General Aviation Turbine Engine (GATE) study

TABLE OF CONTENTS Page SECTION I 1.- 1.0 Summary 1 1.1 Task I - Market Survey 1 1.2 Task II - Broad Scope Trade-Off Studies 2 1.3 Task III - Common-Core Concept Evaluation 4 1.4 Task IV - Technology Program Plan 5 .., SECTION II I 2.0 Introduction SECTION III 3.0 Market Survey 3.1 Market Forecast 11 3.1.1 Market Forecast - Fixed-Wing Aircraft 11 3.1.2 Market Forecast - Rotary-Wing Aircraft 26 3.2 Engine Conceptual Design 32 3.3 Definition of G~s Turbine Power Classes 38 3.4 Screening and Selection 43 3.4.1 Target Turbine Engine Original Equipment 46 Manufacturer's (O.E.M.) Cost 3.4.2 Allowable Turbine Engine Cost 50 3.4.3 Other Engines Considered 61 SECTION IV 4.0 OBJECTIVE 63 4.1 Selection of Candidate Engine Configurations 64 and Candidates Advanced Technology 4.1.1 Candidate Engine Configurations 64 4.1.2 Advanced Technolgy 67 4.2 Baseline Engine Design 78 4.2.1 Turboprop Baseline 79

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iii , TABLE OF CONTENTS (Contd) Page 4.2.2 Turboshaft Baseline 87 4.2.3 Baseline Turbofan Cycle and Performance 87 4.3 Aircraft Sizing and Sensitivity Studies 96 4.3.1 Fixed-Wing Aircraft Sizing and Miss. Anal. 99 4.3.2 Sensitivity Studies 119 4.4 Eng ine Trade-Off Studies 119 4.4.1 Cycle and Configuration Trade-Off Studies 120 4.4.2 Detailed Component Trade-Offs 131 4.4.3 Turbofan and Turboshaft 141 4.4.4 Summary - Engine Trade Studies 143 4.5 Engine/Aircraft Trade-Off Studies 146 4.5.1 Acquisition Cost' 149 4.5.2 Engine Configuration Trade-Off Studies 152 4.6 Benef i t Analysis 162 4.6.1. Current Technology Turboprop 162 4.6.2 Reciprocating Engines 166' 4.6.3 Current Technology Turb~~haft 172 SECTION V 175 5.0 COMMON-CORE CONCEPT 179 SECTION VI 6.0 PROGRAM SCOPE 6.1 Preliminary Design 6.2 Component Technology 181 6.2.1 High-Pressure Turbine 181 6.2.2 Compressor 188 6.2.3 Clearance Control 189 6.2.4 Combustion System 196 6.2.5 Low-Cost, Digital, Electronic Control 199 6.2.6 Hign-Work/Low-Speed Power Turbine 205 6.2.7 Laser-Hardened Gears 206 6.2.8 Gas Generator 209 6.2.9 Experimental Engine 212 6.2.10 Engine System Analysis and Definition 214 6.2.11 Schedule 214 6.2.12 Technology Development - Benefit Analysis 217 iv ,

TABLE OF CONTENTS (Contd) Page SECTION VII 7.0 CONCLUSIONS APPENDIX I 221 APPENDIX II 237 APPENDIX III 251 REFERENCES 261 DISTRIDUTION v , LIST OF FIGURES Title Page Figure 1 Demand Curve - Aircraft Price Versus 15 Quantity 2 Single-Engine Aircraft Shipments 18 Twin-Engine Piston Aircraft_ Shipments 3 19 4 Turboprop Aircraft Shipments 20 5 Forecast of ~ircraft unit Shipments by 21 Market Segment 6 Aircraft Unit Shipments and Pretax 22 Corporate Profits General Aviation Manufacturer Net Billings 7 24 and Pretax Corporate Profits (Piston and Turboprop) Forecast of Manufacturer Net Billings by 25 Market Segment Historical Light Helicopter Sales 28 Light Helicopter Sales Forecast 10 29 11 Forecast with GATE Engine 31 GATE Engine Criteria Rotary Wing 12 33 Applications TPE Conceptual Design 13 34 14 TFE Conceptual Design 15 Eng ine Power 16 Engine Power Requirements 42 17 Estimated Engine O.E.M Cost 1977 Dollars 18 Estimated Engine O.E.M Cost, 1977 Dollars 19 Potential Turbine Enyine Demand (1988) Candidate Free-Turbine Turboprops and 6S Turboshafts and Two-Spool Turbofans vi , LIST OF FI~JRES (Contd) Title Page Figure Candidate Single-Shaft Turboprop and 66 Turbofan Gas Gen~rators Gas Generator Configuration No. 1 Candidate 68 Technology Gas Generator Configuration No. 2 Candidate 69 Technology Gas Generator Configuration No. 3 Candidate 70 25 Gas Generator Configuration No. 4 Candidate 71 Gas Generator Configuration No. 5 Candidate 72 27 Gas Generator Configuration No. 6 Candidate 73 28 Baseline Turboprop Engine Design 80 Turboshaft Baseline Engine - 88 TSE Model 1060-1 Turbofan Baseline Engine TFE Model 1054-2 89 31 Mission Requirements - Pressurized Twin 103 (Designs 1, lA, and 2) Turboprop-powered Pressurized Twin 104 Turboprop-powered Medium Twin Sizing 105 Turboprop-powered Pressurized Twin-Engine 106 Sizing Turboprop-Powered Pressurized Twin-Engine 107 Sizing 36 Relationsh!p of Acquisiticn Cost 109 to Wing Loading for Aircraft Design No. 1 37 Relationship of Operating Cost 110 to Wing Loading for Aircraft Design No. 1 Mission Fuel Consumption 111 vii , LIST OF FIGURES (Contd) Title Page Figure 39 Mission Requirements - Light Twin 40 Turboprop Light Twin 115 41 Turboprop-Powered Light-Twin Engine 116 Sizing 42 Turboprop Light-Twin Engine Sizing 43 Compressor Efficiency 1985 Technology 122 44 Turbine Efficiency, 1985 Technology 45 Turboprop Cycle Selection - 124 Free-Turbine Engines 46 Compressor Efficiency, 1985 Technology 126 47 Compressor Pressure Ratio Trade-Offs 48 Free-Turbine Engines - 129 Compressor Configurations 49 Free-Turbine Turboprops - 130 Turbine Configurations 50 Single-Shaft Turboprops - Turbine Configurations 51 Turbofan Cycle S~lection Engine Cost Reduction 53 Turbofan Pressurized Twin Turbine Inlet Temperature Trade-Offs.

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54 Turboprop Pressurized Twin Turbine Inlet Temperature Trade-Offs 55 Trade Study Summary - Medium Pressurized Twin 56 Trade Study Summary - Light Twi~ viii LIST OF FIGURES (Contd) Figure Title Page 57 Comparison of a Turboshaft Version of ~he 174 GATE Free-Turbine Turboprop with a Current- Technology Turboshaft in a Light, Utility, Single-Engi~e Helicopter 58 Common-Core Concept 176 Frae-Turbine Turboprop 182 60 Cooled Laminated Radial 184 High-Pressure Turbine 61 Sheet Alloy Processes 185 62 PM Titanium Centrifugal Compressor 191 63 Clearance Control 194 Combustor Technology 197 Low-Cost Combustor, 1478°K (2200°F) 198 66 Low-Cost Airblast Fuel Atomizers 201 67 Low-Cost Digital Electronic Control 204 High-Work/Low-Speed Power Turbine 208 69 Gas Generatol Program Plan 210 70 Experimental Engine Program Schedule 213 71 Experimental Engine Test Plan 215 72 Recommended GATE Experimental Program 216 73 GATE Design No. 1 247 74 GATE Design No. lA 248 75 GATE Design No. 2 249 76 GATE Design No. 4 250 Engine Weight Sensitivities, Turboprop 252 Pressurized Twin, Design No. 2 ix LIST OF FIGURES (Contd) Figure Title Page Fuel Consumption Sensitivities, Turboprop 78 253 Pressurized Twin, Des!)n No.2 79 Engine Weight Sensitivity Studies, 254 Turbofan Pressurized Twin, Design No. 1 80 Fuel Consumption Sensitivity Studies, 255 Turbofan Pressurized~win, Design No.1 81 Engine Weight Sensitivities, Turboprop Light Tw~n, Design No. 4 Fuel Consumption Sensitivities, Turboprop 82 257 Light Twin, Design No. 4 x LIST OF TABLES Title Page Table Market Characteristics 12 Fixed-Wing Market Segmentation 27 3 TPE Conceptual Design 35 TFE Conceptual Design 37

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5 Turboprop and Turbofan Comparison 39 Fixed-Wing Turboprop Power Classes 44 7 Engine Production Quantities -Target Cost 49 Allowable Turbine Cost Example 53 9 Single-Engine Fixed-Wing Market 55 Twin-Engine, Fixed-Wing Market 56 Allowable Tutbine Engine Cost (1977 $) 57 12 Baseline Turboprop - Compressor 81 Character istics 13 Baseline Turboprop - Combustor 83 Character isti,:s Turboprop Baseline High-Pressure Turbine 84 Characteristics Baseline Tu:boprop LP Turbine Characteristic~ 85 16 Turboprop Baseline Cycle and Performance 86 Characteristics, uninstalled 17 Turbofan Baseline Fan Characteristicr. 91 Turbofan Baseline Compressor Characteristics 18 92 19 Turbofan Baseline Ccmbustor Ch~r~cteristics Turbofan Baseline High-Pressure Turbine 20 94 21 Turbofan Baseline Cycle and Performance xi * LIST OF TABLES (Contd) Title Pa~ Table . General Aircraft Characteristics-Gate Study 97 Performance Requirements-Fixed-Wing . Aircraft-Gate Study 98 Aircraft Sizing Summary 112 Light-Twin Sizing Summary 118 26 Turboprop Candidate Engine Characteristics.

