Propulsion Study for Small Transport Aircraft Technology (STAT)

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F A , -, u ODA EDR 10470

NASA-CR 165499

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Propulsion Study for Small Transport

Aircraft Technology (STAT)

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FINAL REPORT

(NASA -CR- 16545'0 P&OPU1516b STUL'Y FOa SHALL

N82-1u.i37 T&ANSPURT ASaCRAF T T1CHbG&,:GY (SPAT) Contractor Final uepert (Detzoit lliessi Allison, ludiaAaPolis, iud.) 1d6 P Uuclas

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CSCL 21d GJ/J7 27703

Detroit Diesel Allison

Qirision of General Motors

Indianapolis, IN 46206

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prepared for ' L'

,L AERONAUTICS AND SPACE ADMINISTRATION

NASA Lewis Research Center

Contract NAS3-21995

L^ 3. Recipient's Catelog No.

1. Report No. 2. Governnant Accession NO.

CR-165499 s. Re p ort Dote 4. Tine and Subtitle December 16, 1980 ' Propulsion Study for Small Transport 8. Performing Organiasnon Cods Aircraft Technology ( STAT) 8. Performing Orpsnintion Report No.

7. Autferts) J. C. Gill, R. V. Earls, D. V. Seaton, A. C. Stolp, DDA EDR 10470 D. S. Hualster, B. A. 2olazsi 10. Work Unit No.

Performing Cvysn imtion Nam. and Addrae f.

Detroit Diesel Allison 11 Conuaet or Grant No Division of General Motors Corporation NAS3-21995 Indianapolia, IN 46206 13. Type of Report and Period Covered 12. Sponsoring Aganey Nerve and Address Contractor Final Report NASA Lewis Research Center 14 Sponsoring Agency Cods 21000 Brookpark Road Cleveland, OH 44135 15. Sum Notes Project Manager: William C. Strack MS 501-10, NASA Lewis Research Center, 44135 Cleveland, OH 18. Abet act The objective of the Small Trans port Aircraft Technology (STAT) study was to define future research and advanced technology effort required for propulsion systems for the next generation of small, short-haul type commuter aircraft operational by 1990.

7 Mach aircraft based on studies conducted Propulsion requirements were determined for 0 . 5 and 0 .

Sensitivity studies were conducted on both these by NASA-Ames and Lockheed California Company.

aircraft to determine parametrically the influence of propulsion characteristics on aircraft size Candidate technology elements and design features were identified and direct operating cost (DOC).

and parametric studies conducted to select the STAT advanced engine cycle.

Trade-off studies were conducted to determine those advanced technologies and design features that operation of the STAT engines. These features were incorporated would offer a reduction in DOC fo :• A benefit assessment was conducted comparing the STAT engines to current in the two STAT engines.

technology engines of the same power and to 1985 derivatives of the current technology engines.

Research and development programs were recommended as part of an overall technology development plan to ensure that full commercial development of the STAT engines could be initiated in 1988.

17. Key Words (Suggested by Authorisi ) 18. Distribution Statement Small Transport Aircraft Technology ( STAT) Unclassified Advanced Turboprop 22. Price' 1B. Security Clamif. (of this report) 20. Security Clesif. (of this pope) 21. No. of Pages Unclassified Unclassified ' For sale by the National Technical information Service, Springfield, Virginia 22151 TABLE OF CONTENTS Title ?.So l Summary * .

. .

Introduction

Baseline Airplane and Missions .

.

Mission Requirements . 6 Baseline Airplanes .

10 Baseline Engines . . . . . . . . . .

13 Aircraft Sizing and Cost Model . . . . . . .

16 Overview of STAT Mission Program. . , •

16 Engine / Airframe Sizing. . . . . . • . .

21 ST AT Cost Model . . . . . . . . . • • •

Task I Mission Study Results. . • . . .

27 ters . Aircraft System Sensitivity to Engine Parame

. 28

DOC Breakdown . . . . . . . .

. 28 Propeller Sensitivity Data. . .

Advanced Technology Identification and Evaluation. 33

Cycle Selection. . . . •

Configuration Trades . . . 43

Compressors . . . . . . . . . . . . • . 45

Turbines. . . . . . . . . . . . • . . • 0

• . . . . • . • • • • • • • • • • • • • 71

Bearings. . . . . . . . .

Diffusers Combustors. . . • . . . . 72

/ a . . . . . 74 A ccessories . . . . . . . . . . . . .

Noise Reduction . . . .

Reduction Gear. . 77

Other . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

• • . . • . • . . • . . . • • . . 84

Advanced Technology Engines. . . . .

1790 kW (2400 shp) STAT Engine . . . . . . . . . . . . . . . . . . 88

( 4800 shp) STAT Engine . . . . . . . . . . . . . . . . . . 89

3579 kW

Maintainability . . . • . . . . . • . . • . , . . . . • . . . 92

• • . . . • . • . . . . . . . . 92

Modular Construction. . . . .

Engine Noise Considerations . . . . . 92

Compressor Noise. . . . . • • . . 92

Combustion Noise. . . . . . . . • 6 92

Meeting the STAT Goals . . . . . . . . . 93 Propulsion System Noise- -

Benefit Assessment

Sfc Trends. • 94

Price Trends. . . . . . 96

Maintenance Cost Trends 99

Reliability Assessment.

Maintenance Cost Projections. • . . • e 106

Weight Trends . . . . . . . . . . . . 0 110

• . . 111

Engine Technology Comparisons .

Mission Results . . • . . . . . . . . 0 112

i

r Page T itle Recommendations for Future Research. .

Program Content * • • . . . . • • .

. . . . 141 Basic Research and Development . .

^.

Compressors . .

Turbines * .

• t, Diffusers and Combustors * . .

Engine Accessories. . . . . . 160 .^ Noise Reduction . . . . . . • • • . • . • • . • . • 0 6 163 Reduction Gears . . . . . . • . . . • • • . . • • 166 . . • • • • • • • • • • • • • • . 166 Experimental Engine Program.

• • 170

Conclusions. • • • • • • • • • • • •

i

• 172 References a. .

List of Abbreviations and Symbols. . . • 173 I^ .t ii t LIST OF ILLUSTRATIONS Page Figure Title km (100 nm) 1 Typical DOC breakdown--high speed aircraft--185 . . . 2 block distance .

2 Mission requirements . . .

3 STAT propulsion study baseline aircraft 4 10 Baseline turboprop engine . . . . . .

. . . 12 5 Sfc versus rated power . . . .

6 Engine horsepower-to-weight ratio versus rated power 7 . . . . . . . . . . . 14 Engine length comparison . .

8 max envelope width . . . . . . . . . . . . . 15 Engine . 16 9 Engints max envelope height . . . . . . . . . . .

. . . . 17 10 Engine acquisition cost versus rated power . . . . 18 11 Engine maintenance cost versus rated power STAT mission program . . . . . . . . . . . . . . . 19 flow diagram . . . . . . . . 20 13 Engine/airframe sizing . . . . . . . . . . . . 24 14 Cruise altitude selection .

. . . . . . . . 25 15 Climb study results --LCC aircraft .

16 Climb study results--Ames aircraft . . . . . . . .

breakdown--high speed aircraft . . . . 31 17 Typical DOC breakdown--low speed aircraft 32 18 Typical DOC 19 Methodology using sensitivity data 20 Engine sfc sensitivity to component efficiency- -hign speed . . . . . . . . . . . . . 37 airplane, alternate mission to component efficiency --high speed 21 Engine sfc sensitivity . . . . . . .

airplane, design mission . . . . . . 38 Engine sfc sensitivity to turbine cooling air and buzner . . . . .

pressure drop--high speed airplane . . . . . 39 . . . . . . 41 23 Mission weighted sfc trends--1790 kW (2400 sh;)) size . . . . . .

24 Mission weighted aft trends--3579 kW (4800 shp) size 42 pressure ratio at 1506 K (2250°F) and 3579 kW 25 DOC trends with (4800 shp) conditions 26 STAT advanced compressor flow path--1790 kW (2400 shp) engine . .

-3579 kW (4800 shp) engine . . 47 27 STAT advanced compressor flow path- . . . 56 28 STAT 1790 kW (2400 shp) engine turbine flow path . . . . .

turbine flow path . . . . . . . . 57 29 STAT 3579 kW (4800 shp) engine 30 STAT radial inflow turbine 1790 kW (2400 shp) engine.

K 31 Airfoil impingement cooling . . . . . . 66 32 Active clearance control --turbines. .

33 STAT first-stage blade sensitivity of cooling flow to turbine temperature, pressure ratio, and cooling air temperature . .

cooling flow to life. . 70 34 STAT first-stage blade sensitivity of . .

in wheel weight . . . . . . 71 35 STAT wheel life versus percent change . 73 36 DDA diffusers . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . 73 37 Diffuser pressure loss results . . . . . . . . .

. . . . . . . . .

38 Transpiration-cooled Lamilloy . . . . . . . . . 74 Impact of compressor bleed and engine power extraction on . . . . . .

performance . . . . . . . . . . . . . . . . . . . 83 40 aircraft system parameters to nacelle weight . . . 84 Sensitivity of . . . .

41 Comparison of propeller and engine noise during takeoff 93 iii --- Figure Title Page I 42 Sfc trends and comparisons. 95 } 43 96 Power section price comparisons Engine specific price comparisons 45 107 if Engine maintenance cost versus rated power 46 STAT specific weight trends e . e e . . 110 DOC versus block distance- - high speed aircraft, fuel cost $0.264 / . . . . . . 117 .1 L ($1.00 / gal).

DOC versus block distance- - high speed aircraft, fuel cost

$ 0.396 / L ($1.50 / gal) . . . . . . . . . . . . . . . . . . . . . 118

49 DOC reductions- high speed aircraft . . . . . . . e 119 - 50 5-year TCO versus block distance- - high speed aircraft, fuel coat

$0.264 / L ($1.00 / gal). . . . . . . 120

51 5-year TCO versus block distance- - high speed aircraft, fuel cost

$0.396/L ($1.50 / gal) . . . . . . . . . . . . . . . . . . . . 121

52 122 5-year TCO reductions- - high speed aircraft.

53 5-year TCO savings- - high speed aircraft e . . e . . 123 =^ 54 Percent reduction in fuel consumed- - high speed aircraft . e . . 124 i 55 - . . . . . . . 125 10 year fuel savings- - high speed aircraft . . . . .

56 Block speed versus block distance- 126 - high speed aircraft 57 126 Fuel consumption comparison- - high speed aircraft 58 DOC comparison- - high speed aircraft . . . . . . . . . . . . . . . 127 59 DOC breakdown- - high speed aircraft . . . . . . . . . . . . a e e 128 60 DOC versus block distance- - low speed aircraft, fuel cost

$0.264/L ($1.00 / gal). e . . . . e . . . . . . . . . . e . 131

61 DOC versus block distance- - low speed aircraft, fuel cost

$0.396/L ($1.50/gal) . . . . . . . . . . . . . . . . . . . . . 132

62 DOC reductions- - low speed aircraft . . . . . . . a . . . . . . . 133 63 5-year TCO versus block distance- - low speed aircraft, fuel cost / / $0.264 L ($1.00 gal)* e e e e . . 134 64 5-year TCO versus block distance- - low speed aircraft, fuel cost

$0.396/1, ($1.50/gal) . . . . . . . . . . . . . . . . 135

a 65 5-year TCO reductions--low speed aircraft . . . . e e 136 66 5-year TCO savings- - low speed aircraft. . . . . . . e . 137 67 Percent reduction in fuel consumed- - low speed aircraft . 137 .

68 10 - year fuel savings- - low speed aircraft. . . . . . . 138 69 Block speed versus block distance- - low speed aircraft . 138 . e . . a 70 Fuel consumption comparison- - low speed aircraft . a . e 139 . . . . . . . . . e . e . .

71 DOC comparison--low speed aircraft . 139

72 DOC breakdown- - low speed aircraft . . . . e . e . e e a . . e a . 140

73 STAT advanced technolog y program e e e . e e . . e . _ . e . a . 142 74 STAT technology research programs . . . . . . . . . . . . . e e . 143 75 Dual - property titanium impeller with a wrought Ti-6246 hub HIP bonded to a cast Ti - 6242 airfoil shell. . . . a e e a e . 146 Foreign particle ingestion phenomenon . . . . . . a e e . e . . . 147 e . . .

77 Inlet particle separator concept . 148 . . .

78 Rotating stall phenomenon e 148 79 Axial compressor e . a . 149 .

80 Axial-centrifugal compressor e e . 150 iv -y , 1.

page Figure Title Mar-M247 ^. 81 Dual - property turbine wheal with PA-101 hub and

. . . . . . . . 152

internally cast airfoil ring . . . . . . . .

82 Schematics illustrating candidate cast airfoil cooling

{

^ schemes

83 High temperature turbine seal concept . . . .

. . . 156

84 Typical composite shaft . . . . . . . .

85 Typical application of supercritical design to a power turbine . . . . . . . . . . . . . . . . . . . . . . . . . . .

shaft

. . . . . . . . . 158 86 Failure density as a function of ball latitude

. . . . . . . . . . . . 160

87 Lamilloy construction . . . . . . . . . .

. . . . . . 162 Advanced turboprop propulsion control system . . . .

. . . . . . . . . . . . . . . . 164 N 89 STAT engine condition monitoring f . . . . . . 165 90 Typical advancad high pressure ratio compressor . . .

. . . . . . . . . 165 91 Typical gas turbine am ular combustor . . . . .

. 167 92 3579 kW (4800 shp) STAT engine reduction gear . 169 93 STAT experimental engine schedule . . . . . .

v LIST OF TABLES r Pa A Table Title . . . . . . . . . . . 8 I Baseline aircraft physical characteristics . . . 9 II Baseline aircraft aerodynamic and economic el-aract6 istics III Engine manufacturer-recommended baseline engines IV STAT short haul baseline engine model V 18 Propulsion installation criteria . . . . . . . 21 VI Cost assumptions.

VII Maintenance cost breakdown . .

VIII Rate of climb capabilities . .

IX Summary of DU sensitivity results R Baseline data—DDA study aircraft . . 29 RI Additional sensitivity data--high speed aircraft (SI units) .

XII Additional sensitivity data--high speed aircraft (customary . . . . . . . . . 29 units) . .

%III Additional sensitivity data--low speed aircraft (SI units) 30 Additional sensitivity data--low speed aircraft (customary XIV . . . . . . . . . . . . . . . . . . . . . . . . 30 units) . . . . . . . . . . . . . . . . 32 XV Propeller sensitivity data XVI DDA mission analysis results--current technology high speed . . . . . J . . . . . . . . . . . . . . . . . . . . 34 aircraft XVII Cycle parameter changes . . . . . . . . . . . . . . . . . . . . . 35 XVIII Mission operating conditions . . . . . . . . 35 XIX PD370-37 sensitivity study of cycle characteristics at 6096 m cruise (SI units) . . . . . . . . . . . . . . . . 35 XX PD370-37 sensitivity study of cycle characteristics at 20,000 ft . .

cruise (customary units) . . . . . . . . . 36 XXI PD370--37 sensitivity study of mission fuel changes (SI unite) 36 XXII PD370-37 sensitivity study of mission fuel changes (customary . . . . . . . . . . . . . . 36 units) . . .

XXIII STAT sensitivity study - -cycle parameters or 1790 kW (2400 shp) . . . . . . . . . . 38 engine .# (4800 shp) XXIV kW STAT sensitivity study--cycle parameters for 3579 engine . . . 39 XXV STAT pressure ratio selection 3579 kW (4800 shp). 41 XXVI Candidate advanced technologies and design features . . . . . . . 44 XXVII . . . . . . . . 50 1790 kW (2400 shp) engine fabrication technology XXVIII 1790 kW engine power section prices and weight breakdown (SI . . . . . . . . . . . . . . 0 50 units) XXIX 2400 shp engine power section prices and weight breakdown . . . . . . . . . . . . . . . . 51 ( customary units) .

XXX 1790 kW (2400 shp) engine effect of wheel shaft and rpm .

variables on section weight . . . . . . . . . . . . . . . 51 XXXI 1790 kW (2400 shp) (variation 3) power section average recurring manufacturing price by section . . . . . . . . . . 52 XXXII 3579 kW (4800 shp) engine fabrication technology . . . . . . . . 53 XXXIII 3579 kW engine power section prices and weight breakdown . . . . . . . . . . . . . . . . . . . . . . . . 54 (SI units) .

XXXIV 4800 shp engine power section prices and weight breakdown . . . . . . . . . . . . . . . . . . . . . 54 (customary units) vii f':.EGCi)INt i hA% r;'w : NOT FILMED Table Title Page X 3579 kW (4800 shp) engine STAT study effect of wheal shaft and . . . . . . . . . . . . . . . . .

rpm variables . . . . . . 55

XXXVI 3579 kW ( 4800 shp) (variation 2) power section average recurring manufacturing price by section . . . . . . . . . . . . . . . 55 XXXVII . . . . .

3579 kW shaft study results ( SI units) 60 XXKVIII 4800 shp shaft study results (customary units) : 61 XXXIX 1790 kW shaft study results ( SI units) . . . . 62 XL 63 2400 shp shaft study results ( customary units) XLI . .

ATE STAT propeller improvements for the high speed aircraft . 81 r^ XLII ATE STAT propeller improvements for the low speed aircraft . . . 82 XLVII Advanced STAT turboprop engines ( SI units) . . . . . . . . . . . 85 XLIV Advanced STAT turboprop engines ( customary units) . . : . . . . . 86 XLV STAT advanced technology application-1749 kW ( 2400 shp)

engine . . . . . . . . . . . . . . . . . . . . . . . . . . . 90

XLVI STAT advanced technology application 3579 kW ( 4800 shp)

engine . . . . . . . . . . . . . . . . . . . . . . . . . . . 91

XLVII

Engine price summary. . . . . . . . . . 98

XLVIII Inherent reliability goals for STAT advanced turboprop--major . .

modules . . 101

XLIX Inherent reliability goals for STAT advanced turboprop-- . .

components and accessories . . . . . . . . . . . . . . . 101 L Summary of reliability assessments of STAT advanced turboprop

system- - major modules . . . . . . . . . . . . . . . . . . . 102

LI Summary of reliability assessments of STAT advanced turboprop system- - components and accessories . . . . . . . . . . . . . 103 LII Inherent reliability prediction summary for STAT advanced .

main drive reduction gearbox . . 105 LIII Maintenance cost comparison for baseline and derivative . . . . . . . . . . . . . . .

engines . . . . . . . . . . . 108

LIV Maintenance cost projection of advanced STAT engines compared with baseline and derivative engines . . . . . . . . . . . . 108 LV Maintenance cost projections of advanced STAT engines with and without condition monitoring . . . . . . . . . . . . . . . .

LVI Maintenance cost projections of 3544 kW (4752 shp) advanced STAT engine with and without on-condition . . . . . . . . . 109 LVII STAT turboprop engine comparisons ( SI units) . . . . . . . . . .

LVIII STAT turboprop engine comparisons (customary units) . . . . . 111 LIX Mission results- - high speed aircraft ( SI units) . . . . . . . . . 113 LX Mission results- - high speed aircraft ( customary units) . . . . . 114 LXI Alternate stage length mission results- high speed aircraft - .

(SI units) . .0 . . . . 115

0. LXII Alternate stage length mission results--gh speed aircraft hi

(customary units) . . . . . . . . . . . . . . . . . . .

. . 116

LXIII Technology comparison--high speed aircraft . . . . . . . . . . . 120 LXIV Mission results--low speed aircraft (SI units) . . . . . . . . . 124 LXV Mission results- - low speed aircraft ( customary units) . . . . . . 127 LXVI Alternate stage length mission results--low speed aircraft (SI units) . . . . . . . . . . . . . . . . . . . . . . . . .

viii Page Table Title LXVII Alternate stage length mission results--low speed aircraft (customary units) . . . . . . . • . • • • • . • • • • a • • LXVIII Technology comparison--low speed aircraft . . . . . . . . . . . . 134 LXIX Critical technology elements for STAT research and development * . . . . . . . . . . . . . . . . . . . . . . . . 144 Lx8 Recommended STAT technology programs . . . . . . . . . . . . . . 168 ix SUMMARY Detroit Diesel Allison ( DDA) completed the Propulsion Study for Small Trans- port Aircraft Technology ( STAT) for the National Aeronautics and Space Admin- istration ( NASA) Lewis Research Center ( LeRC), under Contract N.lS3-21995.

This study defined the future research and technology efforts that are most appropriate for propulsion systems for the next generation of small, short- haul type transport aircraft anticipated for the 1990 time frame, along with their expected benefits.

The propulsion requirements for the STAT aircraft were based on studies con- ducted by the HASA Ames Research Center and Lockheed California Company (LCC).

The two reference aircraft were each 50-passenger, twin-engine turboprop + transports designed for 1111.2 km (600 am) range with reserves. The RASA Ames ^,. airplane cruised at 0.47 MN with an initial altitude capability of 6278.9 m (20,600 ft). The LCC airplane cruised at 0.7 MN with 10 , 973 m ( 36,000 ft) initial altitude capability, Typical cruise altitudes for minimum direct op- orating cost ( DOC) for the average route segment of 185.2 km (100 nm) was 3048 m (10,000 ft) and 4572 m (15,000 ft), respectively, at design Mach number. A brief summary of additional aircraft characteristics follows: Low speed High Speed aircraft aircraft Take - off gross weight ( TOGW)- - kg (lbm) 17,927 ( 39,523) 18,291 (40,325) Take - , off distance- - m (ft) 1 , 219 (4 000) 1,219 ( 4,000) Power per engine- - 1,767 ( 3,573 4,790) kW (shp) 2,369) ( Sensitivity studies were conducted on both of these aircraft to determine par- ametrically the influence of propulsion characteristics on aircraft size and cost. Using a distance of 185 . 2 km (100 nm) and =0 .264/1, ( 91.00 / gal) fuel, a 1OX change in the following engine parameters produced a percentage change in aircraft DOC as indicated: A % DOC eX DOC low speed high speed Sensi.rivity parameter aircraft aircraft 4.1 Specific fuel consumption (sfc) 3.5 Maintenance cost 1.3 1.0 Weight 0.3 0.3 ( avg) Length, height 0.3 0.2 i` Thus, sfc and engine maintenance costs were determined to be the engine para- meters that have the greatest effect on aircraft DOC.

A typical DOC breakdown is shown in figure 1, which indicates that the portion / L ($1.00 gal) with engine of DOC attributable to fuel costs is 292 at $0.264 / maintenance costs representing 142. Depreciation and insurance costs for the engine are less than 52 of the total. If fuel is assumed chargeable to the ' e of the total DOC is engine for the purpose of this study, the engine s shn: 482. If fuel goes to $ 0.528 / L ($2.00 / gal), the fuel cost becomes even more dominating and drives the engine ' s share to 602. Thus, improvement in fuel economy has a significant influence on DOC.

Candidate advanced technology elements ( circa 1990) and design features were identified and screened, and parametric studies were conducted to select an appropriate engine cycle. Sensitivity data were expanded to include the ef- fect of engine design variables such as component efficiencies, pressure drop, .'1 and cooling air quantity on engine characteristics of sfc, weight, and cost.

These sensitivities were used with those linking the engine characteristics to aircraft DOC to aid 'n the screening process. The studies were conducted at 1790 ani 3579 kW (2400 and 4800 shp) for the low speed and high speed air- : 1 and a turbine rotor inlet tem- plane, respectively. A pressure ratio of 20 peratvre (RIT) of 1506 K ( 22500F) were selected for both engines for the s . The cycle was purpose of identifying technology improvement requirement selected for minimum DOC for the 185.2 km ( 100 nm; routc segment.

MRS share 50.52)/1 ` 50,264/1 ^` ► ^, ea4i^n share o Isve^n c5t^9at 40`: Fwl fuel MaLatomaaca 45.5% 14% 29.3% 1 1 i 1 1 nasatawaoa I t ^% Dapr. •_ lation 4% ^tepreciatios 1 3% Insurance 0.3% insuraace 0.4% A/C tae A/r dare R% TE80-2129A Figure 1. Typical DOC breakdown--high speed aircraft- - 185 km ( 100 am) block :t:stance.

Two conceptual engine designs were prepared based on the above cycle and in- corporating the candidate STAT advanced technologies. The resultant engines are representative of those with performance levels anticipated for the 1990s, era commuter aircraft market. The engine configurations eventually selected to achieve those performance levels, however, could depart considerably from those described herein.

The advanced STAT engines are similar in that both have a single-spool gasi- Fier and a free-power turbine with front drive. The smaller engine has an "LA-centrifugal compressor with five axial stages. The final stage features a hybrid centrifugal impeller with a cast flow-path ring bonded to a forged hub by tie hot isostatic press (HIP) process. The larger engine uses an all- axial compressor with nine stages. Airflow for the two engines is 6.35 and 11.54 kg/s (14 and 25 pps), respectively. Both engines have two-stage gasi- fier turbines employing impingement cooling in both stages plus the first-vane ( stage. Both turbines have hybrid turbine wheels with cast airfoil rings that are diffus+.on bonded to powdered natal hubs. Abrasive blade tips and abrad- able coatings are used to provide minimum running clearance in the turbines.

r Turbine shafts are fabricated of borsic-titaniuA matrix composite material for required shaft stability with simple rotor support. The combustors are trans- piration cooled for improved temperature profile and increased hot section life.

The 19.1:1 reduction gear for the smaller engine is a close-coupled star/ planetary gear system mounted concentrically on the front of the power sec- tion. The larger engine has an offset-type reduction gear with a dual com- pound idler gear train designed for a 10.4:1 reductior. ratio. This reduction gear is connected to the power output shaft via a mechanical torquemeter.

Both reduction gear cases employ a composite material.

The control and fuel systems for both STAT engines are configured to use an advanced digital electronic controller. An engine condition monitoring system is employed to gather data for component life usage and to assist in fault detection and isolation procedures so that corrective action can be planned to prevent failures. Modular construction is employed for both engines.

A benefit anslysis was conducted comparing the advanced STAT engines with cur- rent technology engines as well as with hypothetical 1985 technology deriva- tive engines. In the high speed airplane, the derivative engine was 3 to 4X better in DOC than the current technology engine for the average block dis- tance of 185.2 km (100 on). Also, the advanced STAT engine improved 15 to 16% in DOC compared to current technology. These results apply for fuel costs from $0.264 to $0.396/1 ($1.00 to $1.50/gal). Fuel consumption was reduced 11X with the derivative engine and 23% for the advanced STAT engine. Corres- ponding benefits in the low speed airplane were similar but with slightly low- er improvement percentages, i.e., 13 to 14% improvement in DOC over current technology and 20% reduction in fuel consumption for the advanced STAT engine.

Research and development programs necessary to advance the state of the art for engine components, and an experimental engine program to provide basic operating data on the advanced technology elements and design features, were planned to achieve a 1988 date for readiness to release for full commercial deve lopment.

INTRODUCTION The increasing need for newer, more efficient aircraft for specialized short- haul sectors of the transportation market has been generated by unprecedented ^ growth in commuter traffic operations. Improved accommodations for the pas- J L + senger and his baggage, similar to those of the major trunk line operations, lity and sound levels that are customary for large, long- are needed. Ride qu a range aircraft are desired. To accomplish these goals requires a small com- muter transport that is efficient in terms of fuel usage. The need for re- duced fuel consumption has never been greater because of its rapidly rising price and decreasing availability. Reduced maintenance, which contributes significantly to lower operating costs and improved on-time performance, is also an important objective.

NASA LeRC sponsored the Propulsion Study for STAT reported herein. This study attempts to identify the advanced propulsion technology that is appropriate for the smaller short-haul type of aircraft. "Appropriate" technology is that which results in reductions in DOC, fuel usage, or total cost of ownership ti ,y (TCO) size, cost, and maintenance cost for the power output required.

The study was divided into three basic tasks: • Task I--Baseline Airplane, Engine, and Mission Definitions • Task II--Advanced Technology Identification and Evaluation • Task III--Recommendations for Future Research Note: "Engine" as used in this report denotes power section plus reduction gearbox. Also note specific fuel consumption (SFC) is equivalent to brake specific fuel consumption (BSFC).

et E BASELINE AIRPLANE AND MISSIONS The objective of Task I was to define two 50-passenger baseline commuter air- craft representing technology for 1980 initial operation. One of these air- craft was low speed, but with at least 463 km/h (250 kt) indicated air speed (IAS) capability at 3048 m (10,000 ft) altitude, and the other high speed with 0.7-MN cruise capability. The airplanes use turboprop propulsion systems, also of 1980 initial operation capability (IOC) technology.

Reference aircraft data conforming to NASA aircraft and mission specifications and guidelines for economic calculations was obtained. Low speed aircraft high speed aircraft information was fur- data was provided by NASA Ames. The nished by Lockheed California Company (LCC), and was generated as a part of their STAT contract with NASA Ames. Details of the airplane requirements and the configuration will be discussed in the following paragraphs.

Hypothetical turboprop engines representing existing modern production tech- nology 19130 IOC, along with appropriate scaling relationships over the horse- power range of interest are also defined and discussed in the subsequent sec- tion on Baseline Engines.

A computerized aircraft sizing, mission, and economic model was developed by DDA to accept the reference aircraft data provided by NAA Ames and LCC. The model is described, and the aircraft system sensitivity to engine parameters is presented in paragraphs on Aircraft Sizing and Cost Model, and Aircraft i System Sensitivity to Engine Parameters.

MISSION REQUIREMENTS The baseline missions conform to the following performance requirements: Full design payload is carried over a range of 1111 km (600 nm) with IFR o reserves for a 185.2 km (100 nm) alternate, and 45 min at maximum endur- ance power at 3048 m (10,000 ft) altitude.

o Field length is limited to 1219 m (4,000 ft) for a hot day 306 k (90°F) at i sea level, per FAR 25.

o Aircraft meets current FAR 36 Stage 3 noise limits, minus 8 EPNdB at all measurement locations. ' o Cruise speed capability is at least 463 km/h (250 kt) indicated airspeed at 1829-3048 m (6,000-10,000 ft) altitudes, standard day conditions for the low speed airplane, and 0.7 MN for the high speed airplane.

o A terminal area speed capability is at least 334 km/h (180 kt) indicated airspeed with gear and flaps extended in order to stay with large jet air - craft.

Stall speed is less than 172 km/h (93 kt) in landing configuration at max- o imum landing weight in order to qualify for operations in Instrument Ap- proach Category B aircraft requirements.

The mission profile shown in Figure 2 describes the features of the design and alternate missions used in the STAT study. The airplane objective was to carry 50 passengers, which was equated to 4536 kg (10,000 lbm) payload. Basis for specification of the mission segments is described herein.

Engine sizing • 91.4ra/min (300-ft/min) rote of climb at wise Mach altitude—max continuous power—initial cruise weight • Tako-o f and landing Reid length of 1219 m (4000 h) (SL- Capabilities anterm ediate power—standard day—full TOGW, • AEO service ailing • OEI service ceiling "'Jed is 50 passengers + boggage-4336 kg (10,000 Ibm) lO /^ ^ O O

I ^ ^ stop• Steger length

Alternate field --^ t_ z-- ( 9. 16011th 1111.2 kss (600 nn) 45 min 185.2 km 000 nm) endurance *Typical stage length s 185.2 km (100 nm) 1. Take-off allowance-1.0 min at takeoff power SISS 2. Climb to wise altitude 3. Cruise at constant Mach number and altitude • Design stage 16ngtFc —design specified Mach number and altitude • Typical stage !ength—design specified Mach number of altitude for min DOC n mcimutn specific range at 3048 m (10,000 ft) altitude 4. Cruise—Mach number for 5. Endutonoo—Mach mrnber for maximum specific endurance at 3048 m (10,000 ft) altitude Note: 10 min of nonproductive maneuvering time was included for all stage lengths, i.e., block time equals flight time plus 10 min.

TEBO-2123 A Figure 2. - Mission requirements.

A fuel allowance of 1 minute at maximum rated power was selected as represen- tative of fuel usage for the takeoff. All descents were assumed zero fuel and zero time segments.

The climb path was important in establishing mission fuel usage, block veloci- ty, and block time, because climb distance constituted a substantial part of the stage length. Its selection has a direct bearing on DOC. Accordingly, a study was done to select a suitable climb path for each reference aircraft.

Several constant IAS climbs were flown with each aircraft, and the resulting mission capabilities and DOC were examined. Climb speed could be no higher than 463 km / h (250 kt) IAS below 3048 m (10,000 ft) altitude. The climb speed selected for the LCC aircraft was a constant 417 km / h (225 kt) IAS, and for the NASA Ames aircraft, a 370 km / h (200 kt) IAS constant climb speed. These climb speed selections contributed a substantial DOC saving, and also provided terminal Mach numbers at the cruise altitude that closely matched the desired cruise Mach numbers.

The design mission stage length ( climb plus cruise) was specified at 1111 km (600 nm) by NASA to represent typical commuter requirements. This range con- stituted the design range for sizing the aircraft. An alternate stage length of 182 km (100 nm) was used to examine the DOC implications of STAT engine 1'.

technology with the sized airplane. The alternate stage length mission was f exercised with full fuel load and payload at take-off; i . e., at design TOGW.

The cruise altitude for the alternate stage length was optimized to obtain minimum DOC within the constraint that the cruise leg was 3 . at least half the climb-cruise stage length.

i

Additional mission segments were used to determine the fuel reserves necessary to meet the NASA specified alternate field capability. They were a take-off, l climb, and IFR cruise at 3048 m (10,000 ft) altitude, maximum range speed for l 182 km (100 nm), and a 45-min endurance at the same altitude at maximum endur- ance speed. (See details in Figure 2.)

Engine sizing conditions specified for the study were at takeoff and the initial cruise point. A take-off and landing field length limit of 1219 m (4000 ft) was required on a +32°C (+90°F) day, and a 1.52 m/s (300 ft/min) minimum rate-of-climb was required at the initial cruise point with the engine at maximum continuous power.

i BASELINE AIRPLANES The baseline aircraft were required to utilize scaled versions of existing turboprop engines or notational turboprop engines representing existing, mod- l ern designs. They were also required to provide the following accommodation features: Customary Units SI Units 200 lbm 90.7 kg • Weight per passenger plus baggage 200 lbm 90.7 kg • Weight per crew member 130 lbm 59.0 kg • Weight per flight attendant 6 ft 1.9 m • Minimum interior aisle height 32 in.

813 mm • Minimum seat pitch 18 in.

457 mm o Minimum seat width between armrests 18 in.

457 mm • Minimum aisle width • Preloaded baggage storage volume 5 ft3 0.14 m3 per passenger 20 in. x 20 in.

508 mm x 508 mm Carry-on baggage storage volume x 11 in.

x 279 mm per passenger Garment hanging storage width 0.8 in.

per passenger 20 mm One lavoratory 5 psi Minimum cabin pressurization 34.5 kPa Maximum cabin interior noise level less than 85 db OASPL Speech interference level of less than 65 db l^ Airframe design life was required to be at least 30,000 h with 60,000 take- off-land cycles.

^s Physical characteristics of the low speed Ames aircraft and the high speed LCC aircraft are presented in Table I.!

I `y TABLE I. - BASELINE AIRCRAFT PHYSICAL CHARACTERISTICS Low speed High speed Data source NASA - Ames LCC 1 Design TOGW- - kg (lbm) 17,927 ( 39,523) 18,291 (40,325) Payload--kg ( lbm) 4536 ( 10,000) 4536 ( 10,000) Fuel load- - kg (lbm) 1731 ( 3817) 1905 ( 4200) jr OEW--kg ( lbm) 11 , 660 (25,706) 11,850 ( 26,125) Wing Span--m ( ft) 25.4 ( 83.2) 21.6 (71.0) ^l Area - _m2 (ft2 ) 61.19 ( 658.7) 46.8 (504) Taper 0.3 0.3 1/4 C sweep--deg 5.0 5.38 AR 10.5 10.0 Loading--kPa (lb/ft ) 2.87 (60.0) 3.83 (80.0) Fuselage Length--m ( ft) 22.8 (74.8) 22.8 (74.7) 2.74 ( Diameter--m (ft) 9.00) 2 . 89 (9.50) Horizontal tail Span--m ( 37.35) 7.181 (23.56) ft) 11.38 ( Area--m 2 (ft 2 ) 25.92 (279.0) 7.733 (83.24) Vertical tail Span--m ( ft) 6.300 ( 20.67) 3.514 (11.53) Area- - m2 (ft 2 ) 24.81 ( 267.0) 6.172 ( 66.43) Propeller No. blades 4 Diameter- - m (ft) 4.51 ( 14.8) 3.66 (12.0) Performance capability, aerod7namics, and economic characteristics oi both airplanes as used in this study are presented in Table U. This data was obtained from the results of the STAT studies at NASA-Ames and Lockheed where engines of the appropriate power class were scaled to specific propulsion re- quirements.

Figure 3 shows the relationship of the baseline airplanes' characteristics to those of several contemporary transport aircraft types. The study vehicles are appropriately placed at the high speed border of the existing commuter equipment, and also provide a higher speed option that will elevate commuter equipment into a performance category competitive with corporate jet aircraft.

fi TABLE ii. - BASELINE AIRCRAFT AERODYNAMIC AND ECONOMIC CHARACTERISTICS Low speed High spee d LCC NASA-Ames Data source 1111 (600) 1111 (600) Design rang e--ka (nm) 3573 (4792) 1767 (2369) Potter per engine- - kW (shp) 0.7 0.47 Cruise Mach number 10,668 (35,000) 6096 (20,000) Initial cruise altitude- - m (ft) 181.8 L/D at initial cruise 14.9 0.49 0.45

Lift coefficient at guise

7132.3 (23,400) 4998.7 (16,400) Engine-out service ceiling--m (ft) 1218 (3995) 1219 (4000)

Take-off distance- - a (ft)

204 (110) 209 (113) Approach speed--km / h (kt) EAS 1219 (4000) 1218 (3997)

m (ft)

Landing distance- -

4,400,000 4,650,000 Aircraft flyaway coat- - 1979 dollars 2.30 (4.26) at 1111 km ( 600 nm) 2.73 (5.06) Operating cost * 4.92 (9.11) 100 nm) 3.82 (7.08) Operating cost * at 185 km ( / /seat nm) with fuel cost - $0.264 1 ($1.00/gal)

*Measured in E/seat ka. (L i

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BASELINE ENGINES The most representative DDA current technology turboshaft engine over 746 kW (1000 hp), for conversion to turboprop applications, is the Model XT701-AD- 700. This engine is a free turbine turboshaft that was developed through safety demonstration testing for the US Army's Heavy Lift Helicopter (HLH) program. This program was cancelled by the Army in December 1975, but had it continued through complete development for the HLH, it would have been oper- ational in the early 19809• DDA continued the engine development as a com- pany-funded effort, directed toward industrial applications. The industrial version of the engine is designated the Model 570-K. It went into production in early 1979.

