 

L. E. Feed and Tip Flow ("Wae) 1.41't 1 .63" 

S/H Back Flow Margin (Location) 1 .065 (90" Span) 1 . 065 (40"' Spon) 

Tmax 1351 K (1 972·F) 1361 K (1991·F) 

Full~ Pluggpd Showerhead 

Critical hc(" of Full Flowing Case) 76" (50" span) 69" (5C" span) 

(at critical location) 

Tm max +450 K (+3S0·F) +43(J ~ (+31S·F) 

Manufacturability 

Castabi1i ty Base More difficult 

(sma 11 er cores 

and very sma 11 

holes) 

Showerhead hole drilling Base ~'ore difficu1 t 

(small er target 

passage) 

44 

-

.t;

"=
J 

r 

r 

~. 

, 

! 

t 

, 

. .. 

5.1.4 Thermal Barrier Coating with Film Cooling 

Current advanced state-of-the-art airfoil cooling designs employ a combination 

of convective cooling with fiim or thermal barrier coating. In this study, the 

desirability of combining advanced film cooling with thermftl barrier coating wa~ considered. The result~ show that the use of film cooling in com~ination 

with thermal barrier coating minimizes the amount of co01ing air required in 

the supersonic cruise aircraft application. 

Airfoil cooling effectiveness and the interrelationship between film cooling, convective cooling and thermal barrier coating were investigated by treating the airfoil as a heat exchanger. Utilizing an idealized one-dimensional 

analysis of the airfoil (Figure 5-13), the overall cooling effectiveness (~) 

is expressible in terms of film effectiveness (nf), heat exchanger efficiency (ryc)' and heat load parameter (~) as: 

This expression was utilized to assess the relative merits of one cooling design versus another and to assess the relative level of difficulty 

associated with a given cooling requirement. Shown in Figure 5-14 is 

e\fectiveness level (0) versus heat load parameter (~) for a film 

effectiveness (ryf) equal to zero. It can be seen for the VSCE blade that: 

1 ) 

2) 

The required average effectiveness of the VSCE blade is significantly higher than the present state-of-the-art turbine blade. The VSCE airfoil 

requires an average,effectiveness of 0.71. Present state-of-the-art blades 

achieve effectiveness levels of only 0.56 for the most advanced design. Thus, the VSCE blade must achieve an effectiveness level 26 percent higher than present state-of-the-art blades. 

The required effectiveness levels can be achieved in the VSCE blade using 

film cooling with flow levels less than that required for convective 

cooling designs having heat exchanger efficiencies equal to 1. If a 

convective ef~ettiveness of 1 was attainable, the heat load para~eter of 

the VSCE blade must increase 28 percent to attain the desired 

effectiveness with no film (as detailed in the following paragraphs). This 

would represent an increase of approximately 0.8 percent Wae in cooling airflow. The VSCE blade achieves a convective effectiveness level of 0.81. 

At this level of convective effectiveness the heat load parameter must 

increase more than 50 percent. This would represent an increase of over 

1.4 percent Wae in cooling airflow. 

3) From the above observations it is concluded that the VSCE blade must use 

film in order to achieve the required effectiveness with a minimum of 

cooling airflow • 

45 

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t.: _ .;.::. .........." .... -....... --------------=--=-------------------- 

"'"'-- __ a .'u •• _ __ _ _ _ _. - ------~---- -- ~----
r f. 

, 

! 

46 

• IDEALIZED AIRFOIL (THERMAL CONDUCTIVITY IS INFINITE) 

GOVERNING EQUATIONG 

• OVERALL EFFECTIVENESS: 

• HEAT EXCHANGER EFFICIENCY: 

• FILM EFFECTIVENESS: 

• HEAT LOAD PARAMETER: 

SOLVING FOR + IN TERMS OF "If' "e' {l: 

"It + ({l - '1t) '1e 

+ = , + ({l - "It ) "Ie 

+ -

'1e • 

'11 .. 

{l := 

T 9 .. T M 

. , Te 1 

Te - Te 2 1 

TM - Tel 

Tg - Tf 

T - T L 9 ;> 

mCPe Tf - TM -- ,: 

hf Ag Te - Te 2 1 

Figure 5-13 Airfoil Cooling Effectiveness One-Dimensional Analysis 
, 

~ 

, i 

. 

~ ,. 

, ~ 

. ~,\ 

§ 

~ 

.@ 

~~'. 

-1::0 

~ 

10 

T ~I§ 

I Ol I' Ol 

t- t-

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w 

Cl 

Z 

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0 

0 

U 

0.2 

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o 

FILM COOLED 

100% CONVECTIVE 'If = 0 - 30'lb INCREASE 

~f ' 0 

T~-TC, 

T/ a T 

C TU - C, BASED ON hgas 

--. ~ 

VSCEBlADE/ 

_____ ¥-8ASED ~ hEFfEgM lCORAECTEo FOR T~ , 0 

0.1 

- ~ __ ---1l6 

~ __ ---- 04 

02 

--

CURRENT AIAfOIl. DESaGHS 

l1e l1f 

1 Cl485 O.lOl 

2 0.675 0 

3 097(; 0 

4 oe 00!i0 

5 0720 0 

e Cl430 0430 

04 08 1.2 1.6 2.0 H 

I I 

2.1 32 

HEAT LOAD PARAMETER me Cp T f - T M , h = 

t

Ag .Tc -Tc 2 i 

Figure 5-14 Airfoil Cooling Effectiveness Benefits Due to Thermal Barrier 

Coating 

. .... 

11 
L 

" 

r. 

, 

The above advantages of film cooling can be 111ustrlted by looking at the vSCE 

first stage blade cooling requirements. The VSCE first stage blade cooling requires an effectiveness level of 0.71 to achieve its design creep life. The 

coolant flow required to achieve this effectiveness is a function of the 

airfoil heat exchange efficiency and average film effectiveness. This is shown 

in Figure 5-15, where cooling airflow level is indlcated by the heat load 

parameter (~). If film ~coling were completely e11minated (ryf • 0), and the 

heat exchange efficiency of ~~ity were achievable, the 100 percent convect1ve1y cooled blade would still require 28 percent more cooling airflow 

than the combined convective and film cooled desit~ 

The utilization of thpnna1 barrier coating serves to effectively reduce the 

external heat load to the airfoil and, therefore, increase the heat load 

parameter (~) for a given level of cooling flow: 

where hg 

tTSC 

KTBC 

Ag 

Cp 

me 

he ff ec the • _.--;.1_r--_ 

1 + trsc 

-n; KTSC 

~. me Cp 

h effective Ag 

• gas !tide heat transfer coefficient 

• thermal barrier thickness 

• thermal barrier conductivity 

• gas side convection area 

• specific heat of cooling air 

• cooling air mass flow 

The reduction in the external heat load or increase in heat load parameter is 

independent of the use of film cooling and depends only on t~e initial heat 

transfer coefficient and the properties of the thermal barrier coating. The 

benefits of the thermal barrier coating to th~ VSCE airfoils are graphically shown in Figure 5-14, where the blade heat 10ad parameters are plotted with 

and without thermal barrier coating. It is $~nn that the thermal barrier 

coating effectively increases the heat load by 30 percent at constant coolant 

flow. 

48 

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i 

t-, 

w 

> 

... 

3.2 

3.0 

~ 2.8 ... 

... 

w 

..c. 

Z 

o 26 

o 

w 

(f) 

.1: 

al 

LI: 

w 

~ 

w 

~ 2.2 

<t 

LI: 

<t 

a.. 

o 

<t 

o 

...J 

~ 

<t 

w 

I 

0> 

<t 

Z 

'-Cl. 

u 

·E 

II 

"'l 

+ • 

'1, t (fl - '1,) "c 

1 t (fl - !) '1C 

'lc = 0.81 

•• d"l l 

1.0 '--____ .&...-____ .&...-____ .&...-____ ...1 

Figure 5-15 

o 01 0.2 0.3 0.4 

"I, - FILM EFFECTIVENESS 

Heat Load Parameter vs Film Effectiveness; Blade ~ = 0.713, 

T} f = 0.34 

f£) 

.l _______________ ... ______ ...i.-___ ....... -..•.. ~~._ •• ~~ ____ ... __ _ 

49 
, 

0-

r 

Increasing airfoil heat exchanger efficiencies as a means of reducing cooling airflow becomes increasingly more difficult as the heat exchar.ger efficiency approaches unity. Heat exchanger efficiency is given by the expression below 

and shown in Figure 5-16. 

Tj • c 

TCout - TC in 

Tmeta1 - TC in 

• 1 - exp -

where hc • cooling side heat transfer coefficient 

In the case of the VSCE first stage blade with a heat exc~anger efficiency of 

0.81. a 10 percent lncrease in heat exchanger efficiency would require a 32 

percent increce in hc • and a 14 percent increase in heat exchanger efficiency (to 95 percent) would require a 76 percent increase in hc when 

compared to the base VSCE deSign case. Considering the level of effort which 

would be required to achieve these higher efficiencies and the relatively small decrease in the resulting coolant flow. it does not seem reasonable to 

pursue further increases in internal cooling complexity. The pursuit of more 

internal cooling complexity becomes even less attractive when the effects of 

higher required supply pressure are included in the analysis. 

5.1.5 low Hf.at load Design 

Historically. the desisn of airfoils for high-pressure turbines was primarily 

conducted according to aerodynamic criteria to optimize performance. External 

heat loads were calculated for aerodynamically optimized airfoils and. 

subsequently. cooling air requirements were estimated to achieve the goal in 

metal temreratures occasionally at the cost of compromises to aerodynamic 

des i gn. 

The initial design of the high-pressure turbine for the supersonic cruise 

aircraft was conducted by utilizing the airfoil design system which optimized aerodynamic performance. A study was next conducted under the present contract 

to establish whether the mea~ section of the rotor and th~ stator airfoils 

cou 1 d be redes i g;..;:d to reduce hea t loads 'Iii thout comp rom is i ng the aerodynami c 

performance. A two-dimensional potential flow solver and a differential 

boundary layer code with a new turbulence model were utilized in the design optimization study. This code. developed at Pratt & Whitney. gives reliable prediction~ of both momentum and thermal boundary layers on a wide range of 

turbine airfoil designs and makes possible optimization of turbine airfoils 

for both ae'~odynamic and heat load consiaerations. 

Attempts to reduce the heat load on the vane were not successful without 

compromising the ae~'odynamic performance. The mean section of the blade was 

then redesigned by reducing the curvature on the airfoil pressure side. The 

redesigned blade cross section is compared to the original in Figure 5-17. 

Comparion of the obtained airfoil pressure distribution is very similar 

(Figura 5-18). Th~ profile losses calculated for the redeSigned airfoils are 

50 

- _ ...Ii..-
I 

, 

~ 

5 

! 

; I 

:q 

: J 

'1 

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U1 

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u 

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1.0 

0.9 

0.8 

0.7 

0.6 

0.5 

04 

0.3 

0.2 

0.1 

DOTTED LINES SHOW LIMITED POTENTIAL FROM CONVECTION COOLING 

1/ C =0.95 • + 76% hCAC- - - - --- - _ -~7TF:::a----------------- ~C-091 + 32°>bh

C

A

C ____ _ 

0.5 1.0 1.5 2.0 2.5 30 35 . hCAt:,m Cp 

Figure 5-16 Heat Exchanger Effici~ncy vs Heat load Parameter (hC Ac/m 

Cp) 

-. 