Sea Level, Static, Standard Day, Takeoff, Power Uninstalled 121 27 Single-Stage Centrifugal Compressor Comparison 134 28 Radial HP Turbine Comparison 136 Axial HP Turbine Comparison 138 O.E.M. Cost of The Twelve Candidate Engi~es (1977 $) 144 Engine Characteristics with Advanced Technology Improvements 147 o o 32 l3l1 K (1900 F) Turbine Engine Character- istics, Sea Level, Standard Day, Takeoff Power, Uninstalled 148 Turbofan-Powered Medium Press~~ized Twin, Turbine Inlet Temperature Trade-Offs 153 Turboprop-powered Medium Pressurized Twin, Turbine Inlet Temperature Trade-Offs 154 35 Pressurized Twin, Engine Configuration Trade- Offs Studies 1478°K (2200 F) Engines 158 Light Twin Engine Configuration Trade-Offs Studies 1478°K Engines 159 37 Current Technology Turboprop Characteristics 164 Advanced Technology Benefits· 165 xii LIST OF TABLES (Contd) Page Title Table comparison of Gate and Current Technology Fng ines Groun~ Rules and Assumptions 169 comparison to Reciprocating Powered .Aircraft 171 Single-Engine Utility Helicopter Sizing and Mission Analysis Results* 173 43 Gate Experimental Program Plan - Program Scope 180 44 Laminated Radial Gas Generator Turbine 133 Required Compressor Technology- 9:1 Single-Stage 190 Clear?nce Control Technology 193 Low-Cost Airblast Nozzle Technology 47 200 Low-Cost Digital Elect~onic Control 203 High-Work/Low-Speed Power Turbine 49 207 Payoff Relative To Current Technology Turbine Engine 218 Two Place Light Single Engine 222 52 Utility High Perform~nce Single Engine 223 53 Fixed Gear High Performance Single Engine 224 Four Place Light Single Engine 54 225 ··;uht Retractables 55 226 Heavy Retractables 227 :·6 Agr icultural 57 228 Light Twins 58 229 xiii

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LIST OF TABLES (Contd) Table Ti tIe Page Cabin Class Unpressurized Twins 230 60 Pressurized Twins 231 61 Turboprops 232 Helicopters - Single Engine Piston 233 63 Helicopters - Single Engine Turbine 234 64 Helicopters - Twin Engine Turbine 235 65 Weight Breakdown, KG (LB) 238 65 Weight Breakdown, KG (LB) (Contd) 239 Power Plant Installation Details 6.

Weight, KG (LB) 241 67 Wing and Empennage Geometric Charcteristics 242 68 Configurations 1 and 2 - Optional Equipment List 243 69 Configuration 4 - Optional Equipment List 244 70 Estimated Drag Polar for the Gate Study Configurations 245 l478°K (2200 F) Baseline Turboprop Sensitivity Of Performance To Component Paramete~s 6100 M (20,000 FT), 389 KM/HR (210 KTAS), STD Day 258 l478°K (2200 F) Baseline Turbofan Sensitivity Of Performance To Component Par~meters 6100 M (20,000 FT), 389 KM/HR (210 KTAS), STD Day 259 xiv ....

SECTION I 1.0 Summary The ov'~rall objective of this study was to provide data on the applicability of gas turbines in the 112 to 746 kilowatt (150 to 1000 shaft horsepower) class to general aviation aircraft. 'l'his information will aid The National Aeronautics and Space Administra- tion in formulating the most effective technology program for small turbine engines. Airframe portions of this study were supported by the Cessna Aircraft Company, Pawnee Division, and the Bell Heli- copter Company.

1.1 Task I - Market Survey The objective of this task was to define t~e 1988 general aviation market including aircraft characteristics, mission requirements, major turbine engine sizes, and engine types.

A detailed market forecast \'/';:5 conducted that character ized the current and 1988 markets and projected the growth of the market utilizing trend and econometric forecasting methods. The present fixt_d-wing market was separated into 10 categories covering the power range up to 447 kw (600 hp) and ranging from the two-place, single-engine category to the pressurized, twin-engine category.

The rotary-wing market was divided into three categories: o Single-engine piston o Single-engine turbine o Twin-engine turbine An additional fixed-wing category, the current bUsiness turboprop market which utilizes engines in the 447 to 746 kw (600 to 1000 shp) class, was also evaluated.

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The market projection for fixed-wing aircraft resulted in the following annual compounded growth rates: ~ Single Engine - 4.3 percent o Twin Engine - 4.4 percent o Current Turboprop - 9.2 percent This fixed-wing market will grow from slightly over 15,000 units in 1977 to almost 25,000 units in 1988. Factory billings in current- year dollars will increase from approximately 1 billion in 1977 to over 3 billion in 1988.

The forecast for the rotary-wing market shows a two-fold increase in shipments from 1000 unit shipments in 1977 to approxi- mately 2000 units i!"'4 1988. u.S. rotary-wing factory billings are fore- casted to grow from 200 million in 1977 to over 450 million in 1988.

I A preliminary analysis, conducted during Task I, of the suita- bility of turbine engines to the various general aviation cate- gories indicated that turbine engines could be superior to reciprocating engines on over 9000 of the 1988 total units of 25,000. The major i ty of the units where turbines would not be superior to reciprocating engines are applications requiring less than 186 kw (250 hp), which represent a large number of total units but only 25 to 35 percent of the total billings.

The applications selected for detailed analyds in Task II were a pressurized twin, a light twin, and a light single-engine utility helicopter.

\ 1.2 Task II - Broad Scope Trade-Off Studies The objective of this task was to identify the combination of engine cycle, configuration, and technology that forms the optimum engine for each aircraft application.

'-_ .. - , The engine trade-off studies evaluated 17 engines that varied in cycle'and configuration and numerous component technology trades for those of the 17 that appeared most promising. The criteria that were used to evaluate the engines included: o Aircraft three-year total cos~ of ownership o Aircraft fuel consumption o Aircraft operating cost o Aircraft acquisition cost The three-year total cost was the i='r imary evaluation cr iter ion.

The study showed that a high turbine inlet temperature [1478°K o (2200 F») was superior in all applications studied and for all engine types. Turboprop en~lnes were shown to be clearly superior to turbofan engines for the class of fixed-wing aircraft because of lower fuel consumption and smaller size.

Technologies that resulted in improved engine performance and, low manufactur ing cost were found to be essential for the GATE engine.

The optimum engine for the fixed-wing application was a single-shaft turboprop comprised of a single-stage centrifugal com- pressor producing a pressure ratio of 9.0, a reverse-flow annular burner and a cooled turbine having one radial and one axial stage.

The engine rated a close second was a free-turbine turboprop com- prised of a single-stage centrifugal com~resscr producing a pres- sure ratio of 9.0, a reverse-flow annular burner, a cooled radial gas generator turbine, and a two-stage uncooled axial pcwer turbine.

The optimum engine for the light helicopter was a turboshaft version of the free-turbine engine.

The engine sizes required are: o Medi pressurized twin - 313 kw (420 shp) o Light twin - 242 kw (325 shp) o Light helicopter - 224 kw(300 shp) A comparison of the above turboprop engines to CUlrent tech- nology turboprops installed in the same aircraft yielded th~ fol- lowing results: o 9 to l7-percent reduction in total 3-year cost of owner- ship o 17-percent reduction in mission fuel consumption a 15 to lS-percent reduction in aircraft acquisition cost o 16 to IS-percent reduction in operat~ng cost o 6 to B-percent reduction in aircraft gross weight.

A similar comparison to'current reciprocating engines showed the following: o 20 to 28-percent reduction in total 3-year case of owner- ship a S to 16-percent reduction in mission fuel consumption a 14 to 20-percent reduction in aircraft acquisition cost a 2B to 3B-percent reduction in operating cost.

o 20 to 25-percent reduction in airplane gross weight 1.3 Task III - Common-Core Concept Evaluation The common-core concp.pt evaluation task attempted to identify a common-corp. engine, which would be compatible with the single- shaft engine identified as optimum for the fixed-wing application5, and the free-turbine turboshaft identified as optimum for the rotary-wing applications.

The results of the study indicated that a common core for these two engines resul ted in larger compromises than would be necessary if the optimum free-turbine engine was s~lected for both the fixed- and rotary~wing applications. The free-turbine engine is also compatible with a turbofan derivative.

1.4 Task IV - Technoloay Program Plan Program plans were prepared for seven technology items identi- fied as critical to the successful development of the GATE gas tur- bine epgines. The ~even technology programs are: o Laminated, cooled radial turbine o PM Titanium centrifugal compressor o Clearance control o Low-cost combustor and fuel nozzle~ o Digital electronic fuel control , o High-work/ low-speed power turLil1e o Laser-hardened gears.

The program plans were limi ted to high-r isk, high-payoff items which would not normally be developed in industry or Government- sponsored programs.

In addition to the component technology programs, an experi- mental engine program was recommended to provide for the integra- tion of the components in an engine environment. NASA sponsorship of the lntegrated development of these components and demonstration of these components in an experimental engine program would provide the impetus for industry to undertake the development and produc- tion of the GATE engines.

SECTION II 2.0 Introduction The recent history of aircraft engines has been characterized by the progressive introduction of turbine engines into small air- craft. The transition to turbine power in each succeeding category has resulted in safer, more comfortable, more reliable, and more productive aircraft. At this time, all segments of aviation have transitioned to turbine engines ~ith two notable exceptions--small general aviation airplanes requiring less than 336 kw (450 hp) and single-engine helicopters requiring 224 kw (300 hp) or less. This segment of the market has been denied the advantages of turbine power because of the sizable cost difference between ~urbine and reciprocating engines.