The turboprop engine that was derived from the Model XT701 power section is shown in Figure 4. It was designated the Model PD370-37 turboprop engine, and was derived during earlier Maritime Patrol Aircraft (MPA) engine studies for the Navy, where it also represented a low risk, current technology approach for application to that aircraft system. The speed reduction gearbox for the propeller was based on a standard T56 current production unit.

The Lockheed STAT short haul study was being conducted simultaueously with the advanced technology propulsion system studies at DDA, Garrett, and General Electric. To define an appropriate scalable baseline engine, and to ensure a degree of consistency with the engine technology studies, DDA, Garrett, and GE were asked by Lockheed to recommend a baseline engine for each of the Lockheed study aircraft and to provide scalable performance for the recommended engines.

Model P0370-37 3.00 m 018.11(n.)

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(13.36 in.)

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1.515 m (59.66 in.)

0.653 m (2b.72 in.)—T 0.314 m ^1 0.660 m (12.38 in.)

(26.00 In.) d10

Di

m 0.518 (20.40 in.)

die

IV

reeo-2042 0.820 m (32.27 in.)

Figure 4. - Baseline turboprop engine.

DDA and the other engine manufacturers responded with the engines described in Table III. In general, the list represented the latest turboprop engine each manufacturer had to offer.

TABLE III. - ENGINE MANUFACTURER-RECOMMENDED BASELINE ENGINES Rating— Engine Model Comsat$ Basis *Aetna kW (INO manufacturwr kaling corrections provided Proposed turboprop version of Model P0370.37 6301 (8450) Allison kW to sole *Agin* to 101 T701-AD-7W MLM engine dew)op*d (2000 np) range through PPFRT (cancelled 19761.

Now in stationary powerplant use.

Scaling corrections providw Production engine. Latest in TPE331-11 745.7 (1000) Garrett kW to scale engine to u37 T76 family.

(30W bp) Proposed turboprop version of T700 1118.6 (1500) CT7-2 GE turboshaft &*in* (preliminary data).

2339 (3137) Production engine 0960 vintage).

CT644820-4 Only provia*a acquisition an.3 None P8M Canada maintenance cost estimates Estimates of engine power class required for the 30 and 50 passenger baseline aircraft were 1491 kW (2000 hp) and 2983 kW (4000 hp), respectively. Since none of the recommender engines was of these sizes, scaling corrections were required to match engine power capability with aircraft thrust requirements.

DDA and Garrett provided estimates for scaling corrections of the recommended powerplants.

Scaling equations that scaled characteristics down to engine sizes of approxi- mately 1119 to 1491 kW (1500 to 2000 hp), were generated and keyed to the PD370-37. These equations represent technology trends and not scaling rela- tions for a specific engine configuration.

The baseline scaled engine performance, weight, geometry, and cost data were based on a synthesis of the recommended engine data and scaling corrections.

The following paragraphs describe the basis for establishing the baseline en- e gine definitions.

Figure 5 presents the sea level static (SLS) uninstalled engine brake sfc for existing production turboprop engines as a function of rated horsepower. The performance level of the recommended engines is included. On the basis of this data, a baseline engine performance trend was established (i.e., dashed E line on the figure). The sfc level of the engines in the two power classes are noted.

t SLS unlaftlled max prover 0.7 k Georim lea included in UFC a haduction engines 0 000 0 0.6 ^^` 30 PAX bowline O V V H 0.3 `^4%,VI:-30 PAX bowline n, P0370-37 t 0.4 I _flange Of it "' — - 30 PAX - - 50 PAX- 0.3 1,000 2,000 3,OPJ 4,000 5,000 6,000 7 SO rated rower—kW 0 2,000 4,000 6,000 8,000 10,000 SLS rated power—shp T ED0-2043A Figure 5. - SFC versus rated power.

Figure 6 presents the engine power-to-weight ratios for the baseline engines and for other production engines. The selected trend of horsepower to weight ratio with rated power, as well as the values chosen for the 30 and 50 PAX aircraft engine power classes, are noted.

Similarly, Figures 7 through 9 present the baselire trend established for en- gine length, overall envelope width, and overall envelope height, respective- ly, based upon the recommended and existing production engines. It is assumed that engine-mounted accessories are located in the horizontal plane.

DDA, Garrett, GE, and P&W of Canada were asked by Lockheed to define engine (including gearbox) acquisition and maintenance cost estimates (1979 dollars) as a function of engine-rated horsepower for engines representative of today's technology and design practices. Figure 10 presents the estimated acquisition cost in terms of dollars per horsepower. Ground rules included the assump- tions that the gearbox price was included, and that the engine was mature. The recommended baseline trend is shown as the dashed line. Costs are assumed representative of OEM levels. Similarly, the burdened maintenance cost esti- mates and selected base-lines for the aircraft are shown in Figure 11.

SU standard day • Productiar anpinn 10 ^ MM70-V s _ e eawline 5D PAX bowline tend 30 PAX baseline ^,.••••• ^,^^• 4 • 2 •^ M • 1 Ronp of intend - 30 PAX-- 50 PAX - SLS rated pawn-kW 0 4,000 2,000 6,000 8,000 10,000 SLS """ Power—shp TE80-20"A Figure 6. - Engine power-to-weight ratio versus rated power.

The variations in engine performance, weight, geometry, and economic trends with size are quantified in equation form, and are tabulated in Table IV.

These equations were used to define the principal characteristics of existing turboprop propulsion systems in the 745.7- to 5965 6 kW (1000- to 8000 hp) .

range. These scaling data, applied to the Model PD370-37, were used to define baseline engine characteristics for evaluation of advanced propulsion system technologies.

AIRCgAFT SIZING AND COST. MODEL This section presents the following items: • Overview of the STAT mission analysis computer program • Engine / airframe sizing philosophy • Cost model • Results from Task I mission studies

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^Il 4.0 ► ► 0 is is on low w end par ee sov.r.to to s.—eor i 6 Wol 0 Nedualen worm 3.0 r.r .a x0370-^7 ^rrr r r r r r 1lateline trend • rr^1

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30 PAX boudiM 2 .0 a • •o 4 1.0k O• f %rpeof Interest—^ 30 PAX 5o P AX 4,000 5,000 6,000 7,000 0 1,000 2,000 3,000 SLS tared pe«.r—kW 1 t I 1 I 4,000 5,000 8,000 10,000 0 2,000 SAS rated power- O p TEB0.2060A Figure 7. - Engine length comparison.

TABLE IV. - STAT SHORT HAUL BASELINE ENGINE MODEL Synthesis of existing engine characteristics representative of today ' s technology and design practices.

Customary Units SI Units Characteristic kW0.11 1.200 shp0.11 - ( lb/hr-hp) 196.32 Engine BSFC- og/W • s including gearbox loss Weight--kg (lb) kW0.54 6.505 12.24 shp0.54 Engine ®'MuN..;;.

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20.0 0 .

0.4 Ilap. d rm.wq rAx' So 's9 ► Ax 1 1 1 1 1 1 1 0 0 0 1,000 2,000 3,000 4,000 $,000 6,000 71000 SU emd "P-•kw 1 1 1 1 1 0 2,000 4,000 6,000 4,000 10,000 SU Im PO+WFr TM-2%VA Figure 8. -• Engine max envelope width.

TABLE IV. (CONT) Customary Units Characteristic SI Units Geometry- - m (in.)

kW0.21 0.4746 17.57 shpO.21 Length-prop mount flange to rear engine flange kWO.26 10225 3.73 shp0.26 Max envelope height 0 .

kWO.04 21 shp0.04 0.4680 18 .

Max envelope width OEM acquisition cost ( 1979 $) 1192 shp0.7 Engine 1463 . 8 kWO.7 flight hour) Maintenance cost ( 1979 $/ kWO.66 0.2949 4.243 ahp0.66 Engine shp--engine max rated power, SLS, standard day.

engine brake specific fuel consumption at max rated power, SLS standard BSFC- - Assumes exhaust nozzle area sized for turboprop application.

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Overview of STAT Mission Program Input, major calculation functions, and output of the STAT mission program have been generalized in the block diagram shown in Figure 12. Mission re- quirements, aircraft, and engine data are used in the engine/airframe sizing aircraft size combination that will calculations to determine the exact engine / meet the specified mission requirements. Tlst resultant aircraft, mission fuel and time data, plus input economic criteria are used in the cost routine to calculate the required cast parameters.

Engine / Airframe Sizing i A general flow diagram of the engine / airframe sizing procedure used in the STAT study is shown in Fi7uro 13.

Definition of the mission requirements, baseline or reference aircraft char- t acteristics, and unity size engine data constitutes the first step in the pro- cedure. Mission range, speed, and payload must be specified along with the airframe critical propulsion requirements necessary to establish engine size.

Baseline aircraft dimensions, geometry, component weight breakdown, aerodynam- ics, airframe cost factors, and installation information must also be speci- fied. Unity installed angine performance is input for those altitudes, velo- cities, and power settings needed to calculate engine size and mission fuel usage. Engine dimensions and weight data are also required.

t 1979 dollars ! 180 L Ervin manufacturing estimates Recotrionended baseline t.

30 PAX baseline 140— `

a 'S` loo-

1 20- S 30 PAX baseline IL I I I ^ ` trend npe of interest F301PAX 90 PAX 0 1000 1000 6001 SLS rated power—kW l I —_1 I ( 0 1000 4M 6000 SLS rated power—shp TM2071A Figure 10. - Engine acquisition cost versus rated power.

Uninstalled engine performance is adjusteo to reflect appropriate installation effects due to inlet characteristics, aircraft services, and propeller char- acteristics. The installation factors used to obtain installed engine per- formance are listed in Table V. This table lists the propeller efficiencies off, climb, and cruise power conditions for all engines (wp) used for take - evaluated in the STAT study. These propeller efficiencies are typical of the P3 propeller performance. Inlet recovery, gearbox loss, Lockheed Electra / bleed, and power extraction factors were also applied to all engines.

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^e 1979 dollar, bone burdMnd Engine manufoc:w-inp ratimates appmKimately 3000 hr T!O Recommended baseline — i

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Baseline trand 50 PAX baseline S01-D13 ` s 30 PAX 6a»line Amp of in :50:P: 30 PAX X 0 1 0 1000 2000 3000 1000 3000 6000 SLS rated power—kW 4000 6 0 20M — shp SLS rated power TE80-2072 Figure 11. - Engine maintenance cost versus rated power.

TABLE V. - PROPULSION INSTALLATION CRITERIA Efficiencies ( typical Electra / P3 propeller) Propeller • Take - off power- - thrust-to - static-ahp ratio 2.50 -rip at 0.1 M N 0.26 0.57 -qp at 0.2 MN ,qp at 0.3 MN 0.73 0.84 • Climb power, 1p • Cruise, ,q p (0.7 MN cruise) 0.81 Cruise, qp (0.5 MN cruise) 0.90 • Installed performance • Inlet recovery 1.00 • Gearbox power loss--% 2 • Customer bleed-ppm - kw (ahp) 55.9 (75) / engine • Customer power extraction-

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TE30-2106 Figure 12. - STAT mission program.

The engine/airframe sizing is initiated with the selection of a "start point" trial vehicle gross weight. The engines are scaled to mbet the most critical power requirement. Scaling equations developed by Lockheed for engine tech- nology trending were utilized for all engines evaluated in the DDA model.

Baseline aircraft dimensions, geometry, and component weights are then scaled

to correspond to the trial gross weight. The power or drag is adjusted as the vehicle geometry varies from the baseline as driven by propulsion influences, thus requiring iteration of the engine sizing power or thrust. It is noted that the resultant characteristics are similar to the reference aircraft, but modified for scale effects. The fuel available is calculated for this trial gross weight and compared to tine fuel required to perform the design mission flight profile. If the fuels are not equal, a new trial gross weight is se- lected and the process repeated. Upon conv-rgence of the fuel weight itera- tion (fuel available a fuel required), a complete description of the scaled aircraft/engine combination is obtained. Dimensions, geometry, weight break- down, engine sizing data, mission time, power setting, and fuel breakdowns are

provided for on-line cost calculations for the complete aircraft including the

engine and airframe.

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f STAT Cost Model . The previously discussed engine/aircraft results, plus the cost assumptions listed in Table VI, are used with the cost equations presented in this section to obtain total aircraft cost (TAC), direct operating cost (DOC), and 5-year total cost of ownership (TCO).

TABLE VI. - COST ASSUMPTIONS t Base year dollars 1979 Fuel cost--$/L ($/gal) 0.264 (1.00) and 0.396 (1.50) Utilization--h/y 2800 Insurance--%/y 1.5 Depreciation Total flyaway price plus spares to 15%, residual in 12 years Spares investment--% Engine 30 Aircraft less engine Crew cost--$/(block hour x number of seats) 2.50 Total Aircraft Cost (TAC) f TAC is calculated as follows: o Engine acquisition price = 1.50 x OEM price o Propeller acquisition price 2 ) 0.12_ $2653.09 x D x (Ep/Dpp -SI units Z ) 0.12 _ $350.11 x Dp Z x (Ep/Dp --customary units where: Dp . propeller diameter---m (ft) Ep = engine power---------kW (shp) o Airframe acquisition price = FC + PC + (AFSC x AFWT) where: FC = Fixed costs (Low speed A/C) FC = $1,927,000 (High speed A/C) FC = $1,734,000 PC = Procurement costs (logistics support, product development, etc.)

(Low b High speed A/C) PC = $503,000 AFSC - Airframe specific cost (Low speed A/C) AFSC = $254/kg ($115/lbm) (High speed A/C) AFSC $196/kg ($89/lbm) AFWT = Airframe weights (fuselage, wing, vertical tail, horizontal tail, landing gear, and nacelle) t.

i i TAC - (Engine + Propeller + Airframe) x 1.10 (Flyaway Price) (Acquisition Price) (Mark up) Direct Operating Cost (DOC) The elements of DOC are fuel and oil, insurance, depreciation, crew cost, air- frame maintenance, and engine maintenance cost. These items are calculated as follows: • Fuel and oil) These elements were calculated • Insurance ) using December 1967 ATA Standard Methods • Depreciation) • Crew Cost - $2.5/(block hour x number of seats) o Airframe maintenance costs--$/block hour x Maintenance cost - 0.3685 x (Empty Wt.--kg)0.569 AF x AF Maintenance cost - 0.235 x (Empty Wt.--lbm)0.569 Where: AF - Adjustment factor (Low speed A/C) AF - 1.0 (High speed A/C) AF - No. of departures/No. of departures for the design stage length mission.

No.of departures - Utilization/Block time Note: For design stage length mission - - - - - AF - 1.0 For alternate stage length missions - - - AF > 1.0 i.e., stage lengths less than 1111 km (600 NM) o Engine maintenance costs -- $/flight hour engine maintenance cost - $0.295 x (power--kW)0.66 - $0.243 x (power--shp)0.66 engine maintenance cost propeller maintenance cost - $4.15/flight hour Further breakdown of both airframe and engine maintenance cost used in the STAT cost model is shown in Table VII.

TABLE VII. - MAINTENANCE COST BREAKDOWN Airframe Maintenance: Direct material cost - 0.5 x direct airframe maintenance cost Direct labor cost - 0.5 x direct airframe maintenance cost Burden - 0.8 x direct labor cost Engine (power section + gearbox + propeller) maintenance: Direct material cost - 0.8 x direct engine maintenance cost Direct labor cost - 0.2 x direct engine maintenance cost Burden - 0.8 x direct labor cost

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Total Cost of Ownership (TCO) The total cost of ownership was defined by DDA to be the cost of financing the purchase of the aircraft plus the cost of operating these aircraft over a per- iod of 5 years. The finance rate was assumed to be 10%.

+ TCO - 5 - y coat of financing aircraft purchase 5-y cost of operation . 00) x TAC + DOC x CF x VB x U x 5 y TCO - (1.10 5-1 where: TAC - Total aircraft cost ( flyaway)--$ DOC - Direct operating cost--E seat km ( seat um) / CF - Conversion factor ( 50 seats / 100 ^E) VB - Block velocity- - km/h (nm/h) U - Utilization rate- - 2800 h/y Task I Mision Study Results Results from studies completed during the Task I effort are presented as fol- lows: o Cruise altitude selection o Climb path schedule o Rate of climb capabilities Cruise Altitude Selection The cruise altitude and Mach number used by DDA for the design stage length, 1111.2-km ( 600 nm) mission were fixed at values approximate to those specified for the baseline aircraft.

Design Cruise Design Cruise Altitude (ft) Velocity (Mn) 20000 0.47 Low Speed Aircraft High Speed Aircraft 35000 0.70 The cruise altitudes for the off-design stage length missions were selected to minimize DOC at the aircraft design cruise Mach number.

Figure 14 shows a typical plot of altitude versus DOC for the high and low speed aircraft flying the 165.2 km ( 100 nm) stage length mission. This figure indicates the minimum DOC altitude for the low speed aircraft to be at 3048 m (10,000 ft) and at 5791 m (19,000 ft) for the high speed aircraft.

185.2 • km (100-nm) alternate stage length mission High speed aircraft I w in NX altitude 2 low speed aircraft 4.00 3.15 4.25 4.50 4.6 5.00 DOC—slseat km W I ^ 0 7.00 7.50 8.00 9.00 8.50 DOC-9/sut nm TE8O-2108 Figure 14. - Cruise altitude selection.

Climb Path Schedule Because of the importance of the climb path in establishing mission fuel us- age, block velocity, etc., a study was completed to select the most suitable climb path for both aircraft types. Several climb paths of constant indicated air speed (IAS) were flown in the design mission (climb to cruise altitude plus climb to alternate field cruise altitude) with a current technology en- gine. The resulting aircraft were then flown in the 182.5 km (100 ran) off-de- sign mission with the constant IAS climb that was used to establish the design aircraft. The 185.2 km (100 am) DOC results from each climb velocity path are shown for the high speed aircraft in Figure 15, and for the low speed aircraft in Figure 16.

It is noted that the initial or lowest climb velocity is approximate to the path that would be flown for a maximum rate of climb schedule. Also, the climb velocity was not allowed to exceed 463 km/h (250 kt) IAS below an alti- tude of 3048 m (10,000 ft). The climb schedule selected for the high speed aircraft was a constant 417 km/h (225 kt) IAS path and for the low speed air- craft a 370.4 km/h (200 kt) IAS path. The selected climb paths were incorpo- rated into the mission and utilized to evaluate each engine technology. They provide a substantial portion of the DOC reduction indicated for thus range of i 9.0 4.83 Eniline at climb power nttln8 0.84 Propeller efficiency a 0.53 ® Climb Mach at the dai8n minim 1 8.9 4.00 10,668 m (35,000 k) ervir altihde Nate: arvbe velocity + 0.7 MN Rewltinp 0.63 DOC for -2.796 a 185.2 km 8.8 4.73 (100 nn) minion 0.70 4.70 { 0.76 4.651 1 1 1 1 8.6 450 500 300 350 400 Climb velocity-4AAA IAS 225 250 175 200 Climb velocity--KIAS TM-2109A Figure 15. - Climb study results--LCC aircraft.

constant IAS climb velocities evaluated, plus terminal climb Mach numbers that closely match the cruise Mach number for the design mission. However, in the case of the off-design missions, this final climb/cruise Mach number match is not maintained due to the optimization of the cruise altitude to minimize DOC.

Rate of Climb Capabilities w.

The rate of climb capabilities for standard day, all engine operative (AEO) " and one engine inoperative (OEI) maximum rate of climb at sea level, and air- craft service ceilings for the current technology engine-powered low and high speed aircraft are shown in Table VIII. Note that the power loading for the high speed aircraft is approximately 70% higher than that for the low speed aircraf -.

I

7.45 Engine at climb power setting propeller efficiency a 0.84

FXXI

Climb Mach at the daisn minion cruise altitude 60% m (20,000 ft) Note: cruise velocity = 0.47 MN 7.40 Resuitinp 3.99 1.

OOC for a 185.2 Ian -0.7% 3.98 (100 nm) minion 3.97 7.35 3.96 3.95 7.30 250 300 450 350 400 /h iAS Climb velocity—km 150 200 225 Climb velocity—KIAS TEOD-2110A Figure 16. — Climb study results--Ames aircraft.

TABLE VIII — RATE OF CLIMB CAPABILITIES • Current technology engine (Intermediate power) • Standard day conditions • Full design takeoff gross weight Low High speed speed aircraft aircraft All Engine Operative (AEO) o SL maximum rate of climb--m/min(fpm) 884 (2900) 1707 (5600) o Service ceiling--m (ft) 8534 (28000) 12192 (40000) (OEI) One Engine Inoperative o SL maximum rate of climb--m/min (fpm) 299 (980) 674 (2210) o Service ceiling--m (ft) 4877 (16000) 7010 (23000) 1.

AIRCRAFT SYSTEM SENSITIVITY TO ENGINE PARAMETERS iThis paragraph presents the sensitivity data developed in Task I for the high and low speed study aircraft powered by a current technology engine.

Table IX summarizes the percentage changes in direct operating cost values (185.2 km ( 100 nm) stage length mission) for IOX improvements in each of the following engine parameters: TABLE IX. - SUMMARY OF DDA SENSITIVITY RESULTS i ^. Reduction in DOC*,_ % Engine Parameter High speed Low speed aircraft aircraft (10% Improvement) i Overall sfc 3.5 4.12 Weight (dry) 0.30 0.28 OEM price 0.40 0.41 Price ( plus maintenance parts) 1.29 1.10 Maintenance 1.28 1.00 Maintenance labor only 0.40 0.31 Max envelope length 0.28 0.19 Max envelope height 0.41 0.27 * 185.2 km ( 100 nm) alternate stage length mission and fuel cost = / $0.264 L'(81.00/gal).

The relative importance of engine sfc, weight, and cost can be determined from the data presented in Table IX, and is illustrated in the following equations which show the percentage improvement in each parameter required to obtain a 1% reduction in DOC: High speed aircraft -- 1% DOC - 2.8% engine sfc = 33.3% engine weight = 25% engine cost Low speed aircraft -- 1% DOC - 2.4% engine sfc = 35.7% engine weight - 24.42 engine cost It is noted that baseline aircraft parameters and propulsion system weights, dimensions, OEM prices, and maintenance casts are listed for reference pur- poses in Table X.

Additional data from the sensitivity studies of the current technology engine (CTE)-powered high and low speed aircraft configuration are presented in Tables XI through XIV. Each table lists changes in gross weight, empty weight, aircraft acquisition cost, block fuel, and DOC resulting from 10% im- provements in each parameter listed.

a 1.

TABLE X. - BASELINE DATA-- ODA STUDY AIRCRAFT High speed Low speed aircraft aircraft TOGW--kg (lbm) 18,299 (40,343) 18,938 (41,730) Engine take-off rating at SLSS--kW (shp) 13,531 (4735) 2139 (2868) Aircraft flyaway cost--1979$ 5,456,000 4,836,000 Weights (dry) per engine--kg (lbm) Power section 367 (808) 279 (616) Gearbox 166 (367) 127 (280) Total 533 1175) 406 (896) Propeller 420 (926) 254 (561) Dimensions Engine max envelope length--m (in.) 2.66 (104.6) 2.39 (94.1) Engine max envelope height--m (in.)

0.85 (33.3) 0.74 (29.2) Propeller diameter--m (ft) 3.66 (12.0) 2.83 (9.3) OEM price per engine--1919 $ Power section 308,213 216,979 Gearbox 13,682 9,632 Total 321,895 226,611

Propeller 50,881 30,817

Maintenance (fully burdened) 1979 $/EFH Engine 64.75 46.50 Propeller .<.15 4.15 DOC Breakdown Figures 17 and 18 present a breakdown of the engine-related DOC elements for each baseline aircraft in a pie graph format. These data are for the 185.2 km (100 nm) stage length mission and a fuel cost of $0.264/L ($1.00/gal). Fig- ures 17 and 18 indicate the largest cost item to be fuel (30 to 34% of the total aircraft DOC) followed by engine maintenance (11 to 14% of the total aircraft DOC). In addition, Figures 17 and 18 show the effect of doubling fuel cost to $0.528/L ($2.00/gal). This increased fuel cost drives the engine related cost share from approximately 50 to 60% of total DOC.

Propeller Sensitivity Data Table XV presents propeller sensitivity data determined from the previously discussed engine sensitivity results. This table shows the effect on DOC for the 0.7 Mach aircraft, of a 1% change in three propeller parameters. These data indicate propeller efficiency to be the most significant parameter.

TABLE XI. - ADDITIONAL SENSITIVITY DATA--HIGH SPEED AIRCRAFT (SI UNITS) ^l.

Mfg eoty walght Flyaway cost Block fuel reduction-- DM reduction-- 101 Improvement TOSN ..'64 L reduction-- reductlon-- i4mot b 1n engine reduction-- !A 19791 . A K_ _ 19 Itm k9 kj_ parameters 11 11 0.166 82 27,129 127 41 0.004 Overall sfc 296 13,269 6 2 0.007 0.014 Weight (dry) 136 146 0.012 0.019 OEN price --- 106,226 --- --- .-- Maintenance 0.040 0.060 --- --- ...

o (fully burdened) --- --- ° Max envelope 6 2 0.007 0.013 48 38 12.875 length height elope 3 0.010 0.019 68 54 18.462 9 baseline Characteristics Block fuel (1111.2 km) n 1190 kg TOU n 18.299 kg 380 kg Mfg empty weight • 11,328 kg Block fuel (185.2 ko) Flyany cost • $5,456,141 n DOC (1111.2 km) 2.310 9/sat km n DOC 08S.2 km) 4.726 9/seat b TABLE XII. - ADDITIONAL SENSITIVITY DATA--HIGH SPEED AIRCRAFT (CUSTOMARY UNITS) 101 1oprovement TOW Mfg empty weight Flyaway cost Block fuel reduction-- DOC reduction-- in engine reduction-- reduction-- reduction-- ibm 9/sat no ($1.00/981) lbm 6W to TO m 300 nm juu m parameters lbo 1979 i Overall sfc 658 181 27,129 281 90 0.156 0.307 Weight (dry) 345 322 13.269 13 0.013 0.026 OEM price 106.226 0.023 0.035 Me i ntaMnce (fully burdened) --- 0.074 0.112 Max envelope length 105 83 12.875 14 4 0.013 0.025 Max envelope 151 118 18,462 6 0.019 0.036 height 20 Baseline Characterittics TON - 40,343 lbo Block fuel (600 no) - 2623 lbo 838 lbm Mfg empty weight • 24.974 )bm Block fuel (100 no) • n Flyaway cost $5.456,141 DOC (600 no) • 4.278 9/sat no DOC (100 nm) - 8.752 9/sat no i' UNITS) TABLE XIII. •- ADDITIONAL SENSITIVITY DATA--LOW SPEED AIRCRAFT ( SI DOC reduction-- Tilt .Mfg amply weight Flyewey cost Block fuel reduction-- 104 iq"mmt j1peat b 0.264 L in engine reduction-- reduction— reduction-- j^ k9^- k9 1979 = ^ - paraaetws , 125 32,025 161 41 0.109 0.164 Overall sfc 359 11,250 7 2 0.008 0.011 Weight (dry) 134 125 ! 74,782 --- 0.012 0.016 001 price »- »- --- Ma l ntmore --. 0.038 0.040 (fully burdened) ... ••. --- --- r Max envelope a 7,586 S 1 0.005 0.008 length 36 29 Max envelope 50 41 10,743 7 2 0.008 0.011 height I Baseline Characteristics (1111.2 ka) • 1466 kg T06U n 18,926 kg Block fuel n 11,828 kg fuel (185.2 ka) • 369 kg Mfg eagty weight Bl ock Flyaway cost • $4,835,535 iI DOC (1111.2 ka) • 2.907 9/seat ke DOC (185.2 ka) • 3.972 9/seat ka TABLE XIV. - ADDITIONAL SENSITIVITY DATA- - LOW SPEED AIRCRAFT ( CUSTOMARY UAITS) Mfg empty weight Flyaway cost Block fuel reduction— DOC reduction-- 104 Improvement T06W in engine reductio M- reducticn-- reduction-• lba 9/seat IM ($1.00 /911) l ba 1 Vu na Dom— 1 uu m pararters 1 ba 1979 f n• Overall sfc 792 27S 32,025 365 91 0.202 0.304 Weight ;dry) 296 215 111250 15 4 0.015 0.021 74,762 0.022 001 price --- »- »- »- 0.030 Maintenance --- 0.071 0.074 (fully burdened) --- --- •-- --- Max envelope 0.010 Length 80 64 7,586 11 3 0.014 Mac envelope 10,743 4 0.015 height 130 91 16 0.020 Baseline Characteristics TOW • 41,730 lba Block fuel (600 m) • 3231 The 813 lba Mfg empty weight • 26.076 lbm Block fuel (100 m) n Flyaway cost • $4.83S,63S

DOC (600 =1 • :.383 9/se4t m l

DOC (100 mm) 7.356 d/seat as n Engine share $0.264/1 48%

($1 /901)

♦ Fuel Maintenance 1496 29.5% ^ 1 4% Insurance 0.5%

A/C share

52%

♦ `

50.52f./I Eagi^ae share ($Z/901) 60% Fwl

45.5%

I

Maintenance

I

11%

r

Depreciation 3% insurance 0.4% A/C shoro TE80-2129A Figure 17. — Typical DOG breakdown--high speed aircraft.

Engine share Engine share 62.396`\ 49.5 (S?/psl $0.264/1 ($1/0W) Fuel 50.7% fuel traaau.\ 34%

i

^ Kaincenenee 8 2% 11. IS, \ Depreciation 4% Insurance 0.5% I)eoreciation 3% Insurance 0.4% A/C shore A/C shore 37.7% 50.5% TE8Q-2DM Figure 18. — Typical DOC breakdown- - low speed aircraft.

TABLE XV. — PROPELLER SENSITIVITY DATA Reduction in DOC--% Propeller parameter improvement (DDA 0.7 Mach aircraft)

Ll A propeller efficienzy

0.350 1% p propeller weight 4.22 kg ( ( 9.3 lbm) / propeller) 0.024 1% p propeller OEM price ( $ 509/propeller) 0.006 ADVANCED TECHNOLOGY IDENTIFICATION AND EVALUATION The objectives of this portion of the STAT program were to conduct the follow- ing analyses as applicable to the turbine engine size requirements for 30- and j 50-passenger commuter aircraft: o Conduct parametric engine/airframe studies to optimize cycle and design arrangement o Identify technology advances o Screen and select those with best payoff potential o Define and desc,ibe candidate advanced technology engines i o Determine the payoff potential by comparing the advanced technology engine with both the current technology engine and its 1985 time frame derivative CYCLE SELECTION Airframe sensitivities to engine parameters were developed in conjunction with vehicle specialists as explained under Baseline Airplane and Missions. The airframe sensitivities, or airframe/mission partial derivatives as they are denoted in Figure 19, are obtained from the airframe mission model and applied in combination with engine performance, weight, and cost partial derivatives from the engine cycle/performance model, engine weight model, and engine cost model to obtain a net variation in the payoff parameter. Figure 21 shows this process in detail, using a compressor advanced technology element evaluated in terms of a net change in DOC for the 185.2 km (100 nm) route segment.

Sensitivities to changes in the baseline PD370-37 engine cycle were determined using major segments of the design 1111.2 km (100 nm) mission for the high speed airplane. These major segments were determined from the DDA mission analysis model. The results from this model correlated closely with those of Lockheed. Table XVI shows the total fuel burned in the mission and the per- cent of the total burned in each mission segment. Climb, cruise, and loiter use the major portion of the fuel. Power, speed, altitude, and power setting are defined for these major segments at the bottom of Table XVI. Since climb and cruise/loiter mean operating conditions for both the design and alternate missions were almost identical, the six operating conditions were reduced to four for the sensitivity analysis.

To develop the sensitivities to changes in the baseline Model PD370-37 engine cycle, the parameters shown in Table XVII were changed individually at each of the four mission operating conditions shown in Table XVIII. The baseline en- gine performance was obtained by running the mission operating conditions at the BOT's shown in Table XVIII.

The effects of these engine cycle parameter variations on engine performance for one of the four operating conditions are shown in Tables XIX and XX. The representative mission fuel used (approximately 81% to 89%) was determined by using the mission segment times shown in Table XVI. The sensitivity of 185.2- and 1111.2 km (100- and 600 nm) mission fuel used to the changes in engine cycle parameters is shown in Tables XXI and XXII. The change in fuel used, from that of the baseline engine, expressed in percent, is the mission weighted percent change in sfc from those of the baseline PD370-37 engine.

These sensitivities are shown graphically in Figures 20 through 22.

t oil - Nip I j Alrfnome#ftion j Uise {1119 PeAW W^ OerNNNes I end Cost = I Partial Engine CyclelPerformem Oerketlws ; Model a IIOC a sic - _ loo NM . ooOC Advanced

qc"

X -- q M A ^siE ^emo IMP Technology IdeMMkeNon ^ooc or r Resulting Comoressor A a en91 _^^epht x a

MNM -

Cheap in aoc t • with Adwnced whh Advanced cow; aenginewitipt Comoressor Efficiency 1 Technology Element 6IRComo Technology Element

114 Como _"/

a engine cost a DOC 100 NIA

• I Doccost

^e X - a Engine C7 a + Engine Weight ► co Como mo Model , • f Engine Costl Mission 1 i (\ Missbn Model Model e • TE-6171 Figure 19. - Methodology using sensitivity data.

XVI. - DDA MISSION ANALYSIS RESULTS--CURRENT TECHNOLOGY TABLE HIGH SPEED AIRCRAFT 185.2 km Alternate mission (100 nm) mission 1111.2 km (600 nm) Design Time-- Fuel-- Time-- Fuel-- Fuel--% min kg (lbm) Fu g el-_% min kg (lbm) Mission phases - - 3.1 Take-off 1.0 - - 1.9 1.0 12.4 15.7 5.1 - - Climb 15.6 - - 17.2 43.3 10.5 - - Cruise 77.8 - - Fuel Reserves 1.9 1.0 - - 3.1 Take-off 1.0 - - 6.2 3.5 2.3 - - Climb 2.1 - - - - 18.8 Cruise 20.7 - - 11.2 20.0 - - 39.2 45.0 - - 22.5 45.0 Loiter 1156.4 100.0 163.2 1932.7 100.0 84.9 Totals (4260.8) (2549.5) power--both missions Climb--6,096 m (20,000 ft),1 0.5 MN /climb Note: Cruise--10,668 m (35,000 ft)/0.7 M N /86% max con y --design mission 6,096 m (20,000 ft)/0.7 M N /83% max conL--alternate mission (10,000 ft)/0.4 MN /27% max cons--both missions Cruise/loiter- -3,048 m t TABLE XVII. - CYCLE PARAMETER CHANCES Change Cycle parameter + 3% and -6% Compressor efficiency + 3% HP turbine efficiency I.

+ 3% LP turbine efficiency + 25% Turbine cooling air + 25% Burner pressure drop TABLE XVIII. - MISSION OPERATING CONDITIONS Baseline engine Baseline engine ( O R) gower-kW (shp) Mach No. BOT--K Altitude--m (ft) Condition 1456 (1953) 0.4 1036 (1865) Cruise 3 , 048 (10,000) / loiter (2664) 3811 (5110) (20,000) 0.5 1480 6,096 Climb 3747 (5025) 0.7 1343 (2418) 6,096 (20,000) Cruise 2174 (2915) 0.7 1311 (2360) 10,668 (35,000) Cruise TABLE XIX. - PD370-37 SENSITIVITY STUDY OF CYCLE CHARACTERISTICS AT 6096 m CRUISE ( SI UNITS) (altitude: 6096 m; standard day; velocity: 0.7 MN; power: cruise) LP turbine q Turbine cooli Burner AP Compqssor HP turbine ' - T + 25 2S Lu uIatiwe• Sensitivity Change--S 99.9 IUU.0 '19.9 $V.v 99.9 99.9 9v.v 100.0 100.1 99.9 99.9 N/.R compressor-4 99.9 12.10 12.85 11.14 11.11 U. a4 11.4u 11.Ju 12.75 -kg/S 12.12 12.87 13.35 12.22 Mi 18.47 19.6.1 19.46 Ib.e4 lv.14 Id.9J 17.1b 18.50 19.65 20.38 18.65 19.46 M /E--kg/s 11.49 11.49 11.49 1.49 11.49 11.47 11.49 11.49 11.48 11.50 11.52 11.48 Rcomp 8J.6.1 ds.dJ 84.61 dJ.6J 80.4b 83.41 83.1[ 1 ^p .-% 86.35 81.29 78.75 83.83 83.81 .15 5.17 J.lu J-lu 4.14 4.13 4.14 4.14 4.13 4.14 4.13 4.14 4.14 /opyrn--t .1.b 1312.7 1311.8 1311.7 131J.J Ul lJlc.b IJI.1..1 1311.4 1312.9 1311.4 RIT--1` 1312.7 1312.5 99.4 99.4 99.4 99.4 99.4 99.4 99.0 99.4 99.5 99.5 99.6 99.4 N/'R HPT-- % 66.19 66.19 dd.19 66.19 d6.11 9U. di 85.54 8d.19 88.18 88.19 88.20 90.83 'IHPT-4 4.0 4.0 4.0 J.0 4.0 b.0 3.0 4.0 4.0 4.0 4.0 4.0 Cooling HP turbine-- % HPC inlet air 101.9 101.9 lul.9 101.9 lul.9 lul.9 lul.9 1C11.9 101.9 102.0 101.9 101.9 N/0 LPT--t 17.77 91I.7U 9u. .1b 9u.bl 9U.sb 91. v4 93.19 90.01 90.97 91.46 90.01 90.97 1LPT--t 1.0 1.0 1.0 i..1a 0.75 l.0 u.75 1.0 1.0 1.0 1.0 Cooling L â turbine-- 1.0 t HPC inlet air 61.ul 61.88 04.33 63.41 64.[4 bJ.bb oy.47 62.89 65.09 66.45 62.67 65.17 SFC--bg/M-s Note: All runs made to constant (base value) shaft power.

4LpT, -255 TC , +3S -25% burner JP.

3% +1c, +3i '1HPT.