40 

45 
10 

0.8 

06 

0.4 

0.2 

a -

BASE DESIGN 

LOW HEA T LOAD DESIGN 

o 0.2 0.4 0.6 0.8 1 0 

X/BX 

Figure 5-17 Blade Geometry for Low Heat Load Design 

fi

\:
~. , 

r 

,. 

r£) 

1.0 ".. ___ .... _~ 

o.g 

0.8 

0.3 ~---...L.--__ "-___ ..J 

o 2.5 5.1 76 

101 1101 12.0 13.01 

a"SE DES~GN __ _ 

LOWHEAT _ 

LOAD DESIGN 

SURFACE DISTANCE FROM LEADING EDGE. CM (IN) 

Figure 5-18 Blade Loading for Low Heat Load Design 

: "~ ______ M ____ " _________ - ____ '''''''_~_ ••• u ••••• _.. • •.• . _______ •• _ .... _. 

53 

~

::
f 

~, 

, 

. 

~ .. 

the same as calculat£-d for the base airfoil indicllting that the optimized heat 

load airfoil has retained the aerodynamic performance of the base design. The 

redesigned airfoil has lower curvature on the airfoil pressure surface in the 

trailing edge region. 

External heat load to the pressure surface of the blade has been reduced to 

reflect aerodynamic tailoring near the trailing edge. Analysis indicates that 

external heat load to an airfoil pressure wall can be lowered by reducing the 

curvature near the trailing edge. Boundary layer analysis of the VSCE blade 

predicts nearly a 50 percent reduction in external heat transfer coefficient 

near the trailing edge (Figure 5-18) for airfoil contour tailoring as shown in 

Figure 5-19. The level of heat load reduction is based on a surface roughness of approximately 50 to 70 microns. This roughness level should be achievable 

with polished thermal barrier coating, and data obtained from commercial 

application indicate that it will be sustainable during operation. Increased 

roughness levels will only lessen, but do not eliminate the heat load 

reduction benefit. Cascade tests will be required to verify the actual level 

of benefit and the sensitivity to contour shape and roughness changes. 

Predicted heat transfer coefficients and the metal temperatures for the 

pressure surfaces of the base and the redesig,.ed airfoils are also shown in 

Figure 5-19 as a function of surface distances. Reductions in heat load and 

metal temperature distribution for the redesigned blade were obtained for over 

40 percent of the length in the trailing edge regi~ns. 

Predicted metal temperatures (Figure 5-20) for the redesigned airfoil are 

lower than the base d~sign by as much as 340 K (120·F) in the critical 

trailing edge region of the airfoil. Experience at Pratt & Whitney has 

indicated that the trailing edge region on the rotor airfoils may have maximum 

thermal distress problems. The reduction in heat load results in lower metal 

temperature in this region and allows reduction of coolant flow. 

The heat load distribution on the suction surface of the base and the 

redesigned airfoils was found to be similar and, therefore, not shown in this 

report. 

5.2 Preliminary Design 

This section describes the preliminary designs of the high-pressure turbine 

vane and blade, making use of the aerodynamic design results, configurational studies, and technology assumed to De available for the VSCE in the time 

period when the VSCE will be in use. The term "preliminary design" in the 

context of this report refers to complete cooling design definition at the 

airfoil midspan section, including all necessary iterations and detailed 

temperature analysis usually included in the final design. However, not 

included are design iterations which do not affect cooling, detailed 

manufacturing tolerance definitions, inspection and acceptance criteria, final 

detailed dimensional drawings, and cost considerations • 

54 

r 

f!, _ ~"",-~o"", t

-

1. _________________ -" ______ "'"--'-___ ••.••.... - .. - .. ' -- ------.--~-- -

t~ 

\;
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f 

Fi gure 5-19 

t. 

, 

700 

600 

PRESENT STATE·OF·THE·ART ---

LOW HEAT LOAD 

200L-----------~----------~------ o 2.5 5.1 . ____ ~ 76 

10) 11.0) 12,0 (3.0) 

SURFACE DISTANCE FROM LEADING EDGE. CM (IN) 

Heat Transfer Reduction on Pressure Surface for Low Heat Load 

Design 

55 
U1 

~ 1233 

(1760) 

1222 

(1740) 

1211 

(1720) 

1200 

U. 

~ 

~ 

~ w 1189 

, 

, a: (1680) 

=> ! le:{ 

a: 1178 

~ " w 

IL. (1660) 

~ 

w , ~ I- , "~I 1167 W .i t 

U (1640) , e:{ r u. 

:~ a: 

UJ 1155 " IZ (16201 

1144 

(1600) 

1133 

! 1580) 

1122 

'15501 

1111 

(1540) 0 

@ 

~~'. 

0,5 

(02) 

F;gure 5-20 

" 

6 PRESENT STATE OF THE ART PRESSURE WAll 

o LOW HEAT LOAD PRESSURE WAll 

i 5 

(0.6) 

25 

( 10) 

36 

11 4) 

46 

11 81 

SURFACE DISTANCE FROM LEADING EDGE, LM (IN) 

Blade Interface Temperature Pressure ~all 

~ - - -,- ... - .. ~-.~ -:---

56 

f2 2) 

66 

2 5' 
I 

r 

I' 

! 

~ f 

~ 

Two preliminary designs are provided for both the vane and blade for 

s",bsequent technology benefits evaluation: one uses technology assumed to be 

available in the 1990 time period and the other limited to technology 4ssumed 

to be available in 1986. These technology projections and detailed design 

criteria used for the durability design are summarized in Sections 5.2.1 and 

5.2.2. Design details and analysis results are presented in Sections 5.2.3 and g.2 4. 

5.2.1 Technology Projections and Design Features 

The following com~onent and materials technology was assumed to be available 

for the supersonic cruise aircraft engine and was used in the durability 

design. Materials technology predictions for 1986 and 1990 are shown in Table 5-11 1. 

o Advanced combustor design with 1.25 gas temperature pattern factor and 

1.10 profile factor (1990) and 1.12 profile factor (1986) at combustor exit. 

o Advanced single crystal materials that are 370 K (175-F) and 356 K (150-F) 

better than that used for the Energy Efficient Engine vane and blade, respecti vely. 

o Advanced casting methods to provide 0.043 cm (0.017 in) nominal thickness 

wulls around 0.043 em (0.017 in) thick core; trip strips skewed 45-

relative to coolant flow directi~n; an~ high density ribs and pe~estals in 

vane and blade trailing edges. 

0 Conical shaped showerhead film cooling holes. 

0 

0 

0 

0 

0 

Advanced thermal barrier coatings for the 1990 airfoils; zirconia based 

thermal barrier coatings for the 1986 airfoils. 

Diffuser shaped film cooling holes through the thermal barrier coating. 

Advanced oxidation resistant metallic coatings. 

Blade root stress to be 20 percent higher than that used in the Energy Efficient Engine high-pressure turbine blade design. 

Advanced disk materials with a 301 K (50-F) higher temperature capability than the current state-of-the-art materials. 

57 

r! ... t: _____ . ___________________ .-_.. _au ..... __ _ ~~_~;-::"f~.~ ._r- ~~ ______ ~ ___ _ 

-fi

~
, 

r 

TABLE 5-1 II 

VSCE MATERIALS PROJECTIONS 

J 986 Tecnnolog,Y.. 

Fi rs t Vane 

Vane MaterUl: SC1000 (single crystal) 

Metallic Coating: 

PWA 286, 0.012 cm (~.005 in) thick 

Thermal Barrier Coating 

(Yttria Stabilized Zirconia) 

0.025 cm (0.010 in) thick 

+329 K (+100°F) Interface iemperature 

Relative to 264/276/PWA647 

Fi rs t Blade 

Blade r~aterial: SC3000 (single crystal) Metallic Coating: 

PWA 286, 0.Oi2 cm (0.005 in) thick 

+301 K (+50°F) oxidation 

+329 K (+lOO·F) Creep Strength 

Relative to PWA 1480 

Fi rs t Vane 

1990 Technology 

First Blade 

Vane'1aterial: SC1000 (single crystal) 131a~e t1aterial: SC3000 (single cryc;tal) 

Metallic Coating: 

PS200, 0.012 cm (0.005 in) thick 

Thermal Barrier Coating 

(No Zi rconi a-Some\,ha t Opaque) 

0.038 cm (0.015 in) thick 

+367 K (+170°F) Interface Temperature Relative to 264/276/PWA647 

5.2. 2 De s i g n C rite ri a 

Metallic Coating: 

PS200, 0.012 cm (0.005 in) thick 

Thermal Barrier Coating 

(No Zirconia-Somewhat Opaque) 

0.025 cm (0.010 in) thick 

+340 K (+120°F) Interface Temp. 

Relative to 264/276/PWA647 

+329 K (+lOO°F) Creep Strength 

Relative to PWA 1400 

The vane and blade cooling levels were set to achieve 10,000 hours operating life 

';n the predicted failure modes of creep, oxidation and thermal cracking. Limiting 

criteria were considered to be: 2 percent local creep or thermal fatigue cracking 

of the base metal, loss of metallic coating with 0.012 cm (0.005 in) penetration of base metal, and/or spa11ing of thermal barrier coating. 

58 

:: 

s 

"~ ______ = __________ ."' ____ .M.-_________ leu_a .. 

-a ___ • ~:; ___ ·""~< ._ _ _________________ .. -I • ~ _ 

fi

\J
, 

f 

The design gas temperature profiles for the vane and blade are shown in Figure 5-21, based on a 1.25 pattern facto~~TVR) for the hot spot vane and a 1.10 

profile factor for the blade. These temperature profiles reflected expected 1990 combustor technology. The 1986 airfoils were designed to the same pattern factor but with a 1.12 profile factor for the blade, which is representative of the expecte1 1986 combustor technology. Temp~rature profiles for the 1986 

design shown in Figure 5-22 include the effect of an additional 3.2 percent Wae cooling airflow to account for the reduced technology level in the 1986 

engine. 

80th Figure 5-21 and 5-22 temperatures reflect nominal engine performance plus design increments explained in Section 5.3. 

5.2.3 Mission life Analysis 

Mission life analysis combines mission operational variables and hot section 

life models in order to predict turbine vane and blade life. The operating variables must include a mission definition, appiicable engine ratings, engine 

performance. ambi ent envi ronm':nt condi ti OilS and deteri orati on due to 

tolerances or we,r. life l~;;-!lting modes applicable to the VSCE hot section 

were creep and c~,ting oxidation and spalling, based on the temperature and 

stress conditions determined in Sections 5.2.3 and 5.2.4. 