The National Aeronautics and Space Administration, Lewis Research Center (NASA/Lewis) spon~oreti the study reported hetein to investigate the feasibility of turbine eng inel~ for the smaller general aviation aircraft, ana to identify the most effective tech- nology program for developing the smaller turbine engines. The challenge of the General Aviation Turbine Engine (GATE) study is to determine if the advantages of turbine engines can be retained, while simultaneously achieving fuel consumption and engine cost levels required in this class of general aviation aircraft. The results of the GATE study provide added insight into the econom- ics and per formance reC:i<1irements of this aviation s£:gment and cll~ar ly shows the categor ies wi thin the general aviation market segment where turbines and reciprocating eng ines have super ior advantages.

'l'he GA'l'E study was a ten-month effort and cor.sisted of the following tasks: pm:c£L>,Nu P.'\G£ aAUK NCT fiLMEO o Task I characterized and projected the 1988 general avia- tion market and selected aircraft applications where tur- bine engines appeared to offer advantages. Task I considered turboprop and turboshaft engines in the 112 to 746 kw (150 to 1000 shp) class and comparably sized turbofans.

o Task II consisted of broad scope trade-off studies to identify the optimum turbine ~ngines for the applications selected in Task I. Task II was lj~ited to consideration of engines in the 186 to 447 kw (250-600 hp) class. The aircraft applications included turbofan- and turboprop- powered medium and light twins and a turboshaft-powered light single-engine utility helicopter. A comparison of the GATE engines with reciprocating engines and current turbines was also accomplished.

o Task III evaluated the feasibility of a common core for the fixed- and rotary-wing a~plications.

o 'l'ask IV defined the technology programs necessary to develop the engines defined in Task II and includes both component development and an experimental engine pro- gram.

The technology level of the GATE engines was consistent with introduction into service in 1988.

SECTION III Task I Discussion - Market Survey 3.0 Market Survet The objective of Task I ~as to forecast a 1988 market scenario for general aviation aircraft powered by engines in the 112 to 746 kilowat t (150 to 1000 horsepower) class. The forecast was to include the effects of regulatory factors such as noise, emissions, and safety, in addition to market needs as influenced by available engine size, performance, and cost. The identification of poten- tial important market applications for gas turbine e~9ines and cor- responding typical mission profiles was the primary output of this task.

The major elements of the market survey task were: o Market forecast o Advanced technology gas turbine engine conceptual design o Definition of gas turbine power classes for all general aviation categories o Screening and selection of potential gas turbine applica- tions o Definition of aircraft characteristics and mission requirements.

The objective of the market forecast was to characterize the general aviation market with respect to category and features, and to ~roject the annual 1988 production.

The conceptual design of advanced technology engines was to provide preliminary data for comparison to other engine types and provide basic data for preliminary aircraft design. The engine conceptual design effort provideu preliminary engine sizing infor- mation to Cessna, the airframe subcontractor, for use in defining airplane characteristics, and was also the basis for economic fea- sibility studies an~ estimates of production volume.

The sea-level, static, power rating required for a partic~lar gas-turbine-powered aircraft is a function of the mission perform- ance requirements. A ga5 turbine engine may be larger or smaller than a reciprocating engine sized to provide the same mission per- formance, depending on the engine sizing point, mission, and whether the reciprocating engine is turbocharged or naturally aspirated. A preliminary definition of the gas turbine power clas- ses that would be required to adequately cover th~ general aviation spectrum was made.

Screening and selection was conducted by consider ing every general aviation category, assuming the availability of gas turbine engines as defined in the conceptual design element of this task.

Performance, safety, and operating cost evaluations were primarily subjective and were influenced by results of past studies. The cost of turbine engines and the effect of this cost on airplane acquisition cost was quantified. It was apparent very early in the program that engine cost was the primary obstacle to the introduc- tion of gas turbines in the smaller general aviation aircraft.

Turbine engine cost goals were established based on, (1) the con- ceptual engine designs prepared earlier in thi:; task, and (2) detailed cost estimates prepared in prior studies for engines simi- lar in size, performance, and configuration. The selection of applications for detailed study was based un a comparison between the engine cost objectives and the allowable turbine engine cost for each genera! aviation category.

Mission requirements and aircraft characteristics were pro- vided by Cessna Aircraft Company for the fixed-wing aircraft and by the Bell Helicopter Company for the rotary-wing aircraft.

3.1 Market Forecast 3.1.1 Market Forecast - Fixed-Wing Aircraft The fixed-wing aircraft market forecast was conducted by the Garrett Marketing D~velopment Department and supported by Cessna Marketing. ~here were two patts to the market forecast: o Market Characterization Categorization of general aviation fixed-wing air- craft Data Collection De~and Characteristics o Market projections Trend Analysis Econometric Analysis 3.1.1.1 Market Characterization The general aviation fixed-wing market can be grouped into 10 categories excluding current turboprops, turbojets, and turbofans.

The ten categories are list~d in Table 1, which also shows some general characteristics that are associated with each category.

The major categories identified, and as further subdivided by power class, cover the range of applications very thoroughly with respect to cost and capability. New categories do not appear likely by 1988. Some features of each category may change such as engine size, and the split between pressurized ~nd non-pressurized *"'C$AACv=e J. ~:;:azw • TABLE 1. MARKET CHARACTERISTICS.

o Two Place Light Single Engine (Cessna 150) o Trainer o Owned by FBO for 2-4 years o Low initial and operating cost o Establishes brand loyalty o Low power, low useful load o Utility High Performance Single Engine (Cessna 207, Piper Super Cub) o Work horse; special duty applications (farms, ranches) o Functional and high . seful load o Reliability and dura.·ility important.

o Price related to usefulness and not highly competitive o Fixed Gear High Performance Single Engine (Cessna 182, Piper Cherokee) o High speed, high useful load, high power to weight ratio o Good aircraft for business or personal ~se o Very price competitive in given power class o Four Place Light Single Engine (Cessna Skyhawk and Cardinal) o Low power, 112-149 kw (150-200 hp) o Low initial and operating cost o Personal and rental aircraft o Light Retractables (Cessna Car6inal RG, Piper Arrow) o High speed o Low initial and operating cost o Functional o FBO, personal and business use o Very price competitive o Heavy Retractable (Beech Bonanza, Cessna Centurion) o High performance (speed and altitude) o High useful load (6 passengers) o Quality and luxury important o Business airplane o Price competitive .i.2 --"'-1J' A .1~S U ..... •• P 4i • p • .'~"~ ". .

R4QF" .AN • TABLE 1. MARKET CHARACTERISTICS (CONTD) o Agricultural (Cessna AG Truck, Rockwell Thrush) o Single engine specialty aircraft o Useful load important o Pric~ related to ability to perform joh o Reliabili ty, durability, and low maintenance cost are important o Light Twins (Beech Baron, Cessna 310)- o Unpressurized o High speed, good fuel economy o Low maintenance o Price competitive o Top of the line for personal owner: popular with FBO's and corporate owners o Twin engine safety o Cabin Class Unpressurized Twins (Piper Chieftain, Cessna 402) o High useful load (No. of passengers) o Unpressurized; operational altitudes under 3658 meters (12,000 feet) o Durability and low maintenance cost important o Commuter aircraft: high priority cargo: FBO use o Pressurized Twins o High performance (altitude and speed) o High useful load o Quality and luxury important o Corporate use ,.

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llii aircraft, and turbocharged and non-turbocharged engines, but these distinctions were not considered imp'octant enough to warrant con- sideration.

Turbine engines will not change the character of the cate- gories su:l.l.ciently to warrant special consideration. A high- speed, single-engine, turbofan-powered airplane is possible but the production potential for such an aircraft would bt: relatively small. Other highly specialized applications would probably be-the result of the introduction of low-cost tur~ines but would not, of themselves, justify the development of such an engine or contribut~ greatly to its success.

In addition to the categorization and the general character- istics of each category, specific data on engine power cl..lss, acquisition cost (1977 average equipped price), number of seats, cruise speed, engine time ~~twecn overhaul, and service ceiling was gathered for most models within p.ach category. Thi3 data is con- tained in Tables 51 through 61 of Appendix :;: along with similar I data for turboprop~ manufactured by General Aviation Manufacturers Association (GAMA) members. For each of these models or, in some cases, categories, production history and estimates through 1985 were available and provided the basis fer market projections. Th~ produ=tion estimates were obtained from manufacturers and from sub- scription forecasts such as Prost and Sullivan, DMS, and Forecast Associates.

The data obtained confirmed that the traditional relationship of price and demand did exist. Figure 1 shows the relationship between price and quantity sold for most general aviation fixed- wing ai rcraft. Ai rcraf t were grouped in 20-percent pr ice j ncre- ments for the construction of thi::; curve. There were thre~ dis- tinct segments along the curve: ,~, k- rURBOPROP

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~ ~ 500 1000 1500 2000 2500 3000 3500 NUMBER OF AIRCRAFT SHIPPED 1977 Figure 1. Demand Curve - Aircraft Price Versus Quantity ~ o Single engine-piston o Twin-engine piston (heavy and light twin grouping) o Turboprops Analysis of Figure I suggested some inelasticity of the market, i.e., price could vary without affecting demand. A detailed elas- ticity analysis was not per formed but discussion wi th industry representatives indicated that price increases of 10 percent or more could be absorbed without affecting demand if the "intangibles" of the buying decision are improved. This factor was important in the selection of turbine-powered applications for fur- ther study.

3.1.1.2 Market Projection Market projections were made using two methods. The first was an analysis of historical unit shipments and a projection of these trends through the forecast per iod to 1988. The second was an econometric analysis based on the observed relationships between h istor i cal aircraft sh ipments and fluctuations in the economy.

Only u.s. production was considered, and GAMA data was used for

consistency. The forecast does not account for the impact of foreign manufacturers, which could become more important in' the future, nor does it account for a change in the export growth rate.

Exports could result in further increases in unit shipments over and above the forecast if the growth in disposable income in devel- oping nations results in more demand for general aviation aircraft.