TABLE XX. - PD370-37 SENSITIVITY STUDY OF CYCLE CHARACTERISTICS AT 20,000 ft CRUISE (CUSTOMARY UNITS) N ; power: cruise) (altitude: 20,000 ft; standard day; velocity: 0.7 M Cosqssor L NP turbine LP turbine Turbine cooli ourner ly ^6 kumulative* Sensitivity Change--t ^ 98.9 99.2 vir.9 100.0 y y9.9 11/4F compressor-4 99.9 100.0 100.1 99.9 99.9 99.9 99.

27.64 27.644 t.4. VI 26.72 28.38 29.43 26.94 28.12 26.68 28.34 26.08 16.91 Mi -lb/sac 40.79 43.32 44.92 41.13 41.92 Q.4 44.25 42.d6 41.11) 41.19 41.7c 46.1)1 N,^/ --lb/sec 11.52 11.48 11.49 1 1 .49 11.49 11.49 11.49 11.49 11.49 11.47 11.46 11.50 8.1.82 83.82 d3.d4 13.61 b9.40 r pp - -s 86.35 81.29 78.75 83.83 83.81 41.41 64.41 ^pp 4.14 Aa /Pd --i 4.13 4.14 4.14 4.13 4.14 4.13 4.13 4.14 5.17 4.lu 4.10 re . 2362.8 2362.5 2360.6 2363.3 2360.6 2361.6 23b3.0 24.1.1 t4W. y e.301.7 13o3.0 10ol.0 RIT -I 99.4 99.6 99.4 99.5 99.4 99.4 99.4 We 99.4 96.4 99.4 99.5 N14TNPT--t 86.19 88.19 86.19 &1.19 W.ld 90.64 " N PT--t 88.18 86.19 88.20 90.83 85.54 88.19 4.0 4.1) 4.0 4.0 4.0 4.0 4.0 4.0 b.0 3.1) 4.1) 4.0 Cooling NP turbine-- % NPC inlet air WNT LPT --t 101.9 101.9 102.0 101.9 101.9 101.9 101.9 101.9 101.9 11)1.9 iUi.V iUI.9 VT- 90.01 90.97 91.46 90.01 90.97 93.19 47.77 90.7U 90.16 9u.ol 9u.3b 91.94 -s 1.0 1.0 l.lb U.7b l.1) l.1) u.7b Cooling LP turbine-- 1.0 1.0 1.0 1.0 1.0 t NPC inlet air • u.34ue u.3757 0.3802 u.37ol u.4b14 SFC--lbm/hr NP 0.3722 0.3852 0.3933 0.37U9 0.3863 0.3070 0.38119 Note: All nuns side to constant (base value) shaft power.

*+3% Tic, +3S'I NPT, +31'ILPT• -251 Tc, - 255 burner 4P.

TABLE XXI. - PD370-37 SENSITIVITY STUDY OF MISSION FUEL CHANC=ES (SI UNITS) Burner %P Com ressor NP^^ tur b ^ inno LP turbine a Tuurrbine^cooling ^ + _ _ +3- T T +Z3 - -95 _ -d b wsulative* Sensitivity change--% Baseline 185.2-km mission--fuel used-49 510.5 4bb.b ,N 518.4 S40.0 564.6 497.8 542.5 503.7 $34.6 Sc1.8 36.3 bll.b Cruise/loiter 500.2 143.9 152.9 149.3 147.3 149.1 147.o 1x7.9 Climb 148.3 145.9 151.2 154.4 145.5 151.6 15u.0 14U.c 150.7 148.5 153.8 156.9 148.1 154.1 146.5 155.6 151.9 151.8 1bu.l Cruise 794.1 843.1 829.0 807.8 d26.1 dlu.c 733.7 Total fuel--kg 817.4 794.6 844.9 876.0 791.3 848.2 3.76 -2.86 3.14 1.41 -1.18 1.Uo -u.bv -lu.14 Change in fuel--% -2.78 3.37 7.16 -3.19 1111.2-km mission--fuel used--k9 570.7 °03.1 548.3 509.1 540.4 533.4 S16.0 530.9 blb.0 4W.o Cruise/loiter 524.0 50S.6 545.9 X5.7 316.7 323.6 304.8 317.6 3013 320.3 312.8 308.7 312.4 309.1 Z69.0 Climb 310.6 667.0 634.0 656.4 625.8 664.6 648.2 641.1 646.1 641.1 buc4 Cruise 644.6 1035.8 654.6 1314.7 1447.1 1517.2 1561.2 1442.0 1522.4 1436.5 1525.4 1494.4 1465.9 1468.9 1468.9 Total fuel--kg 1479.2 -2.17 2.57 5.54 -2.52 2.92 -2.89 3.12 1.04 -0.9u 0.84 -U.7u -0.01 Change in fuel--% ^ *+3t *I C . +3S -255 Tc, -25; burner JP.

114PT. +3S 'ILPT• TABLE XXII. - PD370-37 SENSITIVITY STUDY OF MISSION FUEL CHANGES UNITS) (CUSTOMARY com p ressor L NP turbine 17 LP turbine Turbine^coolI Burner .0 Sensitivity Change-4 Baseline + _ ^7 X35 ^* 5 - : Z^ Lumulative* _ 100-NI mission--fuel used--lbm CruiseA otter 1142.9 1102.8 1190.5 1244.7 1097.4 1196.0 1110.4 1178.6 110!.5 llo.b II".0 i119.v 1U04.0 Climb 326.9 321.7 333.3 340.5 320.8 334.2 317.3 337.1 329.2 3c4.8 US.7 31b.3 341.1 Cruise 332.3 327.4 339.0 346.0 326.4 339.7 322.9 343.0 334.9 34U.6 334.6 330.v 3u9.0 Total fuel--lbm 1802.1 1751.9 1862.8 1931.2 1744.6 1869.9 1750.6 1858.7 1827.6 1780.9 1411.3 17bb.1 1017.o Charge in fuel--t -2.78 3.37 7.16 -3.19 3.76 -2.86 3.14 1.41 -1.18 1.06 -u.69 -10.L4 600-0 mission--fuel used--lbm Cruise/loiter 1155.2 1114.7 1203.4 1109.2 1206.9 1258.1 1122.4 1191.4 117o.0 1137.7 117u.b 1142.1 lulb.4 Climb 684.8 674.0 698.3 713.4 672.0 700.3 664.8 706.2 b89.7 680.6 688.8 681.0 037.1 Cruise 1421.2 1401.7 1443.2 1470.4 1397.8 1447.1 1379.7 1465.3 14c9.0 1413.4 14c9.0 1414.7 14c7.b Total fuel--lbm 3261.2 3190.4 3314.9 3441.9 3179.0 3356.3 3166.9 3362.9 3294.7 3231.7 3288.3 3130.3 Oi.U.4 -s Change in fuel- -2.17 2.57 S.54 -2.51 2.91 -2.89 3.12 1.03 -0.9u U.ba -u.7u -0.01 1 Tc, -2511 burner AP.

43% 11C• +3% *1NPT• +31'ILpT, -251 g M C -6 -4 '2 0 +2 +4 Change in component eHcfency --S TE80.2bl Figure 20. - Engine sf c sensitivity to component efficiency--high speed airplane, alternate mission.

Sensitivities to changes in the baseline Mode! PD370-37 engine cycle were also determined using major segments of the design 1111.2 km (600 nm) mission and of the alternate 185.2 km (100 nm) mission, for the low speed airplane. These sensitivities were determined in a similar manner to the methods described for the high speed airplane. It was found that sensitivities to compondne effi- ciency, cooling air, and burner pressure drop were essentially the same for the low and high speed airplane missions.

A matrix of engines was postulated at the nominal STAT engine sizes of 1790 and 3579 kW (2400 and 4800 shp). Baseline compressor, high pressure turbine, and low pressure turbine efficiencies and cooling air amounts were assumed as indicated in Tables %SIII and X%IV for each engine in the matrix. The matrix included compressor pressure ratios of 5, 10, 15, and 25 over a range of tur- bine rotor inlet temperatures (RIT) as shown.

Compressor efficiencies were specified for the STAT nominal size engines with consideration for clearance efiects using axial staging arrangements.

+6 0 0-- Compressor T -.0 0---0- HP turbine 0-- LP turbine IN "Current technology" baseline +2

N

N 0

^E c 1111.2 km 1600 nm! mission

r

-2

v

-41L -6 -4 +4 -2 0 +2 Change in component efficiency- % 1E80-2063A Figure 21. - Engine sf c sensitivity to component efficiency- - high speed airplane, design mission.

TABLE XXIII. - STAT SENSITIVITY STUDY--CYCLE PARAMETERS FOR 1790 kW (2400 shp) ENGINE Compressor q (polytropic) 88.0% RIT--K °F 1783 (2750L 1 950 Rc 1506 50 16 88.4 HP turbine -9 5 88.8 88 . 6 88.3 87.2 86.9 86.7 (adiabatic)--% 10 87.5

15 86.4 86.1 85.7 85.5

25 85.3 84.8 84-5 84.1 5 90.1 89.9 89.8 89.7 LP turbine-9 10 89 . 3 89.0 88.7 88.4 (adaibatic)--% 89.0 88.6 88.2 87.8 25 89.1 88.5 88.0 87.6 2.67 Cooling air--% 5 6.57 9.52 14.1 4.53 8.8 13.7 10 3.38 6.40 8.80 13.9 15 16.1 12.0 16.7 25.6 27.7 185.2 km 1100nm1 mission +2 v

b-^

Turbine cooling 0-^ Burner pressure drop -2 "Current technology' baseline N 1111.2 km 1600 nm1 mission c +2 -2 • - +20 -10 0 +10 -20 Change in turbine cooling air/burner pressure drop-% r90-20W Figure 22. - Engine sfc sensitivity to turbine cooling air and burner pressure drop- - high speed airplane.

TABLE XXIV. - STAT SENSITIVITY STUDY-CYCLE PARAMETERS FOR 3579 kW (4800 shp) ENGINE Compressor sl(polytropic) 89.0% RIT--K (°F RC 1506 (2250) 1644 (2500) 1783 (2750) 1950 3 92.0 BP turbine 5 92.1 91.9 91.8 90.8 90.6 90.4 90.2 (adiabatic)--% 10 90.0 89.7 89.4 89.1 25 89.0 88.4 88.1 87.7 5 91.8 91.7 91.6 91.5 LP turbine n 10 91.8 91.1 90.9 90.8 (adaibatic)--% 90.8 90.6 90.3 15 91.2 25 91.2 90.8 90.5 90.1 TABLE XXIV. (CONT) RIT--K (°F) 1506 (2250) 1783 (2750) 1950 (3050) Re 1644 _ (2500) Cooling air--% 5 2.68 6.57 9.52 14.1 3.38 4.53 8.8 13.7 6.40 8.80 13.9 16.1 12.0 16.7 25.6 27.7 HP and LP turbine efficiencies and cooling flow rates were projected to re- flect the effects of engine flow rate, cycle pressure ratio, and turbine RIT upon the turbine performance. coolant flow rates and efficiency levels The were the result of the changing turbine physical size, work requirements, and number of stages. Coolant flows were established consistent with common life requirements based upon cycle temperature, cooling air termperature, stage work requirement, and number of stages that require cooling. The associated cooling effectiveness was chosen with a consideration given to the significant impact of initial engine cost on DOC. Exotic cooling schemes and their atten- dant high engine costs were considered only at the highest temperature levels where they are an absolute necessity.

The effect of these cycle considerations on relative mission weighted-sfc is shown in Figures 23 and 24. The results tend to localize the region of in- RIT terest for both engine sizes at 15 to 25 pressure ratio at from 1506 to 1644 K (2250 to 2500 0F). These results were determined by adjusting design point sfc's calculated for the engine matrix for changes in compressor and turbine efficiency and cooling air quantities from the nominal values used in generation of the sensitivity deca to the values shown in Tables XXIII and XXIV. The mission fuel weighted sensitivities were used as previously de- scribed. Minimums occured where improved cycle efficiency due to increase in pressure ratio and turbine temperature was overcome by reduced component adia- batic efficiency and performance penalties due to increased cooling air.

Further study indicated that an RIT of 1506 K (2250°F) was an acceptable de- velopment risk for the 1988 time period. The risk attendant with higher RIT was judged excessive for the small improvement in sfc realized.

At a turbine temperature of 1506 K (2250°F), further parametric studieL were made to determine the effect of compressor pressure ratio condsidering engine weight, cost, and dimensions in addition to mission-weighted sfc. Table XXV presents these results for the 3579 kW (4800 shp) engine in terms of percent change in weight, cost, length, diameter, and sfc from a 4800 shp reference engine. Using the sensitivity data developed for the high speed commuter air- plane, engine characteristic changes were converted to a percent change in DOC at 185.2 km (100 nm) and plotted in Figure 25. These results indicate that minimum DOC is obtained at a pressure ratio of approximately 20:1.

Summarizing the results of the cycle analisis, an RIT of 1506 K (2250°F), and a compressor pressure ratio of 20:1 were judged reasonable selections for both the 1190- and 3579 kW (2400- and 4800 hp) engines.

i.4 i 1.3 Base: 2400 shp reference engine.

V N 1.2 Y %* 1950 K C30 . ► C^2 K gar .. 17 o ^ ^(2750°F) 150 ^ `^ ^^ ..... ........

....^• I M O'F) K 0.9 0.8 5 10 15 Design pressure ratio TEED-2067A Figure 23. - Mission weighted sfc trends- - 1790 kW ( 2400 shp) size.

TABLE XXV. - STAT ENGINE PRESSURE RATIO SELECTION 3579 kW (4800 shp) SLSS 1 2 3 4 5 10 15 25 Re 1506 (2250) RIT--K ( OF) 1506 (2250) 1506 (2250) 1506 (2250) % weight -10.0 -21.9 -23.8 -21.2 % o cost -8.3 -32.1 -20.5 -18.11 % 0 length + 3.98 - 8 .5 -11.0 -i0.4 +2.4 +0.0005

% 0 diameter +14.7 +2.1

% A sfc +29.0 +2.0 -8.0 -7.1 % A D OC ( 100 nm) +9.49 -4.24 -6.467 -5.66 1.4 a Soo: 4800 shp reference engine 1.3 ^.

t I

u

H 1.2 .:

1.0

---`

s apt

1950K00

1506 K (22501) ^'....

1 ' ^1 :le 0.4 ..•^ • - 1 1^ K (2500°F) 1 ; 0.8 5 10 15 25 Design compressor pressure ratio TE80-2065A Figure 24. - Mission weighted sfc trends--3579 kW ( 4800 shp) size.

+20 DOC trends with pressure ratio 1506 K (22507) 3579 kW (4800 shp) +10 "Current technology" baseline c W V er °- -10 Base: 4800 shp reference engine I (see Fig. 241 -20 5 10 15 20 25 Compressor pressure ratio 'tE80-206M Figure 25. - DOC trends with pressure ratio at 1506 K ( 2250 ° F) and 3579 kW (4800 shp) conditions.

CONFIMRATION TRADES An extensive list of advanced technology candidates for incorporation in the STAT anginas was prepared early in Task II. These candidates were screened and correlated in four major areas: o Rotating components o Static components o Propulsors and drives o Nacelles and accessories Potential technology items were first screened to determine those for which sufficient technical information and background experience existed at DDA to conduct a trade study. Another criterion considered in this initial screening was the extent of effort required to obtain baseline data upon which to make tradeoff assessments. In some cases, the effort required was judged to be beyond the scope of this study program. Advanced technologies associated with propellers were furnished to DDA by NASA.

After this initial screening, individual components in each area were investi- gated to determine the applicability of the technology and the estimated change in component characteristics (efficiency, weight, cost, life, ate). At the same time, the associated risk level was noted to aid in the screening of those technology elements with the greatest promise. Emerging advanced tech- nology areas were investigated to determine their potential 'mpact on power- plants sized to meet the STAT commuter aircraft requirements, particularly in areas that were unique such as the shorter duration flight cycle, and emphasis on reduced initial costs and operating costs.

The study engine characteristics were also evaluated in terms of advanced de- sign features that might be incorporated into the STAT engines. The same screening process and study limitation criteria that were used in selecting technology trade study candidates were applied to determine the design fea- tures to be evaluated. Each design feature selected was studied with respect to the measureable improvements in components, cycle, or engine compared to the STAT baseline engines. As with the technology elements, the associated risks were judged and entered into the evaluation.

The baseline engine sensitivity parameters were applied to technology item and design features to generate the resulting partial derivatives of engine per- formance/cost parameters (sfe, initial cost, weight, etc). Those judged to have the most merit were subjected to the airframe sensitivity analysis. Each selected design change (in terms of engine performance parameters) was then checked to determine the resized airframe (DOC), empty masts, acquisition cost, fuel consumed, etc, for the baseline mission. Those design features with the greatest merit were chosen for further evaluation for the STAT engines.

The lists of technology items and design features that were finally evaluated, are shown in Table XXVI. This table indicates the recommendation: applicable to each item indicating whether it was selected for incorporation in the STAT engines or was rejected. The table also sh,-vs whether this decision was based on a DOC evaluation or was one based on judgment since the DOC impact could not be ascertained.

TABLE XXVI. - CANDIDATE ADVANCED TECHNOLOGIES AND DESIGN FEATURES -i--N-N-----N--N---- CANDIDATE ADVANCED TECHNOLOGIES --------- Selected Rejected Judge- Judge- DOC meat DOC sent Technoloty Item Power Section o Compressor X Cycle pressure ratio X Configuration (axial b axial/cent.)

Hybrid centrifugal impeller X High temperature titanium aft wheels X Melded titanium spool X o Turbine X Rotor inlet temperature Configuration - Axial X Radial inflow X Hybrid rotors, composite shafts b X suFercritical shafts Thermal barrier coatings X Ceramic blades S vanes and airfoil coatings X X Cast-in impingement cooling Long-life bearings X o Combustor/diffusers X Transpiration cooled combustor X Vortex controlled diffuser o Engine accessories Electronic fuel control X Fuel pump and metering system X X Engine condition monitoring o Noise reduction Compressor design X Combustor design X Reduction Gear o Advanced design X Composite materials Steel/titanium gears X X Finite element gear analysis Superplastic formed titanium X TABLE MI (CONT) ADVANCED DESIGN FEATURES--- ----•------------------ CANDIDATE Sect*!

Judl DOC sent Desian Feature Power Section o Compressor Erosion resistant design X Inlet particle separator X X Rotor/case response to rotating stall Clearance control - Active Straddle mounts X X Thermally matched rotor/case o Turbine Clearance Control - Active X Straddle mounts Thermally matched rotor/case X Leakage control X Abradable coatings X Advance Bearings - Beryllium-backed bearing races Tapered bearings Other Modular construction X Remote accessories X Advanced propellers X Turbofan engine X Power extraction, mechanical vs air bleed Reduced weight nacelles X The following paragraph describes the analyses by engine section. The impacts on engine weight, performance, maintainability, and cost are given where tradeoffs could be made; otherwise decisions were based on judgement.

Compressors Compressor configurations for both engines were determined and are shown in Figures 26 and 27 for the 1190- and 3579 kW (2400- and 4800 hp) engines, re- spectively. The compressor pressure ratio selected was 20:1, as previously discussed. In selecting these configurations, the experience derived from recent compressor studies, including the small compressor study for NASA/LeRC, was employed. For tradeoff examination of alternate compressor arrangements, trends were studied that show adiabatic compressor efficiency versus design flow for axial, axial-centrifugal, and one- and two-spool, two-stage centrifu- gal compressors. The compressors were selected to be axial-centrifugal and axial, for the 1790- and 3579 kW (2400- and 4800 shp) engine2, respectively.

aerothermodynamic design data showed axial compressors Preliminary compressor 6 0.15 Axial/centrifugal, 70/30 split Rc - 20:1 Wo - 6.411 kg/s 04.135 lb/sec) 4 0.10 E

daoc^=

W .

2 0.05 0 L .4 Axial dimension—m 14 16 4 8 10 12 2 6 Axial dimension—in.

TE80-2045A Figure 26. - STAT advanced compressor flow path--1790 kW (2400 hp) engine.

at the study pressure ratio have higher design efficiencies than other configurations over the study range of flow rates. An axial compressor was configured for both engines initially. The small compressor, however, was found to have extremely small airfoils at the aft end. These small airfoils would be difficult to manufacture as well as to make the compressor efficiency very sensitive to small changes in clearance. For this reason an axial-cen- trifugal compressor was selected for the small engine. An axial compressor was selected for the large engine.

Hybrid Compressor Impeller Rotor An evaluation was made of the hybrid centrifugal compressor impeller rotor.

This item uses hot isostatic press (HIP) bonding to attach a cast rim with blades to a forged bore insert.

A recent value engineering study of the Model 250-C30 impeller rotor was made in which a similar substitution was considered. This ana'.ysis showed that the cost reduction potential was 71%. Using this ratio and estimated cost of the 1790 kW (2400 shp) STAT engine compressor impeller, the resulting change in engine cost would be -2.3% with a corresponding change in DOC of -0.25X.

This item was selected for the STAT 1790 kW (2400 shp) STAT engine.

R = 20:1 C W = 11.57 k4/s (25.5 lb/sec) 6 0.15

q a

O^^DOOOOO

q

4 0.10

0^0

q

c S s I o.o5 oL c 0.40 0.451 0.30 0.35 0.15 0.20 0.25 0 0.05 ^v. 10 n--m L , Axial dimemla 16 is 10 12 14 6 8 4 0 • Axial dimemion—in.

TESO-2046 Figure 27. - STAT advanced compressor floe path--3579 kW (4800 hp) engine.

Hixh Temperature Aft Titanium Wheels In both STAT engine sizes the cycles selected have compressor discharge tem- .7 to 744.3 K (870° to peratures, at rated conditions in the range of 738 880°F). At these temperatures, alloys such as 6-4 Titanium have little creep strength. These alloys are thus inappropriate for wheels to be used in the latter stages of the compressors.

IMI829, a high temperature Titanium alloy developed outside the U.S., appears to be applicable to this aft wheel location, based on preliminary data. How- ever, data on alloy weldability is not available, and this may impact its se- lection for a welded drum rotor construction. The application of IMI829 to the 2400 shp compressor is uncertain at this time due to the hybrid construc- tion of its centrifugal impeller.

Welded Titanium Spools Advanced compressors in development today use welded drums to provide greater stability and reduced weight. Improvements can be made in design and fabrica- tion technology to further reduce weight and cost by simplifying the configu- ration to facilitate welding in remote or blind areas. Based on experience with other DDA turboshaft engines, it is estimated that the following savings may be achieved with the STAT engines: X 1 DOC % I Weight % J Cost Rated power -- kW sh -0.6 -0.093 1790 (2400) -1.0 -0.196 -1.8 -1.1 3579 (4800) Erosion Protection DDA experience with T56 engines operating in desert climates has shown that severe sand and dust erosion damage to an axial compressor can cause the en- gine to be removed for overhaul in just a few hundred hours.

The STAT engines will be designed to be more tolerant of infested dirt par- ticles in the airstream than current technology engines. If the STAT engines operate in an extremely dusty environment, their normal MTBR may be reduced by one-half. The addition of suitable inlet particle separators (IPS) would re- store their normal MTBR, and would provide the following inpact to the cost and operation of the engines: %.1 Maint.

% X .1 Cost X .1 sfc Cost .1 DOC % .1 Weight Rated pow,.r--kW (hp) -14.4 -0.818 +5.0 +2.5 +1.0 1790 (2400) -0.633 -10.4 +3.0 +2.0 +1.0 3579 (4800) Although the IPS adds to the eagine weight, cost and fuel usage, the gain in lowered maintenance coat could more than offset this disadvantage.

Prediction of Rotor/Case Response to Rotating Stall The small diameter, highly loaded compressors of the STAT engines require close blade tip clearances to achieve the desired performance. It is es- sential that the design of these advanced compressors consider the dynamic behavior of the rotors and cases during surge, rotation stall, and rapid ther- mal changes. Without the ability to predict these phenomena, it would be nec- essary to design with greater tip clearance and to provide either an addition- al compressor stage or an additional bearing and attendant support structure.

With the ability to predict the rotor/case response to rotating stall, how- ever, these penalties could be removed with the following typical improvement to DOC: % A sfc X A DOC X O Weight %,& Cost Rated power--kW (hp) -0.8 -1.334 -5.0 -8.0 1790 (2400) -0.8 -1.234 -6.0 -6.0 3579 (4800) -----.^•-....,...^ .. _ ^. w ...

.._„^,.,, ,.-........r^...,.^.,....-.-,.,....-rte _ t Clearance Control The payoff per stage with compressor active clearance control is such that stage efficiency improves 3% for each 1% of blade height change in tip clear- ance. The highest payoff occurs in the aft compressor stages where blade height is smallest. This complicates the design, particularly where a vane actuating system is desirable from an acceleration/surge margin point of view.

Simple mechanical systems are essentially eliminated if a vane actuation eye- tem is already employed. Thermal systems probably would not pay off since the "on" system time is short and because of the penalty to the cycle. Active clearance control was, therefore, rejected for the STAT engines. Straddle l rotor mounts and thermally matched rotors, blades, vanes, and the case were incorporated in the STAT engines, however, to provide a degree of clearance control. It is estimated that blade tip clearances in this way may be held 15% smaller with the following impact on DOC: i Rated power--kW (hp) %.I Weight %A Cost %,j sf c %A DOC 1790 (2400) -5.0 -8.0 -0.4 -1.141 3579 (4800) -6.0 -0.4 -1.094 -6.0 Turbines Advanced technology baseline engines were developed for examination of the trade-offs involved in choosing turbine and shafting arrangements compatible with high pressure ratio, axial, and axial-centrifugal compressors. The tur- bines were air-cooled to operate at 1506 R (2250°F), as was shown desirable in the previous section.

Configuration Axial turbines were considered initially in the study for both size engines.

Engine weights and prices were analytically determined by section, based on cycle descriptions, technology levels, and unique physical features. Cost data were estimated using the Material Index Factor (MIF) method employed by DDA.

1790 kW (2400 shp) STAT Engine A baseline power section and eight variants were defined to compare different combinations of LP turbine shaft designs, Hp turbine materials, and methods of HP turbine blade attachments. The baseline configuration and variants are described in Table XXVII. The baseline configuration has a 20:1 CPR compres- sor at an airflow of 6.35 kg/s (14.0 lbm/sec) at approximately 38700 rpm.

Tables XXVIII and XXIX show the comparative calculated weights by section, and also show calculated comparative recurring manufacturing prices.

^- 49 i TABLE XXVII. - 1790 kW (2400 shp) ENGINE FABRICATION TECHNOLOGY Engine Technology Adv. Tech.

Baseline Composite, or hybrid, HP turbine wheels and blades r (pcwdered metal discs and cast rim with integral blades), beryllium--reinforced LP turbine shaft Forged IN-73.8 HP turbine wheels with dovetails and a 1 beryllium-reinforced LP turbine shaft, 8% lower HP rpm r Near-net-shaped PA-101 HP turbine wheels with dovetails 2 and a beryllium:-reinforced LP turbine shaft, 8% lower HP rpm Composite borsic titanium LP turbine shaft and composite HP turbine wheels without dovetails Composite HP turbine wheels without dovetails and steel LP turbine shaft without beryllium insert (has mutes for damping) Forged IN-718 HP turbine wheels with dovetails and a composite borsic titanium LP turbine shaft Forged IN-718 HP turbine wheel with dovetails and a steel LP turbine shaft without beryllium insert (has mutes for damping) Near-net-shaped HP turbine PA -101 wheels with dovetails and a composite LP turbine borsic titanium shaft Near-net-shaped HP turbine PA-101 wheels with dovetails and a steel LP turbine shaft without beryllium insert (has mutes for damping) TABLE XXVIII. - 1790 kW ENGINE POWER SECTION PRICES AND WEIGHT BREAKDOWN (SI UNITS) Adv. Tech.

1 2 3 4 5 0 7 d Baseline 11.45 12.93 12.93 11.45 11.45 11.45 11.45 11.45 11.4b Forward support-49 13.34 13.34 11.81 11.81 11.81 11.61 11.61 11.81 Compressor rotor--kq 11.81 19.56 22.10 22.10 19.56 19.56 19.56 19.56 19.56 19.56 Compressor case--k9 27.59 27.55 27.59 27.59 t7.b9 17.59 Burner/diffuser--k9 27.59 31.18 31.18 15.56 20.23 15.65 13.08 13.14 16.76 15.56 14.8d 14.3[ HP turbine rotor (5.72) (3.15) (3.11) (6.db) (5.0/) (4.yA) (4.391 (wheels and spacer)--k9 (5.62) (10.30) 10.37 10.37 9.18 9.18 9.18 9.18 9.16 9.18 HP turbine case--k9 9.18 16.61 17.31 16.61 17.31 16.61 11.31 LP turbine rotor (shaft) --kg 18.74 18.74 18.74 (3.83) (3.83) (3.83) (1.25) (1.96) (1.25) (1.96) (1.25) 0.96 ► 9.17 9.17 9.17 LP turbine caso-Ag 9.17 9.17 9.17 9.17 9.17 9.17 11.34 --kg 11.34 11.34 11.34 11.34 11.34 11.34 11.34 11.34 Near burner support 63.75 63.75 63.75 63.7b 63.75 63.7b 6.s.75 Accy 9Nrbox --kg 63.75 63.75 Total--k9 196.13 213.10 208.57 193.53 194.30 197.23 196.71 195.42 195.47 Recurring Manufacturing price $236.904 $60,660 $249.763 (235.547 u35.b47 f249.jd9 $246.439 U4u.u15 $la9.447 L TABLE XXII. - 2400 shp ENGINE POWER SECTION PRICES AND WEIGHT BREAKDOWN (CUSTOMARY UNITS) I Mr. Tech.

8a5alino 1 2 3 4 S 6 7 d Forward support--lbw 25.24 28.5 28.5 25.24 2a.24 25.24 25.24 25.14 25.24 26. u3 comproswr rotor--lbw 26.03 29.4 29.4 26.03 26.03 26.03 26.0o 20.03 Compressor cote--lbw 43.12 46.73 48.73 43.12 43.12 43.12 43.12 43.11 43.11 Burner-diffuser--lbw 60.83 68.74 66.74 60.63 60.81 60.83 60.83 00.63 60.k" NP turbine rotor 34.30 M.60 34.5 28.64 26.97 37.00 34.3 32.80 31.57 (wheels and spacer) --lbm 02.4) (22.7) (12.6) (6.94) (7.07) (15.1) (12.4) (10.9) (9.67) NP turbine Casa-1 bm 20.24 22.87 22.87 20.24 20.24 20.24 20.24 10.24 20.24 LP turbine rotor (shaft)--ibm 41.31 41.31 41.31 36.61 38.17 36.a1 38.17 36.01 36.17 (7.46) (7.46) (7.46) (2.76) (4.32) (2.76) (4.32) (2.76) (4.32) 20.11 LP turbine cases-lbm 20.21 20.21 20.21 20.21 20.21 20.21 20.21 20.11 Rear burner support--lbw 24.99 24.99 24.99 24.99 24.99 24.99 24.99 24.99 24.99

13 19.565 1 140.SS 140.56 140.6a 140.5b 140.bb

Accy gearbox--ibm 5 . 5 140.55 Total--lbm Recurring manufacturing price $236.904 $260.660 $249.763 $235.547 $235.547 $249.stl9 $248.409 $W0.Wb $23y..141 The effect that the LP turbine shaft design changes had on the finished weight (FW) of the LP turbine and the FW of the HP turbine, due to varying shaft and wheel bore diameters, is shown for each variation in Table XXX. The effect that HP turbine wheel and blade attachment technology and reducing compressor rpm to improve LP turbine shaft critical speed had on FWs is also shown in Table XXX. It should be noted that an increase in the radius of the bore in the HP turbine resulting from the different LP turbine shaft designs greatly increases the FW of the wheels and spacer in the HP turbine because of the addition of bore reinforcing material. Note the increase in engine diameters as a result of slowing down the compressor rpm and the resultant vW penalty.

However, the FW of the HP turbine wheels would have decreased with the reduced rpm in variations 1 and 2 were it not for the fact that nonintegral blades (attached by dovetails) increased the centrifugal load, stress, and FW of the HP turbine.

TABLE XXX. - 1790 kW (2400 shp) ENGINE EFFECT OF WHEEL SHAFT AND RPM VARIABLES ON SECTION WEIGHTS Adv. Tech.

2 3 4 5 6 7 d 1 Feature gasoline Wheels NP turbine Yes Yes Yes Yes Yes No 90 Dovetails Yes Yes PA-101 PA-101 PA-101 IN-718 IN-718 PA-101 PA-lU1 Material PA-101 IN-718 +0.6 -44 -43 -22 0 -12.1 -21 0 +63.2 S A Weiyht 43.2 30.7 12.7 30.7 12.7 30.7 11.7 43.2 43.2 Rb--= (in.)

(U.S) 0.7) 0.21) (0.5) 0.21) (0.5) 0.21) 0.7) (1.7) Shafting 8o/Ti Steel 8o/T1 steel Stl/Be So/Ti Steel Material Stl/8e Stl/Be 0 -63 -42 -63 -42 -63 -41 0 S % might 0 Yes No No Yes No No Yes No Supercritical No -8.0 0 0 0 0 0 0 t % speed hp 0 -8.0 0 0 0 0 u +13.3 0 S A NP turbine case FW 0 +13.3 0 0 0 0 0 +13.3 0 .% coressor case FW 0 +13.3 +13.3 0 0 0 0 u 0 S% i % burner/diffuser FW 0 +13.3 0* 0 0 0 u 0 u 0 0* S A compressor rotor FW :elled out for no FW compe red to be seline engine. *Increased diameter and decreased speed cam Table XXXI provides a breakdown of weights and price by engine section for variation 3 of the baseline 1790 kW (2400 shp) STAT engine. This is the win- ning configuration for both weight and price and provides the minimum DOC, TABLE XXXI. - 1790 kW (2400 shp) (VARIATION 3) POWER SECTION AVERAGE RECURRING MANUFACTURING PRICE BY SECTION (1000 engines at 12/month, in 1979 economics) Ba s Engine section FM--ke O bm) n FM-- IIIF kg i (1b I Pric Brice-ro Roark$ Forward support 11.45 (25.24) 6.26 0.156 7,291 Axial compressor rotor 9.13 (20.13) 20.46 0.411 16,932 Axial compressor case 11.13 (24.54) 17.76 0.436 20,072 Centri compressor rotor 2.68 (5.90) 3336 0.196 9,134 Centri compressor case 8.43 (18.58) 15.09 0.280 12,935 Burner/diffuser 2739 (60.83) 8.88 0.640 24,931 MP turbine rotor 13.06 (28.84) -2.48 (-5.46) 21.90 0.631 27,951 -1764 Composite Mhwls HP turbine case 9.18 (20.24) 14.25 0.288 11,788 without dovetails LP turbine rotor 16.61 (36.61) -2.13 (-4.7) 6.94 0.254 11,706 + 407 eoM shaft LP turbine case 9.17 (20.21) 7.09 0.143 6,609 Rear burner support 11.34 (24.99) 9.06 0.226 10,450 Accy gearbox 45.36 (100.00) Accessories 18.39 (40.55) 4.13 0.581 26,790 Total T>19 UTM-ST r-TT_r_M_ - 6T Tq;7TT >r= l>3 .w __W Control s 40,115 Assy. and Test (approx 90 hr) 6,243 Average recurring manufacturing price Mr.

3579 kW (4800 shp) STAT Engine The baseline engine established for the 3579 kW (4800 shp) size is an axial flow engine with a 20:1 CPR and an airflow of approximately 11.57 kg/s (25.5 lbm/sec) at approximately 29,800 rpm.

Five variations of the 3579 kW (4800 shp) baseline engine were selected: two variations of different combinations of HP turbine rpm, LP turbine shaft de- sign, HP turbine materials, and two methods of blade attachments. Three var- iations have similar combinations as the first two but also have in addition an axial-centrifugal compressor and a shorter LP turbine shaft as a result of the combustor/diffuser design change. The technology describing the varia- tions from baseline is shown in Table XXXII.

The comparative recurring manufacturing prices for the 3579 kW (4800 shp) en- gine power section, plus five variations in turbine design, were also prepared.

The calculated weights by section of the 3579 kW (4800 shp) baseline engine and five variations are shown in Tables XXXIII and XXXIV. Also shown are the calculated comparative recurring manufacturing prices of this engine and five variations.

TABLE XXXII. - 3579 kW (4800 shp) ENGINE FABRICATION TECHNOLOGY Technology Engine Adv. Tech.

This engine is a straight through power section with Baseline IN-718 HP forged turbine wheels and individual dove- tailed blades. In addition, the LP shaft is steel with no stiffening (i.e., beryllium). The shaft re- quires several damping mutes to lower the critical shaft speed below the operating range.

1 This variation is scaled from the baseline engine with a 2% decrease in HP rotor speed. The HP turbine wheels are made by powder metallurgy and include dovetails.

The power turbine shaft is borsic titanium. The slight increase in HP turbine rotor weight, despite a reduced rpm, results from an increase in the bore diameter to pass the larger borsic titanium shaft required to get out of critical speeds.

This variation is the same as in 1 with the exception that the HP wheels and blades are composite without dovetails.

This variation is scaled from baseline with a 4% decrease in HP rotor speed. HP turbine wheels are forged IN-718 with dovetails with large ID. The LP turbine shaft is steel without mutes. The compressor is axial-centrifugal scaled from the 1790 kW (2400 shp) baseline engine.

This variation is the same as in 3 but it has an LP turbine shaft of borsic titanium. The HP rotor rpm is the same as that in the baseline.

5 This variation is the same as in 4 with the exception that it has HP turbine wheels and blades that are composite and steel LP shaft without mutes.

The LP turbine shaft design changes and their effects on FW of the LP turbine and FW of the HP turbine are shown in Table XXXV. The effect of HP turbine blade attachment and of reducing compressor rpm on We is also shown in Table XXXV. It should be noted that the baseline LP turbine shaft is "super criti- cal." This condition is corrected by increasing the diameter of the different LP turbine shaft designs, which greatly increases the FW of the wheels and spacer in the HP turbine because of the increased radius of the bore reinforc- ing material. Note that the overall increase in engine diameter resulting from slowing down the compressor rpm, while maintaining the same R c and FW of the HP turbine wheels would have decreased penalty. The Wa , has an FW with the reduced rpm in variations 2 and 4 were it not for the fact that the larger bore diameters increased the stress and FW of the HP turbine wheels.

TABLE MMIII. - 3579 kW ENGINE POWER SECTION PRICES AND WEIGHT BREAKDOWN (SI UNITS) U Adv. Tech.