Initial performance levels for the VSCE were assumed to be repres~~ted by the 

performance of the Study Turbofan engine in Reference 1. The correspondlng nominal temperatures and rotor speeds applicable to the selected mission 

(Section 3.2) are summarized in Table 5-1V. The temperatures and speeds seen 

in a turbine during actual service and subsequently the design temperature levels are higher than reflected by nominal performance. Temperature levels 

increase relative to nominal performance due to the effects of installation, 

deterioration during operation between overhauls, and nonrecoverable overhaul 

losses in the case of a mature engine. Consistent with subsonic commercial 

design practices increments have been added to the nominal VSCE performance as 

shown graphically in Figure 5-23 and summarized in Table 5-V. To determine 

performance for an installed, previously overhauled (mature) engine typical of 

commercial supersonic operation, increments for installation, nonrecoverable 

overhaul and deterioration must be applied to the nominal temperatures and 

speeds as graphically shown in Figure 5-23. Table 5-V summar'hes all such 

increments at the limiting supersonic cruise operating condition, based on 

present JT9D cOlllTlercia1 engine experience, scaled to : ~f1ect the supersonic flight time where applicable. In addition, an increment was included to 

account for performance loss due to vane and blade cooling airflow increase 

over the reference engine estimates. Cooling flow increases would thermally reduce the temperatures through the turbine before work could be extracted 

with a resultant loss in thrust. In order to maintain t~rust with increased 

cooling airflows, the combustor exit temperature must ~e increa ,·d. This inc~ease was taken prior to design in order to more r~alistica11y account for 

the temperature 1eve15 expected in actual operation. The increase was 5.2 

percent for the 1990 engine and affected both combustor exit and rotor inlet 

temperature. fhe resulting temperature profiles for 1990 and 1986 technologies 

are shown in Figures 5-21 and 5-22. 

59 

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l) 
t, 

I 

r 

f 

-

60 

u.. 

0 

lI:: 

w' 

a: 

::J 

J-

< II: 

w 

Cl. 

~ W 

JVl 

< 

t) 

2144 

134001, 

2033 

132001 

192:1 

130001 

1811 

128001 

1700 

126001 

1589 

124001 

1478 

(22001 

1367! 

(200010 

Fi gure 5-21 

HOT SPOT VANE ~ TVA. I ,''> 

;& AVERAGE PROfiLE 

''II. PROFILE FACTOR", 

-

ro; IIlJIK r2!:l3 r) 'II 

I 

10 I I 20 3D 40 50 60 70 80 90 100 

PERCENT SPAN 

Design Gas Temperature Profile • 1990 Technology 

--- 

.~

\::
U. 

0 

2144 

134001 

2033 

132001 

1922 

130001 

HOT SPOT VANE .l TVR • 125 

,.---__ ..r; ________ ~ T • 20951< 13311 "f' 

T!j AVEnAGE PROFILE 

12% PP'JFILE FACTOR 

" r 5 • 18601( '2888"~1 

W 

a:: 

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to-

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'2200) 

131'.7 

12000)O~-~--::I:-~::---:'::--~-...J..--.J--BO.L..-.J90-.....J, 00 

--~-

PERCENT SPAN 

Figure 5-22 Design Gas Temperature Profile • 1986 ,cchnology 

LIII~ __ ~_' _ •. ____________________ . . 'Ua'h __ ~~_::..' .......... ; •• _. _____ ~ ____ ~_ 

61 
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389 (70)r 

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(60) • 

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(40) 

16.7 

111 

OVERALl DETERiORA nON 

---- MISSION AVERAGE DESIGN LEVEL 

-----------------NONRECOVERABLE OVERHAUL 

--------------,NSTALLATION 

o 

(0) r UNINSTALlED ~~OMINAl PERPJRMANn 

-5.5 

(- 10) 

- 11 1 L 

(-20)O~---------~-----------~-----------~-----------~--------__ 1 __ 0.~ 

FLIGHT HOURS 

Figure 5-23 Mission Performance Increments 

." 

.- ... -----------_.-_ ... -...... '"".~.~.::.-::--

-- - -
· 

~. 

Flight Segment 

Taxi (I d1 e) 

Takeoff 

Subsonic Cl imb 

Subsonic Cru'lse 

Supersonic Climb 

Supersonic Cruise(2) 

Dp.cent and Approach Taxi (Idle) 

DESIGN REQUIREMENT 

TABLE 5-IV 

SUPERSONIC CRUISE AIRCRAFT ENGINE 

MISSION LIFE DESIGN PARAMETER.S 

Time T4(1 ) T5 (mi n) K (eF) K ( eF) 

1 n. 0 

0.7 800 (980) 1483 (2210) 17.4 625 (665) 1376 (2017) 18.5 603 (6?6) 1253 (1796) 14.7 719 (835) 1580 (2385) 99.7 915 (1187) 1754 (2698) 24.6 

5.0 

Total Block Time • 190.6 minutes 

Total night Time • 175.6 mi nutes 

(1) Temperatures are for a nominal base engine on hot days. 

N2 

(RPM) 

9482 

8712 

8235 

9435 

10075 

---

(2) Design Point - Temperatures are not rep~esentative of actual design ltvels. 

High-Pressure 

Compressor Ex it 

Temperature K (oF) 

Burner E.~ 1 t 

Temperature K (OF) 

High Rotor 

Speed - RPM 

Study 

Turbofan 

Nomina I (J) 

915 (1187) 

1754 (2698) 

10075 

(1) STF515B nominal performance 

TABLE 5-V 

VSCE DESIGN POINT PERFOR~~NCE 

SUPERSONIC CRUISE 

Nonrecoverable Average Cooliny Deslvn Installat ion Overhaul Deterioration Air (2 Leve 

+1.11 (+2) +1.67 (+3) +1.67 (+3) 0 919 (1195) 

+3.89 (+7) +10.56 (+19) +13.89 (+25) +47.78 (+86) 2130 (2835) 

+7 +28 +36 10146 

(2) Temperature increase due to increased cooling flow relative to initial performance table (IV 1.9S to 6S, IB 2.5S to 3.61) 

63 

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" I 

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r" 

q 

The existing Pratt & Whitney (P&W) mission analysis data base. reflecting International Standard Atmosphere definition. was used to describe ambient 

temperature distributions expected for the VSCE. However. since the 16.154 m 

(53.000 ft) cruise altitude for the supersonic mission exceeds limits of the 

P&W data base. sorn.e extrapolation was necessary. The probable ambient 

temperature distribution for the 16.154 m (53.000 ft) cruise altitude is shown 

in Figure 5-24. The average ambient temperature at this altitude is shown to 

be 216.5 K (-69.7·F) and one standard deviation equal to 8.2·C (14.76·F). The 

engine was assumed to be flat rated in thrust up , t sta~dard day plus +8 K 

(14.4·F). The combustor exit temperature responst. ',I "'mbient temperature variation is shown in Figure 5-25. 

Airfoil lives were calculated for the vane and blade designs. The 1imitir.g damage modes are coating oxidation and/or spa11ing for both the vane and 

blade, and creep for the blade. Figure 5-26 shows the predicted limiting lives 

of the vane as functions of metal temperature for the 1986 and 1990 engines. The me~a1 temperature represents what can be expected in a typical installed. 

average deteriorated engine in hot day (+8 K. +14.4·F) at supersonic cruise 

ronditions. Figure 5-27 shows the predicted blade creep life as a function of 

design average metal temperature. 

A breakdown of damage due to various flight conditions of the mission (Table 

5-VI) shows that cruise is the most severe condition as expected. 

TABLE 5-VI 

BREAKDOWN OF BLADE AND VANE DANAGE 

AT VARIOUS OPERATING CONDITIONS 

Creep (Blade) 

Oxidation (Blade and Vane) 

5.2.4 Vane Durability Design 

Takeoff ('I,) 

o 

o 

Climb 

.J!L 

o 

1 

Cruise 

( 'I, ) 

100 

99 

Both the 1986 and 1990 vane cooling designs utilize advanced single crystal material and thenP:'~ barrier coatings in order to meet the design life 

requirements with a minimum amount of cooling airflow. The cooling configurations incorporate two cavities in order to optimize the distribution 

of cooling airflow and to maintain acceptable levels of bulging stress. 

The leading edge cavity cooling is achieved by using impingement and external 

surface film. Cooling dir is fed into the leading edge cavity from both the 

inside and outside diameters, impinged on the inner wall via an impingement tube, and then vented through the walls as external film obtaining a 

showerhead pressure ratio equal to 1.02 to prevent backflow. The pressures 

64 

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~ 

~ 

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: If; 

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233 . 

( - 41) 

223 

(- 59) 

213 

( -77) 

203 

( 95) 0.01 

iTAMB)AVG ~ 217K(- 69°F) 

,. _ 8 2K(14 8°F) 

+0 

-0 

67% 

0.1 10 50 90 

PERCENT PROBABILITY OF OCCURRENCE 

Figure 5-24 Arrbient Temperature Probability Distribution at 

16,154 m (53,000 ft) 

- ~ .- til. 

99 998 999 

- -
, 

~ 

i 

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If \ 

,f' 

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1721 

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1361 

0 

iOl 

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1- 721 

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(- 1081 

8 -4 

( - 144) 1-7 2) 

Figure 5-25 

.- ~. 

o 

101 

4 

17._1 

~ T AMB' K (OF) 

8 

.1441 

12 

(2161 

16 

1288\ 

Variation of Combustor Exit Temperature as a Function of 

Variation of the Ambient Temperature 

20 

'360, 
., 

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16,000 

14,000 

12,000 \ 

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cO . 8000 

\ \ 

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6000 

4000 

1700 (1700) 1256 (1800) 1311 (19001 

REFERENCE INTERFACE TEMPERATURE, K (OF) (INTERFACE AT DESIGN POINT) 

Figure 5-26 Turbine Vane Thermal Barrier Coating Spalling Life vs DeSign Metal Temperature 

(£ 

I.t _________ . __________ •____ '""'-____ _ ...to __ ~;; ""'''''4''~. ' - ~_~~ ___ ~_ 

._J 

67 

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0- 3 .. 5 MPa 

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~ DI:SIGN REQUIREMENT 3820 HOT HOURS TO U 

W /' ./ 2% CREEP 

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4OOOL-______ , ____ ~------~--~~----~~, 

1033 (1400) 108911500) 114411600:' 1200'(1700) 

DESIGN AVERAGE METAL TEMPERATURE, K (OF) 

Fi gure 5-27 Turbine Blaee Mission Creep Life vs Average Design Temperature 

, 

t 

~. 

r! 

.. 6 ________ ... ____________________ __ ." .. -... _ ,. _', ,. 

68 
it 

f 

t 

,. 

[ 

outside the impingement tube are balanced so that internal dams are not 

required to d1strlbute the flow. This design configuration produces a much 

more reliable cooling than configurations relying on internal flowaams to 

maintain separate cooling regions. Experience with previous vane designs has 

shown that casting tolerances and impingement tube tolerances combined to 

result in leakage around the internal flow dams. The internal leakages create 

cooling maldistribution followed by burning or cracking of the airfoil. 

Midchord and trailing edge cooling is achieved by using a tubeless 

configuration with flow in the midchord region passing radially from the 

outside to the inside diameters. and then flow chordwise through the multiple channel passages as shown in Figure 5-28. Cooling in both the midchord and the 

trailing edge regions is augmented with internal flow turbulators and film. A 

tubeless m1dchord and trailing edge configuration is utilized because it is 

anticipated that a three-dimensional aerodynamic airfoil design with bowing and leaning of the airfoil will be used. This would ma~e insertion of a tube 

into the trailing edge cavity extremely difficult if not impossible. The 

midchord cavity is tapered as depicted in Figure 5-28 to maintain high internal heat transfer coefficients as fl~W is bled off in going from the 

outside to the inside diameter. The degree of taper will be determined during final design of the vane when multiple sections will be analyzed in detail. 