The historical tr~nd analysis was performed for three groups: o Single-Engine Piston o Twin-Engine Piston o Turboprop , In all three groups, unit shipments were cyclical but there appeared to be a consistent rate of growth over 'the 1955 to 1976 time period. Figure 2 shows unit shipments versus year for single- engine piston aircraft. Data was available from 1952 on, but only 1955 and later years were used to determine the trend line. The average annual compounded growth rate for single-engine aircraft is 4.3 percent. Over the same time period, the average annual com- pounded growth rate for twin-engine, piston aircraft is 4.4 per- cent, as shown in Figure 3. Figure 4 shows actual unit shipments and the growth trend for turboprops. Although data is available for 1964, only i965 and later years were used to establish the growth trend. The average annual compounded growth rate is 9.2 percent, which is more than double the growth rate for the other two groups. The overall growth trend in unit shipments is 4.4 percent and shows the strong contribution of the single-engine seg- ment, which accounts for more than 80 percent of total shipments.

A projection of unit shipments to 1988, based on the above growth rates, is shown in Figure 5. Total units shipped in 1988 ,will increase from slightly over 15,000 in 1977 to almost 25,000 units in 1988.

The econometric analysis attempted to correlate unit shipments to an index of the economy. Prior work at Garrett has shown,that general aviation shipments and billings correlate with pre-tax cor- porate profits. A formula was derived to predict ~nit shipments as a function of pre-tax corporate profits for 1955 through 1976. The corre13tion of GAMA historical data and unit shipments predicted from pee-tax corporate profits is shown in Figure 6. The degree of correlation or "goodness of fit" was not sati~factory at an r (1) ti)

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20,000 TOTAL UNITS I """""'" I"""""'" SINGLE ENGINE PISTON

i

10,000 4.3%/YR l- en , t::

-

:J u.

TWIN ENGINE PISTON I- 4.4%/YR a: w CD :E :J

- i"""'" i

~ ~ TURBOPROP I

---

~ , -""".,..

9.2%/YR

..-

~

III Ii _L_ ~-

-- - 77 78 79 80 81 82 83 84 85 88 81 88 YEAR Figure 5. Forecast of Aircraft Unit Shipments by Market Segment.

---

TOTAL UNITS SHI?PED 2,000 ..

,/

~

... ~

'00

Z w -~V" a::

Vr-

a:: U.s. PRETAX CORPORATE PROFITS :J ~ CJ u. ~ U.s. DEPT. OF COMMERCE CI) BUREAU OF ECONOMK; ANALYSIS z Q ..J ..J CD

_1

L ___ l_~

55 57 59 61 63 65 67 69 71 73 75 YEAR Figure 6. Aircraft Unit Shipments and Pretax Corporate Profits.

.... - --_ .. _----- .. _-- ~---. --------

of 0.69. However, the correlation between pre-tax corporate profits and manufacturer's net billings was very good (r = 0.944).

These net billings, however, included turbofan and turbojet busi- neS~i aircraft sales, which use engines outside the size class studied in GATE. A procedure was developed for removing the con- tribution of turbofan and turbojet aircraft based on average unit prices for each segment, units shipped by segment, net billings by segment for recent years, and total year11 net billings for 1955 through 1976. The results of this analysis are shown in Figure 7.

The solid curve shows total ~anufacturer net billings derived from historical data. The dashed line was predicted from the equation developed by regression analysis. The correlation factor, r2, is 0.937.

A forecast of manufacturer net billings was der ived in two ways. The first method was to project manuf3cturer net billings on the basis of pre-tax corporate profits. A fore~ast of pre-tax cor- porate profits to 1986 is availaole from Chase Econometrics. It was extrapolated to 1988 for the study. The results of using pre-tax corporate profits and the correlation shown in Figure 7 is shown as the broken line in Figure 8.

The second approach forecasted net billings by market segment.

This forecast was derived by multiplying the unit shipments fore- cast by the average unit price of each segment. The average unit pr ice was based on 1976 pr ices and inflated by a correlation between average price and the GNP deflator. The GNP deflator was forecast by Chase Econometrics to 1986.

tit ......

z I ~ w 800 JIII"" a: ~,~

sao

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,7

/ ~

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FITTED TOTAL NET BILLING~ ... ~~

en

l7

Z

r 2

r -.937 .~ S! . ~~ ZOO ..I ..I ~TOTAL ACTUAL NET BILLINGS :::E ~ -.

IN MILLIONS ./

1---... .A

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10017- I

80 . ~ ,7 ....

V

- -

Z 'f' "'-"'" w ~ GO a: :.~ a: ~

-......./'

CJ U.s. PRETAX CORPORATE PROFITS u.

o IN BILLIONS en Z o U.s. DEPT. OF COMMERCE ..I

I BUREjU OF jCONOjlC ANjLYSIS

:!

al 10L 55 57 59 61 63 65 67 69 71 73 75 YEAR Figure 7. General Aviation Manufacturer Net ~illings and 'Pretax Corporate Profits (Piston and Turboprop) · $4 LiP •• 4$ i". C. .i iL2 : 51 .au 4 at .4 4,000 i~....--r--"---r--'r-"'T" -r-.,..-...,..--,.---r--, 3,000 en 2,000 a: < -I -I o o 1 000 J---bo.ooF I I ,..""'- .Jl<""::J

... '

e!

800 I I.' I,.f".:aor IA 1 w a: a: GOG ::l c.J u..

o 400 I 1«"'"='" en Z o ::J -I :!: (UNIT TREND FORECAST) (FORECASTED AVG. UNIT PRICE) YEAR Figure 8. Forecast of Manufacturer Net Billings by Market Segment 2S .w 'pW wsw ,_ zw=sr * I' The results of this approach are shown in Figure 8 also, by segment and the summation of the segments. The difference in total billings between the econometric method (pre-tax corporate profits) and the trend method ranges from 100 to 400 million dollars. The difference can be attributed to the method used or can be looked on as a growth potential in the market not predicted by trend anal- ysis.

The remainder of the GATE sturly is based on the lower fore- cast, i.e., the unit-trend/average-unit-price forecast. This fore- cast projects manufacturer net billings of over 3 billion dollars by 1988 (then year dollars).

Forecasting by market segment also allowed an estimate to be made of the market segments in 1988. Table 2 shows the breakdown of unit shipments and billings for five selected years. In terms of unit shipments, current turboprops increase Sllghtly at the expense of the single-engine category. The breakdown of billings changes drastically. Current turboprops in 1988 will account for the largest percent of the market in terms of billings.

3.1.2 Marke~_forecast - Rotary-Wing Aircraft The rotary-wing aircraft market forecast was furnished by the Bell Helicopter Company. Unit shipments of light [under 4,540 kg (10,000 lbs) gross weight] civilian helicopters from 1963 to 1976 are shown in Figure 9. The market share for single-engine turbines, twin turbines, and piston engine aircraft is shown in addition to total shipments. Total shipments in 1976 were ovec 1000 units.

The forecast through 1988, without considering the impact of a GATE program, is shown in Figure 10. In 1988, more twin turbines will TABLE 2. FIXED-WING MARKET SEGMENTATION.

1965 1970 1976 1981 1988 Units Single 84% 92% 84% 82% 80% Twin 15 16 14 15 15 Current Turbines 1 2 2 3 5 Billings Single 44% 38% 39% 36% 30% Twin 48 45 39 40 33 Current Turbines 8 16 22 23 37 I ~---- -- tv .....

N CD I I 1800 f- en TOTAL I- ;v

Z 1000

::>

/

~ ___ SIN( LE TURBINE .- --.

~

-- -_/

~ " ..

" .....

--

/

.. '

-~ , PISTON

~--~7-

f-a _________ ~ ......

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............. -

--

---,.---

--

2QO V ,

---

-n

~l~RB'NE ..

/.'

- --- o

----

--

63 64 65 66 67 68 69 70 71 72 73 74 75 76 77 YEAR (19_) Figure 9. !listc.!; ical Light Helicopter Sales.

TOTAL (I) !:: 1000 z ::> TU!!.§INE , .. ~.

~

--

--,---

--- -_~ .... t.

,-~' ........... -I--~---

---+"'!!'

-,------

200 -, o 75 85 86 87 88 89 70 79 80 81 82 83 84 76 77 YEAR (19_) N I.;) Figure 10. Light Helicopter Sales Forecast.

"

be delivered than either single turbine or piston engine aircraft.

The 1988 forecast for total sh ipments is almost 2000 uni ts per year.

The rotary-wing market is different from the fixed-wing market in that the conversion to turbine power is well on its way. Rotary- wing aircraft with turbine engines of less than 373 kw (500 hp) are common and represent a majority of the market. Tables 62 through 64 in Appendix I list the engine type, power, and the 1977 average equipped price of aircraft in the current light rotary-wing market.

In the fixed-wing market forecast with GATE, it was assumed that the turbine engine would be used on all applications where it was superior to the piston engine and was cost competitive. This assumption was not made in the rotary-wing forecast. Bell assumed an introduction of the GATE Engine in 1987 and forecast that por- tion of the market where it would be used. This forecast is shown :n Figure 11. The forecast accounts for the continued, though declining, production of piston and older technology turbine- powered aircraft. For later use in engine cost estimates, the year 1992 was chosen to arrive at the GATE engine potential production for helicopters. Gate-powered unit shipments in 1992 include: o 400 Singles (400 Engines) o 280 Multi-Engine - 190 Twins (380 Engines) 90 Tri-Engine (270 Engines) ,

--flleC1':I'r7/~C;f~%({~{/@ GATE

~l«/LU$.~

. TWIN :TURBINE

(I) I- z ::l SING~E TURBINE 88 89 90 91 86 87 92 93 94 YEAR (19--1 Figure 11. Forecast \'lith GATE Engine.

, The total GATE engine production is 1050 engines. The GATE engine size recommended by Bell is 261 ~56 kw (350 ~75 shp.) The criteria that the GATE ~ngine would have to meet to realize the forecasted production are shown in Figure 12 and were suggested by Bell.