Basel i ne 1 2 3 4 Foroard support--kg 20.70 21.37 20.76 22.04 20.70 20.70 Axial rotor-kg 24.89 24.89 24.89 20.b7 40.57 20.57 Axial case--kg 30.35 31.34 30.35 27.81 26.10 26.10 Centrifugal rotor--kg 6.03 6.03 6.03 Centrifugal cast-kg 9.66 9.07 9.07 Burner/diffuser--kq 19.16 19.78 54.39 51.07 51.07 NP turbine rotor (wMels)--kq 21.77 22.47 20.57 26.26 22.58 17.20 (9.98) (10.68) (8.75) (14.47) (10.1W) (5.49) NP turbine case--kg 9.75 10.07 9.75 10.36 9.75 9.75 LP turbine rotor (sheft) --kg 21.04 19.29 19.29 20.17 17.OU 20.14 (6.02) (4.27) (4.27) (5.15)* (1.98)* (5.15)* LP turbine case--kg 24.11 24.11 24.11 24.11 24.11 24.11 Rear burner support--kg 16.21 18.21 18.21 18.21 18.21 18.21 Accessory--kq 62.11 62.11 62.11 64.11 62.11 62.11 Accessory gearbox--kg 14.70 14.70 14.70 14.70 14.70 14.70 Total--kg 266.8 268.3 263.8 316.4 302.0 299.8 Recurring manufacturer ( Ice $401,591 $406,098 $397,793 $481,677 $467,943 $458,253 *Approximately 0.23 m shorter than baseline as a result of flow back burner configuration.

TABLE 1IX%IV. - 4800 shp ENGINE POWER SECTION AND WEIGHT BREAKDOWN (CUSTOMARY UNITS) Adv. Tech.

4 5 1 2 3 Baseline 45.63 48.60 45.63 45.63 45.63 47.113 Forward support--lbm {^ 45.34 54.88 54.88 45.34 45.34 54.88 Axial rotor--lbo S7.53 $7.53 69.09 66.91 61.30 66.91 Axial case--lbo 13.29 13.19 13.29 Centrifugal rotor--lbm 20.0 20.0 i 21.3 Centrifugal case--lbo 119.9 112.6 112.0 42.23 43.60 42.23 Burner/diffuser--lbm 57.89 49.79 38.09 47.99 49.53 45.35 NP turbine rotor (wheels)--lba (12.1) (19.3) (31.9) (23.8) (22.00) (13.54) 21.49 21.49 22.89 21.49 21.49 22.19 NP turbine case--lbm 37.48 44.41 42.53 42.53 44.46 46.38 LP turbine rotor (shaft)--tbm (4.31)* (11.35)* (9.42) (9.42) (11.35)* (13.27) 53.16 53.16 53.16 53.16 53.16 53.16 LP turbine case--lbo 40.14 40.14 40.14 40.14 40.14 40.14 Rear burner support--lbm 136.92 136.91 136.94 136.92 136.92 136.92 Accessory--lbm 32.41 32.41 32.41 32.41 32.41 32.41 Accessory gearbox--lba 665.8 661.0 581.6 697.6 588.1 591.6 Total--lbm $467,943 $458.253 $406,098 $397.793 $481,677 Recurring manufacturer price $401.591 now back burner configuration.

*Approximately 9 in. shorter than baseline as a result of l^ i I ' T •a TABLE XXKV. - 3579 kW (4800 shp) ENGINE STAT STUDY EFFECT OF WHEEL SHAFT AND RPM VARIABLES Adv. Tech.

3 4 1 2 Baseline Feature Wheels NP turbine IN-718 PS-404 PA-101 IN-718 1p a) IN-718 PA-101 tem -46 +12 KS +6 +7.0 36.8 $0.8 59.7 S0.8 16.5 g Rb- - ^ i M-) S9.7 (1.46) (2.0) (2.35) (2.35) (2.0) (0.66) Yes Yes No Yes Yes Yes Dovetails Shafting LP turbine Steel Steel T1 cap Ti cap Ti cap Steel { "atom &I -67 -14 -29 -14 0 -29 % 0 weight No No No No Yer No Supercri ti Cal 49.5 35.6 49.S S8.4 S8.4 15.2 - - m (in.)

Ro 0.96) 0.96) (1.40) (2.39) (2.30) (0.6) i 0.61 (32) 0.81 (32) 0.61 (32) 1.04 (41) 1.04 (41) 1.34 (41) Length--is (in.)

0 0 0 -4% 0 -2% % e NP speed +3.25% 0 +6.S% i % A NP turbine case weight +15.87% +23.45% +15.87% +3.25% 0 % compressor case weight +266.6% +283.9% +266.6% 0 +3.25% 0 % burner/diffuser weight +6.83% 0 +6.83% +6.83% % i compressor rotor Table XXXVI provides a breakdown of weights and prices by engine section for variation 2 of the baseline 3579 kW (4800 shp) STAT engine. This combination of weight and price provided the minimum DOC.

TABLE XXXVI. - 3579 kW (4800 shp) (VARIATION 2) POWER SECTION AVERAGE RECURRING MANUFACTURING PRICE BY SECTION (1000 engines at 12/month in 1979 economics) Baseline Engine section F11--kg (lbm) FN-- kg l ( bm_) MIFh s Ic Price f Baseline price--.1L Reearts Forward support 20.70 (45.63) 4.47 0.204 9,406 Saw as baseline Axial compressor rotor 24.89 (S4.88) 18.54 1.018 46,956 witn Via exception Axial compressor case 30.3S (66.91) 16.27 1.090 50,242 of composite MP burner/diffuser 19.16 (42.23) 23.86 1.007 46,493 turoine wneels and NP turbine rotor 10.57 (45.35) -1.20 (-2.64) 32.3S 0.969 43,440 -18U5 blades And borsic NP turbine case 9.75 (21.49) 21.88 0.469 19,573 t!+unium LP turoine LP turbine rotor 19.29 (42.53) -1.74 (-3.83) 15.78 0.671 30,970 + 448 shaft LP turbine case 24.11 (53.16) 13.19 0.706 32,012 Roar burner support (40.14) 18.21 7.89 0.317 14,S22 A gearbox 14.70 (32.41) 4.57 0.148 0,834 cc y Accessories 62.11 (136.92) 4.38 O.S96 27,536 Total Z6,T:7758M -E -R -x337' Tj.W rw $3287W -Rw Controls $4,287 Assy, and Test (approx 180 hr) 12,485 Average recurring manufacturer price 00.767 Turbine Flow Path The turbine flow paths for the 1790 kW (2400 shp) and 3579 kW (4800 shp) STAT engines are shown in Figures 28 and 29, respectively. The appreciable step-up in diameter of the LP turbines` first stage, in both engines, necessitates a relatively long annulus between turbines and precludes the incorporation of counterrotating turbines.

The preliminary turbine designs for the STAT engine trade studies were made for the 1790 kW (2400 hp) engine at SLS intermediate power and were scaled in the radial direction by the square root of the ratio of compressor inlet air- kW (4800 hp) engine. Meridional chords of the gas genera- flows for the 35;9 tor turbine were maintained when it was scaled up to the 3579 kW (4800 hp) sire; however, the power turbine length was increased slightly to maintain an acceptable flare rate at the outer wall of the flow path.

The gas generator turbines were designed as two-stage axial flow units to keep stage equivalent work (Oh/Bcr) at an acceptable level of 54,638 J/kg (23.5 Btu/lbm). A single-stage transonic (high work) turbine would have resulted in a stage equivalent work of 101,603 J/kg (43.7 Btu/lbm) and a stage expansion ratio of 5.323 (supersonic) which is higher than DDA advanced design practice.

The flow paths were designed for nearly constant hub diameter and have cylin- drical, unshrouded rotor blade tips. Turbine average stage loading coeffi- cient (gJ A h/Um 2 ) was 1.6 and stage work was split 55 and 45% for the first and second stages, respectively. The maximum Mach numbers reached in the ve- locity diagrams were high subsonic.

8.0 7.0 i 6.0 s s 5.0 4.0 3.0 10 15 20 23 5 • Length—cm NR ' 23,9W rpm m 39,767 rp , NGOT 7 9 11 4 5 6 6 10 0 1 2 3 Length—in.

TEW-M47A Figure 28. - STAT 1790 kW (2400 hp) engine turbine flow path.

r .. 8 I t s r 10' Is 20 4 L 10 s ltnpth--em Npt 17,780 9w GOT ' 29,823 rpen N 9 11 8 10 4 6 7 3 5 f 0 1 2 Length—in.

TE80.2048A Figure 29. - STAT 3579 kW (4800 hp) engine turbine flow path.

The preliminary design study of the power turbines resulted in two-stage flow paths with constant hub diameter. Turbine exit area was sized to give an exit axial Mach number of 0.43 at the SLS point and the exit hub-to-tip diameter ratio was set at 0.54 to give adequate rotor bade hub reaction. Rotational The speed was calculated to give a turbine exit AN of about 5.4 x 10 10 .

preceding design considerations directed the design to a two-stage flow path with an average stage loading coefficient of 1.6 to give the desired power turbine efficiency level. Power turbine work was split 55 and 45% for the first and second stages, respectively, to minimize turbine exit swirl at the altitude cruise point. A short transition duct was required to diffuse and transfer the gas flow outward from the gas generator turbine exit to the power turbine inlet.

A number of significant design problems are apparent for the advanced STAT engines. The problems believed to be most severe in our study of the power section were: o LP critical shaft speed and its impact on HP wheel design and lubrication system o Dovetail design limitations in small hub diameter engines o Design life (low cycle fatigue (LGF), stress rupture, oxidation/erosion) of short-range mission hardware o Minimum cooling air passage size set by the present casting state of the art o Airfoil size/height from a manufacturing star.ipuint and off design sensi- tivity to tip clearance.

Radial Inflow Turbine In order to investigate the potential performance improvement of a radial in- flow turbine, such a flo g path was configured (see Figure 30) for the 1790 kW (2400 hp) engine.

The results of this investigation show: • Axial engine length would not be significantly reduced (2 in. decre#Aaed) over g hat with an axial turbine.

• Turbine efficiency is approximately the same (84-86X).

• Small exit hub diameter precludes a concentric shaft engine design.

The desired sfc payoff was not realized, and this type of turbine was rejected for the STAT engines. The primary benefits of radial inflow turbines are best realized in small, low pressure ratio engines. A disadvantage for this type of turbine is that high rotor speeds are required for the radial inflow tur- bine which forces the compressor to run at a speed high:.r than optimum.

Hybrid Rotors, Composite Shafts, and Supercritical S hafts In modern, two-spool concentric shaft gas turbines, the critical speed of the LP shaft influences engine size and configuration. This becomes a dominant limiting feature as engine pressure ratios increase, and airfoil hub diameters decrease, to maintain blade span.

I s N • 39,767 rpm NS • K 16 UT • 637 m!s amfVSK) 6.

IT-T S a Y u io p 2 4 6 e alW lenglh^r^ 3 4 0 1 alibi length—in, RID-202A Figure 30. - STAT radial inflow turbine 1790 kW (2400 shp) engine.

t.

I

This is a problem in the design of high pressure ratio engines, such as the 1790- and 3379 kW ( 2400- and 4800 shp) study engines. A trade study on shaft- ' ing/wheel technology was undertaken. It is important to note wheel technology and shaft technology are integrated, because the UP spool wheels must have sufficient bore diameter to clear the LP shaft and any attachment features.

Several terms used in this discussion are defined as follows: f T_ Features or material N r' IN-718 wheels with dovetails Standard wheal Advanced wheal Power metallurgy wheal and dovetails Composite ( hybrid) wheel Powder metallurgy wheel diffusion bonded to a cast blade ring Ti-composite shaft Titanium selectively strengthened with composite

fibers

Supercritical Used to describe shafting that, if simply sup- ported, would have less than 25% speed between the first critical speed and the maximum oper- ating speed. Some type of device, i.e., bearing, mute, etc., would be used to restrain motion.

As modern two-spool engine pressure ratios and RITs increase to improve sfc, the HP spool speed increases and its diameter decreases. The axial length trends to be fixed and independent of diameter.

A problem develops as this trend continues. Soon the ability of the designer to use dovetc.11s for blade retention on the wheels, and to use conventional, simply supported steel shafting for the LP shaft, is restricted. In the pre- liminary flow paths referred to as the advanced baseline engines, it is not possible to configure a conventional technology wheel ( conventional material and dovetails) around the minimum OD subcritical shaft made of steel.

A number of technology and configuration trades were conducted in the course of this study and resulting weight penalties were assessed. The following items were considered: o Wheels Enhanced properties with dovetails - Composite or HIP bonded wheel / blade assemblies

o Shafting

Composite or fiber - reinforced titanium Supercritical shafting o General arrangement and speed Close coupled LP turbine Decreased HP spool speed

^t - --

r441 The final evaluations for critical shaft speed, with resulting impacts on DOC, ( 4800- and are given in Tables XMII through XL for the 3579 kW and 1790- 2400 hp) STAT engines, respectively. It will be seen that the combination of technologies for 1790 kW (2400 shp) engine No. 7 provides the biggest payoff in DOC improvement. This engine incorporates composite wheels and shafts and cast-in impingement cooling.

TABLE XXXVII. - 3579 kW SHAFT STUDY RESULTS (SI UNITS) 1979 economics P0370-37 STAT scaled baseline 1 2 3 4 15 7 Base 16 8 Cycle R^ I I 12.7 20:1 20:1 20:1 20:1 20:1 20:1 2U:1 Lu:l RIT--K 1505 1505 1505 1505 1505 I 1505 I 1505 1506 15Ub SFC--Ng,'M-s 66.23 66.23 I 79.?5 66.23 66.40 66.13 I 61.41 01.4[ 0.4c

I I

Power section !Mss--kg 381.5 266.8 268.3 268.3 268.3 I 263.8 I 293.7 3UL.0 099.8 Cost-4 276,752 262,590 265,514 264,783 264,607 260,128 I 313,170 296,727 291,080 I 1.466 1.466 I 1.466 Length--m 1.466 1.466 1.460 1.237 1.W l.ta7 Diaee • .er--m 0.503 0.495 0.495 0.495 0.495 I 0.495 i 0.759 0.759 U.759 Gearbox 156.0 I I Mass-49 156.0 156.0 156.0 156.0 156.0 156.0 156.0 156.0 Cost-4 59,000 59,000 59,000 59,000 59,000 I 59,000 I 59,000 59,000 59,00u 0.255 0.255 I Length--m 0.255 0.255 0.255 0.255 0.255 0.255 0.165 I Ratio 10.4 10.4 10.4 10.4 10.4 I 10.4 ( 10.4 10.4 10.4 I Totals I Mass--kg 537.5 422.8 424.4 424.4 424.4 419.8 449.8 458.0 4bb.y I I 321,590 324,514 323,783 323,607 319,128 37•[,170 355,727 Cost-4 335,752 350,080 I I Length--a 2.644 2.644 2.644 2.644 2.644 2.644 2.416 2.416 2.416 0iameter--m 0.503 0.495 0.495 0.495 0.495 I 0.495 I 0.759 U.759 0.759 S weight -21.3 -21.1 -21.6 -21.6 I -21.9 I -16.3 -14.d -15.1

e

-4.2 -3.3 -3.6 -3.6 I -4.95 I +10.85 +5.95 +4.27 S p Cost % p length 0 0 0 0 0 ( - 8 .6 - 8 .6 -d.0 I S 0 diameter -1.5 -1.5 -1.5 -1.5 -1.5 ( +50.9 +W.9 +50.9 % 0 SFC -16.9 -16.9 -16.9 -16.7 -1E.9 -15.5 -15.5 -15.5 -7.04 -7.09 -7.02 i -7.21 i -2.67 -3.48 i 0 DOC -1.16 -3.25 Technology i Configuration Axial Axial Axial Axial i Axial Axial- Axial- Axial- cent cent cent HP turbine wheels Forged Net shape Net shape Net shape I Compositel Forged Forged Composite Shafa Super Composite Composite Composite I Compositel Steel Composite Steel Imp/film Imp / Ioplfilm Imp/film First blade cooling film Imp / film Convection Imp/film Imp/ film tube cast - in ; ube cast-in cast - in cast-in cast-in The study of 3579 kW (4800 hp) engines did not encompass as many combinations of technologies as did that for the 1790 kW (2400 hp) engines. Engine No. 5 ent with composite shafts and wheels and tube- exhibited the best M" improve > type impingement cooling. If this engine had cast-in impingement cooling, it would be expected that the percent change in DOC would be improved an addi- tional 0.05%. This is substantiated by comparing the percent change in DOC

for Engine No. 2 and No. 3, which differ in configur,.%tion only in the type of

first blade cooling.

60 f TABLE XM III. - 4800 shp SHAFT STUDY RESULTS (CUSTOMARY UNITS) i 1979 economics STAT P0370-37 scaled baseline Base 1 2 3 4 i5 i 6 7 d Cycle 12.7 20:1 20:1 ' 20:1 20:1 20:1 90:1 Lu:I 2U:1 RiT»'F 2250 2250 2250 1250 2250 2950 ( 2950 22bu 2950 SFC 0.392 0.472 0.392 0.392 0.393 0.392 ' 0.399 0.:199 0.399 Power section i i Weight--lbm 841.0 588.1 S91.6 591.6 591.6 581.6 647.o 065.d oel.0 276,752 262.590 265,514 264,783 264,607 313,170 296,717 Cost-4 260,128 291,UUu t Dia meter-in. 19.81 19.S 19.5 19.5 19.5 i 19.5 t9.9 9.y 9.9 a Gearbox 344 344 344 344 i 1 Weight--lbe 344 344 544 34 344 Cost-4 59,000 59,000 59,000 59,000 59,000 59,000 59,000 59,000 59,0W 10.02 Langth--in. 10.02 10.02 10.02 10.02 10.02 1 10.09 10.02 10.09 Ratio 10.4 10.4 10.4 10.4 10.4 10.4 ' 10.4 10.4 1U.4 Totals ' Weight--lbm 1185 932.1 935.6 935.6 935.6 925.6 991.6 100.6 luub.0 Cost-4 335,752 321,S90 324,514 323,783 323,607 ( 319,128 372,170 355,727 350,UW 104.1 Length--in. 104.1 104.1 104.1 104.1 ' 104.1 95.1 95.1 95.1 Diameter--in. 19.81 19.S 19.5 19.5 19.5 ( 19.5 ( 29.9 99.9 29.9 -21.3 -21.1 -21.6 -21.6 -16.3 -14.8 -15.9 % O weight -21.9 % A cost -4.2 -3.3 -3.6 -3.6 , -4.95 *10.85 +5.95 +4.91 is length 0 0 0 0 0 -8.6 -d.b -8.6 S -1.5 ( S A diameter -1.5 -1.5 -1.5 -1.5 +50.9 +50.9 +50.9 -16.9 -16.9 -16.9 -16.7 ( -16.9 -15.$ -15.* -15.5 % 0 SFC -7.09 %4% 00C -7.16 -7.04 -7.02 I -7.27 ( -2.67 -3.25 -3.48 Technology I Configuration Axial Axial Axial Axial ( Axial I Axial- Axial- Axial- HP turbine noels Net Net Net I Composite I F cent ite Forged shape shape shape orged forged I Shafting Super Composite Composite Composite I Composite Steel Composite Steel First blade cooling Ip/film Ip/film Imp/film Convection ( Ip/file I Ip/film Imp/film Imp/file cast-in cast-in tube I cast-in cast-in cast-in tube cast-in impingement is available for turbine bla.ies with An additional benefit of the %;heel rim is available for cooling passages, as a greater percentage dovetail and resultant in- This is due to the smaller blade blade retention.

dovetail slots in the wheel rim.

crease in spacing of of using dovetail blade attach- As engine size decreases, the practicality decreases. In the case of our advanced ments, in both compressor and turbine, features (i.e., impingement - cooled baseline engines, the presence of cooling dovetail was sized for the picture. A preliminary first blade) complicates impingement tube passage is mini- the first turbine blade; however, space for mal.

1.

d Thermal Barrier Coatings a.

Thermal barrier coatings show greatest promise in very high temperature en- gines. Their effect is greatest where high gas side-to-metal temperature gradients are required.

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p N1 N^^ O Vt0 TNN O 9I -11-Tm C p O T ^ O e E^ E ^_ T O T a 1 M W . JIM ^ W W W O ]..W Y ^ L .. ~L L.+ L.+ y Q Y N Ly^:.J JO ^ Z:J J -= J J O 1 V- 7 N^ ^S JL11r J ^ ' - M N M M M M - i^ An analysis of cooling flows in the 1790 kW (,2400 shp STAT engine indicates the coating would reduce chargeable cooling by 20%; however, the increase in i airfoil thickness reduces turbine efficiency due to the blockage effect. Us- ing sensitivity factors for DOC versus cooling airflow and turbine efficiency, the following negative impact is shown: % A DOC +1.25 -1X 71 HP turbine -1.10 -20% .1 cooling +0.15% DOC Net impact ^I A similar loss characteristic would be experienced by the 3519 kW (4800 shp) STAT engine.

The assumptions used in this analysis are: • The blade life is constant if metal temperature is held constant. Here we th.-2 effect of increased blade load through the ce- chose not to evaluate ramic mass.

• A constant thickness coating was assumed over the entire airfoil.

• The effect of changing coating thickness through erosion was not accounted for in life calculations.

!

• The surface finish of the coating is the same as that for the uncoated blade.

It should be noted that independent industry test programs have shown that turbine airfoil coatings in the unpolished state tend to have a rougher finish than the parent metal. In some cases, this rougher finish can cause the re- verse etfect of making the airfoils hotter with the coating than without.

Ceramics DDA is involved in programs to develop ceramics for gas turbine engines.

These effurts are directed to two areas: solid, monolithic structures, and coatings. The solid, monolithic structures, such as turbine vanes and blades, show the greatest promise in reducing airfoil cooling airflows. For our STAT advanced technology engines, total chargeable cooling airflows could drop by 4.5% of engine airflow. This is a 45% reduction from the advanced baseline cycles.

Within the time frame specified for STAT engines, however, monolithic ceramic components were not considered due tc current state of the art plus normal time for design, procurement and test ahead of full development release in 1988.

r ^i Ceramic coatings have reached a higher level of development than monolithic structures. Coatings for airfoils and end-wall insulation have been used in

^I. some advanced experimental engines at DDA. In addition, an abradable ceramic

coating has been tried for first- and second-stage tip shrouds. The coating used was eroded severely in less than 35 hours of engine operation. The vane airfoil coatings have had some success in experimental engine operation: how- ever, the resistance to environmental particulates needs improvement. Ceramic coatings were, therefore, not considered.

Cast-in-Impingement Cooling Cooling air distribution through turbine blades by means of transpiration via Lamilloy is very effective where the airfoils are large enough for trailing edge blockage effects to be minimal. The airfoils in the STAT engines are too small for this consideration, and employ impingement cooling instead.

Presently, all DDA production air-cooled airfoils (blades or vanes) are made by an investment casting process using cores to form the internal passages.

At present the minimum core size is limited by the strength of the core mate- rial. In addition, some air-cooled airfoils use impingement cooling to in- crease cooling effectiveness. The small size of some airfoils makes it diffi- cult to manufacture and assemble sheet metal tubes small enough to fit within them.

One method of circumventing the tube problem is to cast-in the impingement airflow passage, as shown in Figure 31. This is done by inserting quartz rods between core sections. The rods are removed from the casting chemically, leaving the required impingement holes. This method has the advantage of pro- viding more load-bearing metal area within the same airfoil contour. A simple costing study indicates a $595 savings per stage for the small engine, and $741 for the 3579 kW (4800 shp) engine. Since the first vanes and second vanes will most likely require impingement cooling, the cost reductions become: Rated power--kW (shp) 1790 (2400) 3579 (4800) $ A engiuc cost 1190 1482 % A engine cost -0.580 -0.494 % A DOC -0.063 -0.064 There should be additional savings to the engine via lowered cooling flows and improved life inherent with this configuration. This results from having a structural load-carrying member isolated from the gas side temperatures.

Long Life Bearings The main bearings incorporated throughout the STAT engines will exhibit a much greater load carrying capability and fatigue life than those in current tech- nology engines. This improvement may be brought about in part by research in the characterization and control of forging flow lines and end grain areas in bearing balls.

E^ Cast-in delivery passages (proposed advance technology) r Delivery tube (current technology)

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TEf30 2049A Figure 31. - Airfoil impingement cooling.

In the case of the STAT reduction gear main bearings, DDA studies show that an improvement in the inherent premature removal rate of approximately 57% is possible by this method. This equates to a reduction in engine maintenance . 11% for the 1790 kW (2400 hp) cost of 1.1%, and resulting decrease in DOC of 0 . 14% for the 3579 kW ( 4800 hp) engine.

STAT engine, and 0 Clearance Control If the STAT HP turbines were straddle-mounted, the second stage clearance . 0008 in. The increased weight would reduce by 0.0013 in. and the first by 0 and cost of the required bearing support would penalize the engine. The fol- lowing table assesses the penalties and payoffs: Rated power--kW (hp) Weight %.A Cost %A sfc % A DOC % A +3.6% -.31% +.345 1790 (2400) +3% +.399 3579 (4800) +3.6% +3.1% -.31% the HP tur- therefore, a DOC penalty rather than an improvement, if There is, bines are straddle - mounted.

The most critical feature required of an active clearance control system for a short-haul aircraft is a quick response to transients. Systems that do not respond rapidly to transients, such as the general class of thermally actuated I systems, will not be cost effective in small transport aircraft.

Most active clearance control systems considered previously have mechanically driven tip seals. For this study, consideration has been given to a simple I, system to axially translate sloped tip seals as shown in Figure 32. This de- vice would be activated by the electronic control and would be continuously I adjusted to maintain the safe, minimum clearance during the entire mission.

The turbine flow path for a current gas turbine under development at DDA is similar to both the 3579- and 1790 kW (4800- and 2400 shp) STAT engines with respect to shape and blade height. The blade heights of the first and second turbine stages of this engine are approximately 1.32 and 2.13 cm (0.52 and 0.84 in.), respectively. Uaing the general relationship that for each change in clearance equal to 1% of blade height, a 2% change in turbine stage effi- ciency results, then gasifier turbine efficiency changes 1% for each 0.13 mm (0.0052 in.) change in first-stage clearance, or 0.21 mm (0.0084 in.) change in second-stage clearance.

As uncrosshatched member moves to right or left, blade tip clearances at A and B decrease or increase, respectively.

TE80-2053 Figure 32. - Active clearance control--turbines.

A The clearance of the first blade is smallest during start and warmup • This results in approximately 0.16 me (0.0062 in.) additional clearance during normal running operation, and a 1.2% loss in gasifier turbine efficiency. The second-stage blade clearance is also least during start, resulting in an addi- tional 0.19 em (0.0074 in.) change in clearance. This penalised the gasifier turbine efficiency an additional 0.88%. Total gasifier payoff is thus esti- matedto be 2.08%. If we assume that an active clearance control system would eliminate this penalty, the sfc would improve by 2.3%, and DOC would decrease a, by 0.89% if the weight and cost penalties are not included.

Based on approximations of system weight and cost, the following results would be obtained for the STAT engines: i.

% A Cost % A sfc % 4 DOC Rated power--kW (hp) % A Weight 1790 (2400) +1.7 +5.7 -2.3 -0.284 3579 (4800) +1.6 +4.8 -2.3 -0.137 As an alternate to an active clearance control, thermal matching offers a low- er risk, albeit lower payoff means of clearance control. DDA development ex- perience has shown that increased attention must be paid to transient response of rotor and stationary elements. Specifically, the HP rotor case material was changed for a current DDA development engine, and the transient-induced running clearances were reduced from 0.16 to 0.06 mm (0.0062 to 0.0025 in.)

for the first stage, and from 0.19 to 0.13 mm (0.0074 to 0.005 in.) for the second stage.

Given sfc sensitivity to turbine clearance, such reductions in the STAT en- gines should improve sfc by 1X, reduce DOC by 0.35, and 0.31% for 3579- and 1790 kW (4800- and 2400 shp) engines, respectively.

Leakage Control The STAT engine turbines will use paired step-seals to minimize air leakage and stage bypass flow. This technology promises a 10 to 20% reduction in flow for the same number of seal elements. The primary technology advancement re- quired to achieve this benefit lies in the modeling of the labyrinth seal flows analytically instead of through experimental test.

k, Abradable Turbine Coatin gs Two reasons for using abradable coatings in the turbine blade tip path have been identifi=d. The first is in making the engine more tolerant of turbine rotor offset caused by normal production tolerances. The second is to mini- mize requirements allowing for extreme transient conditions, thus increasing safety because of the ability to have extreme transients and not experience major structural damage. These reasons apply whether the turbine incorporates active clearance control features or not. IN

F

r L A current DDA development engine incorporates abradable turbine seals. This engine would be expected to experience 0.019 negative growth during an emer- gency shutdown. If the blades were running with a nonabradable seal surface, they would most likely be severely damaged. The additional clearance running, to avoid shutdown damage, causes an unacceptable sfc penalty to the engine.

Incorporation of abradable seals, however, could provide the following typical improvement in DOC: DOC Mated power--kW (hp) X 0 Weight % .1 Cost %a sfc X Q 1790 (2400) -5.0 -8.0 -0.8 -1.046 E -0.8 -1.212 3579 (4600) -6.0 -6.0 Two methods are under consideration for the use of abradable coatings--to de- velop a truly abradable coating for conventional blades, or to combine a part- ly abradable coating with abrasive tip blades. The state of development in this area would indicate this is a high risk technology for commercial engine development.

One method currently used to maintain close tolerance turbine blade tip clear- ance is to line-bore the rotor bearing support cavities with the structural members assembled sans the rotor. This is a costly pracess which could be eliminated if abradable seals were available. An additional payoff for abrad- able coatings is the fact that safety improves without sacrificing perform- ance. In both STAT engines, the penalty in performance to design for center- line offset and emergency shutdown could not be accepted.

Design Life The standard practice for turbine blades is to coat them with materials to provide protection from sulfidation and other forms of corrosion, and thereby assure long life. The limiting feature of the hot section, as far as overhaul time, is the life of the blade coating. Based on available materials data, an airfoil will require two to three recoats during its life.

A study of first-stage turbine blade life was conducted. An analysis of cool- ing air requirement sensitivity-to changes in pressure ratio, RIT, and cooling air temperature is shown in Figure 33. The impact of cooling airflow on life of the first-stage blade is shown in Figure 34.

A study was also made of the impact of increasing turbine wheel stren8th (weight) on LCF life. This sensitivity is depicted on Figure 35. It should be noted that no design restraints have been considered in limiting the size of the wheel while increasing its weight.

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+5 1 ^ L .

u } C u av .

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TE80-2050 Figure 33. - STAT first-stage blade sensitivity of cooling flow to turbine temperature, pressure ratio, and cooling air temperature.

u Baseline 6000 hr Stress rupture life Rc 20

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Tc 912OF

s 0

RIT 22501 Mar-M246 Alpack Mcterial Imp/film Cooling technique Percent We 2.58 18.5 KSI Stress at take-off Assume no restriction to range of W e shown -10{ i A i -50 V TIW tLW T3W Percent change, life TE80-Z0: Figure 34. - STAT first - stage blade sensitivity of cooling flow to F..

t r^ I +30 +20 I +10

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Bored on two major cycles r: per hour -10 ^^— Not LCF limited

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30,000 20,000 0 10,000 Wheal LCF life—hr TE80-2073 Figure 35. - STAT wheel life versus percent change in wheel weight.

Bearings Tapered Roller Bearings Tapered roller bearings have been suggested as substitutes for conventional ball thrust bearings used in most modern gas turbine engines. These bearings have calculated life improvements of at least 50% above the ball bearings they replace. However, it is necessary to install matched pairs of tapered roller bearings in the place of a single ball bearing with the attendant increase in cost, parts inventory, maintenance effort, etc. Therefore, it is judged that DOC would not benefit appreciably from this technology.

Bervilium - Backed Bearing Races In our STAT engine sizes, beryllium-backed bearing races would have a small payoff since the primary benefit is weight reduction. In large engine gear- boxes, the bearings have more dominant design limitations, and a stronger pay- off would be expected.

_j Diffusers/Combustors Combustor flow paths were established in both STAT Qngine sizes by scaling

existLM combustors. Both engines have annular combustors produced with *Lam-

illo technology. The large engine uses a straight-through annular combus- tor. The small engine has a fold back design that takes advantage of a de- crease in engine length available with no increase in power section frontal area. This is possible due to its compressor discharge being radially out- ward and not axial.

Several diffuser and combustor advance technologies were suggested as trade study elements. Of these, the -ortex-controlled diffuser (VCD) and transpira- tion-cooled Lamilloy combustor were utilized since they were the subject of recent DDA life cycle cost studies under Air Force contract No. F33657-77-C- 0425. A description of these technologies and their impact on engine weight, cost, performance, etc. is contained in the final report for this USAF con- tract.(1)** The VCD (shown in Figure 36) is a very short diffuser with 3 major flow paths in lieu of 5 utilized by current 3-passage diffusers. Inner and outer bound- ary layer control is varied by selecting optimum ratios of inner and outer bleed. The VCD is applicable to the larger STAT engine and offers an improve- ment in engine performance (0.3Z SFC) as a result of a decrease in total pres- sure drop across the component, as shown in Figure 37. The VCD also requires less axial length (1.9X) than the conventional diffuser, thereby saving engine weight (0.9X) and reducing its acquisition cost (1.1X).

Using the sensitivity values established previously, the following DOC im- provements would be achievable in the LCC 50-passenger transport with 3579 kW (4800 hp) engines incorporating a VCD: sf c Cost Weight Length Tot^l DOC A —% (Engine) -0.3 -1.7 -0.9 -1.9 --- -0.034 & DOC--Z -0.105 -0.068 -0.027 -0..234 Lamilloy is a DDA-developed and patented quatitranspiration cooled structural sheet material. It is fabricated by bonding together two or more layers of material that have been etched to form a complex internal flow path as shown on Figure 38. The incorporation of a Lamilloy, two-stage transpiration-cooled combustor in an advanced technology engine provides both weight (2.2X) and price (1.5Z) savings.

*Lamilloy is a registered trademark of the General Motors Corporation.

**Numbers in parentheses refer to references list#A at the and of this report.

Outer bleed ^ 1 Airflow Inne.- bleed Three-passage diffuser Vortex-controlled diffuser TE80-2059A Figure 36. - DDA diffusers.

a D O / CJ ^ wwur• iy^X DiNuwr ^lph.^r rM. paopel l^ u ^o/ Ddb n^ ^r ^h 0.2 0.3 0.4 0.5 Difkw inlet. MN TEBO-20EIA Figure 37. - Diffuser pressure loss results.

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Laminate No. 2 Laminate No. 1 T E-713C Figure 38. - Transportation-cooled Lamilloy.

Rated pcwer--kW (hp) %-I Weight %A Cost % _I DOC 1790 (2400) -2.2 -1.5 -0.224 3579 (4800) -2.2 -1.5 -0.260 A second-order advantage of the Lamilloy combuPt.or is the reduced requirement for turbine cooling air and improved reliability cf the HP turbine as a result of the improved burner out temperature (BOT) profile. Also, the reduced com- bustor cooiing ai- requirement of Lamilloy makes available more airflow for combustor outlet temperature pattern adjustment.

Control Svstems Preliminar y concept •ial d•-sign studies were initiatea to identify advanced technology control systems for the STAT advanced technology engines. The cri- teria for These systems were improved reliability ; reduced cost and weight, and '_mproved Tsintainability as compared to current product-on engine con- tr)ls. The control systems would cor.aider total propulsion system require- menrs to provide for overall thrust management/protection through all required operational conditiors.

The control and fuel system will be configured to use an advanced technology I digital electronic controller for all required logic and computational re- quirements for the engine and propeller operation.

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The integrated propulsion control system will be designed for high reliability to require less maintenance than current systems, consistent with the require- ments of lower TCO and high dispatch reliability. The control system will include the follwoing features: • Integration of control functions of engine and propeller to minimize num- ber of systems components • Utilization of a full-authority digital electronic controller incorporat- ing advanced, low cost, low power, large scale integration, solid-state components for high reliability for control of both the engine and the propeller • Utilization of advanced, simplified fuel pumping and metering components for low cost and weight reduction • Self-check capability to detect and provide indication of the occurrence of a malfunction of any of the separate control system components.

• Optimum location and mounting of the control system components for easy access for routine maintenance and replacement and in a suitable thermal and vibration environment for long life • Provisions for remotely actuated devices for all adjustments that may be required in service • Provisions for automatic thrust management incorporating ability to select and maintain a number of power control modes for maximum efficiency (take- off, maximum climb, maximum cruise as a minimum) • Provisions of interface with a diagnostic/conditioL monitoring system sen- sors for engine health monitoring • Provisions for interface, through digital data link, with the airframe flight control systems Engine Condition Monitoring The design objective of the STAT propulsion systems will be the achievement of on-condition maintenance whereby scheduled overhauls are eliminated and in- spections are minimized. This alone has the potential of eliminating 40% of the current engine, reduction gear, and propeller maintenance cost. A condi- tion that will facilitate the implementation of this maintenance coucept in commercial aircraft service is improved fault detection and isolation via diagnostics to identify impending problems so that corrective action can be taken prior to failure.

To show the advantages of using condition monitoring, an estimation of the DOC and maintenance cost savings for the STAT engines is shown below.

1749 (2345) 3544 (4752) Engine rating--kW (shp)/SLS -19.2 -13.5 Maintenance cost savings--%A -2.5 -1.4 DOC improvement--%A Y - x A clearly defined on-condition maintenance concept will be developed in con- junction with potential user airlines and the aircraft desi;ners. These con- cepts will take into account maintenance access times, likely available skill levels, and support equipment. Thus the propulsion system, aircraft, and air- line operations can be designed to derive the benefits of condition monitoring equipment. Such equipment can provide an early indication of malfunction and, especially, pinpoint the specific component needing maintenance, thus reducing secondary damage and eliminating shot-gun maintenance of control/accessory components.

.'c Propulsion system condition monitoring provisions will be incorporated in the STAT engines to permit detection of impending malfunctions and to define the required maintenance action. Early detection and correction of potential pro- blems result in improved aircraft safety and reliability. Transducers, which are req _ed to measure component pressures, temperatures, and positions, along wi,a the associated wiring, will be integral parts of the electronic fuel control system.

Advanced technology sensors will be incorporated into the STAT engine condi- tion monitoring systems. These sensors, which are yet to be developed, in- clude those for optical speed and temperature measurement. In addition, ad- vanced low cost compressor discharge pressure sensors and long life gas stream termperature sensors will be incorporated in the STAT control and condition monitoring sytems.