The trailing edge chordwise passages contain 0.025 cm (0.010 1n) trip strips skewed 45 degrees to the cooling airflow as shown 1n Figure 5-29. Trip strips 

are carried as far into the trailing edge as possible until a minir:um core 

thickness (flow channel height minus trip strip height) of 0.050 cm (0.020 in) 

is reached. This minimum core thickness with trip strips is judced to be near 

the limit of acceptable castability. Additionally. extending t;le trips further 

downstream would serve to significantly increase the pressure losses. thereby reducing the flow capability of the trailing edge. 

The uncoated trailing edge diameter is 0.129 cm (0.051 in). consistent with 

the expected technology levels of 1986 and 1990. The VSCE high-pressure turbine has relatively low levels of aerodynamic trailing edge blockage due to 

its large size and reduced number of airfoils. 

1990 Technology Vane Design 

The 1990 technology vane design incorporates an advanced non-zirconia thermal 

barrier coating and shaped film holes in addition to the features common to 

both levels of technology. A non-zirconia thermal barrier coating. being somewhat opaque to oxygen. allows a higher deSign temperature for interface 

spalling criteria than a current state-of-the-art zirconia based coating. 

Additionally. the 1990 coating technology will allow a 0.038 cm (0.015 in) 

thick coating which is 0.012 cm (0.005 in) thicker than current (1906) 

state-of-the-art technology. Shaped film holf.s through thermal barrier coating 

are assumed to be technically feasible by 1990 and are incorporated into the 

desi£,. 

69 

~\ 

l~ 

\::
TRIP STRIPS 

IMPINGEMENT TUBE 

___ -10 

Figure 5-28 ~ane Cooling Configuration 

. 

f 

70 

r 

.. 6 ~ _______________ . ___ ..6.-_. ___ .... u .... - __ __ _ -:!::;,_ ... ~-'1~J-- ._ #.~~~ __ • ______ ~ 