3.2. Engine Conceptual Design The conceptual design of GATE engines was undertaken in Task I to allow preliminary engine cost targets to be set and to provide engine performance and size data to Cessna and Bell to enable them to define aircraft characteristics. Conceptual design focused on three types of engines: o Turboprop o Turboshaft o Turbofan The turboprop is shown in Figure 13 and consists of a single-stage centrifugal compressor developing a pressure ratio of 9:1, a reverse-flow annular burner, a single-stage cooled radial turbine, and a two-stage uncooled power turbine. Performance and cycle chararteristics are shown in Table 3.

The baseline turboshaft engine was the same as the baseline turboprop except that the gearbox was eliminated. The turbofan engine, shown in Figure 14, uses the same core and low-pressure (LP) turbine as the turboprop and incorporates a geared fan, which produces a pressure ratio of 1.5:1. Performance and cycle charac- teristics are shown in Table 4.

A compar ison of the turboprop and turbofan at the cruise design point selected for the conceptual design--6096m, 389 km/hr-- (20,000 ft, 210 kts) showed a significant advantage for the turbo- prop. Based on an assumed propeller efficiency of 0.85 and equal , ENGINE NOISE (iNTERNAL) (85 dBA) (EXTERNAL) (73-dBA) ($20,000 TO OPERATING COST $35,000) (S10/HR) SPECIFIC

.« 82 KG)

GATE FUEL

1------- WEIGHT « 180 LB)

ENGINE CONSUMPTION CJ.30 TO 0.37 KG/HR/KW (0.50 TO O.~ LB/HP iHR) HORSEPOWER (TAKE-OFF) 261 ±58 KW (350 175 SHP) EMISSIONS CONTROL TIME BETWEEfoJ (3000 TO MAINTAINABILITY OVERHAULS 5000 HOURS) RELIABILITY e Figure 12. GATE Engine Criteria - R~tary-Wing Applications.

w w ~.

": e_ •• ~; .

":i··.?- -:,'/ '.

/, .. -.- TABLE 3. TPE CONCEPTUAL DESIGN Standard Day Uninstalled Parameter Value Shaft Power SLS, T.O. 373 kw (SOO shp) 6096m (20,000 ft), 231 kw (310 shp) 389 km/hr (210 kts), max cruise Shaft Specific Fuel Consumption SLS, T.O. 0.0295 kg/hr/kw (0.484 lb/hr/hp) 6096m (20,000 ftl, 0.0283 kg/hr/kw (0.465 lb/hr/hp) 389 km/hr (~lO kts), m"x cruise Cycle Characteristics, G096m (20,000 ftl 389 km/hr (210 kts), max cruise Corrected Airflow, 1.20 kg/sec (2.87 lb/sec) Compressor Pressure ~atio 9:1 Turbine Inlet Temperature 1';78°K (2200 F) Nozzle Pressure Ratio 1. 01 Weight,* 123 kg (271 Ib) *Including gearbox.

TABLL~. TFE CONCEPTUAL DESIGN Perforruancc an~ Cycle Characteri~tics Parameter Value Net Thrust 3275 N SLS, T.O.

(736 lb) 6096m (20,000 ft), 389 kra/hr (210 i<ts), max cnase 1406 N (316 Ib) Thrust Specific Fuel Consumption SLS, 'i'.O. 0.041 kg/N-h (0.404 Ib/hr/lb) !210 ktsj, rr,ax cruise 6096m (20,000 ttl, 389 km/hr 0.060 kg/N-h (0.586 lb/hr/lb) C1cle Characteristics, 6096m (20,000 ttl, 389 km/hr (210 kt~), max cruize Inlet Corrected Airflow 16.21 kg/sec (35.7 lb/sec) Compressor Corrected Flow 1. 30 kg/f ec (2.87 1b/sec) Fan Pressure Ratio 1. 5:1 COll,pressor Pressure katio 9.0:1 Turbine Inlet Temperature 1478°K (2200°F) Bypass Ratio 8.0:1 \·lei<jht 84 kg (185 Ib) - - -- - - --- -- - W -.J core size, the turboprop produces more net thrust at a lower thrust specific fuel consumption (fuel flow/propeller thrust) as sho~n in Table 5.

For equal thrust, agai n assuming 0.85 propeller eff iciency, the turbofan would require a 30-percent greater core flow. If cruise speeds are greater than approximately 500 km/hr (270 knots), the advantag~ of the turboprop diminishes. However, the turboprop retains a fuel consumption adv~ntage and would probably contribute to improved field performance.

3.3 Definition of Gas Turbine Powe~ Classes A preliminary estimate of turbine engine size typical of each general aviation fixed-wing category was required for an ~ppraisal to be made of the suitab~l.ity cf turbine engines. Current air- planes within each of the 10 general aviation fixed-wing catc90ries idcntified earlier can be segregated by engine power class. Engine power class includes the effects of turbocharging, i.e., a 224 kw (300 hp), naturally aspirated engine is in a different power class than a 224 kw (300 hp) turbocharged engine. Therefore, there are different turbine power classes for each airplane category.

It is not rigorous to generalize concerning the correlation between piston engine power required and turbine engine power required. The relationship depends on: o Engine sizing point, e.g., cruise or takeoff o Degree of turbocharging o Turbine engine cycle o Airframe/engine integration To determine turbine power requirements precisely would require a detailed study of each application. However, it is possible to TABLE S. 'i'""llRBOPROP A~D TURBOFAl~ COMPARISON 60S6m (20,000 Feet) 339 k~/hr (2iO k~ots) Maximum power Equal Core Size 0.85 Pro?el1er Lfficlcncy Turboprop Turbofan Corrected Core Flow 1.30 kg/sec (2.87 Ib/sec) 1.30 kg/sec (2.87 Ib/sec) Shaft Power 231 kw (310 shi?)

1815 N (408 11.:1) Net Thrust 1406 N (316 Ib) TSFC 0.036 kg/N-hr (0.3S:! Ib/hr/lb) 0.060 kg/N-hr (0. S86 Ib/hr/lb) -- ----- - w \D

I

generalize sufficiently to allow screening of the various car-di- oates and plck those where turbine engines offer potential.

Figure 15 shows a typical altitude lapse rate for a turoo- chargea reciprocating engine. A maxtmum power of 231 kw(310 np) was arbitr~rily selected. The performance is typical of all fllght speeos. A variation of power with flight speed actually does occur but it is small and depenaent on intake design ana throttle set- ting.

The critical altitude of the englne was selected to be 6096 m (20,000 tt). To match the 231 kw (310 hp) reciprocating engine at 6096 m (20,000 ft), a turbo?rop engine has to provide 350 kw (470 illJ) at se<l-level, static, (SLS), standarci day, maximum power. The :lasned lllie, intersecting 231 kw (310 hp) at 6096 m (20,000 ft) is the turboprop lapse rate at 37ll kra/hr (200 kts) flight speed. At 370 km/hr (200 kts) at sea leyel, the engine produces 402.7 kw (540 hp). 'fne lS-percent increase i:1 t-0\'ler ..,etween 0 and 370 km/hr (0 ane 200 kt~) ~etcrmines the po~er :~ sea-level static, i.e., 351 kw (470 nlJ). It a turooprop is bizeu In ~hl5 manner, it provides equal or higher cruise power at all altituaes and higher takeoff power.

A slightll aifferent situation exists when sizing a turooprop to repl~ce a naturally aspir~tea reciprocating engine. Figure 16 sno'lls a ty~ical altituae lapse rate for a naturally aspirated reci~rocating engine. The lower dashed line shows the altitude lu?se rate at 370 km/hr (200 KtS) of a turboprop sized to match reciprocating engine power at 3048 m (10,000 ft). The sea-level, static, maXlffium power ot the turboprop in this case is 189 kw (253 hp). This ib probably insufficient power to match takeoff perform- ance ot tne reci.procat:.ing-englne-powered aircraft. The altitude lapse rate at 370 km/hr (200 kts) of a turboprop sized to provide 231 kw (310 hpj at sea level, static, takeoff is shown by the upper I ~,

"

"'"

( , ......

~SLS TURBOPROP (SIZED FOR ~350 KW

- ~~ 350 KW (470 HP), SLS,MAX)

I >- (470 HP) , ......

1', I Q..

~ 260 ....

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I ri· 240 a: "' w w 1"0...

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" :no KM/HR

2 ~

~

200 I

i"... Y (200 KNOTS)

""TURBOCHARGED

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180 » RECIPROCATING ENGINE I'.... ............

160 231 KW (310 HP), SLS, MAX

ALL~'

200 .............

~ FLIGHT SPEEDS

""-

r---.......

9 10 11 5 6 7 8 0 2 3 4 PRESSURE ALTITUDE, M I I I I I I I I 0 5 10 15 20 25 30 35 PRESSURE ALTITUDE, FT ~ ......

Engine Power Requirements Figure 15.