Integration of the control system and engine condition monitoring system into a single electronic system will save engine weight and cosC to an extent that could make the following impact on DOC.

% % ADOC Rated power--kW (hp) &Weight &Cost -.460 1790 (2400) -1.0 -4.0 -4.0 -.546 3579 (4800) -1.0 Noise Reduction The STAT engines with their advanced technology, high pressure ratio cycles will have to be designed so as to minimize their fore and aft noise signature: in order to meet federal regulations and satisfy public demands. If suitable noise suppression technology was not available when the STAT engines were de- signed, they would f.ave to incorporate inlet and exhaust duct acoustic treat- ment. With the noise reduction technology available, however, the following impact on DOC might be obtained by removing the duct treatment: i % % A DOC % 0 Weight % ACost Asfc Stated power--kW (hp) -0.3 -.624 1790 (2400) -14.0 -1.0 -.804 3579 (4800) -19.0 -1.0 -0.3 The above analysis gives an insight into the payoff to be realized by develop- ing the technology required to quiet an advance technology engine without re- sorting to acoustic duct treatment.

Reduction Gear In considering the various ways to mount a turboprop reduction gear to a power section, the following alternatives are available: Remote mounted versus integral with engine frame and concentric versus off-set.

There are four possible combinations with the noted alternatives. The choice of the "right" one depends on an assessment of its impact on the following: o Weight o Complexity--effect on reliability and maintainability o Effect on engine inlet; i.e. an integrally mounted concentric gear box could be high in air losses and therefore not be the best choice.

o Air frame structural requirement; i.e. an offset remote mounted gearbox allows the main wing structure to be aligned with the propeller thrust center line. Another consideration is the desired location of the landing gear. If it were to be in a wing-mounted nacelle, this would impact the desired propulsion system arrangement.

o Location of, and access to, air frame-required engine driven accessories.

The scope and timing for this STAT study did not permit the full evaluation of the gearboxipower section combinations and their impact on the above factors.

It was, therefore, decided to select different arrangements for the two STAT engines which were plausible and would permit full evaluation of both at a later date.

1790 kW (2400 hp) STAT Engine The 0.5 Mach number flight speed of the Ames aircraft was suitable for a con- ventional propeller. From the NASA-furnished reference (2), a design propel- ler speed of 1250 rpm was obtained. The 1790 kW (2400 shp) engine LP turbine speed of 23,900 rpm established a reduction gear ratio of 19.1:1. A split- torque planetary reduction gear system was well suited for this ratio. The small overall diameter of this system fit well with the integral, concentric design and permitted a short, compact and lightweight engine design. A thorough study of the inlet configuratioc would be required to provide the irn desired engine pressure recovery, to avoid a high surface - to-flow - area rela- I^, tionship for the annular scoop, and to improve the level of propeller super- charging.

Aircraft secondary power systems and the engine accessories will be separated to allow engine removal without removal of the aircraft systems. This is ac- complished through the use'of a remote gearbox with a simple mechanical coup- ling for the aircraft sytems drives. The propeller control is located on the gearbox forward housing.

3579 kW (4800 hp) STAT Engine 1.^ The 0.7 Mach number flight speed of the LCC transport led to the selection of an advanced propfan propulsor ( 3) for this application. The propfan, with !'

higher disk loadings and rotational speeds than a propeller, had a beneficial impact on the gear train. The power turbine speed of the 3579 kW (4800 shp) engine is 17,7o0 rpm, which with the propfan speed of 1710 rpm establishes a reduction: ratio of Th is lowered ratio permits the use of a dual- 10.4:1.

compounc taler gearbox. This configuration has fewer gears and bearings and a load-sharing capability to equalize gear and bearing loads. also, the gear ratio can be altered within a given gear case by simple replacement of the idler set.

This engine incorporates a remote-mounted offset gearbox which provides maxi- mum reliability and installation flexibility. The offset output shaft design permits locating the propfan thrust axis near the wing centerline. Propfan blade tip ground clearance is improved by the offset arrangement. The offset also affords the use of a scoop inlet, which provides higher pressure recover- ies than the full annular type.

The propfan controls are located on the aft side of the gearbox facilitating propfan/gearbox/engine integration. A drive is provided for a remote aircraft accessories drive module.

High reliability and low maintenance coats are projected based upon the low number of powertrain gears and bearings, the simplified accessory gear train, and the remote aircraft accessories module.

Composite Materials DDA designed and produced three composite T56-type reduction gear cases in the late 19609, and demonstrated the technical feasibility of this approach. The DDA cases were produced by inserting composite preforms with directionally oriented fibers in a closed-form die. The case thus had local stiffening.

The most significant problem with this type of material was the excessive cost resulting from the hand labor involved.

Epoxy-based materials have severe temperature limitations (generally 256 K (400 0 F)). Within this constraint, few engine components could use this tech- nology. The only locations acceptable would be within, or forward of, the compressor inlet housing, and the payoff is a small weight reduction.

t Attention has recently turned to composite cases since the cost of fiber has dropped by a factor of 3 since 1970. It has been estimated that the follow- ing advantages may be achieved with composite gear cases: • Cost: 0 to 102 reduction • Weight: 0 to 20% redcction Based on DDA's T56 composite reduction gear case experience, the most optimis- tic DOC payoff for a composite reduction gearbox appears to be: ine s Gearbox only Eagine Gearbox only Enxine % 0 cost -2.4 -0.43 -2.4 -0.29 % 0 weight -4.7 -1.63 -4.7 -1.48 % 0 DOC --- -0.291 --- -0.20 Steel Gear Teeth With Titanium Web And Shaft This technology has its payoff in an approximate 20% saving in gear weight.

Except for V/STOL aircraft where weight is a significant DOC driver, this technology does not have an appreciable payoff. The gears would be much more expensive and probably show an increase in DOC, if a detailed analysis were possible.

Finite Element Gear Analsys This technology, or design tool, should r3duce redesign effort of gearing and case designs. This is especially true where computer graphics systems supple- ment the use of three-dimensional finite element techniques. Development of this design tool is a natural extension of systems now in use by industry, and little development should be required. Although engine development effort is slightly reduced by the technology, no appreciable DOC benefit can be identi- fied.

Super-Plastic Formed Titanium Housings Super-plastic formed titanium housings represent a small potential weight sav- ings and, therefore, would have a marginal DOC benefit to the STAT engines.

Advanced Lubricants Primary lubricant development efforts are ! irected toward higher temperature capability. In the STAT commercial engines, nigher temperature capability is not an important driver.

Other Modular Construction z; The entire propulsion system will be designed using modular concepts so that failures and resulting removal and repair can be restricted to small equipment' packages with little or no disturbance to the rest of the propulsion systems, thus avoiding additional maintenance/shop costs and the opportunity for main- tenance errors.

The benefits of modularity include ease of line maintenance, shorter line and i shop maintenance repair times, and reduced spare parts requirements. These factors, in turn, reduce aircraft delay times necessitated by component re- placement. The price paid for concessions to achieve modularity will, in some cases, include a small weight increase due to incorporation of quick-discon- nect features. Another penalty to consider in determining the extent of mod- ularity attainable is the increase in turbine blade tip clearances due to the elimination of line boring major structural members. This was discussed earlier under Clearance Control. Both of these considerations, though not quanitifed during this study, are judged to be more than offset by the advan- tages of modularity to maintainability.

Remote Accessories Accessory drives will be isolated and modularized so that the engine or reduc- tion gearbox can be removed without removing most accessories. Also, required maintenance to such modules as accessory drive gearboxes will be performed without removal of the engine or reduction gearbox. The objective will be minimal equipment removal and disturbance, to perform a maintenance action.

Advanced Propellers Technical evaluation of advanced propeller concepts was accomplished by Hamil- ton-Scandard and McCauley under contract to NASA for future commuter/local :service aircraft. Advanced propellirs were studied for application at cruise baseline current technology speeds of 0.7 MN and 0.47 MN . At 0.7 YI N propellers were referred to the Lockheed Electra application data base using two types of airfoil construction, namely, solid aluminum and the more recent spar-shell technique. Advanced STAT propellers for the 0.7 MN application featured advanced composite construction, proplets, and precision synchroniza- tion with an increase in the number of blades from four to six for the same power input and diameter. The advanced propellers operated at to%er tip speed. Cruise effeciency wac improved from 78.6% to 85.6% (an improvement of 8.9x), while the fuselage acoustic weight penalty was reduced from 6% of air- plane empty weight for a cabin noise level of 98 dB OASPL, to 3.3% of airplane empty weight at a reduced cabin noise level of 85 dB OASPL. along with these improvements, propeller weight was reduced 28% from the solid aluminum and 3% from the spar-shell current technology propellers. Advanced propeller OEM costs were higher by 17% than the current technology spar-shell type, which is 52% !sore costly than the solid aluminum type. Evaluating these factors in the DDA 0.7 MN airplane model results in the improvements for the advanced STAT propeller, compared to the current technology with solid aluminum airfoils, as listed in Table XLI.

HIGH SPEED AIRCRAFT TABLE XLI. ATE STAT PROPELLER IMPROVEMENTS FOR THE Empty Flyaway Block Use X Weight Cost Fuel DOC % A %A %A Value Change %a Prop fuel average efficiency .797 +6 -0.5 -0.4 -6.5 -2.4 (Avg. of climb and cruise) -45 -3.8 -0.8 -1.5 -0.9 Acoustic weight fraction 6.0 (% of empty weight) -0.8 Propeller weight fraction 4.6 -28 -2.9 +0.6 -1.1 (% of gross weight) +78 +0.4 Propeller OEM price fraction 1.8 --- +2.3 --- (% of aircraft cost) -3.7 Net Change* --- --- -7.2 +0.5 -9.1 *Refer to Tables X, XI, and XII for base values of DOC, Block Fuel, Flyaway Cost, and Empty Weight for the high speed airplane.

Thus the advanced STAT propeller for th 0.7 M N airplane results in net im- provements in DOC of 3.7%, in block fuel of 9.1%, and in empty weight of 7.2%, while increasing flyaway cost 0.5%. In addition, cabin noise level was re- duced from 98 dB for the current prop to 80 dB OASPL for the advanced STAT prop.

At 0.47 MN , current technology propellers are the standard general aviation type employing solid aluminum blades with circular shanks. Current technology also include an improved type using lighter weight spar-shell construction and more efficient airfoil shaped shanks. The improved propeller efficiency is 3 percentage points better at cruise than the standard, 15% lighter in weight, but costs about 44% more. The advanced STAT propeller for the 0.47 M N ap- plication (compared to the improved general aviation propeller) features prop- lets and precision synchronization with an increase in the number of blades from three to six for the same horsepower. Cruise efficiency was improved from 87.5% to 92.3% (an improvement of 5.5%) while the fuselage acoustic weight penalty was reduced from 4.5% of the airplane empty weight to 0%, while holding the 85 dB cabin noise level. These improvements resulted in an in- crease in weight of the advanced STAT propeller compared to the improved cur- rent type, as well as an increase in cost. Weight increased 20.3% and cost increase 155%. Evaluating these factors in the DDA low speed 50 passenger airplane results in changes in DOC, block fuel for the 185.2 km (100 am) seg- ment, flyaway cost, and empty weight for the advanced STAT 0.47 M N propeller compared with the improved current type, as listed in Table XLII.

TABLE XLII. ATE STAT P PELLER IMPROVEMENTS FOR THE LOW SPEED AIRCRAFT u l Empty Flyaway Block Base 2 Weight Cost Fuel DOC

Value Ch anze %e

%o %e %o 0.811 Prop fuel average efficiency +7.3 -0.8 -0.6 -8.0 -3.1 (Avg. of T.O. + 4 x cruise) 4.5 Acoustic weight fraction -100 -7.1 -1.7 -3.2 -2.1 (% of empty weight) Propeller weight fraction 2.7 +20.3 +1.3 +0.2 -0.7 +0.3 (Z of gross weight) Propeller OEM price fraction 1.3 -- +155 --- +3.2 +0.8 (% of aircraft cost) Net Change * --- --- -6.6 +1.1 -10.5 -4.1 *Refer to Tables X, XIII, and XIV for base values of DOC, Block Fuel, Flyaway Cost, and empty weight for low speed airplane.

The advanced STAT 0.47 MN propeller resulted in significant gains over the improved current technology type when considering the desired cabin noise level of 85 dB OASPL. The new propeller resulted in improvements of 4.1% in DOC, 10.5 % in block fuel, and 6.6% in empty weight while increasing acquisi- tion cost 1.1%.

Turbofan Enz ine The possibility of using turbofan powerplants for STAT was reviewed. The most likely opportunity was judged to be for the high speed airplane. Combining the performance of the advanced STAT propeller previously discussed with the performance of a high pressure ratio turboprop engine at 0.7M N , 6096 in (20,000 ft), maximum continuous power resulted in a thrust sfc of 0.499. This result was compared with the performance of an advanced 6.0 bypass ratio tur- bofan which had an sfc of 0.631 at the same condition, thus indicating a 21% sfc advantage for the turboprop engine. Based on this indication of the fuel savings potential of the turboprop engine combined with advanced propeller I^ technology, it was decided to concentrate on the turboprop power plant for the purposes of the STAT study.

Bleed and Power Extraction An assessment was made of the effect of aircraft services provided by the en- gine on engine performance. Lockheed determined in their STAT studies of 50 passengar a..,irt haul aircraft that aircraft requirements amounted to 52.2 kW (70 shp) per engine for electrical, hydraulic, and cabin environmental re- quirements. A parametric study was conducted using an advanced turboprop en- gine designed for efficient cruise at a compressor pressure ratio near 20:1 I I at a rotor inlet temperature of 1366 K (2000°F). The flight condition chosen i was 0.7 M N at 6096m (20,000 ft) altitude. Accessory drive power was ex- tracted from the low pressure spool. Compressor bleed was taken from compres- sor discharge, and using gas state conditions, an equivalent power was calcu- lated assuming an 80% efficiency for an air driven turbine. Results obtained are shown in Figure 39 for the baseline with no bleed or power extraction, and for cases where the required horsepower was furnished (either all by power extraction or all by bleed), and for the case where half the required horse- power is provided by bleed and half by power extraction. Relative shp is plotted vs relative sfc. Cruise power available to the propeller is reduced 2.1% if 512.2 kW (70 shp) is extracted from the LP spool. This percentage re- flects an arithmetical subtraction of the aircraft requirement. For the split case, power loss is 4.6%, and for the all bleed case, power loss is 7.5X. At a given horsepower of 96% of the referet±ce value, relative sfc is 1.008, 1.028, 1.032, and 1.035, respectively, for the baseline, 52.2 kW (70 shp) (shaft), split shaft and bleed, and 52.2 kW (70 shp) (bleed). Thus, for the cases examined, direct power extraction offers the best alternative with re- spect to minimizing engine performance losses.

Other factors should be evaluated before drawing conclusions on the optimum system to provide aircraft services. These would include a better definition of the power extraction and bleed requirement, the engine power range over which this must be furnished, whether the requirement is continuous or vari- able, and the trade-offs involved in weight and volume of the environmental conditioning system for variation in air turbine efficiency.

1. OF Altitude • 60% m (20, 000 ft) Mach No. • 0.7 Reference power - 2446 kW (3280 shp) 1.06 Reference shaft sfc - 0.332 Bleed • 52.2 kW (70 shp) equiv.

Bleed • 26 kW (35 shp) equiv.

PX • 26 kW 05-shp) Relative 1.04 SFC 1.02 PX • 52.2 kW (70 shp) 0 bleed, 0 PX 1.00 0L 0.8 0.9 1.0 Rel'ative SHP TE81-719A Figure 39. - Impact of compressor bleed and engine power extraction on performance.

Reduced Weight Nacelles

1j

A trade study was conducted to determine the airframe sensitivity to nacelle weight reduction. Figure 40 shows these sensitivities as applicable to both the high speed and low speed aircraft where the nacelle weight fraction ap- proximated 0.2% of the take-off gross weight (actually 0.21% for the high speed aircraft and 0.172 for the low speed aircraft). The lines shown closely approximate the sensitivity of both the high speed and low speed airplanes to change in nacelle weight as it affects empty weight, flyaway cost, block fuel, and DOC.

A significant consideration in the design of the nacelle for the advanced tur- boprop engine is the inlet configuration, especially for the high speed air- plane. The flow field behind the rAvanced 0.7 M N STAT propeller should be examined. A screening evaluation of candidate turboprop engine core inlets 1 , including concentric types (namely, annular and bifurcated) and scoop types should be undertaken to explore inlet pressure recovery and pressure distor- tion characteristics.

ENGINES ADVANCED TECHNOLOGY Two configurations evolved for DDA's STAT advanced technology engines. The basic difference in the engines stems from tae two study aircraft power re- quirements. Engines for the 0.5-Mach Lockheed transport were studied at the nominal size of approximately 1790 kW (2400 shp), and the 0.7-Mach Lockheed transport at 3579 kW (4800 shp). The engines are similar in that both have a •oith front drive. The overall single-spool gasifier and a free power turbine design and performance features of these engines are shown in Table XLIII and XLIV.

.2

Nacelle weight Empty

Percent fraction . .002 weight

reduction i ^ 1

Flyaway

system

cost

parameter

Block fuel

and DOC

0 10 20 30

Percent reduction in nacelle weight

TE81-720

Figure 40. - Sensitivity of aircraft system parameters to nacelle weight.

ti

_A

TABLE XLIII. - ADVANCED STAT TURBOPROP ENGINES (SI UNITS) High Speed Low Speed Aircraft Aircraft Shaft power--kW 1749 SLSS SLSS Condition 2491.00 Jet thrust--N 1401.19 Equivalent shaft power --kW 1843 3711 Cycle Re 20 20 ' RIT--K 1506 1506 71.6 64.5 sfc--og/w • s ESFC--μg/W • s 68.1 61.7 11.12 Airflow--kg/s 6.19 100 100 Inlet recovery, % 1.114 1.114 Nozzle pressure ratio % 99 99 Gearbox efficiency, Overboard seal leakage, % 1.0 1.0 Technology Compressor (type-stages) Axial-5/cent-1 Axial-9 Turbines (type-stages) Axial hybrid; Axial hybrid; HP-2, LP-2 HP-2, LP-2 Composite Shafting Composite Components Compressors: 82.29 83.75 n Adiabatic, % N, rpm 39767 29823 Axial: Tip speed ( U t /0) m/sec 411.5 457.2 Rc aver./stage 1.35:1 1.4:1 AR blades, aver. 1.2 1.07 Centrifugal: 0.75 0.5 ) N s 28.7 - - (rpm m /sec 0.545 - - N; Tip speed ( U t AW , m/sec 660.8 - - Rc _ 4.47:1 - - Blade to blade shroud loading,LD 0.24 - - Combustors: -4 comb., 2 99.9 99.9 e P/P burn. 0.04 0.04 Turbines: High pressure.

% 87.78 91.52 nAdiabatic, Aver. Stage loading coef.

(gJeh /U mean) 1.60 1.55 Equivalent work (oh/9cr) WI Ng 0.102 0.100 4.727 Expansion ratio 5.32 Inlet temperature, K 1505.6 1505.6 8.3 8.3 Cooling Airflow, % film, Type of cooling (1st blade) Imp. film, Imp.

cast -in cast-in i TABLE XLbi:. (CONT) High, Speed

Low Speed of

Aircraft Aircraft Low pressure: 23,900 N, rpm 17,780 aAdisbatic, Z 89.0 91.2 Aver. stage loading coef.

(gJAh/U mean) 1.50 1.68 0.090 Equivalent work (Ah/Bcr), Btu/lbm 0.076 Expansion ratio 3.183 3.584 Inlet temperature, K 1063.3 1070.0 % 0.5 0.5 Cooling airflow, Weight--kg Power section 189 250 Gearbox (ratio) 130 (19.1:1) 149 (10.4:1) 319 399 Total Power/weight 5.48 8.88 Length--m Power section 1.247 1.285 1.793 2.215 Overall _9ax height--m 0.668 0.734 Price*--$ 243,100 407,600 Power section I Total 286,000 468,500 164 132 $/kW *OEM price at 12 /mo, 1000 units, 1979 economics UNITS) TABLE XLIV. - ADVANCED STAT TURBOPROP ENGINE (CUSTOMARY Low Speed High S — ed Aircraft Aircra.._ SHP 2345 4752 SLSS SLSS 1 Condition 7 315 560 Jet thrust--lb ESHP 2471 4976

,I

Cycle R 20 20 C RIT-- oF 2250 2250 0.424 0.382 sfc--lbm/hp-hr ESFC--Ibm/hp-hr 0.403 0.365 13.65 24.69 Airflow - -pps % 100 100 Inlet recovery, IA 1.114 Nozzle pressure ratio 1.114 Gearbox efficiency, % 99 99 1.0 1.0 Overboard seal leakage, X TABLE XLIV. (CONT) Low Speed High Speed

Air craft Aircraft

Technology .Axial - 5 /cent.-1 Axial-9 Compressor (type-stages) Axial, Hybrid Axial, Hybrid; Turbines (type-stages) HP-2, LP-2 HP-2, LP-2 Composite Shafting Composite Components Compressors: 83.75 82.29 ,1Adiabatic, % 29,823 31,761 N, rpm Axial 1500 t /M, fps 1350 Tip speed (U 1.4:1 Rc aver./stage 1.35:1 1.07 1.20 AR blades, aver.

Centrifugal: 0.75 Ns , rpm ft /secO.5 Ns (dimensionless) .545 fps 2168 Tip speed (Ut /M, Rc 4.47:1 _ .24 Blade to blade shroud loading,LD Combustors: 99.9 99.9 '9 Comb., % .04 .04 J% P/P burn.

Turbines: High pressure: 91.52 87.78 'q Adiabatic, % Aver. stage loading coef.

1.55 (gJ&h/U 1.60 mean) 42.9 43.7 Equivalent work (Ah /Bcr), Btu/lbm 4.727 Expansion ratio 5.32 2710 2710 Inlet temperature, OR 8.3 8.3 Cooling airflow, X Imp f ilm, Imp film, Type of cooling (first blade) cast-in cast-in Low pressure: 17,780 23,900 N, rpm 89.0 91.2 Adiabatic, % Aver. st a loading coef.

1.50 1.68 (SJAh/^ mean) 32.8 38.8 Equivalent work ( Ah /Bcr), Btu/lbm y.183 3.584 Expansion ratio Inlet temperature, OR 1914 0.5 0.5 Cooling airflow, % a TABLE XLIV. (CONT) ^i Low Speed high Speed Air craft Aircraft Weight--lb - Power section 416 287 (19.1:1) 328 (10.4:1) Gearbox (ratio) 703 879 Total 3.34 5.41 Power/weight Length--in.

49.1 50.6 Power section 70.6 87.2 Overall Max height--in. 26.3 28.9 Price*--$ S 243,100 407,600 Power section s 286,000 468,500 ` Total . $/SHP 122 99 .Y . *OEM price at 12 mo, 1000 units, 1979 economics 1790 kW (2400 shp) STAT Engine This engine has an axial (5) centrifugal (1) compressor with welded titanium drum and high temperature titanium aft wheels. The final stage of the com- pressor is a hybrid impeller with cast flow-path ring HIP bonded to a forged hub. The compressor rotor is straddle mounted for dynamic stability and clearance control. The compressor materials will be selected for the rotor, blades, case, and vanes to reduce thermal mismatch and thereby improve clear- ance control.

The combustor has a fold-back design and features a Lamilloy, transpiration- cooled liner. The two-stage gasifier turbine employs cast-in impingement cooling in both stages plus the first vane stage. Both turbines feature hy- brid wheels with cast airfoil rings diffusion bonded to powder alloy hubs.

The turbine rotor/case/blades/vanes materials will be selected to minimize thermal growth mismatch. The gasifier turbine incorporates an active clear- ance control device actuated by the electronic fuel system control.

The turbines incorporate abrasive blade tips and abradable coatings to provide minimum running clearances. Paired knife step seals will be used to reduce seal leakage losses.

Turbine shafts are borsic-titanium matrix composite for light weight and for the required shaft stability in the small diameters required of advance tech- nology, high pressure ratio engines.

The reduction gear for the 1790 kW (2400 shp) engine is close-coupled, coax- ially inf root of the power section. It incorporates a split-torque planetary gear system with 19.1:1 speed ratio, and features a composite gear case.

I The control and fuel systems for both STAT engines shall be configured to utilise an advanced technology digital electronic controller for all required logic and computational re^..:,irements for the engine and propeller operation to accomplish the following control functions: o Automatic start sequencing o Turbine inlet temperature limiting during all operation, including start, for turbine protection and long life o Control for acceleration and deceleration fuel flow, bleed, and compressor geometry for smc+oth and rapid operation without surge or flameout o Control for gas generator speed as a function of power lever input posi- tion to provide modulation of engine power from max rating-to-idle-to max reverse o Control for propeller/power turbine speed over the required operational range o Independent back-up control function that limits maximum power turbine overspeed o System for autofeather based upon operating through the propeller control system o Provisions for torque limiting for gearbox protection o Provisions for automatic mode selections for optimum thrust control (take- off, maximum climb, maximum cruise as a minimum) o Provisions for digital link interfacing with flight control system for automatic propulsion control throughout all regimes of engine operation o Propeller synchronizing/synchrophasing 3579 kW (4800 shp) STAT Engine This engine has a nine-stage axial compressor with welded titanium drum and high-temperature titanium aft wheels. The materials for the rotor, blades, case, and vanes will be selected to reduce thermal mismatch and thereby reduce clearance control.

A VCD forms the transition between the compressor and the flow-through, Lamil- loy-cooled annular combustor. The turbine section of this engine features the same number of stages, materials, and cooling technology as that of the 1790 kW (2400 shp) STAT engine.

The reduction gear for the 3579 kW (4800 shp) engine is a remote-mounted, off- set type featuring an advanced dual-compound idler gear train with a 10.4:1 reduction ratio. The gear case is made of composite material.

Tables XLV and XLVI show the changes in engine and aircraft parameters for the advanced 1790 kW (2400 shp) and 3579 kW (4800 shp) engines respectively, which incorporate the advanced technologies and design features judged beneficial as described in the previous sections. The total benefit of a group of advance technologies may differ from the sum of the individual benefits attainable from each technology item when considered separately. Therefore, the results presented in Tables XLV and XLVI may differ from the sum of those technology benefits when described individually.

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w i H M M M 1.1 Maintainabilit Propulsion system condition monitoring provisions incorporating advanced tech- nology components will be incorporated in the STAT engines to permit detection of impending malfunctions and to define the required maintenance action. Ear- ly detection and correction of potential problems will result iu improved air- craft safety and reliability. Transducers which are required to measure com- ponent pressures, temperatures and positions, along with the associated wir- ing, will be integral parts of the electronic fuel control systea.

Modular Construction The entire propulsion system will be designed using modular concepts so that failures and resulting removal and repair will be restricted to small equip- ment packages, with little or no disturbance to the rest of the propulsion system, thus avoiding excessive maintenance/shop costs and the possibility of maintenance errors.

Enzine Noise Considerations The DDA advanced technology STAT engines, like conventional engines, will rad- iate noise from the compressor inlet, and combustion, turbine, and jet noise from the engine exhaust. The engine case and reduction gAarbox will also rad- iate noise: however, they and the jet will be minor noise source&• Turbine tones will be substantially above the 10 kHz analysis range (18 kHz at take- off) so that thq compressor and combustor remain as the major engine noise sources affecting aircraft far-field noise.

Compressor Noise The primary impact of the advanced technology will be on compressor noise.

The advanced engine cycles that yield low fuel consumption require high pres- sure ratios, which, in turn, lead directly to supersonic blade tip speeds and increased pressure ratios in the compressor initial stages. The net effect is multiple pure-tone generation and enhanced bladepass tones during takeoff and, depending upon the aircraft power requirements, on approach. Engines incorporating advanced technology compressors will require inlet noise sup- pression in the form of swept first-stage blades, sonic attenuation in the inlet guide vanes, or inlet duct acoustic treatment.

Combustion Noise Combustion noise levels for the STAT engines are comparatively unknown, since the engine parameters used in the conventional predictions are outside the data base used to generate the prediction empiricisms. In any event, if the STAT engines product 2 to 4 dB more combustion noise than current engines, as predicted, combustion noise will be a major contributor to propulsion system noise. The understanding of combustion noise generation is not sufficiently developed to permit noise reduction either by combustor design or engine cycle bias. Until these controls are developed, combustion noise reduction can be achieved by exhaust duct treatment and/or using shielding.

t ^ - oals Propulsion System Noise--Meeting the STAT ( Current turboprop aircraft in the STAT-size category just meet or exceed the FAR Part 36 noise requirements, indicating that a propeller noise reduction of 8 dB is required to meet the STAT goal of Part 36-8 EPNdB. A modern, thin, elliptical tip propeller or propfan used in concert with increased take-off performance cal provide the required noise reduction. However, unless engine compressor and combustion) is also reduced, the net aircraft gain will noise ( be substantially less than the 8 dB required. Figure 41, which is based on measurements of the Electra engine, illustrates this situation. This figure ' s technology is the dominant shows clearly that while the propeller of today noise source, engine combustion noise provides a floor effect that will, in large measure, negate the propeller propfan noise reduction. The addition of tones to the engine spectrum will raise the EPNL frequency sen- multiple - pure - sitive portion of the spectrum as well as provide an increase in forward rad- iation to increase the time during flyover when a given noise level can be perceived on the ground.

Takeoff

N

Lockheed Electra--CV580 propeller technology Z i STAT propeller R Multiple pure-tone I ' frequency range m

'K i

o

*` +' t

i

Propeller reduction Engine Noire required to meet STAT pool ror 100 200 1000 2000 4000 1/3 octave bond center frequency--Mz TE80-2074 Figure 41. - Comparison of propeller and engine noise during take-off.

In the absence of a well developed noise source reduction technology for tur- boshaft engines, duct treatment may be used to control engine noise but with an associated increase in fuel consumption, and decrease in power, on the A oid this performance penalty by order of 0.5 to 42. The STAT engines will giving proper consideration to noise reduction during design and development phases of the components and the engine.

BENEFIT ASSESSMENT The benefit assessment completes the evaluation of the advanced technology STAT engines by quantifying their effect on the critical airframe and system parameters. In addition to the advanced technology engines, baseline deriva- tive (1985 time frame technology) engines were also analyzed. The resulting data permit comparisons of the potential payoff between: o 1990 advanced technology u:d 1980 baseline technology o 1990 advanced technology and 1985 derivative technology o 1985 derivative technology and 1980 baseline technology The evaluation of potential payoff for advanced technology was based on re- sults from a final airframe parameter analysis using the baseline mission de- scribed earlier. The airframe and engines were sized to match the required mission, and the resulting operational criteria were quantified. These in- clude: • Aircraft DOC for stage lengths of 92.6, 185.2, 277.8, 370.4, 740.8, and 1111.2 km (50, 100, 150, 200, 400, and 600 nm) • Fuel used • Acquisition price • Total cost of ownership for 5 years based on the following parameters: o Fuel cost $0.264 and $0.396/L ($1.00 and $1.50/gal) 2800 hour/year utilization Sfc Trends Sfc trends for current, derivative, and advanced technology turboprop engines are shown in Figure 42 in terms of uninstalled brake sfc as a function of rated shaft horsepower. The trend for current technology turboprop engines •resulted from compilations of engine manufacturer's data analyzed by the Lock- heed California Company in performance of their STAT Short Haul Study as described earlier in this report. A turboprop version of the XT701 engine, designated PD370-37, established the sfc at the upper end of the horsepower range shown, and otter Allison engines such as the Model 501 In the 3729 kW (5000 shp) class and the Model 250 in the 377.9 kW (500 shp) class, influenced the placement of the trend line.

L: 1 The "derivative" engine technology trend line is based on that for the current technology engines and is keyed to a variant of the PD370-37 in which tha coar pressor pressure ratio is increased from 12 . 7 to 17.7 and RIT is increased from 1506 to 1533 K 12250 to 2300 e F). The derivative engine incorporates a scaled version of a compressor demonstrated in advanced technology programs at DDA and uses the basic shaft, bearing, and turbine arrangements from the XT701 turboshaft engine. The reduction gearbox for the full-scale derivative engine is a now simplified design icorporating a dual-compound idler arrangement and increased design life compared to the T56 series of gearboxes. The new gear- box design is based upon a study of the reliability and maintenance cost his- tory of past turboprop systems and incorporates the recommendations of that study for a gearbox with high reliability, easy maintainability, and low main- tenance costs. Realistically, the derivative engine postulated is a major modification involving significant changes in the cold and hot sections of the J engine as well as in the gearbox. Performance and maintainability improve- ments and cost estimates reflect these changes. The sfc improvement is 11% compared to current technology, as shown in Figure 42.

I l E 0.6 ^ GumeM technology

I

l 90 i i Derivetlw technology — — —^ 0.5 -1171 I _ I XT701 tu&Wop s0 ` ^ -- — 1 _ ^XT701 nMboprap 0.4 I I _ _ d^riva lv.

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M I } N U `^ W r Advanced technology I 0.3 , -SO $hoft power, SLSS, take-elf power—kW II1 l 0 2000 4000 6000 6000 10000 Shaft power, SUS, toke"N pewee slip TM-207SA Figure 42. - Sfc trends and comparisons.

The advanced technology engines at 1790 and 3579 kW (2400 and 4600 shp) have an additional improvewnt in of c over the derivative engines of 62 and 9X re- spectively, for a total improvement over current technology of 17Z and 19X9 respectively.

Price Trends Historical turboprop production engine prices were reviewed based on DDA data for 3729 kW (5000 shp) and 5966 kW (8000 shp) class engines . The approach was to develop a consistent set of data for extrapolation to STAT sizes. The economic baseline for STAT was: o 1979 dollars o 1000 engines produced at a rate of 12/month o cumulative average price for 1000 engines (with and without inherited learning) Figure 43 shows relative prices for the T56 and XT701 power sections based on actual DDA experience adjusted to the standards as previously stated, includ- ing consideration of the engines as having "no common production base." The learning parameter is considered in order to reflect the price of an available Derivative technology PD370-41 [8183 kW (10,974 shp)] • Common economics • Common production rate • All with cannon production base advantage except * a no cannon production base Current technology PD370-37 T56 Series 1 [2581 kW (3460 shp)] I u T56 Series II T56 Series IV [ (2800 kW (3755 shp) ] T56-A-100 [4157 kW (5575 shp)] s a 3o 1956 1961 14,000 4,000 6,000 8,000 10,000 12,000 16,000 2,000 Total production TE80-2076 Figure 43. - Power section price comparisons.

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current technology engine at the power size required (thus inherited learning applies in pricing) versus a case where a completely new engine is constructed using current technology (thus, in the extreme case, no inherited learning applies).

The model T56-A-100 engine noted in Figure 43 is a current technology Series IV T56 engine with 13,000+ production engines and a 25-year historical back- ground. This engine has redesigned compressor blading to produce a 11.9:1 compressor pressure ratio (CPR) versus the 9.6:1 ratio of the model T56-A-15.

The turbine blading has been redesigned to provide 4157 kW (5575 take-off horsepower). The power section weight is 571 kg (1259 lbm). The power sec- tions of the Series 1, 2580 kW and 524 kg (3460 shp and 1155 lbm), the Series II, 2800 kW and 540 kg (3755 shp and 1190 lbm), and the Series III, 3424 kW and 553 kg (4591 shp and 1220 lbm) engines all were single spool, axial flow power plants driving an offset reduction gearbox. T56 Series III engines have a 9.6:1 CPR, a 15 kg/s (33 lbm/s) airflow, and a RIT of 1350 K (19700F).

The projected relative price line shown for the Series IV engine is based on accumulated learning of the T56 production experience. This engine, when factored to "no common production base" is also shown.

The model PD370-37 current technology engine is a turboprop conversion of the XT701 turboshaf t engine, where the LP turbine was rematched for better turbo- prop characteristics. It represents a minimum change from the XT701 engine, which is a free turbine turboahaft engine that was developed through safety demonstration for the U.S. Army's HLH program. The Model PD370-37 is a single-spool, axial flow power unit with a free turbine that is connected by shafting and supporting structure to an offset reduction gear assembly, which is based on an improved T56-A-15 gearbox design. The baseline PD370-37 engine is in the 5966 kW (8000-hp) class. Compressor pressure ratio is 12.7:1, the compressor rpm is 15,049, and RIT is 1506 K (2250°F) for take-off on a stan- dard day. Airflow is 20.8 kg/s (44 lbm/sac), and the power section weight is 510 kg (1125 lbm). Price trends for this engine are also shown with and with- out inherited learning.

The Model T56-A-100 and the KT701 current technology engines are considered to have the "inherited" learning of approximately 7000 engines, and were adjusted to the 1000 engine basis applying a 90Z learning curve consistent with DDA experience.

Derivative engine technology is represented by the Model PD370-41 engine, which is a X1701 turboprop derivative engine with a 17.111 compressor pressure ratio. It incorporates a scaled ATEC.0 demonstrated compressor with combustor and turbine arrangements from the new XT701 turboshaft engine. It is an axial flow engimn, having a single-spool core and a front drive free power turbine connected by shafting and supporting structure to an offset reduction gearbox.

The engine is in the 8203 kW (11,000 hp) class. At the 17.7:1 pressure ratio at sea level take-off (SLTO), RIT is 1533 K (2300'F), airflow is 24.5 kg/s (54 lbm/sec), compressor rpm is 16,400, power is 8183 kW (10,974 shp), and the power section weight is 682 kg (1503 lbm). This engine has an advanced design gearbox of the dual-compound idler concept, which is lighter in weight than the current technology gearbox.

Data from Figure 43 is used to establish power section OEM comparative prices in Table XLVII. Gearbox price estimates are without inherited learning, shown added to the power section price to obtain a total engine price comparison without inherited learning. The gearbox costs have been factored from histor- ical costs similar to the power section. However., a 94% learning curve was applied for current technology gearboxes.

TASLE XLVII. - ENGINE PRICE SUMMARY T56, Series IV XT701 XT701 derivative Engine base (Model T56 -A-100) (Model PD370-37) (Model PD370 -41) Current Derivative Technology Current 8184 (10975) Size --kW (shp) 4157 (5575) 6301 (8450) *Power section price $582,200 $782,800 $1,079,200 $114,900 $88,000 0 *Reduction gear price $102,400 $1,167,200 *Engine price $684,600 $897,700 *Cumulative average OEM prices of first 1000 engines at 12/mo, 1979 economics, without learning or benefit of previous production base.