SEe FIGURE 6-2 
o 130 

100611 

0043 

100171 

~~~~~~-==~~ 

UNCOA TED DIMENSIONS. CM (IN) 

0.254 

10100) 

... 0.178 0.178 

10,070) 10,070) 

0.432 

10.170) 

Fi gure 5-29 Typical Vane Trailing Edge Geometry 

.. ~ ____ ... c ________ .... __________ ~ _ •• u:~ ___ .~:;,.'~;.,-;-F;' .. ' 

T 

.. 

71 

.~

\:;
1 

r 

r 

, 

t 

r 

Utilizing 1990 technology. the VSCE first-stage vane requires 5.28 percent 

Wae cooling airflow. distributed as shown in Figure 5-30. to achieve the 

design 8-5 life goal of 10.000 hours unfcheduled engine removal (UER). The 

resulting fnterface temperature distrfbutfon for the pressure surface and 

suction sur~ace fs s~own in Figure 5-31 and 5-32, respectively. 

Shaped film holes (Figure 5-33) are ut'iHzed on the second and third suction 

side film rows and allow the removal of two rows of suction sfde film holes 

and a 0.6 percent Wae reduction in required cooling flow. Film effectiveness 

for the shaped hole 1s based on recently obtained data for advanced 

state-of-the-art film cooled airfofls (Figure 5-34) and represents a 

signiffcant increase relative to standard cylindrfcal holes. 

The showerhead holes are designed to be conical, tapering from 0.076 cm (0.030 in) on the surface to 0.038 cm (0.015 in) at the inner surface (Figure 5-35), 

Tapering the showerhead holes produces the same leading edge coverage and film 

effectiveness at reduced coolant flow levels as closer spaced cylindrical holes would with higher flows, based on an extrapolation of the existing cylindrical showerhead hole data base. Confirmatfon of the effectiveness level 

as well as the resistance to plugging needs to be veriffed through testing. Cooling flow from the conical shaped holes are a function of the 

internal-to-external hole area ratio and the pressure ratio across the hole, 

as shown in Figure 5-36. In the VSCE vane design, the conical shaped holes 

reduce the cooling airflow by about 10 percent (0.6 percent Wae ). The flow 

characteristics of the actual machined hole shapes need to be confirmed by 

test. 

A summary of the initial internal and external hole diameters and spacing is 

shown in Figure 5-37. This geometry is based on an analysis of the mean 

section and may vary somewhat when a detailed analysis at multiple sections is 

undertaken for the final deSign. 

1986 Technology Vane Design 

The 1986 technology vane incorporates the 0.025 cm {0.010 in) thick yttria 

stabilized zirconi~ thermal harrier coating and casting and manufacturing technologies consistent with the 1986 time range .. The zirconia based thermal 

barrier coating is porous to oxygen so that the allowable interface 

temperature is 21 K (70·F) lower than that for the 1990 technology. This lower 

allowable interface temperature coupled with the thinner thermal barrier 

coating results in a significantly higher cooling requirement in 1986. 

Utilizing 1986 technology, the VSCE first-stage vane requires 8.66 percent Wae cooling airflow, distributed as shown in Figure 5-38. The resulting interface temperature dist~ibutions for the pressure side and suction side are 

shown in Figures 5-39 and 5-40, respectively. The internal and external 

cooling hole geometries based on the 1986 technology are shown in Figure 5-41. 

Relative to 1990 technology, the 1986 technology vane requires 3.38 percent 

more cooling airflow and reduces the turbine efficiency by approximately 0.3 

percent. 

72 

~ ~.;; -..,r.;' .. 

t;~. ___ -----------oA-.---. _ ........ -.' . 
0,72% 

r We A TOTAL 528% Waf) 

~P 

PT TOTAL -032% 

0,85% \ 

Figure 5-30 Vane Cooling Air Flows (1990 Technology) 

, 

r 

73 

r' 

E 

: 

.. ~ ________________ ..-'-____ •• u •••• __ , -.., ___ _ 
u.. 

0 

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Figure 5-31 

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O~SIGN LIFt AEOUIREM~Nl 

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5 1 

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6 I 

12 41 

PRESSURE SURFACE OIST ANCE. eM (IN) 

I I 

12 HI 

H I IJ 1 

Jill 

Vane Interface Temperature Distribution on Pressure Surface 

(1990 Technology) 

""-
u.. 

0 

~ 

W 

cr:: 

::J 

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<X: 

cr:: 

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Cl. 

~ 

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Figure 5-32 

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DESIGN lI~E flEOUIREMENT 

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4 1 

11 61 

5 1 

12 01 

6 1 

1241 

7 1 

12 81 

PRESSURE SURFACE DISTANCE. CM (IN) 

8 1 

13 21 

9 1 

13 61 

102 

14 01 

11 2 

14 41 

If 

Vane Interface Temperature Distribution on Suction Surface (1990 Technology) 

75 

~ ""'< -~':., . .. 6 __________ • ____ ......... ____ .. u_ •. _. __ .. _ 

'~

'=
I 

A 

[ ;:::; 

----\-

, 

V!EW A·A 

Figure 5-33 Geometry of Suction Surface Shaped Holes 

~. F 

76 t :' 

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fl,.. 

~ ________ . _.0li0.-____ .... _, ...... " ..... ..-. b __ _ 

J 

A 
~ 

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o CYLINDRICAL 

C 5'"iAPED (14") 

j 1 0 (HOLLOW) 

l'cVc/PgVg ~ t 20 (SOLIDI 

005~ ________________ ~ __________ ~~~ __ ~ ________________ ~ 

10 X/MS 100 1000 

Figure 5-34 Film fffectiveness vs Nondimensional Distance Cylindrical and 

Shaped Holes 

7· /-

> \ . . / 

/ ./ ./" / <, /" 

/ / J..-/" / ", / 

~ ~~~-. J~o. - ~-----? ~------~"""",~------------------------------­ .... 

i-igure 5-35 Geom~try of Conical Showerhead Holes 

77 

h - • _ .. =~ .r-..r::-:r:;~. ~. ..... _________________ .Ii-. ___ • ........ __ . ___ ~ ____ ~ ______ ~ __ ~_ _ ~ _____ ~~ 
, 

, 78 ~. 

~. 

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1 O~-Y-'---r-----'~----------------------------------------------

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O~ __ ~ __ ~ __ ~ __ ~ __ ~ __ ~ __ ~ __ ~~ __ ~ __ ~ __ ~ __ ~ __ ~ __ ~ __ ~ 

10 1 1 1 2 1 3 

Figure 5-36 Conical Holes Cooling Air flow Rates 

e: l - ~~'~~~3;'· ,,' 11 _____ =. _____ .. _________ ..-___ ..&-__ .-__ _ ...... ~ __________________ ~ _______ ~ 

,f

\:
3 

2 

---. ---. 

·SHOWERHEAD HOLES ARE 

30° RELATIVE TO 

RADIAL SURFACE' ALL 

OTHER GROUPS ARE 

HOLE DIAMETER PITCH 

GROUP CM(IN) CM(INI 

-'1 0038 (0015) 040610.1601 - 2 0.03810.015) 0.137 (0.0541 3 003810.015) 0.10210040) 

4 003810015) o 14210.0561 

5 0028 (0.015) 0"'5710.1801 

6 0038 (0015) 0.381 (0.150; j 7 006410025) 0.262 10 103) 

8 006410.025) 0.29510.116) 

9 007610030) 0345 (0 1361 

A 0064(0025) 0.122 10 048) 

B 0086(0034) 0.173 (0.0681 

C 0056 (0022) 0089 (00351 0 0.056 (00221 o 26210.1031 

E 0.05610.022) o 1 ciS (0061) 

F 006410025) 0201 (0079) 

0043 CM 

It- (0017 IN) --'""' 

Fi gure 5-37 Vane Hole Geometry (1990 Technology) 

1 79 

I" ____________________ "'"-____ ...... _____ ~ ~ 

.. --.-

.fi

'-
. 

~ r 

~.£ 

r I 

80 

10.66~ 

-~-__ -....,,_....:.- ----. 074% 

,.,,%-~~~-t---*+----

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1.'4% \ 

Figure 5-38 Vane Cooling Air Flows (1986 Techno'logy) 

! 

~ - - ~"'''-~'''',:.;~ . -' .. Ij..... ___ • _________________ d.-___ • .......... _ .. __ ... __________ . __ _ 
, 

,t 

r 

t. 

i 

f 

. 

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r 

1278 

118401 

1256 

118001 

1233 

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u.. 1211 ~ 111201 ~ 

w 

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w 

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I- (16401 w 

U 

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a:: 1144 w (16001 IZ 

1122 

(15601 

1100 

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1078 

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0 

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LE 

Figure 5-39 

10 

(041 

DESIGN ALLOWABLE TEMPERATURE 

-------------------

20 

(0.81 

3.0 

I· 2) 

4 1 

(1.6) 

5 1 

(2.01 

6.1 

(2 4) 

PRESSURE SURFACE DISTANCE, CM (IN) 

7 1 

(28) 

8 1 

13 21 

9 1 

13 61 

IE 

Vane Interface Temperature Distrib~ti'l~n Pressure Surface (1986 Technology) 

81 

r:t 

r~: ______ ~. -,., __ . b- OM cnc __ .A-. ___ ' •• u ••••• _ ••••.• 

(fi

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cr 

UJ Illig fZ 11680) 

11 "18 

116601 

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11640) 

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Figure 5-40 

82 

DESIGN ALLOWABLE INTERFACE TEMPERATURE 

\ :~ I " : \ 

I f \ ~ 

I 

til 

\ , 

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t!I 

rl!f j ~ ~ 

1 U 2 ° 30 4 1 5 1 b 1 7 1 8 1 9 1 102 1 1 2 12 2 

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SUC, 'Ol~ SURFACE DI5T ANCE, eM (IN; 1£ 

Vane InterfaCl~ Temperature Distribution on Suct.ion Surface (1986 

Technology ) 

E 

,.. 

COII!:- ~ _____ ~;; '>~:<"~(~ . -. ..... _ .. ___ .. __________ • ____ ~ __ ~ ..... Jou_ •• ~ __ • ___ •• _. ________ ._~~ _____ _ 

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Figure 5-41 

m Err r 

3 

·SHOWERHEAD HOLES ARE 

JO- RHA TIVE T ,) 

RADIAL SURFACE AU 

OTHER GROUP!, ME 

AT 90-

4 

Vane Hole Geometry (1986 Technology) 

*Z577 eD'? r 

83 

,tV 

\;
.. 

5.2.5 Blade Durability Design 

The VSCE first-stage blade operates at a design AN2 level approximately 50 

percent higher than present state-of-the-art second-stage turbine blades. This 

high AN2 (high stress) level, coupled with a hot operating time that is ten 

times greater than that for a typical subsonic commercial applic~tion, results 

in a cooling requirement that is significantly greater than the present state-of-the-art airfoils. 

In order to attain the design life goal of 10,000 hours B-S UER (f.e., 5 

percent of engines will experien\~e un .. cheduled engine removal) with a minimum 

coo11ng f)ow, both the 1986 and 1990 I!irfoil designs incorporate advanced 

single crysul material, a low heat-load pressure wall, stress reduc1rjg ribs, 

augmented internal convection, and advanced state-of-the-art film cooling. The 

1990 airfoil only incorporates a 0.025 cm (0.010 in) yttria stabflized 

zirconia thermal barrier coating in addition to these features. 

The blade cooling configuration (Figure 5-42) is a three-feed mUltipass design with showerhead and pressure and suction side film co01ing, internally 

augmented convection, and a high dens~ty pedestal trailing edge desi9n. 

leading edge cooling is accomplished through internal convection and 

showerhead film. Cooling airflow is passed from the inside di~D?ter of the 

blade through a leading edge cavity containing intern~l trip s'rins. The 

cooling air is vented to the main stream gas through two rows of showerhead in 

the leading edge, two rows of film holes on both the pres~ure and suction 

surface, and th~ remaining air to the tip where it is used for cooling the 

pressure si de and the tra 1'1 i ng edge ti p. A two row showerhead was found to 

achieve adequate cooling for both the 1986 and 1990 technology levels. The 

trip strip heignt and spacing ha~ been established to supply adequate cooling 

to the mean section, based on state-of-the-art commercial engine design practices. Trip strip geometry has not been optimized to refle~t any spanwise cooling variation. AnalysiS during detailed final design would be used to d~termine the optimum spanwise spacing. 

Midchord cooling is accomplished through a three pass configuration with trip strip augmented convection and external film. Cooling air is fed from a 

separate midchord feed, flowed radially upward in a rear midchord passage, down the midule region of the airfoil, and then up the last midchord passage 