.I_ I", 400- 300

I TURBOPROP, SLS, MAX, 231 KW (310 HP)

~~~~ (S/~€ I-L ... _-

2~~ ~/J~4 r 'r'O~ <'3,1

~:::.."....~....,..I(~ 1.1.., ~I(~ - ( 0 4S (370 3' P1 CL.

........ ........ liP). :4 .,..~O ~ I!:P), SI.S

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a:: 150 - ~k±-- I~ I -t-I --1----1

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189 KW (253 HP)

100 I I . I· I I-'---~-----

~ 370 KM/HR (200 KTS) TURBOPROP (SIZED FOR .

189 KW (253 HP), SLS,MAX) -

I

__ -+ __ - ALL FLIGHT SPEEDS_ W 1~--~~------~---4 0,,- 0 1 2 3 4 5 6 8 9 10 11 PRESSURE ALTITUDE, M I I I I I I I I 0 5 10 15 20 25 30 PRESSURE ALTITUDE, FT Figure 16. ~ngine Power Requirements da~hed line of Figure l~. Here it is apparent that the turboprop will have a higher cruise _~wer for equal sea-level, static, takeoff power.

These two examples provided the correlation between the required power for reciprocating and turbine engines. For turbo- charged engines, the turboprop must be 50-percent larger at sea- level, static, takeoff power and the turboprop must provide the same take-off power as a naturally aspirated engine. Table 6 shows the resulting equivalent turbine power for current power classes in each of the 10 general aviation categories. In some cases, a range is given to account for possible future changes in mission perform- ance. It is emphasized that the correlation is only approximate and was accomplished solely to allow screening and selection of candidates for Task II, Trade-Off Studies.

As mentioned earlie" the rotary-wing engine size recommended by Bell is 280 ±56 kw (375 ±75 shp).

3.4 Scr~enin9 and Selection The objective of this element of the market survey was to identify the domain of super ior i ty of the var ious eng ine types, particularly turbine engines.

Screening was limited to the 112 to 447 kw (150 to 600 hp) size class. Engines producing more than 447 kw (600 shp) were not screened because turbine engines are universally used in general aviation applications in this size class because of their superi- ority and the lack of competition from other types of propulsion systems. Also, U.S. engine manufacturers are heavily committed to the 447 to 746 kw (600 to 1000 shp) turboprop and turboshaft mar- ket, and will continue to dev~lop the technology =equired for its growth. Finally, the U.S. Army's program to develop a demonstrator engine in the 447 to 746 kw (600 to 1000 shp) class will provide " TABLE 6. FIXED-WING TURBOPROP POWER CLASSES.

Equivalent Current Turbine Power Power Airplane Category Class Class kw hp kw hp - 2-Place 75 100 75 1 . ...

.1.~ 150 112 150 utility 112 150 112 149 200 149 200 224 300 224-280 300-375 2' • ~'t 300 TC 366-410 450-550 r'ixcci GC.:lr lIigh 186 250 168-205 225-275 Pert:ormancc 224 300 224-280 300-375 224 300 TC 336-410 450-550 4-Place 112 150 112 150 149 200 149 200 Llyht Rctract~bles 149 ~OO 149 200 149 224 200 'rc 300 heavy Retract~ole~ lU6 ::50 168-205 225-275 224 300 224-280 300-375 2~4 300 TC 335-410 450-550 i\yr icultural 186 250 168-205 225-275 i : .

224 300 224-280 300-375 336 450 336-410 450-550 447 600 447 Li~ht Twin 149 200 149.1 200 IG6 250 168-205 225-275 224 300 224-280 300-375 149 200 TC 224 300 2;';4 300 'rc 336-373 450-500 C~bin Class Twin 224 300 TC 336-410 450-550 298 400 TC 410-485 550-650 Pressurized Twin 224 300 TC 373 500 298 400 TC 447 ----- ._- - ___ L..- TC - Turbocharged \ much of the required future technology advancements. GATE tech- nology development effort should be focused on the under 447 kw (600 shp) size class since the larger engine technolo~y 447 to 746 kw (600 to 1000 shp) being developed by industry and the Army is not universally applicable to smaller engines. The under 447 kw (600 shp) class requires a primary emphasis on engine cost, which cannot be compromised for performance or weight.

Screening of candidat~ turbine engines was accomplished pri- marily on the basis of engine cost. Previous studies (Ref. 1, 2, and 3) had shown that performance and operating cost of gas tur- bines could ~e competitive with reciprocating engines but that the compar ison must be made on a system basis, i.e., airplane and engine. This comparison is part of Task II. A method was derived \/hich a:"lowed a preliminary assessment of the feasibili ty of gas turbine engines with respect to engine cost and its effect on air- plane cost.

The method derived required that target costs be established for advanc~d GATE Engines and allowable turbine engine costs be established for each ai rplane category. The comparison Of the target costs and allowable costs will show those categories ~here turbine engines can compete.

Allowable turbine engine costs need definition because turbine enqines can cost more than reciprocating engines and remain compet- itive for the following reasons: o Based on earlier market survey results, gas-turbine- powered aircraft may command a la-percent or greater premium o Lower engine weight and decreased vibration and noise will result in lighter, less expensive airfram~s.

,

~~

~~1

~~

~/~

..- ta~

The lO-percent premium, a figure based on the judgment of ~r· AiResearcn and Cessna marketing personnel, is justified because of the recognized superiority of turbines in the following areas: o Lower interior noise and vibration o Higher reliability and safety o Improvea takeoff/altitude/speed performance Pr ior studies have shown that lower engine weight and decreased vibration and noise can result in airframe weight savings of 10 percent or more.

An aaditional factor that was considered in developing the allowable engine cost was the potential increase in reciprocating engine cost, becuuse of technology advancements for improved per- tcrmance ana durabili ty, lower we ight, tip-creased vibratIon and noise, ~nd lower emissions. Subsequent to completing this portion at the <..i/\T1:: stuay, the EPA pul;lisncc. their intent to remove all emission requirements for small engin~s. Study results were not moditied to retlect this and can oe viewe~ as a necessary adjust- ment or a provision for future regulatory action.

Target Turbine Engine Orisinal Equipment Manufacturer's 3.-1.1 (D.E.M.) Cost The conceptual design studies indicated that turboprops are superior to turbotans in tile aircraft categories being studied, in terms at fuel consumption ana required engine size. This finding was not by any means based on a detailea ana rigorous analysis.

However, it suggested that engine cost screening could be done on the basis of the turboprop engine for fixed-wing aircraft. Turbo- Luns may offer lower system cost than a turbo?rop engine plus pro- peller (or a given core size but turbofans will r~uire a lar~er core.

Screening for helicopter applications also follows the devel- opment of target costs for the turboprop. Turboshaft engine cost for helicopters should be lower than the turboprop cost at equal power, due to the elimination of gearbox cost.

Signif icant potential cost improvements were identified fer GATE turboprops, relative to current-technology turboprops. Pro- jections for 1983 component and manufacturing technology indicate improvements in performance that result in lower cost, and new fab- rication techniques that promise dramatic decreases in labor and material requirements. The GATE turboprops can afford lower power- to-we ight ratios than current turbine engines, and on a r"!lati v.e basis can have a lower quality cycle than larger engines of comparable technology. This flexibility in weight and performance is the basis for a successful Design-to-Cost (OTC) program. Many OTC programs are ineffective because little flexibility is allowed due to hard requirements for high performance and low weight.

I Another major factor in cost improvement is the high volume production typical of the general aviation market segment being stud ied. The potential for large production releases, automated machining, and dedicated equipment offers significant cost reduc- tions. Based on the above factors, GATE turboprop target costs were established as shown in Figure 17. The production quantities associated with these target costs are shown in Table 7. The data assumes the cost benef i t associated wi th these high-production levels.

A 90 percent learning curve is assumed.

The variation in production quantity and specific cost with power is a result of matching target and allowable engine cost, and is an itera~ive process. Target costs were initially based on a constant production volume. As the complrison between target and allowable engine cost was completed, estimates of production volume were made for those applications where the target cost was equal to or lower than allowable cost. Additional discussion of these pro- duction quantities is contained in subsequent paragraphs.

1,-\ I I I I CURRENT C1.

TURBOPROPS J: 10, - 80 ........

~ ~ ~ ~ w ~ ~ u.

u'

w 55 ~ 361 I ....... uaaI""""" CURRENT RECIPROCATING ENGINES 19~1--~----r---~---r---+----~--+---~

o~, ____ ~ ____ ~ ____ ~ ____ ~ ____ ~ __ ~~ __ ~ ____ ~

200 ~OO 400 500 600 700 800

o

kw I ___ ~ ___ ~.~ 600 . 800 1000

o 200 400

HP Figure 17. Estimated Engine O.E.M. Cost 1977 Dollars TABLE 7. ENGINE PRODUCTION QUANTITIES - TARGET COST.

Power I Hp Annual Production Quantity Kw 200 1250 ~ 186 250 2000 261 350 6200 373 2100 447 600 Another source of variation in the specific cost versus power relationship shown in Figure 17 is the exponential scaling law issued by AiResearch for engine cost, namely~ Power 0.75

Engine cost = (c ..... ~ c"' .......... ) (Base Engine Cost)