Price trends taken from the data points shown in Table XLVII were then con- structed (using the scaling relationship: scaled price • unity price x (scaled power/unity power] 0.7 ), and are shown in Figure 44. Note that the T56-A-100 falls very close to the current technology specific price versus take-off rated shp "trend line." Also note that the derivative engine trend curve par- allels the current technology trend line, and is 5 to 7% higher than the cur- rent technology lines as a result of the new compressor and turbine design and 0F).

the increase of RIT to 1533 K (2300 The advanced STAT technology 3579 kW (4800 shp) engine is a high pressure ratio conceptual engine with a compressor pressure ratio of 20:1, a W a of 11.6 kg/s (25.5 lbm/sec), an RIT of 1506 K (2250'F), and a 29,800 compressor rpm. The baseline engine power section has been calculated to weigh 250 kg (551 lbm). The engine is a single-spool, axial-flow power unit with a free turbine connected to an offset reduction gear assembly.

This STAT engine is approximately 16% lower in cost than current technology engines of the same power. The small compressor diameter and overall low weight of the power section plus the new design gearbox contribute largely to this low cost. The costs were derived parametrically from known costs on DDA production and demonstrator engines and DDA MIF methodology and have been factored to a common basis.

The advanced STAT technology 1790 kW (2500 sh,)) engine has an axial compressor centrifugal outlet stage. This engine has a 20:1 combined with a single-st.Age CPR, 6.35-kg/s (14.0-lbm/sec) airflow, 38,700 rpm, and a 1506 K (2250 °F) RIT, fI

i

1910 ecattomi a engines Cuawlotive average price of 1000 i Production rote of 12 o elno Per month ► production boo No conma z—r—^'1 -1> % I at

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/_ P0370-41 Derivative technology ' — Cwront technology %— P0370-37 Advance STAT technology SO 40006000 =Smr^ 0 0 Take-cN rating, SLSS, t , *@-eff powoo W I -- ' - -f 8000 10,000 12,000 4000 6000 Take-off rating, SLSS. take"N power—thp TM-2077A Figure 44. - Engine specific price comparisons.

with a power section weight of 188.7 kg (416 1bm). This STAT engine is a single-spool, axial-centrifugal power unit with a free turbine connected to an in-line reduction gear assembly of advance design.

Maintenance Cost Trends, Maintenance and Reliability Recommendations Advanced Turbooro p A study of actual turboprop reliability and maintenance costs in commercial operation, conducted by DDA for NASA and reported in NASA CR 135192, showed that turboprop maintenance costs could be reduced in future propulsion systems by incorporation or improvement of the following elements: • Incorporation of on-condition maintenanca • Improved modularity • Incorporation of total system management to integrate all elements of the propulsion system.

• Improved reliability and durability Or.-Condition Maintenance w!

A prime cost driver in the maintenance of aircraft propulsion systems can be scheduled overhauls. It was found in the reliability and maintainability study for RASA that scheduled overhauls accounted for 40% of the maintenance - 580s. A poten- costs of the turboprops on the Lockheed Electra and Convair CV tial for most of the 40% reduction may, therefore, be realised for the STAT engines by adopting on-condition maintenance. Scheduled overhauls are elimi- nated, and periodic inspection or equipment malfunction are the only sources to justify equipment removal. The on-condition maintenance concept can be facilitated by improved fault detection and isolation via diagnostics ( condi- tion monitoring) to identify impending problems so that corrective action may ^6 be taken prior to failure. Most of the large turbofan engines currently in are maintained on-condition, al- use with the aircraft of the major airlines though some section of a given engine, such as the high pressure turbine, may Iy have a time - limited removal requirement based upon operating experience. All- up condition monitoring systems are still in development and have not been adopted by the major airlines, but they should be operational by the 19908.

The estimated reduction in maintenance costs for the adoption of on-condition `l maintenance and condition monitoring is discussed later iu this section.

Total System Manazement The management of the total propulsion system must be centralized, including the propeller, engine, and nacelle. This would ensure the proper definition and control of interfaces between the three major system modules, which would minimize independent approaches to major module design. Maintenance access to all components and total system condition monitoring would be well integrated.

Reliability and Life Goals The reliability and life goals for the advanced STAT engines were established consistent with safety, a minimum maintenace cost for the mature system, mini- mum delays and down time, and the expected service life of the engines in short haul commuter airline service.

Life Goals Design life was defined as the time or life that the propulsion sytem would operate satisfactorily: • With routine sche'Auled maintenance • Without scheduled replacement of parts or components • With unscheduled replacment frequencies ao more than are consistent with the stated MTBF k_:an time between failure) values Using the guidelines of the contract work statement and other direct operating ground rules supplied by NASA, in which airframe design life of at least off and landing cycles, and a spares factor of 1.3 30,000 hours, 60,000 take - for engines were definvcd as requirements, the following life goals for the engines were established: • Design life of high pressure turbine of 7500 hours • Design life of all other parts of 25 , 000 hours Reliability Goals Reliability goals were established for the power sections and gearboxes of the STAT engines. These goals were based on the incorporation of all advance technology items and are considerably improved over those of current technolo- MTBR), based upon propul- gy engines. Mean time between unscheduled removal ( sion system inherent events, should be no less than those shown in Table XLVIII for major modules, and Table XLIX for campon±nts or accessories. For comparison, this table also shows removal races for nearly 2.5 million flight- hours of experience on the Series I T56 engines.

r I TABLE XLVIII. - INHERENT RELIABILITY GOALS FOR STAT ADVANCED TL'RBOPROP - r MAJOR MODULES Inherent Corresponding removal Ref. T56, Series I MTB R - rate/1000 h removel ratk / 1000 h Major module h Core engine 6,250 0.160 0.426 0.020 LP (power) turbine 50,000 Main drive reduction gearbox 33,333 0.030 0.060 Total for major modules 4,762 0.210 0.486 TABLE XLIX. - INHERENT RELIABILITY GOALS FOR STAT ADVANCED TURBOPROP-- COMPONENTS AND ACCESSORIES Components and Inherent Corresponding removal accessories MTBR--h rate/1000 h Power section major accessories 50,000 0.020 (oil, scavenge, fuel pumps,

' ignition)

Engine accessory drive gearbox 50,000 0.020 0.150 Power section minor accessories 6,667 f Control System 2,500 0.400 Start system 3,700 0.270 Total for major components 1,163 0.860 ^. Maintenance Philosophy Maintenance philosophy is defined as a characteristic of equipment design that + facilitates maximum maintenance effectiveness and minimum TCO. Cost-effective maintenace concepts must be designed i:Ato the end product during the concept- ual and design phases. An on-condition maintenance approach, as described earlier in this section, is a result of this philosphy and is facilitated by r the improved durability, condition monitoring with improved diagnostics, in- creased modularity, and simplified hardware. Other specific design requir- ments for improved maintainability are methods of mounting for easy installa- tion and removal, accessibility, repairability, inspection / verification con- cepts, serviceability, and component detail design for ease of maintenance.

Reliability Assessment A reliability assessment was made of the STAT advaocad turboprop engines to provide the predicted improvements, in contrast to current systems and to provide a basis for maintenance cost predictions.

Inherent Versus Noninherent Reliability Inherent events / removals are those caused primarily be propulsion system equipment failures. Noninherent reliability is based upon those events that are primarily not caused by propulsion system equipment. In order to make direct comparison with current technology rate and cost data, which reflect total or " operational" removal rates and cost, an allowance had to be provided for noninheront events. This allowance has the effect of increasing the re- moval rates beyond the inherent value or decreasing the MTBR values below the inherent value. The noninherent allowance provided for the effects of the following: • Unsubstantiated / unnecessary removals • Improper maintenance - caused failures • Foreign object damage (FOD) • Convenience to perform nonpropulsion maintenance • Accident damage Summary of Total Operational Reliability Assessment The total operational reliability assessment is shown in Tables L and LI.

These values were used to determine maintenance costs. Also shown are the values of the inherent and noninherent rates used to arrive at the predicted operational rates. The inherent and noninherent reliability assessments are discussed in the following subsections.

TABLE L. - SUMMARY OF RELIABILITY ASSESSMENTS OF STAT ADVANCED TURBOPROP SYSTEM- - MAJOR MODULES Name of major Inherent removal Noninherent removal Operational removal Operational module ^ r-ate11000 h rate/1000 h rate/1000 h MTBR--h Core engine 0.160 0.040 00200 5,000 LP turbine 0.020 --- 0.020 50,000 Main drive 0.030 0.010 0.040 25,000 reduction gearbox Total for 0.210 0.050 0.260 3,850 major modules _A TABLE LI. - SUMMARY OF RELIABILITY ASSESSMENTS OF STAT ADVANCED ACCESSORIES SYSTEM - -COMPONENTS AND Name of Operational component or Inherent removal Noninherent removal Operational removal MTBR--h rate/1000 h rate/1000 h rata/1000 h accessory Engine acces- 0.020 0.005 0.025 40,000 sory gearbox

0.333 3,000

Start system 0.200 0.133 2,000

Control system 0.400 0.100 0.500

0.036 27,777

Major accessories 0.020 0.016

(oil pumps, fuel pumps, ignition) 0.200 5,000 Minor accessories 0.150 0.050

Total for 0.790 0.304 1.094 914

components and accessories Core Engine and LP Turbine Inherent Reliability Assessment The history of DDA turboprop experience was studied from two important aspects: • Whether reliability problems existed that were uniquely related to turbo- prop operation or application • Identification of principal engine problems to evaluate solutions from current technology No current engine problems were identified as being unique to turboprop oper- ations or application. Some of the problems resulted from the engines being originally designed and developed for much shorter life in military use and without the comprehensive design criteria of today. Thus, one aspect of the core engine and LP turbine assessment recognizes the beneficial results of comprehensive design criteria, which include clearly stated reliability and life requiremets for commercial operation.

Some of the corrective actions to historical problem areas are straightforward configuration and processing changes to better adapt to commercial maintenance plans and commercial overhaul periods. Others were based on improvements from technology and analytical improvement programs. It is assumed that this trend will continue to produce improved materials, coatings, and analytical tech- niques available for the advanced STAT engines.

Based upon the reliability of current mature engines, the projected inherent reliability of the advanced core is 6250 hours MTBR and of the LP turbine, 50,000-hour MTBR. The corresponding inherent PRA are 0.16/1000 and 0.02/1000 hours, respectively.

1' Noninherent Reliability Assessment An assessment of 0.040 removals for 1000 hours was made for the core engine.

The causes of these would be expected to be FOD to the compressor, improper maintenance and operation, and unsubstantiation (no fault found). With fault diagnostic systems employed and the beneficial effect on improper maintenace resulting from a lower maintenance action rate, the estimate of 0.040 seems reasonable.

Main Drive Reduction Gearbox Inherent Reliability Assessment The assessment of the STAT advanced main drive reduction gearboxes, each a major module, was based upon current engine experience. For example, the study of turboprop reliability and maintenance costs for NASA, which were men- tioned earlier in this section, found that 36% of the reduction gearbox remov- als were caused by failures in the main drive system or in other characteris- tic turboprop functions such as propeller brake, safety coupling, and negative torque signal (NTS). The advanced main drive reduction gearboxes contain those functions performed by the 36% represented previously, less the safety coupling and NTS, plus the limited accessory drive gears, bearings, and asso- ciated hardware necessary to drive the lube pump, propeller high pressure pump, and the power takeoff for the remote aircraft accessory gearbox.

Therefore, the advanced main drive reduction gear system can be compared di- rectly with the corresponding portion of the current gearboxes. In the study, it was found that the main drive system (the 36%) accounted for: • An inherent premature removal rate of 0.048 per 1000 engine flight hours (EFH) (equivalent MTBR 0 20,770 hours) • A total operational premature removal rate of 0.061 per 1000 EFH (equiva- lent MTBR - 16,400 hours) The advanced main drive reduction gear systems: • Contain less than half the number of powertrain bearings of the current main drive system • Contain proportionately fewer other hardware parts • Must be expected to operate for 25,000 hours with no scheduled overhaul (compared to about 7000 hours average for the current system) The first two points produce a favorable effect on premature removal rates, whereas the third tends to produce an unfavorable effect. To control and min- imize this potential unfavorable effect, one feature that the advanced designs will incorporate is bearing sizes selected for much longer life. Some further reduction in the failure rate is expected as a result of the improved oper- ating conditions afforded by use of helical gears and improved mounting and Seat support.

Improved alignment, concentricity, and reduction of torsionals in the remain- ing accessory drive train will contribute to attainment of full calculated life at these locations. These factors, however, cannot be easily quantified to produce a further improved failure rate value.

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Consideriag all these factors and conducting a reliability analytical analysis with the record of the current system as a baseline, a premature removal rate of 0.01 per 1000 hours was determined for the advanced main drive systems.

A premature removal rate of 0.01 per 1000 hours was assumed for the accessory drive system within the main drive reduction gearbox. This assumption was based upon the complexity of the accessory drive system being comparable to that of the main drive reduction system and that the same reliability design criteria would be used for the accessory drive as for the main drive.

More frequent internal inspection opportunities exist for the current tech- nology reduction gearboxes than forecasted for the advanced main drive reduc- tion gearboxes. Current technology reduction gearboxes are disassembled for repair and inspection after each premature, and TBO, removal. During the in- spection performed at the repairs and overhauls, distressed parts are re- placed. With an "on-condition" operation and the simplified, more reliable gearbox there will not be the frequent opportunity for replacement of parts.

The effect may be a slight increase in the premature removal rate (PRR) that was predicted solely from the improvement in the designs. An additional rate of 0.01 per 1000 hours was assumed for this effect.

The inherent reliability predictions for the STAT advanced main drive reduc- tion gearboxes are summarized in -able LIi.

TABLE LII. - INHERENT RELIABILITY PREDICTION SUMMARY FOR STAT ADVANCED MAIN DRIVE REDUCTION GEARBOX.

Inherent premature removal rate Gearbox system per 1000 h Main reduction 0.010 system Accessory drives for 0.010 two pumps and remote gearbox Effect of less 0.010 frequent disassembly inspections Total predicted for 0.030 main drive gearbox Equivalent MTBR 33,000 hours Noninherent Reliability Assessments

S

The principal noninherent removal & of the main drive reduction gearbox are for: • Improper maintenance • Unsubstantiated ( no failure found) • Accident damage Accident damage to the gearbox results from unusually large forces transferred to the gearbox from the propeller in case of an accident. The rate of acci- dent damage removals plus precautionary removals after an accident are a func- tion of the propeller accident rate. Projected improved propeller accident V rates supplied by Hamilton Standard were used to estimate noninherent removals of the gearbox for this cause.

I The noninherent removal rates for improper maintenance and unsubstantiated causes were estimated from studying the historical data base and the relative complexity of current technology and advanced gearboxes.

.i The estimated noninherent rate for the STAT advanced main drive gearboxes is one-third of the inherent rate, or 0.010 / 1000 hours.

Power Section Accessory Gearbox, Components, and Accessories Inherent Reliability Assessment The basis for the reliability assessments of power section accessories were detailed studies made for the XT701 engine ( most in conjunction with the sup- pliers), listing of the Allison Model 501 including the more recent Series III - / (Model 501 - D22 and the T56-A 14 and -15), and DDA supplier estimates for elec- tronic control systems.

Noninherent Reliability Assessment Estimates were made for component and accessory removal rates for noninherent causes using the same background data, studies, and supplier experience as discussed for the inherent reliability assessment.

Maintenance Cost Proiections Baseline Engine The determination of the baseline maintenance cost of current technology en- gines and reduction gearboxes was described previously in the Baseline Engines section, and the trend with horsepower is shown in Figure 45. This trend was established by Lockheed using cost data and scaling characteristic supplied by DDA, General Electric, and Garrett for current turboprop engines.

^f

.i0 b

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f l Rine burdened I Current and derivative — sdaduled overhouis, no condition nonfarm#

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Advanced-on co idiNan, condition monitoring 1979 dollars

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s0 Currant +admolapy

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Derivative tedmolooy A ~^ Advanced $TAT *cJ%W py

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0 •1

0 300 1000 13D0 2000 2500 3000 3500 Rated power—kW i 1 t I 0 1000 3000 4000 Ram pav'K—dV TEED-2103 Figure 45. - Engine maintenance cost versus rated power.

Derivative Engine DDA derivative engine technology was defined as that equivalent to a deriva- - 41 engine tive of the Model PD370 - 37 engine and designated as the Model PD 370 in the Navy MPA studies. Scaled characteristics of this derivative engine were derived for use in this STAT study.

The derivative engine differed from the baseline engine by having a booster stage added to the compressor and a new main drive reduction gearbox that was based upon the results of the study of turboprop reliability and maintenance costs. The added booster stage increased the pressure ratio from 12.5:1 to 17.7:1.

The following assumptions were made in establishing the maintenance costs in comparison to the current technology engine: o The main drive reduction gearbox would be designed to the "on-condition maintenace philosophy with no scheduled overhauls o A condition monitoring system would not be available on the derivative engine o Improvements in the core engine would increase mean time between overhaul ` (MTBO) by 8.5% and 4752 shp) derivative The maintenance costs of the 1749- and 3544-kW (2345 engines were estimated to be 8.4% lower than the current technology engines.

A comparison is shown in Table LIII.

TABLE LIII. - MAINTANANCE COST COMPARISON FOR BASELINE AND DERIVATIVE ENGINES 1749 (2345) 3544 (4752) Engine rating--kW (shp)/SLS Maintenance cost--$/EFH 40.73 64.91 Baseline engine 37.38 59.57 Derivative engine engines is primarily the 3 The reduction in costs attributed to the derivative The gearbox reduced the result of the improved main drive reduction gearbox.

by 0.8%. The in- cost by 9.2%, while the power section increased in cost net effect of increased creased cost for the power section resulted from the parts cost and decreased overhaul removal rate.

Advanced STAT Engines The maintenance cost projections for the advanced STAT engines were estimated by multiplying the line and shop labor and material charges per maintenance action by the corresponding rate of maintenance action, or repairs. Labor costs expressed in 1979 dollars were based on a direct labor rate of $10.00/h and a burden labor rate of $18.00/h. Material costs per repair were developed using estimated acquisition costs and historical data relating per-repair ma- terial costs to acquisition costs on a percentage basis.

The results of the maintenance cost projection is shown in Table LIV and Fig- ure 45, where they are also compared with the baseline (current technology) and derivative engines.

TABLE LIV. - MAINTENANCE COST PROJECTION OF ADVANCED STAT ENGINES COMPARED WITH BASELINE AND DERIVATIVE ENGINES 3544 (4752) Engine rating--kW (shp)/SLS 1749 (2345) Maintenance cost--$/EFH 40.73 64.91 Baseline engines 37.38 59.57 Derivative engines 24.64 Advanced engines 17.87 The improvement in maintenance cost of the derivative engines over the base- line engines is due primarily to incorporation of advanced reduction gear- boxes. The further improvement of the advance engines is due to the incorpo- ration of advance technologies which improve reliability and facilitate main- tenance actions. The cost projections of advanced engine maintenance over that of the baseline engine, are such that the 1749 kW (2345 shp) engine cost will be reduced by about 56% and the 3544 kW (4752 shp) engine cost will be reduced by about 62%. These reductions are the net effect of the design, re- liability, and maintainability features described previously.

f To show the advantages of using condition monitoring, an estimation of the maintenance costs without the use of condition monitoring was also made. Com- parison with and without condition monitoring are shown in Table LV.

TABLE LV. - MAINTENANCE COST PROJECTIONS OF ADVANCED STAT ENGINES WITH AND WITHOUT CONDITION MONITORING Engine rating--kW (shp)/SLS 1749 (2345) 3544 (4752) Maintenance cost--J/EFH With condition monitoring 17.87 24.64 Without condition monitoring 20.67 30.50 These costs were estimated by adjusting the rate of removals and the increased damage (material costs) that would occur before detection of a failure could 4 be made. Instead of the 56% for the 1749 kW (2345 shp) and 62% percent for the 3544 kW (4752 shp) maintenance cost sa vings, they would be reduced to about 49 and 53%, respectively. Thus, the net effect of condition monitoring is about 7 to 9% in cost savings compared to the baseline. Comparing a given engine with and without condition monitoring, this feature reduces maintanance costs by about 14% for the 1749 kW (2345 shp) to 19% for the 3544 kW (4752 shp) engines.

To show the effect of on-condition maintanance, an estimation of the mainte- nance costs of the 3544 kW (4752 shp) engine was made by assuming scheduled overhauls of 10,000 h on the cold section and 5000 h on the hot section. Com- parisons are shown in Table LVI.

TABLE LVI. - MAINTENANCE COST PROJECTION OF 3544 kW (4752 shp) ADVANCED STAT ENGINE WITH AND WITHOUT ON-CONDITION Scheduled overhauls-- On-condition-- 10,000-h cold section 7500-h TBO-HP turbine 5,000-h hot section Maintenance cost--$/EFH 24.64 With condition monitoring 39.82 Without condition monitoring 30.50 46.21 These estimations were made by adjusting the rate of unscheduled removals downward and inserting the rate for scheduled removals. Material and labor costs per action were ales adjusted to account for the effect of increased inspection during scheduled overhauls on the extent of damage at failure.

Scheduled maintenance (overhauls) increased :he cost about 62% with condition monitoring and 52% without it. Compared to the baseline engine, without con- dition monitoring, the cost reduction of the advanced engine with scheduled overhauls is about 29%.

i (l M+ , This discussion shows that maintenance cost improvements from the baseline i ` engine can vary from about 29 to 62%, depending upon whether the maintenance r concepts of on-condition and condition monitoring are used. The 29% improve- ment for the Advance STAT engines results from the incorporation of advance technologies which improve the reliability of the engines. Refer to Tables +1 XLV and XLVI for a breakdown of these maintenance costs.

k Weight Trends Specific weight trends estimat4d for the STAT engines, as compared to current and derivative technology trends, are shown in Figure 46. The current tech- nology line was based on results from Lockheed's studies for N"A Ames. De- rivative technology was constructed to parallel the current technology trend line and was keyed to the weight of the XT701 turboprop derivative engine dis- cussed earlier. The 1790 kW (2400 shp) STAT engine was 14% better in specific weight than the current technology reference. The 3579 kW (4800 shp) engine was 34% better. The 3579 kW (4800 shp) engine has greater improvement primar- ily because the reduction gear ratio is much lower as the result of driving a propfan rather than a conventional propeller.

Gear I ratio I 10.4* — XT701 turboprop Advanced 0 8 technology I 134% Derivative technology I Current technology I 0 `o C»ar ratio 19.1: 1 I l +14% I 13 Y S Gearbox lou included I I I I 0 6000 8000 Rated power—kW 0 6000 2000 4000 8000 10,000 Rated power—" TE80-2078A Figure 46. - STAT specific weight trends.

C

Engine Technology Comparisons Current (CTE), derivative (DTE), and advanced (ATE) STAT engine comparisons 1 terms of sfc, weight, price, maintenance cost, and maximum envelope length ai height at the rated condition are shown in Tables LVII and LVIII. (Note that the current technology engine 'CTE' is the same as CTE, except that the 'CTE' denotes "no inherited learning".) Engine prices, at STAT baseline economics, are shown on a common basis with respect to learning, i.e., without benefit c common production base. Maintenance costs are for a mature engine. Note thi the STAT engines achieve significant improvements in sfc and maintenance costs, which are the major drivers in DOC. Tables LVII and LVIII also show the price of the current technology engine with learning in a footnote to the table.

TABLE LVII. - STAT TURBOPROP ENGINE COMPARISONS (SI UNITS) 'CTE' DTE ATE Rated power - 1749 kW *sfc--Pg/W • s 86.19 76.46 71.69 Percent change 0 -11.3 -16.8 Weight--kg 364.7 358.3 318.9 Percent change 0 -1.7 -12.6 286,000 OEM price--$ 352,800** 371,200 Percent change 0 -18.9 +5.2 Maintenance coat--$/EFH 40.73 37.38 17.87 Percent change 0 -8.2 -56.1 Max envelope length--m 2.292 2.217 1.793 Percent change 0 -3.3 -21.8 Max envelope height--m 0.705 0.675 0.668 Percent change 0 -4.2 -5.3 Rated power - 3544 kW 70.78 *sfc--og/W's 79.74 64.68 Percent change 0 -11.2 -19.0 Weight--kg 533.9 524.8 398.7 Percent change 0 -1.7 -25.3 OEM price--$ 560,100*** 599,200 468,500 -16.4 Percent change 0 +7.0 24.64 Maintenance cost--$/EFH 64.91 59.57 -62.0 Percent change 0 -8.2 2.572 Max envelope length--m 2.658 2.215 Percent change 0 -3.3 -16.7 Max envelope height--m 0.847 0.811 0.735 0 -4.2 Percent change -13.3 *Gearbox loss included **CTE OEM price - $285,500 ***CTE OEM price - $430,800 TABLE LVIII. - STAT TURBOPROP ENGINE COMPARISONS (CUSTOMARY UNITS) i ATE 'CTS' DTE

Rated power • 2345 shp

0.4243 *sfc--lbm/hp-hr 0.5101 0.4525 -16.8 Percent change 0 -11.3 Weight--lbm 804 790 703 -1.7 -12.6 Percent change 0 286,000 OEM price--$ 352,800** 371,200 -18.9 Percent change 0 +5.2 EFH 40.73 37 . 38 17.87 Maintenance coat-- $/ -56.1 Percent change 0 -8.2 70.6 }^ Max envelope length--in. 90.23 87.29 0 -3.3 -21.8 Percent change . 58 26.28 Max envelope height- - in. 27.74 26 0 -4.2 -5.3 Percent change ,i Rated power 4752 shp 0.4189 0.3823 *sfc--lbm / hp-hr 0.4719 -11.2 -19.0 Percent change 0 1157 879 Weight--lbm 1177 -25.3 Percent change 0 -1.7 OEM price 560 , 100*** 599 , 200 468,500 -4 Percent change 0 *7.0 -16.4 64.91 59 . 57 24.64 Maintenance cost-- $/ EFH -62.0 Percent change 0 -8.2 Max envelope length- - in. 104.66 101.25 81.2 Percent change 0 -3.3 -16.7 Max envelope height- - in. 33.34 31 . 94 28.92 -4.2 Percent change 0 -13.3 *Gearbox loss included **CTE OEM price a $285,500 ***CTE OEM price a $430,800 Mission Results The engine technology comparisons will present results from the mission and aircraft cost analysis of the current, derivative. and advanced technology ed and presented in previous sections engines having characteristics as develo p of this report. It is noted that in this study, aircraft characteristics and technology level are fixed; i.e., each engine technology is evaluated without change in aircraft technology, even though it is probable that advancements in areas such as aerodynamics and structures could be incorporated with advanced engines to provide additional savings. This is also true with respect to the propeller characteristics; i.e., each engine technology incorporates current conventional propeller efficiencies, weight, acquisition, and maintenance cost values. Agaia, it is likely that advanced technology propellers would provide additional cost savings.

Engine performance data for the derivative and advanced engines Were simulated by vain& sfc characteristics of the DDA current technology engine (Model PD370-37). The simulation consists of adjusting the sfc level of the CTS by the ratio of the design point sfc values for the respective derivative and advanced angina to the design point sfc value of the CTE angina.

Blab Speed Aircraft The following section presents a general overview of the results obtained from aission evaluations 4f the CTE, DTE, and ATE in the high speed aircraft. This overview will be followed by a more detailed presentation of the technology comparisons obtained from the 185.2 ka (100 nm) alternate stage length mission results.

General mission results, i.e., aircraft design TOW, design fuel load, engine take-off rating, and total aircraft cost (TAC) are summarised in Tables LIZ and LX along with fuel consumption, DOC, and 5-year total cost of ownership 185.2 km (TCO) results for the 1111.2 km (600 am) design stage length and the (100 am) alternate stage length missions. Tables LXI and LXII lists the DOC and TCO results for each of the alternate stage lengths, i.e., 97.6-, 185.2-,.

277.8-, 370.4 and 740.8 km (50-, 100-, 150- 0 200-, and 400 um) missions.

TABLE LIX. - EISSION RESULTS--RIGH SPEED AIRCRAFT (SI UNITS) CTE 'CTE' DTE ATE Design aircraft TOM1--kg 18,299 18,299 17,895 17,193 Fuel--kg 1940 1940 1700 1500 Fuel fraction 0.106 0.10b 0.095 0.087 Engine TO rating at SLSS--kW 3532 3532 3467 3341 Aircraft power loading--kW/kg 0.386 0.386 0.388 0.388 Airframe acquisition--Billion $ 4.226 4.226 4.198 4.148 Propulsion system acquisition--million $ 1.609 2.042 2.144 1.666 Total aircraft cost--million $ 5.835 6.268 6.342 5.814 1111.2 ka stage length (design) Design cruise speed--MN 0.7 0.7 0.7 0.7 Cruise altitude--m 10,668 10,668 10,668 20,668 Block fuel--kg 1190 1190 1041 920 Fuel consumption--L x 10-3 aircraft/year 2:96 2496 2184 1931 DOC at Q1 --E/skm 2.355 2.406 2.286 1.932

DOC at Q -- 1 /skm 2.718 2.769 2.603 2.213

5-y TOO at--million = 14.249 14.745 14.245 12.320

5-y TCO at (Q --million 15.897 16.393

3 15.687 13.595

Wt

Ij

TABLE LIX. (CONT) CTE 'CTE' DTE ATE 185.2 ka stage length (altitude) wi 1.1 Cruise altitude--a 5791 5791 5791 Block fuel-kg 380 380 336 293 Fuel consumption--L x 10-3/aircraft/year 3155 3155 2807 2435 DOC at --E/skm 4.794 4.871 4.643 4.066 DOC at --!/slaw 5.490 5.567 5.258 4.603 5-y TCO at - - million = 17.902 18.398 17.857 15.717 5-y TCO at - million # 19.986 20.461 19.711 17.324

i{

(D fuel cost - $0.264/L Q fuel cost - $0.396/L TABLE LX. - MISSION RESULTS--HIGH SPEED AIRCRAFT (CUSTOMARY UNITS) CTE 'CTE' WE ATE - Design aircraft TOCW - -lbm 40,343 40,343 39,451 37,905 Fuel--lbm 4276 3747 3308 Fuel fraction 0.106 0.106 0.095 0.087 Engine TO rating at SLSS--shp 4736 4736 4649 Aircraft power loading--shp/lbm 0.235 0.235 0.236 0.236 Airframe acquisition- - million $ 4.226 4.226 4.198 4.14F Propulsion system acquisition - -million 1.609 2.042 2.144 1.666 Total aircraft cost- - million $ 5.835 6.268 6.342 5.814 600 -NN stage length (design) Design cruise spend-- MN 0.7 0.7 0.7 0.7 Cruise altitude- -e ft 35,000 35,000 35,000 35,000 Block fuel--lba 2623 2295 2029 Fuel consumptior.--gal x 10-3 /aircraft/year 659.3 659.3 576.9 510.1 DOC at ---s; s:m 4.361 4.455 4.233 3.578 DOC at Q -- ! /sna 5.033 5.128 4.821 4.098 5 -i TCO at - million $

T 14.249 14.745 14.245

12.320

5-y TCO at 0 -

million * 15.897 16.393 15.687 13.595 100 -M stage length (altitude) Cruise altitude--ft 19,000 19,000 19,000 19,000 Block fuel--lbw 838 838 741 Fuel consumption--gal x 10 -3 /aircraft/year 83 ?,•4 833.4 741.4 643.1 DOC at --E/sum 8.878 9.021 8.598 7.530 DOC at --!/anm i 10.168 10.311 9.738 8.524 5-y TCO at - -rillion = 17.902 18.398 17.857 15.717 5-y TCO at - - million

I 19.986 20.481

19.711 17.324 fuel cost - $1.00/gal fuel cost - =1.50/gal TABLE LXI. - ALTERNATE STAGE LENGTH MISSION RESULTS--HIGH SPEED AIRCRAFT (SI UNITS) 'CTS DTE ATE CTS 92.6 km stage length Cruise altitude- a 3048 3048 3048 3048 Block fuel--kg 258.1 258 . 1 225 . 199.1 DOC--J / skm-- (1) 7.326 7.431 7.118 6.321 DOC--E/skm ® 8.271 8.409 7.943 7.050 5-y TCO--million S-- 595 19.452 17.387 19 . 20.091 5-y TCO--million ^-- Q 21.603 22.158 21.259 18.984 185.2 km stage length Cruise altitude --m 5791 5791 5486 Block fuel- - kg 380.1 380 . 1 336 . 1 293.5 4.871 DOC--J/skm-- ® 4.794 4.643 4.066 DOC--J/skm-- (2) 5.490 5.567 5.258 4.603 5-y TCO--million 17.902 18.398 17.857 15.717 5-y TCO--million 19.986 20.481 19.711 17.324 277.8 km stage length Cruise altitude --m 7315 7010 7315 477.2 477.2 421.4 367.9 Block fuel--kg DOC-E/skm-- 3.873 3.940 3.750 3.254 DOC--1 -- 4 .455 4.523 4.264 3.704 5y TCO--million-- 16.860 17.356 16.828 14.729 5-y TCO--million $--(^ 18.860 19.356 18.606 16.272 370.4 km stage length Cruise altitude--m 8230 8230 7925 8230 Block fuel--kg 565.6 565.6 499.4 436.4 DOC-iE/skm-

® 3.390 3.452 3.282 2.832

DOC--J/ (b skm-- 3.908 3.970 3.739 3.231 5-y TCO--million $-- 16.188 16.166 14.098 16.683 17.878 5-y TCO--million $-- ^ 18.177 18.612 15.587 740.8 km stage length Cruise altitude--m 10363 10058 10058 10053 Block fuel--kg 877.2 884.5 772.9 683.1 DOC-j /skm-- (1) 2.624 2.678 2.545 2.165 DOC-i/skm- (2) 3.026 3.083 2.898 2.478 5-y TCO--million 14.773 15.316 14.788 12.839 5-y TCO- million $-- 16.489 17.052 16.306 14.180 ® fuel cost - $0.264/L fuel cost - $0.396/L i; W, r.1 HIGH SPEED AIRCRAFT TABLE LBII. - ALTERNATE STAGE LENGTH MISSION RESULTS- - (CUSTOMARY UNITS) ATE 'CTE' DTE CTE !j 50-NM stage length 10,000 10,000 10,000 10,000 Cruise altitude-ft 569 569 497 Block fuel--lbm ^t a DOC-J/snm-- 0 13.567 13.763 13.182 11.706 13.056 15.317 15.513 14.711 DOC-i/sna-W 19.452 17.387 19.595 20.091 5-y TCO--million $-- 21.259 18.984 5-y TCO- - million $-- 21.663 22.158 wn 100-NK stage length , 000 19 , 000 18 , 000 19,000 Cruise altitide - f t 19 838 838 741 Block fuel-lbm 9.021 8.598 7.530 DOC-i/snm-- (I) 8.878 10.311. 9.738 8.524 DOC--E/snm-- (2) 10.168 17.857 15.717 5-y TCO - -million $-- 17.902 18.398 20.481 19.711 17.324 5-y TCO- - million $-- 19.986 150-NM stage length 24,000 23 , 000 24,000 Cruise altitude-ft 24,000 1052 929 811 Block fuel- - lbm 1052 6.027 7.172 7.297 6.945 DOC -i/snm-- 6.859 8.251 8.376 7.897 DOC-E/snm-- 14.729 16.860 17.356 16.828 5-y TCO- - million $- 16.272 18.860 19.356 18.606 5-y TCO- - million $-- 200-NM stage length 27 , 000 27,000 26,000 27,000 Cruise altitude- - ft 1247 1247 1101 Block fuel--lbm 6.079 5.244 DOC-i/snm-- (I) 6.278 6.394 5.984 7.238 7.353 6.925 DOC-i/snm-- (2) 16.166 14.098 5-y TCO- - million $- 16.188 16.683 17.878 15.587 18.117 18 . 612 5-y TCO- - million $-- 400-Nth stage length , 33,000 33,000 33,000 Cruise altitude - ft 34 000 1934 1950 1704 1506 Block fuel - lbm 4.713 4.010 4.860 4.960 DOC-E/sam-- 4.589 5.604 5.710 5.368 DOC-E/sum-- 14.788 12.839 5-y TCO--million $-- 14.773 15.316 14.180 17 . 052 16.306 5-y TCO- - million $-- 16.489 fuel cost - $1.00/gal fuel cost - $1.50/gal Note that Tables LI% through LXII list DOC and TCO results for fuel costs of $0.264 and $0.396 / L ($1.00 and $1.50/gal).

i i The DOC results listed in Tables LIX through LXII are illustrated in Figures 47 and 48 as plots of DOC versus block distance (mission stage length) for the $0.264 and $0.396/L ($1.00 and $1.50/gal) fuel costs, respectively. Figure 49 shows the percent reductions in DOC obtained with the DTE and ATE, relative to the CTE, plotted against block distance. The DOC savings for the DTE are in- dicated in Figure 49 to be relatively insensitive to block distance with an approximate 4% reduction at the $0.396/1, ($1.50/gal) fuel cost. The DOC re- sults for the ATE indicate 15.5% reduction at the higher fuel cost increasing to 18.5% at greater block distances.

i I 1 Fuel cost • W6411.

i I ; `•\ 'No common production base j ! g \^^ \\^•`' ^^^ '^ ---:^ CTE ......... ? ^^ ------ DTE ATE 0 200 400 600 800 1000 Block distance—km 0 100 200 300 400 500 600 Block distance—nm TE80-Z0T9A - high speed aircraft, fuel cost Figure 47. — DOC versus block distance- $0.264/L ($1.00/gal).

3I I 1 Fuel cost • 10.3961101.501ge11 *No common production base ^ S T 45 8 Z5 CTE * ATE 800 1000 1200 0 200 400 600 Block distance—km 0 100 200 300 400 500 600 Block distance—nm TEM-2080 Figure 48. — DOC versus block distance--high speed aircraft, fuel cost $0.396/1, ($1.50/gal).

CTE baseline engine

--- ATE

^—r^—

c Fuel cost • $0.264/1.31.001gal l '9 1-7 Fuel cost • 10.396/1.111.50/ga11 B 10 T: ---I OTE .^ 0 • 0 20D 400 600 800 1000 Block distance—km 300 400 500 600 TE80-2081A Block distance—nm - high speed aircraft.

Figure 49. — DOC reductions- w P The 5-year TCO results listed in Tables LIX through LXII are shown in Figures 50 and 51 as plots of TCO versus block distance. Figure 52 shows the percent reductions in TCO obtained with the DTE and ATE, relative to the CTE, plotted against block distance. The percent reductions in TCO for both the DTE and ATE are indicated to be relatively insensitive to block distance with the DTE having an approximate 1% reduction and the ATE a 14% reduction in TCO at the $0.396/L ($1.50/gal) fuel cost. Figure 53 shows the 5-year TCO improvement obtained with the DTE and ATE in terms of dollar savings at each block dis- DTE tance. The indicates an approximate $250,000 savings at each block dis- tance with the $0.396/L ($1.50/gal) fuel cost. The ATE indicates an approxi- mate $2.3 to $2.7 million savings dependent upon block distance, with the $0.396/L ($1.50/gal) fuel.