where it is vented out through the pressure side and suction side film holes. 

Utilizing internal convection in conjunction with external film significantly increases the cooling effe~tiveness of the airfoil. Internally: the full 

potential of the cooling air is essentially used because the cooling 

temperature is approaching the w&ll temperature of the airfoil. Venting spent air over the airfoil reduces the effective gas temperature to the rear 

portions of the airfoil; since the spent air still has approximately 537 K 

(lOOO°F) cooling potential re1ative to the gas, it can still be used 

effectively to film cool the blade. Suction side film hole~ in the blade are 

shaped similar to the vane holes (Figure 5-33) in order to maximize the film 

84 

-- ...... cr .... ~ • ...a-t 

-~~---~---~------------­
r· 

, 

SUCTION SIDE FILM --::~;;t:;::~ 

SHOWERHEAD 

PRESSURE SIDE FILM 

TRIP STRIP GEOMETRY 

PaPITCH, E-HEIGHT,4-ANGLE TO FLOW 

PIE 76-204' .46° 

EO 025 -0051 eM 10 01_002IN) 

E 0025 eM 1001 INI-O,051 CM (0.02 INI 

PRESSURE SIDE TIP M 

STRESS REDUCTION 

Fiqure 5-42 

THREE FEEDS 

HIGH BLOCKAGE 

~DESTAL TRAILING EDGe 

...... u .......... INTER"/AL AUGMENTOR 

!TRIP STRIPS) 

Blade Cooling Configuration 

L~t.: __________ """-___ ........ . 

85 

(

.~

'=
l 

r r t, 

t 

r 

~. 

~ 

effectiveness. The trip strip height and spacing depicted in Figure 5-42 have 

been established to supply adequate cooling to the mean section and are not 

optimized to reflect any spanw1se cooling variation. Analysis during the 

detailed final design again will be required to determine the optimum spanw1se distribution. 

Trailing edge cooling is achieved through a single-pass cavity with trip strips venting to a high dens1~ pedestal trailing edge (Figure 5-43). The 

high density pedestal trailing edge is used to reduce the momentum mixing 

losses of the airfoil cooling air. The high density pedestals provide increased heat transfer coefficients fo~ a given flow as well as increasing 

the exit I~ach number of the coo11ng airflow which reduces the momentum mixing losses. Fi gure 5-44 shows the relationship between trail i ng edge pedest.al 

dtns1ty and momentum mixing losses for an initiaj estimate of 1 percent W~e 

trailing edge cooling flow. As can be seen from Figure 5-44. the increased 

density pedestals significantly reduce the cooling losses relative to a 

typical state-of-the-art pedestal trailing edge. The pedestal density and 

trailing edge gap of the final design wOIJld be opt1mited. dependent on the 

cooling airflow requirement of the trailing edge and the overall supply 

pressure requirement of the blade. 

Stress control in the VSCE blade is accomplished through tapering of the 

airfoil wall thiLkness and through the use of stress redu~tion ribs. Wall 

thicknesses are tapered. At midchord. the wall thickness is 0.203 em (0.080 

in) at the root and 0.088 cm (0.035 in) at the tip. Three stress reduction 

ribs are added to the midchord cooling passages (Figure 5-42) in order to 

further reduce the stress of the airfoil. The resulting pull stress 

distribution of the foil at design speed is shown in Figure 5-45 for both the 

1990 design and the 1986 design. Stress levels in the 1990 design are higher due to the added pull load of the thenllal barrier coat'lng. 

Tapering of the thermal barr-ier cOclting to reduce airfoil trailing edge thickness and improve turbine efficiency was not pursued in this blade design. Turbine efficiency could be increased by 0.12 percent with the thermal barrier 

tapered to the base undercoat (Figure 5-46}. However, the resulting metal 

temperature would be approximately 1244 K (1780·F) which. although acceptable 

for coating durability. would necessitate a blade coolant flow increase to 

maintain adequate creep life. Tapering of the thermal barrier coating should ~e reconsidered as part of a final design in terms of cost. durability an1 

efficiency trade factors. 

1990 Technology Blade Design 

The 1990 technology blade design r'equires a 2.77 percent Wae cooling airflow. (Figure 5-47) with 0.96 percent to the leading edge. 0.99 perc~nt to 

the midchord region. and 0.82 percent to the trailing edge. Creep is the 

limiting failure mode of the airfoil rather than oxidation or spalling of the 

thermal barrier coating. Cooling airflows are set to result in uniform creep within the airfoil and thus minimize the required airflow. Detailed creep ana1ysis of the airfoil mean secticn shows that the limiting location of the 

1990 airfoil design is at the trailing edge and the creep distribution at the 

design creep level as shown in Figure 5-48. 

86 

r r£) 

6.· _________________________ ~." .. ____ ,~::._'~~~;~'.- . ______ ~ ____ .. 
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o 10 

!O 04) 

o 3 

101: 

o 130 0043 

10 0511 10 01 71 

--1_ + I I 1 I '! I; 

ALL DIMENSIONS ARE IN eM IINI! t 

Figure 5-43 Blade Trailing Edge Cooling Configuration (Representative Geometry) 

b h 

C" ________________ __ _ -...11.... ___ • _ ..... n ______ _ 

87 

1.~

'-
04 

WC/A • 1% W •• 

TEOAP ~ 0 068 CM 10 028 tNt 

03 

Figure 5-44 Blade Trailing Edge Cooling Losses vs Blor.kage 

88 

-
414 

1601 

345 

1501 

TURBINE SPEW 10 146 RPM IAVERAGl DETfRIORA fEDI 

.-1988 TECHNOLOGY 

INO TaC) 

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fi4 

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o 

101~--~----~--~.---~ ____ ~ __ ~ ____ ~ __ ..... ~ __ ~ ..... __ .J 

o 40 80 90 100 

PEACl:N1 SPAN 

Figure 5-45 Blade Average Pull Stress vs Span fc 

(Average Deterioration) ,", ne ~~eed • 10,146 RPM 

f 

r 

t+) 

~ -:'::-;0 ""..,,-r;. ' .. 6_________________ ..... ____ . __ .u •. ___ .. __ 

89 

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90 

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- 0.3 

0100130.15018020 0230.25028 

(0041 10051 (0061 10 071 (0.081 (0091 (0 i 01 (0 111 

TRAILING EDGE THICKNESS. CM (IN) 

Figure 5-46 First Blade Trailing Edge Thickness Efficiency Sensitivity 

r£> 

t - ~~"?I'~~.'p~' 

.. -... -----------------------......... . - -. -- - -- ------
,. 

L 

, 

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rr 

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Figure 5-47 1990 Cooling Flow Distributio~ 

0.82% 

91 

~ .~

v:
18% 

1166K (163BOFI 

_2% 

Figure 5-48 Blade Creep at Critical Locations (1990 Technology) 

f 

[C' 

r 

r 

92 

® 

~ _____________________ ..L...A-____ ...... -______ ~~.-~~.- _ __ _ -' _~~ _____ ~ __ _ 
I 

f 

t 

. - r 

= 

The interface temperatures on both the suction and pressure surfaces (Figures 5-49 and 5-50, respectively) are significantly below the design allowable 

limit thus providing a margin of safety from coating spalling. A1rfoil metal 

temperatures at m1dsectfon are shown in Figure 5-51. 

Showerhead holes and film holes are sized to provide adequate cooling while 

maintainfng at least a minimum internal to external pressure ratio of 1.1 to 

prevent backflo~. Final hole sizes for the 1990 blade are shown in Figure 5-52. The resulting supply pressure required for the blade 15 68 percent of 

PT3 (Ps~pply/P3 • 0.68) (compressor discharge pressure), set by the 

m1dchord region (Figure 5-53). 

The trailing edge cooling airflow requirement of 0.82 percent Wae results 1n 

a design pedestal blockage of 77 percent. The geometry of the trailing edge 

was chosen to minimize momentum mixing losses by maximizing coolant exit 

velocities. Pedestal blockage in the trailing Podge is maximized with the 

restriction that the supply pressure requirement of the trailing edge circuit 

does not exceed the supply pressure requirements of the limiting midchord 

serpentine as depict!d in Fig~re 5-53. Present airfoils are limited by casting restrictions to 50 percent blockage or a 2 diameter pedestal spacing. Advanced 

castings can attain greater blockages by uti1izinp oblong pedestals with a 

minimum of 0.102 cm (0.040 in) of spanwfse openin(i between adjacent pedestals. As can be seen from Figure 5-53, this represents a significant reduction in 

cooli n9 10n. A practi cal 1 imi t of 80 percent blockage is set due to a t'isk of 

temperature mal distribution in the wake of the last pedestal row. 

Final design of the blade would require consideration of spanwise cooling requirements in addition to these requirements at the mean section. 

1986 Technology Blade Design 

The 1986 technology blade design differs trom the 1990 technology design in 

that it does not utilize thermal barrier coating. Additionally, less advanced 

burner technology in 1986 and increased first-stage vane and blade flow result 

in a peak gas temp~rature of 23.3 K (42-F) hotter than the 1990 technology design. In order to obtain the design life goal of 10,000 hours, B-5 UER 

(unscheduled engine removal), the 1986 deSign requires 3.8 percent Wae 

cooling air, with 1.1 percent at the leading edge, 1.4 percent at the 

midchord, and 1.3 percent at the tra"ling edge. The resulting metal 

temperatures at this mean cross section are shown in Figure 5-54, and the 

surface ~tal temperatures are presented ir. Fi gures 5-55 and 5-56. Creep in 

the airfoil is the life limiting failure mode with a creep distribution ~t 

failure as shown in Figure 5-57. 

Showerhead holes and film holes are sized to provide adequate cooling while 

again maintaining at least a minimum internal to external pressure ratio of l.~ to prevent backf10w. The cooling airflows and hole geometry for the 1986 

blade are shown in Figures 5-58 and 5-59. The resulting supply pressure requirement for the blade is 77 percent PT3. Here, as in the 1990 design, spanwise tailoring of the film has not been considered but sho~ld be 

considered during final detail design analysis. 

.. - -

93 

·-~

\:
I 

r 

r 

r 

t 

f' 

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94 

1256 

(1800) 

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1233 

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UJ 

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cr (1660) UJ 

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(1640) 

1156 

(1620) 

1144 

(1600) 

1133 

(1580) 0 

(0) 

• LE 

Figure 5-49 

(04) 

DESIGN LIMIT - 12781( (1840UF) 

(1.2) (24) (2 8) 

8 1 

(32) 

Sue nON SURFACE DIST ANCE.-CM (IN)------

Bl~de Interface Temperatur0 Distribution on Suction Surface (1990 Technology) 

! 

r.~ _____________ ..&-- ____ ... _. _____ _ 

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DESIGN LIMIT = 1278K 11840 0 F) 

1211 

117201 

1200 

(1700) 

\ 

1189 

(1680) 

1178 

(1660) 

1167 

11640! 

1156 

(1620) 

1144 

(1600) 

1133 

11580) 

I I 

0.5 10 1,5 20 25 3,0 35 4,1 46 51 5.6 61 6.6 71 

(02) 10,4) 10.6) (0.8) 11.0) (12) (14) (16) 11.8) (20) 1221 (24) 1261 128) 

1122 ! 

(1560)0 

(0) 

-=---=:-=--:- ---------. PRESSURE SURFACE DISTANCE. CM (IN) 

Figure 5-50 Blade Interface Temperdture Distribution on Pressure Surface 

(1990 Technology) 

95 

® 

I!- _ ~~ • ..,.--- -=---------------------""----- 

........ ..-. 
I 

rr 

f 

.. 

j 

t-

• 

r" 

r 

Figure 5-51 

96 

Blade Mean Section Isothermals Based on 1990 Technology (All temperatures 1°F are below Design Limit) 

S 

~~i:rr _______________ "",-- ____ ... ____ _ 
r 

f 

j. 

. 

~ 

r· 

r' 

S 

DIAME'-ER PITCH 

GROUP 

- CM IINI CM IINI 

1 0076100:;.J) 005010 177) 

2 0064 10025) 025410 100) 

3 005 ' 10020) o 12710050) 

NOTE. 

1 SHOWERHEAD HOLES ARE 25' RELATIVE TO RADIAL 

2. SUCTION SIDE HOtES ARE SHAPED 