As engines are scaled down, the specific cost increases, assuming that cycle quality and engine configuration remain the same.

Also shown in Figure 17 are the Original Equipment Manufac- turer's (OEf.I) specif ic cost for current turboprops and an estimate of the specific OEM cost of current reciprocating engines. The GATE turboprop cost target represents a cost reduction of over 50 percent when compared to current production turboprops. Compared to the cost of reciprocating engines, the GATE turboprops are 25 to 100 percent higher. On a specific cost basis, turbines will prob- ably be high~r than reciprocating engines until common cores, high parts commona,lity, and product maturity of gas turbines increase to levels comparable to reciprocating en~ines.

3.4.2 Allowable Turbine Engine Cost Given the difference in specific cost between the turbine and reciplocating engines, can the higher cost of turbines be justified and absorbed such that turbine-powered aircraft price is competi- tive with reciprocating-engine-powered aircraft ptice? To answer this question, it was necessary to determine the t~rbine engine OEM cost which would allow a competitive situation between gas turbines and reciprocating e~gines.

A simple procedure was developed to determine the allowable turbine engine OEM cost. In this procedure, the current aircraft dealer cost is first adjusted to reflect a 20-percent increase in current reciprocating engine OEM cost, to allow for reciprocating engine technology.

ADJUSTED CURRENT RECIP) (RECIP ) CURRENT ENGINE COST DEALER

=

DEALER + OEM INCREASE

~~KUP)

COST ( COST COST Current dealer cost - Factory price with standard equipment Reciprocating engine OEM cost Estimated on basis of Figure 17 Reciprocating cost increases - 0.20 selected for increa~es due to noise, emissions, and advanced technology OEM markup - Airframe markup factor on engine cost for direct and indirect cost, overhead, and profit (Factors over 2.0 were s~ggested. A factor of 1.5 was selected. The lower factor is conservative.)

I For .:l single-engine airplane with a current dealer's cost of $36,000 and OEl-I engine cost of $5,000, the increase in current dea 1 cr·.3. cost due to technology improvements in the reciprocating engine would be:

($5000) (0.2) (1.5) = $1500

The adjusted current dealer's cost is: $36000 + 1500 = $37500 The second step adjusts the airframe cost to reflect the lower gas turbine engine weight and decreased airframe weight due to lower noise and vibration.

Sl

AIRFRN-IE CURRENT) (RECIP) ~ )] AIRFRAME

AIRCRAFT ENGINE OEH COST COST = DEALER - OE!-I HARKUP X REDUCTION [( nj/TURBINE) COST COST FACTOR In this procedure, the airframe cost is the cost of the air- ~lane less enqine. The airframe cost reduction factor was assumed to be O~90 or a lO-percent reduction in cost for turbine enqines.

A new airplane cost or adjusted dealer cost with turbines is computed b~sed on increasing the adjusted dealer cost by 10 per- cent, which is the assumed premium for turbine power.

D~Gcd on these three steps, the allowable turbine engine cost may be computed: l'.l HCR/WT DE/\LER AIRFRAHE ) ( ) (OEM ) = COST . • ALLOWABLE TURBINE ENGINE COST ( (HITII TURBINES) (U/TURBI~ESl ENGINE COST MARKUP or AIRFRAME AIRCRAFT DEALER COST ( ALLOWABLE TURBINE _ COST (W/TURBINES) \o,J/TURBINES) - 1.5 ENGINE OEM COST The ~11nwable turbine engine cost must be divided by two for twin- engine aircraft.

1\ specific example of this procedure is shown in Table 8. The reciprocating engine cost was obtained from Figure 17 for a 231 ~w (310 hp) engine. The current dealer cost is an average of all models in the light-twin cat~gcry.

'l',\BLE a. ALLG;'iABi..E ~URBINE COS'l' EXAl·lPLC

Light '1'\;in

Reciprocating Engine Cost = 8550 (17,100/(2) ~n£ines)

Turbine Engine Prerniuffi = 10%

Current Dealer Aircraft Co~t 136,496

=

Aojuste~ Current Dealer -

Aircraft Cost = 136,496 + 0.2 (17,100) (1.5) = 141,626

Dealer Aircraft Cost with Turbine Engines ('l'EDC) ~ (1..:) (141,626)

= 155,789

Airframe Cost (AFe) : (~.)~,"96 - (17,100) (1.5)] 99,761

0.9 =

, I i Allowable Turbine Engine I Cost (2 engines) - i'.EDC. -Al'C) /1. 5 37,352

=

I

Allowable Turbine ~n~ine Cost (each) 18,075

=

I.

I

I I I j

I

~ Vl W

.. ~---.~-- .. ------------

Tables 9, la, and 11 list the data required to calculate the allowable turbine engine cost for each power class in all general aviation categor ies. The turbine power classes listed were dis- cussed earlier. The average dealer cost is a unit shipment weighted aver~ge of the 1977 average dealer cost for every model.

rhe current OEM engine cost is the 1977 cost to the airframe manu- facturer for presently used reciprocating engines and was estimated by the cost/kilowatt relationship shown in Figure 17. The 1977 unit slllpments were estimated in mid-1977 from available data for every ~oael and were totaled by category. Final 1977 shipment dat~ was conservative by approximately 10 percent. The 1988 unit ship- ments are ~rojected from the 1977 shipments using the growth rates previously uetined for single- and twin-engine aircraft. Since the forecasted trends were made tor the general groupings of single-and twin-engine aircraft, proJections bi power class for each of the more specific categories are only approximate.

~he data shown in Tables 9 and 10 was used to calculate the allowable turbine engine cost and the results were grouped by power ' class. These resllits are shown in Table 11. The allowable turbine engine cost assuming a la-percent premium for turbine power and the 1988 annual production is shown. In addition, the cumulative pro- auction for ea~h power class is shown. The total figures for both the single- and twin-engine categor ies differ slightly from the forecasts shown earlier. PreVlOUS data was based on G~\A data for the single and twin categories. The data shown in Tables 9 and 10 are based on forecasts :or each manufacturer's model.

The results of the comparison between allowable and estimated engine cost is shown in Figure 13. This figure shows the GATE tur- ooprop target cost, and the range of allowable engine cost in vari- ous power classes is super imposed. All categor ies in two power classes, 224 to 280 kw (30e to 375 hp) and 410 to 485 kw (550 to 650 hp), have allowble turbine engine costs that are greater than the TABLE 9. SINGJ ·I;;-ENGINE FIXED-WING HARKET Turbine Average Current* Airplane Power Dealer OEM Engine 1977 1988 Category Class Cost Cost Shipments Shipments (1977 $) (1~77 ~/) kw Chp) ~~ 2 Place 75 13,223 100 3700 1864 2964 112 150 18,875 4500 424 674 Utili ty 112 150 17,300 4~00 150 239 200 19,700 5500 70 111 168-205 225-275 32,400 6750 149 237 224-280 300-375 38,563 8550 349 555 336-410 450-550 53,500 8990 30 48 Fixea GecH 168-205 225-275 33,573 6750 1071 1703 High Performance 224-280 300-375 43,843 8550 460 731 336-410 450-550 47,900 8990 240 382 4 Place 112 150 20,525 4500 2833 4504 149 200 26,629 5500 1260 2003 Light retractables 149 200} 39,331 5500 788 1253 224-280 300-375 49,600 5900 100 rueavy retractables 168-205 225-275 53,800 6750 170 270 224-280 300-375 59,759 8550 1142 1816 336-410 450-550 58,959 8990 358 56~ AgrIcultural 168-205 225-275 30,500 6750 400 636 224-280 300-375 40,942 8550 418 665 336-410 450-550 52,400 250 398

--

447 600 59,600 84 134

--

Total single engine aircraft 12,610 20,051 *Specific Cost Estimate U1 Ul U 0> 'i'i\3LE 10. ri;I!;-i:.I;GIl:~, I:'r;.;LD-\Ht~G HARKe'!'

I Iwcr <:9c Current* I 'furtane Dealer G!~H Ens inc Airplc111!! POhcr Cost Co~t (1::.:::) 1977 1988 c..1c;.s!: Cilte;orr (.:.977 $; (E77 $) S 10 i t-'11.cn ts Shipmer.tz >-- kw (1:2) Light t\,iins 168-205 2:5-275 112,037 6750 344 554 224-280 :)OD-375 136,496 8550 555 894 336-410 450-550 157,692 8990 127 204 Ciltin cl<l!:>s 33b-41U (50-550 1!33,978 8990 91 147 (llnprcssurizec) 410-465 550-6:)0 253,(;00 14,600 200 322 PrCSSUrlZeG twins 336-410 450-550 199,048 8990 439 707 410-485 550-650 319,90 ..

14,600 257 414 Total twin cngi~~ aircraft 2013 32~2 'fot .. l en9ines 4026 64b4 I ---- --- *~pecitic Cost Estimate , 'rABt.E l.I.. .;z.:'O;;AB!.E TURBINE ENGINE COST (1977 $).

-------- Allowable 1988 Annual Production Turbine Engine Cost Aircraft CUMulative 10\ Premium, S Category Category 112 ~i1owatt (150 hpl Class FOllr-Pl;lce 7777 4504 4'l04 Two-P13ce 674 5178 2'~ Utilitv 7H7 ~7 5417 149 K llowat t (200 hpl Class l!l 3367

~

Fcur-Pi~-:(' 9710 Z003 3256 Light Rct:;lct~hle 11404 1253 1253 168-205 Kilowatt 1225-275 hpl Class Aqr iC'~1 t UP 1 636 636 lSO::!9 1108 !.ight T"d~ 1144 14733 "~~~¥ Rctr3cta~le 270 2014 Fixed Gear Hiqh Perf. 12036 1703 I 11880 Unlit\· 237 39';4

I

224-280 Kilowatt (300-375 hpl Class t(~t3r:' \'"inc; 10SO ;., ~ 0 .'ar iC'llltU~,;l 665

I

18675 Li<:t'.'. Twin 1788 3503 Hcavv Retr3ct3ble 17544 1816 16189 t.lc:ht Retractable 159 :'478 i..

Fixed Gear High Perf. 15';21 731 1;209 t.:ti lity In18 555 676'; I 336-410 Kilowatt (450-550 hp) Class , A'lricultural 398 398 Pressurized'T'\."in ::!3319 141'; '812 C3bin C13s~ Twin 22314 294 2U6 Light Twin ::!0582 408

I 25H

3eavv Retr3ct3hle 17930 569 3(183 I Fixed Gear High Perf. 17202 382 3H5 i:tility 16456 J9 3513

I

i

I

410-485 Ki!owatt (550-650 hpl Cl~ss Ac:ricultural 134 134 Pressuri~ed Twin 37678 828 Cabin Clas~ Twin 33258 644 1601;

J~ ___

I ! - - I

o A~~ur.cs helicopters and ;lgriculturJl ~ircraft use turbine engines when avai!able o 1968 ~ronuction based on !orccasteo grcwth '4.3% single: 4.4\ :win~l 5; , • SEA LEVEL STATIC. STANDARD DAY. TAKEOFF • 1971 ('lnLLAR~ • PREMIUM FOR TURBINE POWER -10 PERCENT 0..

:I: ......

== fh ~ 107 fh t; 224-280 KW t- Il) GATE (,) (300-375 HPI -\ .S~ I I I Y-TURBOPROP----i :E w