The fuel burned or fuel consumption improvements obtained with the DTE and ATE are illustrated in Figure 54. This figure shows percent reduction in fuel consumption, relative to the CTE, plotted against block distance. The reduc- DTE and ATE are indicated to be essentially constant over tions for both the the range of block distances, with the DTE obtaining an approximate 11 to 12% reduction, and the ATE an approximate 22 to 23% reduction in fuel burned per mission. The fuel savings that could be achieved over a 10-year period of operation with a fleet of 100 aircraft is shoe-a in Figure 55. This figure plots DTE and ATE fuel savings, relative to the CTE, versus block distance.

Figure 55 indicates relatively higher savings to the shorter block distance with the DTE savings ranging approximately from 341 to 303 million liters (90 to 80 million gal). The fuel savings obtained with the ATE range approximate- ly from 719 to 568 million liters (190 to 150 million gal) of fuel.

The variation in block speed across the range of block distances examined in this study is shown in Figure 56. This figure shows a change in block speed from 315 km/h (170 kt) for the 92.6 km (50 nm) distance to 648 km/h (350 kt) for the 1111.2 km (600 nm) design stage length with the block speed for the 185.2 km (100 nm) distance indicated to be approximately 426 km/h (230 kt).

A summary of the percentage changes in critical aircraft and cost parameters resulting from engine technology improvements is shown in Table LXIII. These percentage changes are relative to the CTE powered aircraft values. The cost parameters were calculated for the 185.2 km (100 nm) stage length mission 'CTE' (185.2 km (10C nm) block distance) only. (Note that is the same engine as CTE except for price, which is higher. 'CTE' reflects in all new engine with no inherited learning.)

Fuel cost • $k2WL *I.00igN1 i ^ 18 V N `` -'^^----^— CTE•

I

_'' CTE DTE o 14 *No common production base ATE 400 600 800 1000 1200 — Block distance—km I __1 I 1 1 1 1 0 100 200 300 400 500 6M TE80-20R Block distance-nm Figure 50. - 5-year TCO versus block distance- - high speed aircraft, fuel cost $0.264/L ($1.00/gal).

TABLE LXIII. - TECHNOLOGY COMPARISON- - HIGH SPEED AIRCRAFT 185.2 km (100 nm) block distance change from CTE Percent CTE 'CTE DTE ATE TOW Base 0 -2.2 -6.0 TAC +7.4 +8.7 Base -0.4 Fuel consumption Base 0 -11.0 -22.8 DOC at Base +1.6 -3.2 -15.2 ® DOC at Base +1.4 -4.2 -16.2 ® 5-y TCO at Base +2.8 -0.2 -12.2 G 5-y TCO at Base +2.5 -1.4 -13.3 ® Q fuel cost - $0.264 / L ($1.00/gal) ® fuel cost - $0.396 / L ($1.50/gal) Utilization 2800 h/y • Fuel cost - f0.39611141.501ge11 b.

s b CTE• X 14 ATE •No common production base 0 200 800 1000 1200 400 600 Black distance—km 0 100 200 300 400 600 Block distance—nm TEW2083 "W.

Figure 51. — 5 —year TCO versus block distance--high speed aircraft, fuel cost $0.396/L ($1.50/gal).

• CTE baseline engine ATE r fS -- Fuel cost • WWI.01.001ga11 --- Fuel cost • $0.3WL lt1.501ga11 u° 5 DTE 0 200 400 600 800 1000 1200 Block distance —km 100 200 300 400 500 600 Block distarte—nm TE80-2116M Figure 52. — 5 —year TCO reductions--high speed aircraft.

i..l .

The combination of the sfc, weight, and envelope dimensions of the ATE is shown to have reduced aircraft TOGW by 6%. However, the total aircraft cost is shown to be essentially the same as the CTE-powered aircraft. This is a result of the fact that the lower airframe less power section cost was offset by the higher acquisition cost of the ATE compared to the CTE. It is noted" that the significantly higher cost of the DTE with respect to the ATE con tributed significantly to a 9% increase in total aircraft cost over the CTE.

ij The fuel consumption percentage changes basically reflect the sfc improvements ;± associated with the DTE and ATE plus the reduced rated engine power associated with the CTE and ATE aircraft, which in turn resulted from the gross weight reduction. The 12% difference between the DOC and TCO reductions shown for the DTE and those shown for the ATE is a result of the combination of the ATE's larger sfc, engine acquisition, and maintenance cost improvement rela- t tive to the DTE. It is noted that increasing the fuel cost from $0.264 to $0.396/1 ($1.00 to $1.50/gal) produces a 1% larger reduction in DOC and TCO for both the derivative and advanced engines compared to the current engine.

A comparison of the 10-year fuel requirements of an assumed fleet of 100 air- craft flying the 185.2 km (100 nm) stage length exclusively is shown in the bar graph presented in Figure 57. This figure indicates a savings of 348 mil- lion liters (92 million gal) for the DTE, and 719 million liters (190 million gal) of fuel for the ATE.

A comparison of the 10-year DOC for the fleet and mission prescribed in the preceding fuel comparison is shown in Figure 58 for both $0.264 and $0.396/ liter ($1.00 and $1.50/gal) fuel costs. This bar graph indicated a savings in the order of $116 million for the DTE, and $529 million for the ATE at the $0.396/liter ($1.50/gal) fuel cost.

A breakdown of the cost element in DOC for each of the technology engines is shown in Figure 59. Figure 59 indicates that the DOC reduction for the DTE was essentially a result of the sfc improvement. Note that the sfc improve- ment was offset to a small extent by the higher depreciation and insurance costs associated with the higher engine acquisition cost. The DOC reduction noted for the ATE was essentially a result of a combination of significant sfc reduction, engine acquisition, and maintenance cost improvements. This figure also illustrates that significant impruvement in efficiency is required to maintain current DOC levels in the face of rising fuel costs.

C

• CTE baseline engine 3.0

i .j

} ATE N V 2.0 Fuel cost •WWII. 01.001981) ^ Fuel cost • WWI. 41-Mal) 1.0

E

I

N V 1000 1200 600 800 200 400 D Block dlstance—km 400 500 600 0 100 200 300 TE80--2085A Block distance—nm Figure 53. - 5-year TCO savings- - high speed aircraft.

Low Speed Aircraft The following section presents a general overview of the resins ubtained from mission evaluations of the CTE, DTE, and ATE in the DDA low speed aircraft.

This overview will be followed by a more detailed presentation of the technol- ogy comparisons obtained from the 185 . 2 km (100 nai) alternate stage length mission results.

General mission results, i.e., aircraft design TOW, design fuel load, engine take - off rating, and TAC, are summarized in Tables LXIV and LXV along with fuel consumption, DOC, and 5-year TCO results for the 1111. 2 km (600 nm) de- sign stage length and the 185.2 km ( 100 nm) alternate stage length missions.

Tables LXVI and LXVII list the DOC and TCO results for each of the alternate stage lengths.

Note that Table LXIV through LXVII list DOC and TCO results for fuel costs of $0.264 and J O.396/L ($1.00 and $1.50/gal).

e CTE bowline .nline ATE OTE 1000 1200 0 200 400 600 800 Black dfstma-W 0 100 200 300 400 500 Black dishna-nm ^^ Figure 54. - Percent reduction in fuel consumed--high speed aircraft.

TABLE L%IV. - MISSION RESULTS- - LOST SPEED AIRCRAFT (SI Ub1ITS) CTE 'CTE' DTE ATE

Design aircraft

TO(W--kg 18,928 18,928 18,468 18,083 Fuel- - kg 2080 2080 1822 1684 Fuel fraction 0.110 0 . 110 0.099 0.093 Engine TO rating at SL3S--kv 2139 2139 2095 2057 Aircraft power loading- - kW/kg 0.083 0.083 0.084 0.084 Airframe acquisition- - million $ 3.986 3.986 3 . 954 3.926 Propulsion system acquisition- - million $ 1.116 1.421 1.468 1.138 Total aircarft cost--aillion $ 5.102 5.407 5.064 5.422 1111.2-ka stage length ( design) Design cruise speed--IN 0.47 0.47 0.47 0.47 Cruise altitude- - m 6096 6096 6096 Block fuel- - kg 1466 1466 1279 1180 Fuel consumption- - L x 10-3/aircraft 2298 / year 2298 2005 1851 DOC at - - E/slm 2.949 2.997 2.846 2.544 DOC at ^`--i/skm 3.396 3.444 3.237 2.904 5-y TCO at --million $ 13.118 13.467 12.966 11.722 5 - y TCO at --million

8 $ 14.636 14.985 14.290 12.944

• Utillution d 2100 hly • 100 aircraft flo • CTE basell ne engine 0 200 400 wi 800 1000 1200 Block dfsbnce—km L_ i 1 1 _ 1 1 1 1 TE80-2087A Block distance—nm Figure 55. - 10-year fuel savings--high speed aircraft.

TABLE LXIY. (CONT) CTE 'CTE' DTE ATE 185.2-km stage length ;aiC:L•:^e) Cruise d'!itude--tn 3048 3048 3048 3048

Block fuel--kg 369 369 321 296

Fuel consumption--L x 10-3/aircraft/year 2523 2523 2199 2028 4.030 4.096 3.884 3.522 DOC at 0 --E/skm

DOC at Q - Vskm 4.705 4.771 4.472 4.064

5-y TOC at --million $ 13.063 13.412 12.900 11.789 5-y TCO at --million $ 14.730 15.078 14•.352 13.128

fuel cost - 1$0.2644

fuel cost - $0.396/L I N1

OL

m

0 1 0` 400 600 800 200 1000 1200 Block distance—kln 0 100 200 300 400 500 600 TE80-200 Block distance—nm Figure 56. - Block speed versus block distance--high speed aircraft.

e 15-km 1100-nml block dlsbnce • Utllitation of 2l00 hly • 10-y period of operation • 100 aircraft flog CTE UMAMAW ^wwww^wwwwrww^ OTE ATE 1500 2000 2500 300D 9500 'Millions of liters X 400 J 500 do0 700 so 900 Millions of 011ons TEW2099 Figure 57. - Fuel consumption comparison--high speed aircraft.

• 185-km 1100-nm1 block distance

• Utilization of 2800 hly

• 10-y period of operation

• 100 aircraft fleet

:0.2641101.0010 l

CTE :0.34611 01.50gal l '^, ;^, 0TE ATE

- million

^

2.0 3.0

2.5 3.5 4.0

Billions of dollars

TEW2M

high speed aircraft.

Figure 58. - DOC comparison- - - LOW SPEED AIRCRAFT ( CUSTOMARY UNITS) TABLE LXV. - MISSION RESULTS- CTE 'CTE' DTE ATE Design aircraft TOGW- - 41 ,730 41,730 40, 715 39,866 lbm 4586 4586 4016 3713 Fuel--lbm Fuel fraction 0.110 0 . 110 0 . 099 0.093 Engine TO rating at SLSS- - shp 2868 2568 ZBiu 2758 Aircraft power loading- shp/lbm 0.137 0 . 137 0 . 138 0.138 - Airframe acquisition million $ 3.986 3 . 986 3 . 954 3.926 1.116 1.421 1.468 1.138 Propulsion system acquisition- - million t Total aircraft cost- - million 5.102 5.407 5.422 5.064 600-NM stage length ( design) - .47 0.47 0.47 Design cruise speed- MN 0.47 0 Cruise altitude- - ft 20 , 000 20 , 000 20,000 20,000 Block fuel-ljo 3231 3231 2819 2602 607.0 607 . 0 7 488.9 Fuel consumption- - gal x 10 -3 /aircraft/year 529 .

5.461 5.550 5.271 4.711 DOC at (I) --i/sum 5.378 DOC at Q --E/snm 6.289 6.378 5.994 5-y TCC at CD --million $ 13.118 13.467 12.966 11.722 5-y TCO at Q --million $ 14.636 14.985 14.290 12.944 t b i Of OTC Aif $ Nouns) 7"D•atl Figure 59. - DOC breakdown--high speed aircraft.

TABLE LXV. (CONT) CTE 'CTE' DTE ATE 100-NM stage length (altitude) Cruise altitude--ft 10,000 10,000 10,000 10,000 Block fuel--lbe 813 813 708 Fuel consumption--gal x 10-3 /aircraft/year 666.6 666.6 581.0 535.8 DOC at Q --i/snm 7.463 7.585 7.193 6.523 DOC at 0 -- !anm 8.713 8.835 8.282 7.527 5-y TCO at 0 --ailli.on 13.063 13.412 12.900 11.789 5-y TCO at (Z --million S 14.730 15.078 14.352 13.118 (j) fuel cost • $1.001gal (2) fuel cost $1.50/gol c ; LOW SPEED AIRCRAFT TABLE LXVI. - ALTERNATE STAGE LENGTH MISSION RESULTS- - (SI UNITS) t ATE CTE 'GTE' DTE 92.6-lam stage length `rl 3048 3048 3048 Cruise ^ ltitude--m 3048 219 219 191 Block fuel- - kg 4.614 5.261 5.352 5.092 DOC-i/skm-- 790 5.258 DOC-i / 6 .063 6.153 5.

ska--(`, 12.629 12.978 12.520 11.438 5-y TCO--million s-- 14.078 14.427 13.782 12.602 5-y TCO--million $-- ® 185.2 - km stage lengtl•.

3048 3048 3048 3048 Cruise altitude- • - ,v,.

369 369 321 Block fuel- - kg 3.522 4.030 4.096 3.884 DOC-i/skm-- (3) 4.705 4.771 4.472 4.064 DOC-i/skm-- ® 13.412 12.900 11.789 5-y TCO--million $--Q 13.063 14.352 13.128 14.730 15.078 5-y TCO--million $-- Q 277.8-km stage length 3962 3962 3658 3962 Cruise altitude- - m 437 398 495 495 Block. fuel--kg 3.606 3.666 3.476 3.152 DOC-i/skm-- Q 4.270 4.010 3.638 DOC-#/skm-- 4.211 ® 13.415 12.961 11.791 5-y TCO--million-- ® 13.066 14.444 13.132 14.734 15.082 5-y TCO--million $-- ® 370.4- km stage length 4572 4572 4572 4572 Cruise altitude --m 611 611 533 Block fuel-kg 3.363 3.007 DOC-E / 3 .494 3.549 skm-- m 3.850 3.457 4.053 4.109 DOC-i/skm-- ® 13.718 13.181 11.918 5-y TCO--million-- 13.369 15.360 14.613 13.239 5-y TCO--million 15.011 Q 740.8-km stage length 6096 6096 Cruise altitude- - m 6096 6096 1029 1029 897 828 Block fuel--kg 2.667 3 .093 3.143 2.984 DOC-i/ skm-- ® 3.564 3.614 3.395 3.046 DOC-E/skm-- (:) 12.981 11.734 13.137 13.486 5-y TCO--million 14.664 15.013 14.312 12.963 5-y TCO--million $--(D ® fuel cost - $0.264/L (Z) fuel cost - $0.396/L .. z TABLE LVII. - ALTERNATE STAGE LENGTH MYSSION RESULTS--LOW SPEED AIRCRAFT (CUSTOMARY UNITS) 'CTE' DTE ATE CTS 50-NM stage length 10,000 10,000 10,000 Cruise altitide--ft 10,000 387 t ► 482 482 420 Block fuel-lbs 9.744 9.911 9.430 8.545 DOC-E/snm-- (D 9.737 11.395 10.723 DOC--E/anm-- ® 11.228 12.978 12.520 11.438 5-y TCO--million 12.629 Q 13.782 12.602 14.078 14.427 5-y TCO--million ^-- ^?

100-NM stage length 10,000 10,000 10,000 Cruise altitude--ft 10,000 708 653 Block fuel--lbm 813 813 6.523 7.463 7.585 7.193 DOC-J/snm--(D 7.527 8.713 8.835 8.282 DOC--!/snm-- 2) 11.789 13.063 13.412 12.900 5-y TCO--uillion-- 14.352 13.128 5-y TCO--million 14.730 15.078 150-NN stage length 12,000 13,000 13,000 13,000 Cruise altitude--ft 964 877 1092 1092 Block fuel--lba 6.437 5.838 6.679 6.789 DOC--i/snm-- 7.426 6.738 7.799 7.908 DOC--i/snm-- 13.066 13.415 12.961 11.791 5-y TCO--million 3-- 14.444 13.132 5-y TCO--million S-- 14.734 15.082 200-NN stage length 15,000 15,000 15,000 Cruise altitude--ft 15,000 1174 1083 1347 1347 Block fuel--lbm 6.228 5.569 6.470 6.573 DOC--i/snm-- 7.609 7.131 6.402

DOC--E/snm-- 1 7.506

13.369 13.718 13.181 11,918 5-y TCO--million t-- 15.360 14.613 13.239 5-y TCO--million S-- 15.011 400-NM stage length 20,000 20,000 20,000 20,000 Cruise altitude--ft 1978 1825 2269 2269 Block fuel--lbm 4.939 5.728 5.821 5.526 DOC--E/scam-- ® 6.287 5.641 6.600 6.693 DOC--E/snm-- (2) 13.486 12.981 11.734 5-y TCO--million s-- 13.137 12.963 5-y TCO--million 14.664 15.013 14.312 fuel cost - =1.00/gal fuel coat a 31.50/gal t ^_ s

i

? The DOC results listed in Tables LXIV through LXVII are illustrated in Figures

! 60 and 61 as plots of DOC versus block distance ( mission stage length) for the $0.264 and $0.396/L ($1.00 and $1 50/gal) fuel costs, respectively. Figure 62 .

shows the percent reductions in DOC obtained with the DTE and ATE, relative to the CTE, plotted against block distance. The DOC savings for the DTE are in- dicated. in Figure 62 to be relatively insensitive to block distance with an / gal) fuel cost. The DOC re- approximate 5% reduction at the $0.396/L'($1.50 sults for the ATE indicate slightly lower percentage reductions for the shorter block the ATE indicate slightly lower percentage reductions for the shorter block distances. A range of 13 to 15% reduction at the higher fuel cost is shown.

z The 5-year TCO results listed in Tables LXIV through LXVII are shown in Fig- ures 63 and 64 as plots of TCO versus block distance. Figure 65 shows the { percent reductions in TCO obtained with the DTE and ATE, relative to the CTE, plotted against block distance. The percent reductions in TCO foi both the .

DTE and ATE are indicated to be relatively insensitive to block distance with the DTE having an approximate 2.5% reduction and the ATE a 11% reduction in TCO at the $0 . 396/L ($1.50%gal) fuel cost. Figure 66 shows the 5-year TCO improvement obtained with the DTE and ATE in terms of dollar savings at each block distance. The DTE indicates an approximate $ 200,000 savings at each block distance with the $ 0.396 / L ($1.50 / gal) fuel cost. The ATE indicates an ` approximate $ 1.5 to 1 . 7 million savings, dependent upon block distance, with the $0.396 /L ($1.50/ gal) fuel.

Fuel cost • $0.26411(81.001gal )

c .Q 5

I

T 8 Y TV CTE uie ATE 4 *No common production base . __ - t 1 1 — F_ 200 400 600 800 0 1000 1200 Block distance-km 0 100 200 400 500 600 Block distance-nm TEW2092 Figure 60. - DOC versus block distance--low speed aircraft., fuel cost $0.264/L ($1.00/gal).

E Full cast • (0.3961101.sOl 0 ^ 8 AM *No common production base 0 2pp - 400 600 " M DOC-tlslet iun . .

1 , 1 .

400 500 600 0 100 200 TE80-21M DOC-0sest nm Figure 61. - DOC versus block distance--low speed aircraft, fuel cost $0.396/L ($1.50/gal).

The fuel burned, or fuel consumption improvements, obtained with the DTE and ATE are illustrated in Figure 67. This figure shows percent reduction in fuel consumption, relative to the CTE, plotted against block distance. The reduc- tions for both the DTE and ATE are indicated to be essentially constant over the range of block distances, with the DTE obtaining an approximate 13% reduc- tion, and the ATE an approximate 19.5% reduction in fuel burned per mission.

The fuel savings that could be achieved over a 10-year period of operation with a fleet of 100 aircraft is shown in Figure 68. This figure plots DTE and ATE fuel savings, relative to the CTE, versus block distance. Figure 68 indi- cates relatively higher savings for the shorter block distances, with the DTE ( 85 to 75 million gal). The savings ranging from 322 tc 284 million liters fuel savings obtained with the ATE range from 492 to 454 million liters (130 to 120 million gal) of fuel.

The variation in block speed across the range of block distance examined in this study is shown in Figure 69. This figure shows a change in block speed h (140 kt) for the 92.6 km ( 50 nm) distance to 482 Im / h (260 kt) from 259 km / 2 km (600 nm) design stage length, with the block speed for the for the 1111 .

100 nm) distance indicated to be approximately 352 km / h (190 kt).

185.2 km ( j

^q

^I -- Fuel cost • $0.2611L ti1.OWpl1 .

--- Fuel cast 0%3WL 1:1.501ge11 CTE - Baseline engine is ATE e OTE 200 400 600 800 1200 Block distance-km 1 1 1 1 1 1 1 1 1 1 400 500 600 I. 0 100 200 300 Black distance—nm TE80-209411 Figure 62. - DOC reductions--low speed aircraft.

A summary of the percentage changes in critical aircraft and cost parameters resulting from engine technology improvements is shown in Table LXVIII. These percentage changes are relative to the CTE-powered aircraft values. The cost parameters were calculated for the 185.2 km (100 nm) stage length mission (185.2 km (100 nm) block distance) only.

The combination of the sfc, weight, and envelope dimensions of the ATE is shown to have reduced aircraft TOW by 4.5%. However, the total aircraft cost is shown to be essentially the same as the CTE-powered aircraft. This is be- cause the lower airframe less power section cost was offset by the higher acquisition cost of the ATE compared to the CTE. It is noted that the signif- icantly higher (with respect to the ATE) acquisition cost of the DTE con- tributed significantly to a 6% increase in total aircraft cost over the CTE.

The fuel consumption percentage changes basically reflect the sfc improvements associated with the ME and ATE plus the rated engine power reduction asso- ciated with the gross weight reduction. The approximate 9% difference between the DOC and TCO reductions shown for the DTE and those shown for the ATE is a result of the combination of larger sfc, engine acquisition, and maintenance cost improvement associated with the ATE relative to the DTE. It is noted that increasing the fuel cost from $0.264 to $0.396/L ($1.00 to $1.50/gal) ;p roduces a 1% larger reduction in DOC and TCO for both the derivative and ad- vanced engines compared to the current engine.

Fuel cost •:0.26411.($1.00lgalf CTE

^rrr

CTE E N x DTE A bATE NC O . No common production base 0 200 600 800 1000 1200 Block distance—km 0 100 200 300 400 500 600 Block distance—nm TESO_2 Figure 63. - 5-year TCO versus block distance--low speed aircraft, fuel cost $0.264/L ($1.00/gal).

TABLE LBVIII. - TECHNOLOGY COMPARISON--LOW SPEED AIRCRAFT 185.2 km (109 nm) block distance Percent change from CTE CTE 'GTE' DTE ATE TOGW Base 0 -2.4 -4.5 TAC Base +6.0 +6.3 -0.7 Fuel consumption Base 0 -12.8 -19.6 DOC at Base +1.6 -3.6 -12.6 DOC at Base -?.4 -4.9 -13.6 5-y TCO at Base +2.7 -1.2 -9.8 5-y TCO at Base +2.4 -2.6 -10.9 Q fuel cost - $0.264/L ($1.00/gal) ® fuel cost - $0.396/L ($1.50/gal) Utilization - 2800 h/y i Fuel cost • =0.39611 t61.50NMI1 C I EO N ^MNN^^NM^M^^ CTE DTE

N

ATE

V

N

E

v

t- • No common production base 0 400 600 800 1000 Black distance-km 200 300 400 500 600 0 100 TEW20% Block distance-nm Figure 64. - 5-year TCO versus block distance--low speed aircraft, fuel cost / $0.396 L ($1.50/gal).

A comparison of the 10-year fuel requirements of an assumed fleet of 100 air- craft flying the 185.2 tom ( 100 nm) stage length exclusively is shown in the bar graph presented in Figure 70. This figure indicates a savings of 326 mil- lion liters ( 86 million gal) for the DTE and 496 million liters ( 131 million gal) of fuel for the ATE.

A comparison of the 10-year DOC for the fleet and mission prescribed in the foregoing fuel comparison is shown in Figure 71 for both $ 0.264 and $0.396/L ($1.00 and $ 1.50/gal) fuel costs. This bar graph indicates a savings on the $ order of 115 million for the DTE, and 316 million for the ATE at the $ $0.396 / L ($1.50/gal) fuel cost.

A breakd oirn of the cost element in DOC for each of the technology engines is shown in Figure 72.

Fuel cast • $0.26111 ($1.0gel) Full cost • $0.99611(61.50 gill CTE - BesNi M engine 1S O ATE ^` S OTE } 0 200 1000 1200 400 600 800 Black distance-km 0 100 300 600 200 400 500 TE80-2091A Black distance-nm Figure 65. - 5-year TCO reductions- - low speed aircraft.

This figure indicates that the DOC reduction for the DTE was essentially a result of the sf c improvement. Note that the sf c improvement was offset to a small extent by the higher depreciation and insurance costs associated with the higher engine acquisition cost. The DOC reduction noted for the ATE was essentially a result of a combination of significant sfc reduction, engine acquisition, and maintenance cost improvements.

Full cost - ti0.26411 SLODgall ---- Fuel cost - !0.39611. $1.501ge11 CTE - Usellne engine V ^O } ATE .L--..—L 800 1000 Black distance—km

I I I I I I I ^I

i I l I l

100 200 300 400 500 600 Black distance—nm TEED-2096A Figure 66. — 5 — year savings- - low speed aircraft.

a CTE Baseline engine 29 F

I

z ATE g w CC O 15 $ DTE to Boo 1200 0 - 200 400 600 1000 Block distance—km 6 a I -..L- A . I I 400 600 200 300 500 0 100 Block distance—nm TE80-2099A Figure 67. — Percent reduction in fuel consumed- - low speed aircraft.

^F • Utilization of 2800 hly • 100 aircraft fleet e CTE baseline engine

N

'a 120 `a ATE a^ i

E

E I Lip OTE I ii 0 200 600 800 1000 1200 Block distance—km ^I 400 500 0 100 200 300 Block distance—nm TE80-Y100A Figure 68. - 10-year fuel savings--low speed aircraft.

t m N 3W m m .

0L a 0 200 400 600 a00 1000 1200 Block diSbna—km 300 400 500 610 0 100 2m TEW2101 nm 91m disbnu — - low speed aircraft.

Figure 69. - Block speL. d versus block distance- .I C1E I e 85-km (100-nmI block disance 9TE e Ut ilkatl on of 28M Aly I e 10-y period of operation e 100 aircraft fleet ATE I 2000 2500 3400 Millions of IRers 400 600 700 500 800 900 1E^-2102 Millions d gallons Figure 70. - Fuel consumption comparison —low speed aircraft.

• 18S-km (100-m) block distance • utilization of 2000 kh • 10-V period of operation • 100 aircraft fled $0.264/L ($1.00%901) C CT $0.396/1 (51.30 901) TIE Q A TIE 316M n t . e n ?_S 3.1 Nllion of dollars TEW2103 Figure 71• - DOC comparison- - low speed aircraft.

15-km QWnmI block ON= 1.2 w.1 rw1 3 f^) 1 0.6 ID.2Wl(E1.l0 p1,oaMn i i _ o.e AwnMMns.

0.4 O"M"" OW IOWA"" 0.2 Csw Mrs ATE 6001100) TEW2104 Figure 72. - DOC breakdown- - low speed aircraft.

RECOMMENDATIONS FOR FUTURE RESEARCH NASA should address basic research and development needs for the near term to broaden the industry data base for design of advanced gas turbine engines for small transport aircraft. A general program directed toward specific areas where an investment in R&D dollars could produce critical data needed in de- sign of new turboprop engines for commuter application will be presented.

These results are based on a generalized preliminary design study of commmuter aircraft turboprops at 1790 kw (2400 shp) and 3579 kw (4800 shp) sizes for two different 50-passenger aircraft differing principally in initial cruise alti- tude and design Mach number.

As advanced transport needs are resolved through STAT vehicle studies and user requirements, a useful purpose could be fulfilled by implementing an experi- mental engine program in a size consistent with 1990 projections of vehicle requirements. Although most new technology voids can be filled by component and material programs, the application of certain high risk design innovations must be assessed by experiment to develop sufficient confidence to proceed with full development.

PROGRAM CONTENT The overall program content is shown in Figure 73. This program includes the STAT propulsion study, basic R&D effort, and component R&D in the area of com- pressors, turbines and shafting, advanced structures, combustors, and controls and systems, leading to an experimental engine program that includes a design study, further component tests, test of a gas generator core, and finally the experimental core with the grower turbine and propeller gearbox added.

The experimental engine program is a four-year program led by a design effort in which results of the component R&D efforts are integrated into the experi- mental engine design. Aoout 200 hours of developmental engine testing are included in the program.

R&D on high risk The experimental engine provides a means to continue basic components in an engine environment, and also provides a vehicle to gather data on high risk mechanical systems such as a full-time turbine active clear- ance control. This is an exsmple of a system that would not be released for development on an engine program without some experimental engine experience.

Certain undeveluped technology elements have been identified during this study program as essential in the full-scale development of the STAT engines. Re- search programs leading to the fulfillment of these element requirements are described herein, and are listed in Figure 74.

BASIC RESEARCH AND DEVELOPMENT STAT studies show that the largest improvements in commuter aircraft DOC, as i:.fluenced by the engine, are achieved through reduced engine fuel consumption and reduced maintenance cost.

I CY I so 1 ei a STAT Study

p

sale Rao Twbinas Diffusers and canbmtas Sttuetures Etpine system b 4mrimantei bons Dnion fabrication Component t vt &pine tact TESO-2113 Figure 73. - STAT advanced technology program.

Fuel consumption can be reduced through improved thermodynamic cycle effi- ciency. Compressor pressure ratios around 20:1 and turbine RIT's near 1506 K (2250°F) appear to offer the minimum DOC. Advanced engine component physical size is greatly reduced compared to current technology, which brings about a new set of challenges to realize the potential efficiency gains, and opportun- ities to reduce engine cost through application of new technology. As engine pressure ratio and RIT increases, more emphasis must be placed on development of technology applcable to small high pressure ratio compressors of various types including axial and axial-centrifugal, both single and dual spool. New turbines, featuring advanced construction methods and new materials, are need- ed to improve cooling techniques and permit efficient blade design in the rel- atively small flow passages available. Bearing and shafting technology must be advanced to permit increased rotative spuds with improved dynamics and increased bearing life. Engine reliability and maintainability must be im- proved to achieve low maintenance costs and on-time performance. Basic re- needed to achieve these gains are search and development programs which are presented on Table LUX with the benefit/cost ratio, rank, and probability of success shown for each program. Benefits are based on the reduction in total DOC assignable to each technology as applied to a fleet of 100 aircraft power- ed by current technology engines, operating at 2800 hr per year for a 10-year period on a typical route segment of 100 nautical miles and using fuel priced at $0.264/L ($1.00/gal).

.T Technology Items Schedule Benefits F%1 $ ix 83 84 85 Compressors Hybrid centrifugal compressor impeller Reduce cost 4 Compressor errosion protection I ncrease durability I Net particle separators Increase durability Compressor noise reduction Reduce noise Axial compressor aft stage study i Improve performance Centrifugal compressor adv. high hubltip study 1 I mprove performance Turbines Hybrid turbine wheels Increase durability Cast-in impingement cooling Improve performance Abradable coatings Improve performance Diffusers-combustors Vortex-controlled diffuser I mprove performance Transpiration cooled combustor-sheet fabrication Reduce cost Transpiration cooled combustor - fabrication Improve durability Combustion noise reduction Reduce noise Structure and shafting I mprove design Rotor/case response to rotating stall Composite gear cases Reduce weight and cost I mprove design Bearing fatigue life I ncrease durability Composite shafting I mprove design Supercrit ical shafting e Engine systems Improve reliability Electronic control I mprove reliability Fuel pump and metering system Improve maintenance Engine condition monitoring system *Rank on a basis of 1 to 411 - most urgent or beneficial!

TE80-2116A Figure 74. — STAT technology research programs.

Compressors Compressors chosen for STAT engines feature high pressure ratio, single—spool configurations with reduced numbers of stages and reduced blade count. Split — - spool configurations appear to offer similar gains in performance with gen erally increasing complexity and somewhat lower risk.

centrifugal compressors tend to result in shorter engines when used with Axial — foldback combustors. This, in turn, reduces shaft length and eases dynamic problems at the expense of greater diameter at the engine midsection and pro - bably lower overall compressor efficiency. Axial compressors tend toward higher efficiency; however, small blade sizes in the latter stages can result in performance penalties that may offset fundamental gains. An expanded com - pressor data base is required to make the proper choice in advanced engine design.

I ^r TABLE LXIX. - CRITICAL TECHNOLOGY ELEMENTS FOR STAT RESEARCH AND DEVELOPMENT 5 ` Cost benefit benefit/ Probability $000's $000,000's Cost Ratio Rank* of Success Engine Technology Proaram Power Section • Compressor $60 5.851 10.45 12 Likely Hybrid Cost. Compr. Impeller 620 16.046 25.88 5 Likely Axial Compr. Aft Stage Study 680 7.26' 10.60 11 Likely Cent. Compr. Adv. Hi-Hub/Tip Study 620 10.503 16.94 8 Likely Compr. Erosion Protection 500 12.253 24.51 6 Likely Inlet Particle Separators 300 15.170 50.57 2 Rotor/Case Response to Rotat, Stall • Turbine 560 1.459 2.61 19 Likely Hybrid Rotors 400 2.042 5.11 17 Likely Composite Shafting 350 1.750 5.00 1s Likely Super Critical Shafting 250 1.750 7.00 16 Likely Cast-In Impingement Cooling 680 13.128 19.31 7 Likely Abradable Coatings 280 2.917 10.42 13 Assured bearing Fatigue Life • Combustor/Diffuser 490 8.169 16.67 9 50/50 Vortex-Controlled Diffuser 530 0.875 1.65 20 Likely Transpiration-Cooled Comb,-Sheet Fab 0.583 0.90 21 Likely 650 Transpiration-Cooled Comb. Fab • Engine Accessoried 250 2.917 11.67 10 Likely Electronic fuel Control 400 3.501 8.75 14 Likely Fuel Pump S Metering System 400 29.174 72.94 1 Likely Engine Condition Monitoring • Noise Reduction 250 6.710 26.84 Likely Compressor Noise Reduction 500 4.084 8.17 15 Likely Combustor Noise Reduction Reduction Gear 300 38.90 3 Likely 11.670 o Composite Cear Case bank s ed on^basis of benafit/Cost Ratio In order to reduce maintenance cost, a primary STAT engine requirement, it is beneficial to protect the engines from dust-laden air common to many airports used by commuter transports. This protection may be obtained by faltering dust particles from the inlet air, making the compressor more tolerant of the dirty air, or a combination of both.

The state of the art for advanced propellers and propfans is such that they will exhibit much lower noise generation than do current propulsors. For this reason, it is essential that advance turboprop engines be considerably quieter than current engines of comparable power. The STAT engines, in particular, due to their high cycle pressure ratios and resulting supersonic compressor blade tip velocities, require some form of compressor noise reduction.

The high pressure ratio of the STAT engines is achieved with compressors which are designed with minimum rotor tip clearances and have highly loaded stages, compared with current technology engines. One of the problems encountered in designing such compressors is predicting tip clearance during periods of dy- namic structural response due to surge, rapid thermal gradients, and the rela- tively unknown phenomenon of rotating stall.

The small diameter, high speed requirement of the 1790 kw (2400 shp) STAT en- ,)reesor impeller.

Sine presents a problem for the design of a practical e t The bore loads for such an impeller preclude the use of a conventional cast impeller. The solution selected for the STAT engine is the use of an advanced hybrid impeller with a forged hub diffusion bonded to a cast outer shell with blades.

Compressor research programs recommended are: o Hybrid centrifugal compressor impeller o Compressor erosion protection o Inlet particle separators

F o Axial compressor aft stage study

o Centrifugal compressor-advanced high hub/tip ratio study t o Rotor/case response to rotating stall Hybrid Centrifugal Compressor Impellers The STAT 1790 kw (2400 shp) engine incorporates a centrifugal compressor im- peller in its last stage. The high speed and small size of this impeller is

such that a conventional cast impeller fill not meet the bore loading require-

ments. It is necessary to develop a hybrid impeller for this application, which has a high strength, forged hub diffusion bonded to a cast outer shell with blades. An example of this type of impeller is shown in Figure 75.

This STAT program will address to the heat treat response of selected alloy combinations, the generation of a data base for a single alloy combination and component cyclic spin (LCF) testing.

Compressor Erosion Protection The STAT engine requirement for minimum maintenance cost will be met in part by designing its compressor to be tolerant of dirt particles ingested in the airstream, as shown in Figure 76. A need exists for a computer model that will enable the designer to predict the dirt tolerance of any given compressor configuration, and thereby enable him to select the optimum design from sev- eral candidates.

The elements required to develop such a computer model include: o Modification of an existing solid particle trajectory calculation model to account for particle impacts with the rotating and stationary blading within a compressor.

• A new mathematical model for the mechanics of- Particle rebound from compressor surfaces Particle shattering upon impact Compressor blading and side-wall material loss • Determination of performance and stability of an eroded compressor from the effects of- Clearance changes Increased roughness Decreased solidity Blade shape change Figure 75. - Dual-property titanium impeller with a wrought Ti-624b hub HIP bonded to a cast Ti-6241 airfoil shell.

o Test validation--A test program will be conducted with controlled contami- nant ingestion to ascertain compressor performance less, stability loss, and erosion damage resulting from contaminant ingestion. The test program will include two single-stages test compressors, one representing current technology and the other designed to minimize erosion effects.

Inlet Particle Separators In meeting the STAT requirement for minimum maintenance cost, one of the areas of concern is preventing ingestion of dirt-laden air, which is the primary cause of erosion damage causing Ferformance loss and frequent repair. An in- let particle separator (IPS) may be required for the STAT engines to eliminate this problem. Basic test data and analytical studies are required to design an efficient and cost-effective IPS for the STAT engine commuter transport aircraft.