Figure 5-52 Blade Hole Geometry (1990 Technology) 

• ~ _________________ .. _____ .010.-.- ___ ."u ••. _______ .. 

97 

@ 
98 

-

90 

i IJC SUPPl y PR~SSUR~ 

, RfOUIAEO FOR 

# MIOCHOAO COOLING 

u.i 70 

-~- -1. a:: 

::::> -- - - ~ 80 , w 

f 

>- ~I ( 50 rRAILING EDGE 

1,,) -" !I SUPPl Y PRESSuRE - I 

~I REOUIREMENT 

I 40 

;1 I ~ fl I DESIGN 

BlOCKAG~ J..-PRACTICAl LIMIT 

.04 1 

*-

to 3 III 

It 

;:, 

~ 

I::' 

.., to :2 

en 

en 

C,.) 

..J + 0.1 

(,!) 

~ 

X 

~ 0 

:2 

::> 

..... -01 Z 

UJ 

:2 

0 

:2 - 0 :2 

- 03 

Figure 5-53 

c. 

50 70 80 90 

PEDESTAL DENSITY, % 

Blade Trailing Edge and Midchord Cooling Air Optimization (1990 Technology) 

~_ ft ___ .," el 

-- ------------

lfi

~
I 

f 

r 

. 

l 

~ 

~ 

f 

f 

r 

Figure 5-54 Blade Mean Section Isotherma1s Based on 1986 Technology (All temperatures 1°F are below Design Limit) 

99 

B 

-

L' 

~ _____ .. __________ --...11-. __ ~_, .... _____ - -_ .. .

-

;~~

\;;
100 

u.. 

o 

1278 

118401 

12!i6 

118001 

1233 

~ 117601 

UJ 

ex: 

:::> 

r-

;:. 1 2 1 1 

~ ,IUOI 

~ 

w 

rUJ 11 B~ 

';i I 16!JOI 

u.. 

ex: 

:::> 

<.f) 

1167 

116401 

1144 

PRESSUR~ SURFACE 

v 

11600) '------';::' ---....... --~~---..L...-__ .J.... ___ .L __ -..J 

o 10 30 5 1 

1201 

Figure 5··55 

10) 104) 11 2) 

L~ PRESSURE SURFACE DISTANCE, CM (IN) 

7 1 

1< HI 

T£ 

Blade Surface Temperature Distribution on Pressur~ Surface (1986 

Technology) 

,.

\::
1 JOQ 

I 1 tl tll! I 

I}1H 

11tl401 

L4: 

0 

1 }',I) 

~ I 11ll!l!1 

w 

ex: 

~ 

~ 1233 

~ 117hOl ex: 

w 

0.. 

~ w 1:7 11 

~ 117201 W 

U 

ct 

L.I.. 1189 ex: 

~ 116801 

~" 

1167 

116401 

1144 

(16001 

0 

101 

,.. Figure 5-56 

.. 

® 

~ 

1 0 

1041 

30 

(1 21 

SUCTION SURFACE 

5 1 

(2 01 

SURFACE DISTANCE. CM (INI 

7 1 

12 81 

9 ' 

13 6) 

Blade Surface Temperature Distribution on Suction Surface (1986 Techno logy) 

101 

------------------------- -- 
o 7!1'1j, 

TM - 1160K 1162!!"FI 

096% 

Figure 5-57 Blade Creep at Critical Locations (1986 Technology) 

.. 

f 

f' 

~. ~ 102 

r 

! 

j.. ~~-..;p~ •• iIIb __________________ ito.--~_~.~_~~ __ H • h' ___ •• _ ___ _ __ ~ 
~p t p-O.13~ 

WC/A-3.n~ 

O.4&~ 

1.27% 

Figure 5-58 Cooling Flow Distribution (1986 Technology) 

103 

... ___________________ .,u_u. ___ __ ._ 
104 

DIAMETER PITCH 

GROUP CMIIN) CMIIN) - 1 007610030) 045010 - 177) '} 006410025) 005610 100) 3 007610030) o 19010075) 

NOTE 

1 SHOWERHEAD HOLES ARE 25" RELATIVE TO RADIAL 

2 SUCTION SIDE HOLES ARE SHAPED 

Figure 5-59 Blade Film Hole Geometry (1986 Technology) 

~_...;;;."....:. _________ ._.. _ ••• u .... __ 

.~

'=
SECTION 6.0 

BENEFITS EVALUATION 

6.1 Benefits for Supersonic Aircraft 

The high-pressure turbine for the supersonic cruise aircraft engine represents the incorporation of considerable experience and anticipated technology. Tt.e 

predicted efficier.cy has exceeded the original most optimistic goals by 0.3 

percent to provide a 92.3 percent cooled efficiency. This level was obtained 

without precoolers and with advanced thermal barrier coating, film cooling, and highly augmented internal cooling. The total cooling air f~r vane and 

blade was 5.28 percent of Wae and 2.77 percent of Wae , respectively (Table 6-1). The 1986 design, with less advanced technology, has a lower cooled 

efficiency of 91.9 percent, and requires 8.66 percent of Wae cooling air for 

the vane and 3.77 percent W,e cooling air for the blade. Both designs (1990 

and 1986) have 10,000 hour lives. 

Performance Benefit 

TABLE 6-1 

PRELIMINARY DESIGN CONCLUSIONS 

o 1990 Design ry = 92.3%, 10,000 Hours -- 5.28% Wae Vane -- 2.77% Wae Blade 

o 1986 Design ry .. 91.9%, 10,000 Hours 8.66% IJae Vane -- 3.77% Wae Blade 

6.2 Benefits for Subsonic Aircraft 

The advanced technology concepts developed in the high-pressure turbine of the 

supersonic cruise VSCE also are applicable to other advanced engines. A ~tudy 

based on a commercial subsonic transport engine in the 267,000 N (60,000 lb) 

thrust class was used to evaluate the overall impact of VSCE technology in 

subsonic engines. The study engine has a 38:1 overali pressllre ratio and a 

two-stage high-pressure turbine, considered repres~ntative of general ~ngine 

technologies of the 1990's for subsonic aircraft. 

The VSCE technologies outlined In Table 6-11, along with the 1963 

technologies, were applied to the '.;ommercial subsonic transp'lrt engine study. The overall benefits of these tc<.;nrlologies were determined to be 2.68 percent in high-pressure turbine efficiency and 1.94 percent in thrust specific fuel 

consumption. Table 6-111 shows the detailed breakdown of predicted benefits by individual technology. 

- - *""'--- ---... ... __ .. _~_~ __ - __ A __________________ _ 

~05 

.~

\:
I 

f 

f 

r 

~ r . 

r 

TABLE 6-II 

1990 TECHNOLOGY FEATURES COMPARE~ TO CURRENT SUBSONIC ENGINE 

Aerodynami c Technology 

Load Coefficient 

Profl1 e Loss 

End Wall Loss 

Tral1i ng ldge Thickness: 

Fi rst Vane 

First Blade 

Second Vane 

Second Blade 

AN2 

Bl ade Materi al 

Mechanical Technology 

Burner Technology 

TVR (l V/2V) 

Profile Factor (lB/2B) 

Vane Base t1ateri al (1 V/2V) 

Thermal Barrier Coating: Vanes 

Blades 

Improved Cooling Fi rst Vane 

Fi rst B'I ade 

Second Vane/Second Blade 

106 

1983 (Base) 

0.198 cm (0.078 in) 

0.198 em (0.078 in) 

206 em (0.081 in) 

0.183 em (0.072 in) 

4.2 x 1010 

P!iA 1480 

1. 4/1. 35 

PWA 647/PWA 1422 

None 

None 

1990 (Design) 

+17':, 

-10':, 

-15':, 

0.155 em (0.061 

0.155 em (0.061 

0.155 em (0.061 

0.155 cm (0.061 

6.0 x 1010 

SC3000 

1.25/1.2 

-4':, 

SC1000 

in) 

in) 

in) 

in) 

0.038 cm (0.015 in) Semi 

Opaque 

0.025 cm (0.010 in) Yttria 

Stabil i zed 

Flared Holes 

Flared Holes/Low Heat Load 

Low Heat Load 

..... ~ ______ .... ______________ _.o.iio_....a..-, ____ .. ......... -_______________ """"-~.,, ~ _______ _ 
I 

, 

r ." 

, 

.. 

f 

r' 

r 

TABLE 6-111 

1990 TECHNOLOGY - SUBSONIC APPLICATION BENEFITS 

AERODYNAMIC TECHNOLOGY 

4WC/A ~ 4TSFt • LOAD COEFFICIENT .17~} I" WAE) 

PftOFIl,! LOSS - 10,. +0.580" END WAll LOSS -15,. -0.310" 

• TAAlUNQ EDGE THICkNESS REDUCTION +0.140" -0.078" • AW-14.2x'0'O _e x '0'0 MAX) - AERO + 1.220" -0.lIII0'' • AERO COOUNQ (WfO SC3000IW SClOOO' 1+ 2.24"'.0.38,., 1-0.35,.'-0.053,., 1 + 0.118,., + 0.1,., TOTAL AN2 

1 + 2.24,., +0.38'" 1.0.87"'.'.17,., 1-0.02,.'-0.51,., • TOTAL AERO TtCHNOLOGY IREFLICTS SClOOOI +0.38" -1.880,. -0.870" 

MECHANICAL TECHNOLOGY 

• BURNER TECHNOLOGY -2.M" +0.330" -0.330" • VANE BASE MATERIAL 

IV IPINA 847 .... SC10001 

2V IPINA 1442 -1.06" +0.130" - .,.0911,. .... SC10001 -0.35" +0.025" -0.034" TOTAL BAtii! MATERIAl -1.40" -

+O.IM" -0.130" • THERMAL BARRIER COATING 

IV - 2.115" +0.320" lB - 0.240" -0.74" -0.010" 2V -0.0911" -11.011" 2B 

~ 

-0.0112" +0.030" 

-0.025" -0.020" - - TOTAL TaC -3.51" +0.223" -0.330" • IMPftOVED COOUNG 

1 V FLARED HOLES '-0.38" +0.062" 1 B FLARED HOLlSIlOW HEAT LOAD -0.037" -0.110" ·0.024,. 2V LOW HEAT LOAD -0.083" -0.07" +-O.natI,. 2B LOW HEAT LOAD -0.007" -0.11 ,. +0.035" -0.053" - - TOT At IMPROVED COO'~ING -LOll'll. +0.1111" -0.1110" • TOTAL MECHANIr.',AL TECHNOLOGY -8.59% +0.820" -0.870" 

OVERALL BENEFIT -8.21 ,. + 2.lIII0'' -1.940" 

107 

® 

F~ ____________ ~ ___ .......... _» ----~~~.,.:,. .. -
t. 

r 

, 

f 

e 

, 

~ 

~ 

t 

~ 

~. 

r 

r£ 

~ . , 

As shown in Table 6-111, turbine aerodynamic 1mp~ovements provide 1.86 percent of the efficiency increase and 0.97 per~ent of the thrust specific fuel 

consumpt,i')n decrease, wit~ increased AN contributing a major portion of the 

benefits. However, the AN benefit obtained from the use of advanced 

materials is partially offset by the increased coolant flow requirement due to 

higher AN2. With current state-of-the-art mate~ials, the proportionately higher coolant flow increases due to higher AN would offset. to a larger degree, the perfcnllance gains which would be otherwise at:hieved. 

Figure 6-1 illustrates the net fuel cl.,nsumrti~n improvement due to AN2 using 

current and advanced materials. 

Mechanical technologies extension beyond that of the 1983 base further 

improves engine performance by 0.82 percent and reduces fuel consumption by 0.97 percent. Individual contributions due to burner technology, vane 

materials, thermal barrier ,oatings (TBC), and cooling technology are also 

listed in Table 6-111. 

The T~C benefits are greatest where the external airfoil heat loads are 

highest, generally in upst~eam airfoils. The net benefits attributed to the 

TBC are negated by the 1!"!cr~ased pull 10,lds due to TBC on the blades and a 

maximum temperature limit of bond coating temperature for coating spalling on 

blades and vanes. Consequently, no net benefit is realized on the second vane. 

TBC probably wov~d not be used on such a second vane unless dictated by new 

thermal problems identified during a final detailed design. 

In summary, the advanced technologies utilized for the VSCE supersonic cruis~ 

application offer substantial potential for both subsonic and supersonic aircraft engines and warrant pursuing. 

108 

- -=-.. .' .>;.. ... 

tr 

__________________ --..110..."'"-___ ., ..... ______ _ 
I 

r 

" 

! 

. , 

f' 

rr 

~ ~ 

~ 

Z 

w 

:I 

w 

> 0 

cr: 

Q. 

~ 

A TSFC, % 

0.7 

-0.8 

0.5 

·-04 

-0.3 

-0.2 

-01 

0 --

0.1 

02 

0.3 

30 

Fi gure 6-1 

BASE 19110 LEVEL 

- --~~-SAJE 1983 MATERIALS 

I 

40 5.0 60" 

'°'0 

AN2 (SECOND STAGE) 

A TSr:C vs AN2 

____________________ ... ___ _ •• 'u .. _________ .. 

109 

--.~

~
~ f; 

rl 

f 

, 

I 

SECTION 7.0 

CONCLUSIONS 

The high-pressure turbine preliminary design conducted under this program shows that it is feasible to obtain a 10.000 hour airfoil life and a 92 

percent cooled efficiency in the supersonic cruise engine environment. 

However. incorpor«tion of technology advances relative to the p,-esent state-of-the-art turbine practices is critical. The improved technologies include advanced aerodynamics. improved airfoil materials. advanced thermal 

barrier coatings. new burner technology and improved cooling/casting methods. 

These advanced technologies also have a significant payoff in subsonic 

engines. offering almost a 2 percent improvement of fuel consumption. A 

breakdown of the improvements by individual technology is summarized in Table 7-1 below. 

Program 

Ar.2 

Airfoil Aerodynamics Thermal Barrier 

Improved Coo 11 ng Vane Material 

Trailing Edge 

Burner Technology 

TABLE 7-1 

CONCLUSIONS 

Thrust Specific 

Fuel Consumpt·: oJ! 

-0.58% 

-0.31% 

-0.33% 

-0.18% 

-0.13% 

-0.00% 

-0.33% 

-1. 94<,;:. 

In order to make possible an advanced supersonic transport engine by the 1990s 

or to achieve the 2 percent fuel consumption reduction in subsonic engines. it 

is imperative that technology development programs be initiated now. These 

programs would allow several years of research and development of these new techno 1 ogi es. 

Specific programs should be directed at m~nufacturability of airfoils 

cot'!tai ni ng advanced cool i ng concepts. These ai rfoil manufacturabfl ity efforts ~nou1d investigate methods of casting complex internal cooling concepts. 

controll ing wall thicknes,es. forming diffuser shaped holes in thermal barrier 

coatings. and reducing the thickness of cooling passages for trailing edges of airfoils. 

110 

r., _ ~*-..<!flj-" ." .... 

------...--------...&---  .....  ... - .~.---. 
'. 

, 

~c 

SECTION 8.0 

RE COt+1E NOAT .IONS 

The preliminary design of the high-pressure turbine for the second generation supersonic cruise aircraft engine has reinforced the conclusion of previous 

studies that high-pressure turbine technology is crucial. 

This preliminary design study was confined to the high-pressure turbine. A 

detail design should be started to examine more completely the interactions 

with other components. A large benefit has been obtained by increasin~ the 

turbine annulus area and reducing its speed. Preliminary examination in~icates 

these modifications to be acceptable but, their impact should b~ pursue~ to 

the level of emphasiS in a detailed design. 

In addition, a level of manufacturing technology was assumed whic~ is beyond the current state-of-the-art. Ongoing technology programs will prc''!ide some of 

these advances, but specifiC programs must be conducted for fc1IJC1flg traiiing edge thickness, controlling wall thicknesses, casting internal complex coo1in~ 

surfaces, and producing shaped holes through thermal barrier systems. 

lhe application of this technology to subsonic commercial engines demonstrated 

this technology to be germane to all advanced engfne designs. The initiation 

of the technology programs for high stress turbines (AN2), advanced airfoil 

aerodynamiCS, thermal barrier systems, improved airfoil cooling techniques, 

and burner technology in a timely fashion will produce broad benefits in all 

future engine applications. 

111 
APPENDI X - HIGH TURB INE AIRFOIL COORDINATES 

A. BASE BLADE 

X em y TOP em ELLIPSE em y BOT em ELLIPSE em 

0.0 1.7221111 1.1171135 1.4911114 1.1171135 

0.011lZ4 1.75546 1.751011 1.514114 1.112571 

0.01111111 1.71151S 1.7114911 1 .529911 1.b11116 

0.05472 1.111214 1.54463 1.1105011 

0.072911 I.U611. 1.5511110 1.110220 

0.09120 1.115956 1.571116 1.110207 

0.10944 1.1111045 1.51144Z 1.11010010 

0.12ibll 1.89971 1.591125 l.b07711 

0.14592 1.91749 1.60736 1.61l00 

0.1610111 1.91390 1.61 77Z 1.1119112 

0.111240 1.94905 1.112734 I. b2751 

0.22&00 1.911195 1.1110806 

0.2n60 Z.00851 1.1163911 

0.119Z0 Z.OZ949 1.67494 

0.364110 Z.04544 1.611095 

0.41040 2.05675 1.68195 

o 105600 2.01117Z 1.67794 

0.50160 2.06659 l.b61S94 

0.510720 2.06552 1.65500 

0.59280 2.06061 1.11111 19 

0.113840 2.05192 1.b1262 

0.1111400 2.03949 1.56440 

0.72960 2.0ZlZ;; 1.551611 

0.77520 2.00321 1.51462 

0.82080 1.97932 1.107338 

0.8b6100 1.95129 1.42013 

0.91200 1. 9111911 1.37907 

I 0.957110 I .882010 1. 3211 17 

1.00320 1.8100110 1. '7':.! 1 

r1.0108110 1.79274 1. zo083 

1.0910100 1.73912 I. 1"~31o 

1.110000 1.117829 1. 062911 

1.165110 1.1109211 1.01434 

1. 231 ZO 1.53\ '5l 0.94415 

~ 1.27:'80 1."'.440 0.67105 

t 1.3Z240 1.34895 0.795119 

1.311800 1.24S93 0.71620 

1.1013110 1.1311100 0.113873 

1.45920 1.0ZI77 0.55739 

I .504110 0.902113 0.47431 

1.55040 0.711052 0.38959 

1.51/bOO 0.tS554 0.3CJ33 

1.641110 0.!:t842 0.21Sb3 

1.('5~84 0.47707 0.16017 

1.67805 0.4ZS48 0.14450 

1.69632 0.37365 0.10863 

1.71456 0.321111 0.07255 

1.732110 0.269.58 0.03626 

1.75104 0.21697 -0.00018 

1.769ZS 0.111439 -0.0368<' -0.02795 

1.78752 0.11165 -0.073(,5 -0.03444 

1.11057" 0.058711 -0.11065 -0.03043 

1.82400 0.00576 -0.00000 -0.14783 -0.00000 

112 

-~ 

~
I 

r ,-

X em 

0.01721(-

0.011240 

0.0'6ltI0 

0.054720 

0.072~60 

0.091200 

0.10~4ltO 

0.127610 

o.lltSnO 

0.16ltl"0 

0.182400 

0.2Z1000 

0.273600 

0.119200 

0.3"4800 

0.410400 

0.456000 

0.501600 

0.547200 

0.592800 

0.618400 

0.684000 

0.729600 

0.775200 

0.820800 

0.866400 

0.912000 

0.957600 

I. OOlZOO 

1.048800 

1.094199 

1.119999 

1.185599 

1.211200 

1.276800 

1.322400 

1.368000 

I. 4 n600 

1.459200 

1.504800 

1.550400 

1.5'16000 

1.641600 

1.659840 

1."78080 

1.696320 

I. 714560 

1.732800 

l. :51040 

1.769279 

I. 787519 

1.805759 

1.8Z4000 

HIGH TURBINE AIRFOIL COORDINATES 

B. LOW HEAT LOAD BLADE 

·4 

Y TOP em 

1.6U7Zf> 

I.U5571 

1.611741 

1.604946 

1.60209Z 

1.601968 

1.601953 

1.607686 

1.612938 

1.619562 

1.627467 

1.648063 

1.663960 

1.674945 

1.680954 

1.681954 

1.677~·'] 

1.668 I 

1. 6549~7 

1.636189 

1.61U16 

I. 587000 

1.555000 

1.525000 

1.4eOOO 

1.44000 

1.3900 

1.32000 

I.Z55 

1.178259 

1.101506 

I. 024752 

0.9479'18 

0.871244 

0.794490 

0.717736 

0.64098 

0.564229 

0.487475 

0.410721 

0.333'167 

0.257213 

0.180459 

0.149757 

0.119056 

0.088354 

0.057653 

0.02692 

-0.004000 

-0.027954 

-0.034443 

-0.030428 

0.0 

Y BOT em 

1.62.,2b 

1.750953 

1.784879 

1.8121lt] 

1.81685'1 

1 .85955 

1 . 880ltlt6 

1.899707 

1.91748b 

1.93390Z 

1.948053 

1.977949 

2.00151lt 

2.018lt94 

2.0314] 

2.03603 

2.03926 

Z.0\90b 

2.03481 

Z.OZ719 

Z.016b8 

1.00l1Z 

1.9854b 

1.96n8 

1.93G8Z 

1.90898 

1.87079 

1.83161 

1.78632 

1.733n 

1.67327 

1.61500 

1.54648 

1.47212 

1.39165 

1.30480 

1. :!l1Z8 

1.11195 

1.00570 

0.89338 

0.71584 

0.1>5391 

0.52843 

0.47743 

0.421>05 

0.37431 

0.32226 

0.26994 

0.21740 

0.16466 

0.11179 

0.05681 

0.0 

113 
HIGH TURBINE AIRFOIL COORDINATES 

C. VANE COOR 01 NATES 

X em Y TOP em ELLIPSE em y BOT em ELLIPSE em 

0.0 3 63289 3.50924 ) Z71 n 1 'ion<o 

0.02097 ).64927 3 59959 ).2770) ) <OZ101 

0.04194 ).USI4 1./»)/)00 ) 28222 J.1MbO 

0.OU91 3.U050 ] bblZl 1 2851011 1 3bl00 

0.00300 3."509 1 baSH ) 287<01 1 li, 171 

a .IOlt85 3.70'117 1.701012 ).2111100 ) 32!>0? 

0.IU02 ].7ZZ4C ).72041 ).Z8120 1 310'14 

0.14'79 3. n5Z& ) HIo74 ),2851'1 1 2 ')6 75 

0.1677' 3.74731 ).747101> ) 1817'1 3 26111'1 

o .Ioon 3 75803 ~. Z770' 3 27'100 

O.toCno 3.7"13 ).27102 3.:7103 

0.ZU12 3.7UIl 1. UOH 

0.311055 ).01211 3.22148 

O.lUH ).02710 3.18507 

O.lt19ltO 3.0371' ).14Ill 

0.117102 3.04228 3.090/)0 

O.52lt25 3.84254 3.oBn 

0.57U7 ).037

tI2 '.'1b'l?1 

0.6"10 3. au. 'IZ :.'1001lO 

0.'8152 3.01080 2 OU17 

0.73395 3.78On 2.74707 

Q 7M}? 3 l'-~bZ 2.663~f. 

o .83MO 3 n·.58 2 575n 

0.89122 1.685a7 Z.40J5? 

0.'141b5 1.61aZ6 2.18041 

0.99,o7 1.58371 2 19(103 

I 

1.04050 1.5H05 2.1111176 

1.10092 1.4SZ92 2.0tllt8Z 

r 

1.15335 1.3756) I. 'J7M7 

1.20577 3.290ltO 1.869a7 

1.25820 1.19b6Z 1.75925 

1.11062 3.09387 l.b

t

.b7b 

l.lb305 2.901H I 51Z51 

f 1.41547 Z.05H1 I. ttl b 7Z 

r 1.4b790 2.12780 I. ''1'146 

1.520'32 2.51110711 1.leOM 

I. 57Z75 2.41025 I.Obl06 

1.62517 2.Z'1105 o 9',0 II 

1.'77'0 2.0034' i).elMO 

1.73002 1.&&911t 0.69507 

1.702lt5 1.67995 0.57117 

1.01407 1.45391 0.44641 

I.oono 1.20955 0.32089 

f 1.90027 1.IOb27 O. ?JO

t,7 

r 1.92924 0.99962 0.Z19'14 

r 

I. '15021 0.00942 O.lb'l]O 

1 .97110 0.77552 0.11055 

, 1.99215 0.65776 0.067''1 

Z.01312 0.53500 0.01674 

Z. 03409 0.40970 -0.03432 -0.02b43 

2.05506 0.27897 -0.08547 -0.0!bb8 

2.07b03 0.14319 

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27 

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0.00. 

0.005 

0.001> 

0.007 

0.006 

0.009 

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0.014 

0.016 

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0.022 

0.024 

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0.029 

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0.037 

0.041 

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0.073 

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0.123 

0.136 

0.153 

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0.184 

0.::01 

0.216 

0.219 

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(ft/sec) 

UGIHI 

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73.100 

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102.92'1 

115.40? 

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130.05:: 

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132.243 

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124.557 

121.562 

119. '145 

118.316 

116.667 

114.55'. 

111. ~eb 

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130.611 

146.578 

164.145 

200.'164 

250.846 

316.365 

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1011.181 

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42.964 0.004 

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69.715 0.006 

79.483 0.007 

88.724 0.008 

90.593 0.010 

11l0.063 0.012 

102.929 il.014 

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109.347 

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103.879 0.021 

913.sn 0.024 

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95.062 0.031 

97.100 0.035 

10!.87? 0.039 

110.230 0.Ot.4 

127.803 0.048 

142.902 0.057 

160.545 0.068 

1n.796 0.080 

260.965 0.094 

347.664 0.109 

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0.140 627.717 

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1108.487 0.186 

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15 

16 

17 

16 

19 

20 

21 

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23 

24 

25 

26 

27 

28 

29 

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32 

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36 

37 

38 

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42 

43 

44 

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0.274 

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(ft/sec) 

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34.146 

83.656 

113.283 

13".631 

156.550 

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291.066 

338.951 

468.891 

570.720 

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804.174 

996.586 

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1366.364 

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2073.163 

2236.695 

2261.839 

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List of Symbols and Abbreviations 

best cruise altitude 

heat e)(changer carb,n mono)(lde 

a)(lal flow velocity 

diameter 

~nvlronmental Protection Agency fnvlronmental Protection Agency Parameter 

film effectiveness 

net thrust 

specific work 

Inner diameter 

I nve;'ted Flow Eng I ne 

Low Bypass Engine 

leading edge Mach number 

mechanical speed. rpm Ndtlonal Aeronautics and Space Administration 

oxides of nitrogen 

outer diameter 

pressure 

static pressure 

total pressure 

pressure ratio 

Pratt & Whitney 

revolutions per ~Inute 

single crystal Supersonic Cruise Aircraft Researrh 

showerhead 

temperature 

ambient temperature 

thermal barrier coating 

cooling temperature trailing edge 

gas temperature 

metal temperature 

tangential on-board Injection thrust specific fuel consumptIon 

temperature variance ratio. pattern fbctor 

tangential wheel speed 

unscheduled engine removal 

velocity 

Variable Cycle Engine 

Variable Stream Control Engine 

total cooling airflow 

total engine airflow 

difference 

efficiency 

overall ccoling effectIveness 

heat load parameter 

stress 

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119 
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1. 

2. 

3. 

4. 

REFERENCES 

Hunt rLB •• et al: IINoise and Economic Study fllr Supersonic Cruise Airplane Research - Phase 11111. P&WA Final Report. NASA CR-165612, March 19b2 

Howlett. R.A .• Hunt R.B .• :IIVSCE Technology Definition Study". P&WA Final Report. NASA CR-159730, Aug 197~ 

Hunt R.B .• et al: IINoise and Economic Study for Supersonic Cruise AirplanE Researchll. P&WA Final Report. NASA CR-165423. June 1981 

Howlett. R.,I\ .• et al: IIAdvanced Sup~rsonic Propulsion Study _ Phase III", P&WA Final Repvit. NASA CR-13514~. December 1976 

5. Nicholson. T.H. et al. IIHeat Transfer Optimized Turbine Rotor Blades _ An 

Experimental Study Using Trans;ent Techniques. II ASME 82-GT-304 

120 

L 

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