Q "..4 i t.

i ,i a

7 Typical particle trajectory

TE80-2031

Figure 76. - Foreign particle ingestion phenomenon.

An existing particle trajectory calculation will be used to assess the poten- tial separation efficiency of various IPS concepts suitable for use with STAT engines. Since the calculation is for axisymmetric or two-dimensional ducts, the no:uxisymmetric cases will be modeled as piecewise two-dimensional or axisymmetric.

The two most promising candidate concepts (one axisymmetric and one nonaxisym- metric) will be selected for detailed design and testing on a flow rig. A typical concept is shown on Figure 77. The testing will allow determination of pressure drop through the separation devices, inlet total pressure distor- tion at the exit of the device (which is the compressor inlet plane), and sep- aration efficiencies of the device for various types of test contaminants such as aircraft coarse dust and Mil-C-sand.

This STAT research program will provide the basic data required for the devel- opment of a practical IPS for the STAT engines.

Rotor/Case Response to Rotating Stall Increased performance and reduced life cycle costs desired for compressors of advanced gas turbine engines, like the STAT engines, are being achieved by designs with fewer and more highly loaded compressor stages, while maintaining tight rotor tip clearances. This simultaneous achievement of higher aerody- namic loading and minimum tip clearance is extremely difficult because of the dynamic behavior of the rotor and case during surge, rotating stall (Figure 78), and periods of rapid thermal gradients.

Dirty inlet air TE80-2032 Figure 77. - Inlet particle separator concept.

Case Rotor Face on ce.- TESO-2033 Stall zone speed Figure 78. - Rotating stall phenomenon.

A predictive capability which will yield accurate information regarding tip clearance variation during periods of dynamic structural response is vital to the success of the STAT advanced compressor designs. Current analytical tech- niques do not account for Asymmetric effects in rotor/case coupled, structural response. These effects, due to such common features as compressor horizontal flanges or bleed manifolds, can be very important in the structural dynamic response of compressors.

i This STAT research program would provide an analytical capability recognixing these asyms<etric effects, and thus yield an improved predictive capability compared to the current state of the art. Additionally, this program will yield experimental data describing the structural response of an advanced tur- boshaft compressor.

Axial Compressor Aft Stage Study The STAT engines have a high pressure ratio, compared to that of current tech- nology engines, in order to achieve the desired fuel saving performance. This NASA research study would be concerned with potential problems associated with the very small blade rows comprising the latter stages of high pressure ratio/ low flow axial compressors, as shown in Figure 79. To improve the erosion resistant characteristics of these small airfoils, nonoptimum distributions of thickness-to-chord and leading edge radius will be needed. A systematic analytical/experimental investigation into the performance characteristics of these nonoptimum airfoil sections will add credance to performance estimates of potential STAT compressors. A determination of the sensitivity of these small blade rows to production tolerances including fillet geometry, angle variation, and surface finish, is needed. The compromises between close tol- erance (higher manufacturing cost) and aerodynamic performance with the re- sulting effect on DOC can be more accurately assessed with the data from this program. Reduced sensitivity of these stages to tip clearance will be a third part of the program work plan. Possible aerodynamic improvements may include blade track trenching and low loss end-wall loading distributions.

TEW-203-7 Figure 79. - Mial compressor.

Centrifugal Compressor-Advanced High Hub/Tip Astio Study p' 1790 kW (2400 shp) STAT Derivation of the optimum cycle performance for the engine resulted in the selection of an axial - centrifugal compressor as de- scribed earlier in this report. Designing the centrifugal stage for this com- pressor under all operating conditions will require certain data not currently ^, 9 available.

This NASA research study would deal with problems associated with designing high hub / tip ratio centrifugal compressors to operate behind an axial compres- sor, as shown in Figure 80. In general, the centrifugal stage of an axial-, centrifugal compressor will be characterized by high inducer hub/tip radius .!

ratio and low specific speed. Operating behind an axial compressor, the in- ducer would have to accept distorted inlet conditions, which would vary with speed and loading. The high inducer hub/tip ratio results from geometric matching to the axial compressor and the desire to reduce engine length (re- duced transition from axial exit to centrifugal inlet). The low specific speed (i.e., reduced mechanical rotative speed for a given flow and pressure ratio) results from tradeoffs between axial and centrifugal compressor effi- ciencies.

The combination of high hub / tip ratio and low specific speed results in an impeller flow path that looks quite different from most modern high perform- ance centrifugal compressors, which are designed to the lower hub/tip ratios and higher specific speeds to maximize performance potential. It is not sur- prising, then, that the bulk of available data defining state-of - the-art cen- trifugal performance is based on these "more optimum" designs.

TESO-2040 Figure 80. - Axial-centrifugal compressor.

F___ i } High impeller exit back curvature is generally accepted as having positive benefits on compressor performance, and will probably be incorporated in ad- vaned axial - centrifugal designs. The effects of high inducer hub/tip ratio on stage performance of these advanced compressors will be needed for future configuration studies.

Dual - stage centrifugal compressor testing has shown reduced performance levels for the second stages. These reduced performance levels are currently attri- buted to distorted inlet conditions being delivered by the intrastage cross- over duct. Similar inlet effects would be expected for operation behind an Tip Centrifugal Compressor Program axial compressor. The proposed High Hub / clean" and "distorted" inlet would address the inlet effects by testing with " conditions.

Turbines

With the higher pressure ratio and higher turbine temperature chosen for the STAT engine cycle, the gasifier spool speed increases, and its diameter de- creases, relative to current technology engines. Turbine configurations chosen for STAT engines feature hybrid turbine wheels that can meet the re- sulting requirement for bore load-carrying capability within the size con- traint.

The small size of the turbine wheels makes it virtually impossible to use in- dividual turbine blades with conventional dovetail attachment. The hybrid turbine wheels provide the solution to this problem by incorporating blades cast onto a ring, which, in turn, is diffusion bonded to the wheel hub. Also, the small size of the turbine blades make necessary the use of cast-in iz- pingement cooling air passages.

The small size and high speed of the main rotor shafts in the STAT engines make it difficult to design the main bearings with satisfactory load - carrying capacity and fatigue life. Considerable promise of improving these character- istics exists in better understanding the formation and control of forging flow lines in the bearing balls.

The small diameter of the STAT engines makes it impractical to use subcriti- ^i1, conventional steel shafting. Metal matrix composites, however, offer improved mechanical properties and can meet the desired shaft strength and stiffness within the size constraints. As a backup to composite shafting (should that material development not be achieved in the STAT program time " second" choice.

frame), supercritical steel shafting offers an alternate Basic design information is required in the ability to predict dynamic re- sponse, and to control vibrations in supercritical shafts in order to obtain their benefit the the STAT engines.

Another area, in which the small size of the advance technology STAT turbines presents a challenge is in maintaining minimum acceptable interstage leakage with practical production tolerances. The solution selected for the STAT tur- binas is the application of abrteive blade tip coatings and abradable tip seal coatings.

The turbine research programs recommended are: • Hybrid rotors • Composite shafting • Supercritical shafting • Cast-in impingement cooling • Abradable coatings • Bearing fatigue life n Hybrid Rotors The small diameter of the STAT gasifier turbine, coupled with its high speed, makes impractical the use of separate turbine blades with conventional dove- taLl attachment. It is necessary to use a hybrid turbine wheel with blades can onto a ring as one piece, and the ring, in turn is diffusion bonded to a high strength powder metallurgy hub.

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The feasibility of using hybrid, or dual-property, turbine wheels in gas tur- bine design has been demonstrated in several previous programs at DDA. An example of such a turbine wheel is shown in Figure 81.

Bond Wunmry r ,. lam'' 1180-2051 Figure 81. - Dual-property turbine wheel with PA-101 hub and .4ar-H247 internally cast airfoil ring.

In all of these previous programs, however, the alloy combinations used were PA-101 hub and Mar -M247 airfoils, and little effort was allocated to bore l: load-carrying improvements.

In the proposed STAT research effort, alternate hub alloys (i.e., IZRL76, AF95, AF115) will be evaluated for compatibility with Mar -M247. These alloys, along with a baseline PA-101 composition, will then be evaluated for post-HIP bonding response. Heat treatments with rapid cooling rates from the bonding and airfoil-coating diffusion cycle will be emphasized.

With establishment of preliminary heat treat response, a single alloy will be selected and heat treatments further refined. Following establishment of a fixed process, additional rotors will be produced, a mechanical property data ' base generated, and cyclic spin tests performed to evaluate LCF capabilities.

The resulting data will enable the selection of materials for the STAT an- ! Sines, which would provide an appreciably longer fatigue life than those of today's gas turbine engines.

Cast-In Impingement Turbine Cooling The small size of the turbine blades on the STAT engines makes impractical the use of conventional separate impingement tubes through the blades for the de- livery of blade cooling air. The obvious solution is to cast the cooling air supply passages into the blade as it is cast. The technique for this process is not currently available on a cost-effective basis.

This program would explore two possible core fabrication techniques in an ef- fort to develop a reliable, cost-effective foundry process for producing air- cooled turbine airfoils with cast-in impingement tubes: o Assembled core o Fabricated core insert The assembled core approach would be to inject the feed cavity core (marked A in Figure 82) and the collection cavity core (8 in Figure 82) separately. The cores would then be assembled by inserting quartz rod (marked C in Figure 82) into holes formed in the cores during the injection process. "gill" discharge holes (marked D in Figure 82), or other film cooling holes, would be formed with quartz rods inserted through the wax shell prior to ceramic mold dipping.

The fabricated core insert approach involves the placement of an insert with final blade internal geometry into the core die prior to injection. The core material would then be injected around this insert forming a one-piece core with the insert embedded within it; the core would than be processed to remove the insert. Subsequent processing would then be similar to present blade ring fabrication.

The principal problem anticipated for the development of cast-in impingement cooling would be core fragility. SJ.ugle casting trials could be made prior to core design in order to establish hole diameter and wall thickness constraints.

I I Figure 82. - Schematics illustrating candidate caat airfoil cooling schemes.

Abradable Coatings The small diameter of the STAT engines presents a problem in maintaining ac- ceptable interstage turbine leakage with pra..tical production tolerances. To comhst this problem, the STAT engines will use abradable surfaces on the tips of the turbine blades. In order to select the beat material for this use and the process for its application, an investigative program is required.

i_ This program will result in the development of the materials and processes required for direct application of abrasive particles to turbine blade tips, as shown in Figure 83. The abrasive elements will provide an augmented abrad- ability capability for enhanced gas path sealing and improved cycle effi- ciency. A primary advantage of the direct application process is its appli- cability and affordability for small turbine rotors with integral blades.

This concept has been taken satisfactorily through the proof-of-principle dem- 1 onstration stage, and represents a development effort with reasonable techni- cal risk.

Composite Shii tit4 The high pressure ratio of the STAT engines dictates high rotor speeds and small engine diameters, as compared to current technology engines. This sit- uation compounds the probl" of designing shafts to transmit torque from the power turbine through the gasifier rotor to the output shaft. i*, becomes im- possible to use conventional, subcritical, forged steel shafts for such appli- cations.

The best way to overcome the problem appears to be in the use of composite shafting, such as shown in Figure 84, which offers greatly inaroved mechanical properties over steel shafts. Composite shafts offer payoffs in the areas of engine weight and co-_iplaxity on the basis of having available to the designer (tensile the potentially high E/p (Young's Modulus to density ratio), Cr/P stress to density ratio), and rip ;shear stress to density ratio) engineering prcperties. Before composite shafts can be successfully applied to a commer- cial engine, however, the difficulty of manufacturing such a shaft with ac- companying biased-ply layups and end fittings, and of achieving the assumed high goals of engineering properties, must be overcome.

A ► brodable coating

Abrasive blade tip coating

TE80-2034

Figure 83. - High temperature turbine seal concept.

r L s4 ^j [I l^ aI i^ FIBERS TE 8797A Figure 84. - Typical composite shaft.

This STAT technoligy research program provides a finite element analysis pro- gram capable of computing the engineering properties ;E, p, o-, r) for lami- nates representative of large, metal matrix composite shafts. Sample shafts of this material will be fabricated and tested to verify analytical predic- tions made with this progri,:: This verified methodology will be an essential tool for the design of advanced STAT engines using composite shafts.

Supercritical Shafting If the composite shaf is are not developed sufficiently by the time they are needed for the STAT engine prograu, a satisfactory alternate may be available in the form of supercritical steel shafts. These shafts operate above bending critical speeds, as shown in Figure 85. As can be seen, the shafting is con- figured to operate above two bending critical speeds. Pasuge through the bending criticals in the transient range is attained through the use of squeeze film dampers located at non-mode locations.

.1 Max I nd Response Max Rotor speed TE80-2121 Figure 85. — Typical application of supercritical design to a power turbine shaft.

This STAT technology program will use a preliminary STAT engine design and perform rotor dynamics analyses. Various methods for vibration control of the supercritical shaft will be identified. A rig will be fabricated and tested to evaluate the design and verify the analytical methodology. The results of this program would provide a technique for the application of supercritical shafting to advanced $TAT engines if composite shafting is not available.

Bearing Fatigue Life The high rotor speed and small diameter, which is characteristic of advanced technology, high pressure ratio gas turbines such as the STAT engines, present a problem in designing main bearings with sufficient fatigue life and load — carrying capability. Significant improvements must be made in ball bearings over that which is available for today's gas turbine engines.

The primary contact fatigue failure site on bearing balls has been correlated with end grain concentrations in the polar and equatorial areas, as shown in Figure 86. Reports indicate that ball life also varies with end grain area.

End grain varies with grain flow angle relative to the surface. Also, grain areas may well change in size with different ball forging practices and die designs. The equatorial band width may vary with die design and with grind stock allowance.

^i

Polar area `'-

90 deg

u

t

60 dog. i

I'l l

50 deg

•w

Failure

40 deg

.density

30 deg

Fiber flow

^ deg

pattern '

"_/ ^ _'-i ; X10 deg

Equatof

- 0 deg equator i`

.471 TE 80-2056

Figure 86. - Failure density as a function of ball latitude.

Thus, standard practices in the forging process and acceptance criteria are needed to obtain maximum ball bearing fatigue life. Fatigue testing of balls with varying end grain characteristics must be performed to achieve fatigue life improvements.

The objectives of this STAT research program are to develop and validate meth- ods of characterizing forging flow lines in bearing components, and to sub- stantiate improvements in fatigue life and load-carrying capability. Au as- sessment would be made of current forging practices for balls and races.

" .'ree-dimensional characterization of parts forged by various methods will be performed. Improved forged bearings will be fabricated, and both static tests and fatigue tests will be conducted to quantify the fatigue life improvement.

Diffusers and Combustors The STAT study showed that one of the biggest factors in reducing commuter transport DOC was the reduction of engine fuel consumption. The advanced VCD, used in the STAT 3579 kW (4800 shp) engine, offers a significant reduction in pressure drop and at the same time a reduction in engine length and weight.

158 ,

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The STAT engines incorporate combustors fabricated from Lamilloy, a DDA-devel- oped, transpiration-cooled structural material. The use of Lamilloy combus- tors has proven benefits in the areas of cooling air reduction and combustor outlet temperature pattern profile. There is a need to reduce the cost of fabrication of Lamilloy material, as well as to simplify combustor construc- tion techniques, for this material to achieve its full potential for commer- cial engine production.

The state-of-the-art advances being made in propellers and propf&no will re- sult in much lower propulsor noise generation in the STAT transports. It is essential that the STAT engine core noise signature be reduced below that of current technology engines in order to stay below that of the propulsors.

The diffuser and combustor research programs recommended are: • Vortex-controlled diffuser • Transpiration-cooled combustor Vortex Controlled Diffuser The STAT engines must achieve significantly lower fuel consumption than cur- rent technology engines. A VCD offers an appreciable improvement in this area but requires further development to ensure the potential gain.

This STAT research effort is aimed at providing needed technical information to assist in the application of the VCD technology to diffuser-combustor sys- tems of advanced S 4 T engines. The VCD concept features a trapped standing vortex to achieve I;;:. pressure loss diffusion in a short length. In order to apply the advanced technology of the VCD concept to future engine designs, it is necessary to investigate and identify the geometric parameters important to system performance. Because of the separated flow nature of the VCD system, analytical modeling of the diffusion process requires elliptic Navier-Stokes numerical schemes, which are in the development stage and require experimental data for verification and improvement.

The proposed research effort will consist of an empirical program to obtain critical design information concerning the influence of selected geometric parameters upon VCD system performance. Results from the test program will be compared to analytical model results, and where discrepancies exist, the model will be improved. the payoffs from this effort will include improved diffuser system performance, reduced diffuser system length, reduced development and manufacturing costs, and improved analytical tools for future design of the STAT engines.

Transpiration-Cooled Combustor One factor in achieving a low acquisition cost for the STAT engines, and in obtaining longer engine life through a more uniform RIT profile, is in the use of Lamilloy (transpiration-cooled material shown in Figure 87) combustors.

The benefits of using Lamilloy for combustor walls have been proved on several Allison gas turbine engines. For this material to be used in commercial en- gine production, however, it is essential that the cost of fabricating the sheets of Lamilloy material be reduced.

l --f Cooling cir flow TE80-2036 Figure 37. - Lamilloy construction.

This STAT research program would reduce fabrication costs of the Lamilloy ma- terial without sacrifice of quality. An expected reduction of 25% in labor is projected with improvements in etching, hole production, sheet cleaning, and bonding procedures. An added improvement would be increased sheet size from 25.4 x 63.5 cm (10 x 25 in.) to 61.0 x 91.5 cm_ (2 1 4• x 3o in.), thus eliminating welds and details in combustor assembly.

Lamilloy combustors are fabricated by forming and joining flat Lamilloy sheets. This additional STAT research program will improve combustor durabil- ity, and therety increase STAT engine life, through assembly improvements.

These improvements will be made in the areas of welding (to produce a narrow bead), forming (to reduce flow restriction), and nondestructive testing meth- ods. These fabrication improvements will produce lower stress and temrerature values in joints and at radii.

Engine Accessories The STAT studies gave a clear indication that the two major areas, wherein engine improvements would be most beneficial to reducing commuter aircraft DOC, were in reduction of fuel consumption and maintenance cost. With regard to lowering turboprop maintenance costs, an area with great potential is in the application of an all-electronic fuel control system integrated with an engine condition monitoring system permitting full on-condition maintenance.

Three areas requiring state-of-the-art advances for the practical application of such a system to STAT engines are: r'- urboprop engine : o Electronic control for advanced, variable geometry, t o Fuel pumping and metering system o System integration of electronic control, fuel pump and metering system, / propfan control condition monitoring system, and propeller The engine accessories research programs recommended are: ^.

k o Electronic control r o Fuel pumping and metering system I o Engine condition monitoring E_ Electronic Control To meet the STAT program goal of a significant reduction in maintenance cost - authority, digital elec- over current technology engines, an advanced, full tronic fuel control system is a must. Such an integrated control system, as shown in Figure 88, would provide for constant optimum operation of the engine and propeller plus provide condition monitoring of both.

The proposed STAT research program is the first step in a total control devel- opment program. It is structured to address the advance turboprop require- ments associated with failure modes, manual control, and total system integra- tion, including the engine condition monitoring system.

Fuel Pump and Metering System An all-systems integrated, electronic control system is mandatory on STAT en- gines to achieve the low maintenance cost and high reliability desired. The objective of this STAT technology research effort is the design and partial development of an advanced fuel pump and metering system 'specifically suited for use in a full-authority digital electronic control system.

In order to realize maximum benefits in performance improvements, while reduc- ing overall system cost, reducing weight, and improving reliability, the fuel handling portions of the system must have design features and characteristics fully compatible with the overall system requirements.

DDA is currently evaluating five new fuel pump and metering system concepts specifically configured for use in future applications such as the STAT en- - 22046, has defined gines. This current program, funded under Contract HAS3 the requirements for such a fuel-handling system and developed possible ap- proaches for tradeoff studies. These studies define the following factors for making a comparative assessment of the candidate systems: o Reliability o Cost • Weight and size • Maintainability o Performance • Back-up operation • Development risk Engine Parameter Sensors

I ^

r.

Air Signals I DIU Prop Pita I^ Computer Control N Digital .: I l E le c t ro nic i on tro l \C ^r Digital Data links I

I ^ ^

t Aircraft Fuel Pump : r Electronic Power Automatic { and Control Signals and I .

Flight I Control Control System i Signals j I AFCS Metered Command Relay SAS Fuel Flow Signals Box Stator Vane Compressor Actuators Variati e Stator Vanes Engine-Aircraft Interface ` I TEED-1038 Figure 88. - Advanced turboprop propulsion control system.

From these assessments, a system is to be selected and a preliminary design

The recommended STAT research program would be a direct follow-on

developed.

to this currently funded work to proceed with hardware development and proof i tests.

This STAT program would include the detailed design of an advanced fuel pump Fabrication of hardware and bench testing would be per- and metering system.

After bench evaluation, the formed to determine performance capabilities.

t system would be evaluated on a Model 250 gas turbine engine during test stand operation, and would be tested in conjunction with an existing full-a•lthority digital electronic controller. i

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Flight testing of the total system model on a 250 engine would be performed to ` evaluate the dynamic performance under the actual aircraft operating environ- l ment .

Engine Condition Monitoring

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The STAT commuter aircraft requirements for propulsion, high reliability, and low maintenance cost would be met in part by the complete integration of an t ^, engine condition monitoring system with the engines' basic electronic fuel and prop control systems.

This STET engine condition monitoring program will develop improved methodolo- gies in order to support total on - condition maintenance concepts.

An engine condition monitoring system, such as that shown in Figure 89, will continuous- ly monitor all engine and propeller operation to determine engine life usage, and to provide maintenance directives and safety warnings. Life usage based on actual operation permits vastly improved usable life within acceptable risk limits. Automatic maintenance directives result in timely repairs, while re- ducing unrequired maintenance activity. Accurate safety warnings lower the risk of aircraft operation.

In particular, this program will develop and demonstrate improved techniques relative to: • Mechanical LCF accounting • Turbine thermal stress analysis • Performance degradation analysis LCF and turbine thermal stress analysis enable part - life prediction to be based on actual usage rather than worst - case time estimates. This feature permits safely running parts nearer their theoretical life, and therefore greatly improving service life. Performance degradation analysis complements the true on-condition maintenance philosophy.

The techniques resulting from this STAT research program will enable incorpo- ration of an on-condition monitoring system for the next generation STAT en- gines.

Noise Reduction Compressor Noise Reduction High cycle pressure ratio advanced gas turbine engines use high speed compres- sors, such as shown in Figure 90, with supersonic blade tip velocities. The resulting multiple pure - tone (MPT) generation would add significantly to the noise signature of the STAT commuter aircraft equipped with a propfsn or ad- vanced, low noise propeller. This effect is expected to be particularly acute during approach, when propeller thrust is low.

INLET VARIABLE COMBUSTOR INLET GEOMETRY PRESSURE. & TEP'PERATURE PRESSURE & TEMPERATURE POSITION PROP BLADE VIBRATION ANGLE TORQUE TURBINE PRESSURE & TEMPERATURE LUBRICATION FUEL SPOOL SYSTEM CONTROL SPEEDS PARAMETERS PARAMETERS x•

0-

AIRBORNE COMPUTER REPORT GROUND OUTPUT DATA ANALYZER TE8G-2119 Figure 89. - STAT engine condition monitoring.

This propobed STAT research program will validate the theory that leading edge sweep materially -educes tha effective Ma-ii number at that location, and thereby reduces the noise generated. This theory can be proven by test of either an axial or centrifugal compressor. An existing single-stage centrifu- gal compressor rig will be used for this research program. The inducer will be redesigned to provide leading edge sweep and used to fabricate a test unit.

A test wiii he conducted of both original and modified configurations.

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Small diameter, high speed ^r 1, ## ^ 1 TESO-2118 Figure 90. - Typical advanced high pressure ratio compressor.

The results of this test should verify the theory in question and provide a data base for designing advanced gas turbine compressors without MPT noise generation.

Combustion Noise Reduction generated noise ( see Figure 91) usually referred to as "core" Combustion - noise, has been a limiting factor in achieving large turbofan noise reductions in the low thrust approach condition where jet and fan noise are lowest. For many turboshaft engines, the combustor is the dominant noise generator, and combustion noise represents the major portion of the total sound power radi- ated at both high and low power settings. Engine noise has generally been assigned a minor role in assessing noise generated by turboshaft propulsion c noise generated by propel- systems because of the obvious high level harmoni / rotor broad lers and rotors, and the difficulty in separating engine / propeller band noise. Studies at DDA indicate that the contribution of engine combus- tion noise is not minor, and in fact, may be a major obstacle to designing new turboprop transport aircraft to meet certification noise requirements. This will be particularly true for the STAT transports where advanced propulsoos are being developed with very low noise signatures.

[e nbuStion noise source

TE80-2117 Figure 91. - Typical gas turbine annular combustor.

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► s The recommended STAT research program addresses the problem of low noise com- bustor development. A Model 250 engine will be tested to determine its noise •s signature in combination with select performance characteristics. An analysis of this data would be made, and combustor design modifications formulated for reduced noise generation. These design modifications would be tested on a .` Model 250 combustor rig to obtain correlating experimental data. The optimum combustor configuration would be tested on a Model 250 engine to verify the reduced noise signature.

The information gained from this STAT research program would enhance the abil- ity to design advanced gas turbine engines with lower noise combustors.

Reduction Gears The metal matrix: composite materials now available for the reduction gear and other major engine castings, offer significant gains in weight reduction while providing increased thermal stability and greater rigidity. Before these ma- terials can be used advantageously in the STAT engines, however, much informa- tion is required relative to their material properties and machinability. The reduction gear research program recommended concerns the use of composite gear cases.

Composite Gear Cases One way to reduce the DOC of STAT engines is to lower their weight, compared with current technology engines. A considerable weight reduction potential exists in the use of new metal matrix composite materials for the reduction gear cases as shown in Figure 92. Composite gear cases will also provide greater rigidity and increased thermal stability for the reduction gear as- semblies.

In this STAT research program, DDA would investigate two metal matrix compo- site materials for this application: SiC whiskers in magnesium or aluminum castings, and chopped polycrystalline alumina (FP) fibers in magnesium or alu- minum castings.

In order to use this material in the design of the STAT engines, material characterization of each composite must be performed. Mechanical properties must be determined, including stiffness, strength, thermal coefficient, LCF, HCF, and corrosion resistance. Also, the machinability of the new materials must be determined in order to successfully fabricate the advanced engine gearboxes. Machining characteristics such as turning, boring, threading, etc., will be evaluated using test specimens.

EXPERIMENTAL ENGINE PROGRAM Full-scale development of a STAT engine i- viewed as a commercial risk venture dictated by market requirements. Technical and economic decisions are re- quired to establish engine size. Such items as passenger load, unrefueled range with reserves, flight speed, cruise altitude, and field length strongly impact engine size. The R&D programs recommended will provide design data to permit selecting the proper components and arrangement for the size ultimately .t Figure 92. - 3579 kW (4800 shp) STAT engine reduction gear.

required. Technology needs can sometimes be developed in larger sizes; how- ever, it is generally recognized that risk is reduced appreciably if suitably constructed experiments can be accomplished in the size and environment of an appropriately sized experimental engine.

Such an experimental engine program would also offer an unparalleled opportun- ity to test unique technology features and systems being newly incorporated in the STAT engines. An example of this would be the active turbine clearance control system described in the subsection entitled, "Configuration Trades."

In this instance, the technology for such a system exists now, but the mechan- ical integration is best accomplished on an experimental engine prior to full commercial development.

Table LXX lists the 21 STAT technology research programs recommended to NASA as a result of this study effort. This table indicates 10 research programs which would benefit from the experimental engine program, since the applicable technologies are best developed on the unity size engine.

Upon completion of basic R&D programs of the type described in the previous section of this report, it will then be feasible to conduct an experimental STAT engine program. The results of this program will provide assurance that full commercial development of the STAT engine may be initiated within an ac- ceptable risk.

^. TABLE LXX. - RECOMIENDSD STAT TECHNOLOGY PROGRAMS TTgftlo*r research item Momt +tabu k1loble COMM Compresears Mybrid centrifugal compressor impeller X Centrifugal impeller used only in sell engines Compressor ores on protection X Inlet particle operators X Engine sine related to Problem X $in and hi gh spead related to problem Compressor "else reduction Axial compressor aft sups study X Aft stage bleding su=e critical X in engines small Centrifugal impeller used only Centrifugal high hub/tip compressor-adwnced Rotor/case response to rotating stall X Turbines X Used in ell engi"ea only sea sits is critical lybrid voters Cast-in impingement cooling X Problemisolution occurs in small sits only X Minim*) cl earances relate to small also@* Abredable coatings Diffusers-combustors Vortex-controlled diffuser X X Trenspiretion-cooled combustor-shoot fab.

X Trenspirstion-cooled combustor-fab.

X +^ castor noise reduction Structure and shafting X Composite gear cases tl X large engine shafts use different composite Composite shafting gearing fatigue life X X Applicable to high spot only. dampers sin critical Supercritical shafting Engine systems X Electronic control Feel pump and metering system X - X Engine condition monitoring The basic elements and timing of such an experimental engine program are shown in Figure 93. After six months of design effort, a "long lead time" release of selected engine hardware will be made. The full detailed release of all hardware will be achieved nine months from go-ahead.

Three serialized test engines would be built for this program, with the first ' s worth of spare parts would delivered in 22 months. Three equivalent engine be fabricated to support the component and engine test programs.

The component development required for this experimental engine program would entail three basic types of rig operation. A full-scale combustor rig prograA would be initiated 10 months from go-ahead. The combustor testing would ro2- quire approximately SO hours of rig time, and would cover cold flow aerodynam- ic and pressure drop tests and burning tests of the following types: • Thermal paint ( metal temperatures) • Sea level and altitude relight characteristics • Fuel distribution • Power calibration Approximately 100 hours of compressor rig tests would be conducted to deter- mine aerodynamic performance in terms of airflow, pressure ratio, and effi- ciency, as well as interstage data to define the operating characteristics of each individual sicage.

ONion Fa6rt^otian ^J^ tnra En41m qtt M &qim ► a ► tiros. hr too MFGA TESO-2122A Figure 93. - STAT experimental engine schedule.

The reduction gear assembly would be tested on a back-to - back rig, wF _h is capable of simulating full-design torque at rated speed. Approximately 200 hours of testing will be conducted to determine gear profile characteristics, lubrication capability, and structural stability of the gearbox.

The engine testing consists of three types: gas producer, power section, and full engine. The gas producer tests would provide the initial performance verification of the HP turbine. The power turbine performance would be deter- mined from the power section tests. The power section tests would be con- ducted on a dynamometer to provide full power absorption and checkout of all mechanical and lube systems. The engine testing would be conducted on a prop stand to provide for complete fuel and control system checkout. A total of 200 hours would be accumulated during these tests.

This program, with a parallel prop fan program and aircraft studies, will pro- vide the basis for launching an integrated demonstration effort, which would then be followed by a commercial aircraft system development program.

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t CONCLUSIONS STAT studies show that the largast improvement in commuter aircraft DOC, as influenced by engine characteristics, is achieved through reduced fuel con- sumption and reduced maintenance cost. A 10% reduction in sfc results in 3.5 to 4.0% reduction in DOC for the aircraft used in these studies when fuel is priced at $0.264 / L (=1.00 / A 1OZ reduction in maintenance cost was found gal).

to be worth 1.0 to 1.3Z in DOC, Fuel consumption can be reduced through improved thermodynamic cycle effi- ciency. A compresor pressure ratio of approximately 20:1, and an RIT near 1506 K ( 2250 ' 7), appears to be best for minimum DOC.

Compressors chosen for STAT engines feature single - spool configurations with a reduced number of stages and reduced blade count, compared to current tech- .^.

related research programs recommended for technol- nology engines. Compressor - ^.

ogy development include: o Hybrid centrifugal compressor impeller o Compressor erosion protection 1 o Inlet particle separators i o Compressor noise reduction •^ o Axial compressor aft stage study o Centrifugal compressor advanced high hub/tip ratio study I^ Hybrid turbine wheels are needed with greatly increased bore load capability.

Composite shafting with increased stiffness is required to permit subcritical Dynamic problems associated with rotor/ operation with simple - rotor support.

case response to compressor rotating stall must be understood to optimize com- pressor vane and blade tip clearance and ensure structural rigidity. Inter- o pment must be continued if desired turbine ef- stage and blade tip seal devel ficiency levels are to be attained. Turbine, bearing, shafting, and dynamics research programs recommended for technology development include the following: • Hybrid rotors • Cast-in impingement cooling • Abradable coatings • Composite shafting • Bearing fatigue life • Rotor / case response to rotating stall Fuel system maintenance is a mayor contributor to problems in achieving con- sistent engine availability, and is also a significant cost factor. The elec- tronic control system concept offers the potential of solid improvement in t reliability and maintainability, as well as interfacing with an on-condition !

maintenance data system to predict timely maintenance actions.

Maintenance costs can be reduced by incorporating on-condition maintenance, improved modularity, improved reliability and maintainability, and an effi- cient maintenance management system using engine condition-monitoring data for _ F key input.

I^ { R&D programs recommended in this area include: o Electronic control o Fuel pump and metering system o Engine condition monitoring recommended for the STAT engine will Incorporation of the technology advances result in the following improvement in engine characteristics referred to the current technology base used in the study: 1790 ( 2400) 3579 (4800) Power size, kW (shp)

-16.8 -19.0

sf c--Z

-25.3 Weight--Z -12.6 Price--% -18.9 -16.4 -56.1 -62.0 Maintenance coat--Z Z -21.8 -16.7 Envelope length- - in significant fuel savings and re- Achieving these improvements will result as shown below: duction in DOC compared with the current technology base, High speed Aircraft _. ow speed Fuel consumption- - Z -19.6 -22.8 -13.6 -16.2 DOC *--% Total cost of ownership *-- Z -10.9 -13.3 *$ 0.396 / L ($1.50/gal) M REFERENCES 1. Reduced Cost Turbine E ine Concepts Program, Final Rep^rt No, AFAPL-TR- eptember 1978.

2. Small Trans ort Aircraft Pro eller Studv. Parametric Proceller Data Pack e f or Current Techaology M—w— ut-e-r-I prepared for S a NASA-Lewis by Hamilton t ndard Div., Corp., under Contract NAS3-22039.

3.

Corp., under Contract NAS3-22039.

LIST OF ABREVIATIONS AND SYMBOLS ISymbol Meaning AFSC Airframe specific cost ^• A Axial A-C Axial-centrifugal ATE Advanced technology engine AEO All engines operative AR Aspect ratio Aw Area wetted ( aircraft) ATA Air Transport Association ALT Altitude AN2 Blade stress parameter AFWT Airframe weights ' AF Adjustment factor ATEGG Advance turbine engine gas generator BOT Burner outlet temperature Burner ,gyp Combustion pressure drop BSFC Brake specific fuel consumption CPR Compressor pressure ratio 'CTE' Current technology engine ( without inherited learning) CTE Current technology engine ( with inherited learning) CF Conversion factor C D Coefficient of drag Cent Centrifugal DOC Direct operating cost DDA Detroit Diesel Allison dB Decibel Dp Diameter ( propeller) DTE Derivative technology engine EAS Equivalent airspeed EFH Engine flight hour E s modulus Young ' ESFC Equivalent specific fuel consumption EPNL Effective perceived noise level E/p E/rho ( Young ' s modulus-to-density ratio) Ep Engine power EPNdB Effective perceived noise-decibels EMDP Engine model derivat l 4ve program Net jet thrust FNET FTOT Total engine thrust FOD Foreign object damage FAR Federal Aviation Regulation FW Finished weight F Fahrenheit Fixed costs FC

I

SSmbo1 Meani . nz Gallon gal GE General Electric GBL Gearbox loss GB Gearbox gJeh Turbine average stage loading / O2mean HIP Hot isostatic press HLH Heavy lift helicopter f HP High pressure i High pressure turbine HPT HCF High cy c le fatigue h Hour(s) Initial operation capability IOC IFR Instrument flight rules IAS Indicated air speee Imp Impingement Ic Index of cost ID Inside diameter Inertial particle separator •^ IPS T KIAS Knots indicated air speed K Kelvin ;C L Litre LCC Lockaee:? California Company L/D Lift / drag ( ratio) LCF Low cycle fatigue LP Low pressure LPT Low pressure turbine LD Blade-to-blade shroud loading MN Mach number Maritime partrol aircraft MPA Kfg Manufacturing MIF Materials index factor MTBF Meantime between failures MTBR Meantime between removals j MPT Multiple pure tone NA Not available NTS Negative torque signal NPR Nozzle pressure ratio N14T Corrected rotational speed NpT Power turbine rotational speed NGGT Gas generator turbine rotational speed N Rotational speed NS Specific speed i OASPL Overall sound pressure level OEW Operating empty weight U6 Original equipment manufacturer OEI One engine inoperative OD Outside diameter , PC procurement costs PPFRT Prototype preliminary flight rating test PAX Passenger PS Power section P&W Pratt and Whitney PRR Premature removal rate Px Power extracted Rotor inlet temperature (turbine) RIT Compressor pressure ratio Rcamp Compression ratio R C Rb Radius of bore (wheel) Radius of outer diameter (shaft) Ro RED Research and development Specific fuel consumption (uninstalled) sfc-un Specific fuel consumption sfc Sea level static, standard day SLSS Scale factor SF Sea level static SLS Sea level takeoff SLTO Shaft horsepower shp Shaft horsepower (uninstalled) shp-un snm Seat nautical mile Seat kilometer skm Takeoff gross weight TOGW Total cost of ownership TCO TAC Total aircraft cost (flyaway price) Ti Titanium Tc Temperature of cooling air Time between overhaul TBO TO Takeoff Thrust specific fuel consumption Tsfc U Utilization (rate) Ut Tip tangentrial velocity

Ut /47 Corrected tip speed

Block velocity VB VCD Vortex controlled diffuser Vertical/short takeoff and landing V/STOL Corrected airflow

W08

Cooling airflow We Wa Airflow y W f Fuelf low

i

Greek Alphabet IQ Eta ( efficiency)

"Ip Propeller efficiency

tensil strength) Sigma ( Q T Tau ( shear strength) Combustor pressure drop P/Pburn Stage equivalent work A h/ecr Total - to-total adiabatic efficiency 14 T-T Rho ( weight density) P Q/P Tensile stress - to-density ratio z/p Sheer stress-to - density ratio I ^ I; 1.

r is >x

i