I TARGET

:: IoU (,)

72 IIIW~~410lw7 K~

u. (,) (550-650 HP) (,) u.

W u 0..

II) w

+--+-1 1H-'1-~336-373 kw -+---1

0..

II) (450-550 HP)

I

I ..... ~. I

-CURRENT RECIPROCATING ENGINES RANGE OF ALLOWABU~ TURB'INE

O

o I EN~INE C?ST I I , I I I o 100 200 300 400 500 600 700 800 POWER. KW I I I I I I o 200 400 600 800 1000 POWER. HP Figure 18. l::.stimatcG ~nsine O.B.tL Cu5t, 1977 [;o}.inr5.

GATE tnrget costs. Two power classes, 112 kw (150 hp) and 149 kw (200 hp) do not have any categories that have allowable turbine engine costs equal to GATE target costs. GATE target costs would have to be decreased an additional 15 to 25 percent before these categories would be attractive for turbine engine propulsion. The two remaining power categories, 168 to 205 kw (225 to 275 hp) and 336 to 410 kw (450 to 550 hp) have some categories where the allow- able costs exceed the target costs. Over 50 percent of the poten- tial pro(luction in these categor ies could be powered by turblne cnlJlnes. 'l'he potential turbine engine production for each power class is shown in l-'igure 19. In all cases, rotary-wi ng and ag r i- clll-tural applications are included because the results ot the market survey indicate the applications would use a turbine engine if it were available at the GATE target levels. It was assumed that gas turbines developed as a result of GATE would be used in lieu of reciprocating engines based on allowable cost. This approach docs not ac~ount Lor a retrofit market, ~or does it allow for a change in the nl..1rket growth rate as a ConsC(luenCe of the availability of GATE gas turbines. Immediate 100 percent penetration of the gas tur- bines in 1988 W..1S ..1150 assumed, i.e., there is no start-up perioa during which production gradually builds.

The analysis of projected versus allowable cost was performed in Task I. In Task II, detailed cost estimates and more precise determination at power requirements were made. In general, the Task II results showed that engine cost was slightly lower than the target ~nd required engine size was lower than estimated in Task I.

'l'herefore, the potential turbine demand and number of categor ies where g..!s turbines are competitive arc larger than predicted in Task I. Task I results as presented herein have not been updated based on the results of Task II.

B..1sed on the market survey results and particularly the case analysis, the applicatlons selectee tor study in Task II were: C· G 90DOr-

I D RECIPROCATING ENGINE PRODUCTION

a: <t

80DO r rrm TURBINE ENGINE

w TOTAL ~

I 'I..11JJ PRODUCTION (W/GATE)

CI) ENGINE !:: PRODUCTION z 7000, ~ z 6000- Q I- U ::> TURBINE 0 5000 ENGINE a: PRODUCTION Q..

WITH GATE L'J ...

. -

t!)

Z W -J <t ~ Z 2000 Z <t

-----

o I....... ",.. IV" lY 168·205 KW 224·280 KW 336-410 KW 410-485 KW (225·275 HP) (300·375 HP) (450-550 HP) (550-650 HP) POWER CLASS Figure l~. Potelltial Turbiue EI.gil.e Dema .. d (198d).

---_._----._ ....... .

o Pressurized Twin o Light Twin o Single-Engine Utility Helicopter The two fixed-wing applications chosen and the heavy retrac- tac table single-engine category had the highest potential produc- tion volume of all categories where turbine engines showed promise.

It is recommended that the heavy, retractable, single-engine candi- date be investigated in follow-on programs. The single-engine utility helicopter was chosen for study primarily because this seg- ment of the market is currently dominated by reciprocating-engine- powered helicopters.

3.4.3 Other Engines Considered The other types of engines that were considered in addition to turbines were: o Reciprocating engines Gasol i ne Diesel o Rotary engines Only current and advanced gasoline reciprocating engines were retained after initial screenitig. Available information on advanced diesel and rotary engines indicates that they are consid- ered potential rropulsion systems for future general aviation air- craft and offer advantages in performance, weight, and durability.

There is, however, very 1 i ttle specif ic information about their characteristics, cost, or how advancements will be made. A compar- ison including these engines would be desirable but without more specific data, a fair comparison cannot be made.

SECTION IV TASK II BROAD SCOPE TRADE-OFF STUDIES 4.0 OBJECTIVE The objective of this task was to determine the optimum engine for the aircraft applications chosen in Task I, the Market Survey.

The applications chosen were: o Pressurized Twin o Light Twin o Light Single-Engine Utility Helicopter The tasks performed to define and select the optimum engine were: o Selection of candidate engine configurations and appli- cable advanced technology o Baseline engine definition o Aircraft sizing and sensitivity studies o Engine trade-off studies o Benefit analysis r:~~ECEDiNG PAGE BlANK NOT FlLMEO 4.1 Selection of Candidate Engine Configurations and Candidate Advanced Technology 4.1.1 Candidate Engine Configurations The gas generator configurations selected for consideration during Task II are shown in Figures 20 and 21. Figure 20 shows gas generators compatible with free-turbine turboprops and turboshafts and two-spool turbofans. Figure 21 shows candidate single-shaft turboprop, turboshaft, and turbofan gas generators.

All the configurations shown in Figures 20 and 21 use reverse- flow annular combustors. Consideration was given to studying in- line, radial, and can-type combustors. They were eliminated because AiResearch has generally found that the reverse-flow annu- lar combustor is competitive with or superior to the alternate con- figurations in th~ 186 to 447 kw (250 to 600 hp) class and when I r~dial flow components are being used. The in-line combustor could be competitive with the reverse-flow combustor if turbine inlet temperatures considered exceeded 1589°K to 1644°K (2400 F to 2500°F). At turbine inlet temperatures higher than 1664°K (2500°F), cooling of the r~verse-flow annular transition section is difficult. For specific applications, the radial or can-type com- bustors may offer some cost advantages and acceptable performance.

However, they have a large e tfect on eng i ne envelope. The GATE e .gines must be compatible with a variety of aircraft and the envelope of engines with radial or can-type burners could restrict the number of applications and/or affect aircraft design and per- formance.

Gas generator configurations utilizing all-axial 9omprcssors were eliminated from consideration. For core flow of less than 5 pounds per second, prior experience has shown axial-centrifugal or centrifugal compressors to be superior. A front drive, concentric

CD

(2A) ..

n nnn I I I I I I I I I I I II I I I I I I

11t (6A)

~ (6A)

L l (5B)

• (5C) (60) C CI Figure 20. Candidate Free-Turbine Turboprops and Turboshafts and Two-Spool Turbofans ~ <..

o (9C) :

i

,

!

I Figure 21. Candidate Single-Shaft Turboprop and Turbofan Gas Generators

.j

shaft, low-cressure spool was the only configuration evaluated for the two-spool engines in Task II. This arrangement offers common- ality among the turbofan, turboprop and turboshaft configurations and does not require special installation considerations. Low- pressure spool arrangements were limited to one- and two-stage tur- bines and, in the case of turbofans, to single-stage fan designs.

4.1.2 Advanced Technology Thf' advanced technology considered for the GATE engines is listed for each of the gas gener~tor configurations in Figures 22 th;:ough 27.

4.1.2.1 Co~cressors Three types of com~res~ors were chosen for investigation, namely: o Sinqle-Stage Centrifu~al o Two-Stage Centrifugal o Axial-Centrifugal The single-stage centrifuga~ was evaluated over a pressure ratio range of 6 to 10. Mater ials and fabr ication processes evalu'lted were: o Cast Steel o Cast Titanium o Powder Mp.tal Titanium (PM Ti) o Powder Metal Titanium Aluminide o Machined Titanium The cast and powder metal approaches would allow use of sophis- ticated 3-D blading, while maintaining low cost.

cr· a:> TURBIN£ COMIU.lSTOB.

DifFUS£B COMl'B£SSOB o 1255-147B K (lBOO-22000F) (COOLED • ANNULAR REVERSE CAST STEEL

PIP = 6-10 •

•

•

ANO UNCOOLED) FLOW SHEET METAL CAST STEEL

•

• o

CAST SUPERALLOY BLADE RING • 1255-1478 K

•

PM Ti CAST TITANIUM

• + PM SUPERALLOY HUB

• naOO-22000F)

SINTERED PM PM TITANIUM INTEGRAL CASTING

•

• CeRAMIC

•

VANES BqAZED HI •

PM Ti·AI -Af2-IDA

• • THERMAL BARRIER

CAST END WALLS -IN192 + Hf MACHINED Ti COATINGS

•

DIE CAST STEEL

•

• PM SUPERALLOY NET SHAPE • ODS SHEET ALLOYS CLEARANCE

•

LAMINATED SUPERALLOY • PHOTOETCHEDI

CONTROL •

LAMINATED • THERMAL BARRIER COATING CONSTRUCTION • CERAMICS • AIR BLAST NOZZLES CLEARANCE CONTROL (ACTIVE

•

AND PASSIVE) .

INTEGRAL-CAST OS BLADES

•

EQUIP.XED HUB Figure 22. Gas Generator Configuration No. 1 Candidate Technology COMPREssOR DIFFUSER COMBUSTOR TURBINE ~ D.LADfS SEE FIGURE 22 SEE FIGURE 22 SEE FIGURE 22 • LAMINATED SHEET • INTEGRAL ALLOY CAST SUPERALLOY • DOS EXTRUSIONS BLADES + PM SUPER- ALLOY HUB • CERAMIC CAST OS BLADES EOUIAXEO HUB • SUPERALLOY - CAST + HIP CASTINGS LAMINATED • INr,IVIDUAL BLADES - OS BLADES - CONVENTIONAL CASTING • SINGLE CRYSTAL • CERAMIC Figure 23. Gas Generator Configuration No. 2 Candidate Technology \0

'"

....: c.

COMPRESSOR QIFFUSER COMBUSTOR TURBINE 1ST STAGE 2ND STAGE 1ST STAGE 2ND STAGE PIP = 2.5-5.0 • PIP = 3.0-5.0 DIE-CAST SEE FIGURE 22 SEE FIGURE 22 SEE FIGURE 22

•

ALUMINUM CAST STEEL SEE FIGU.1E 22

•

CAST TITANIUM

•

PM TITANIUM

•

PM Ti -AI

•

CAST ALUMINUM

•

MACHINED Ti

•

Figure 24 Gas Generator Configuratio~ No. 3 Candidate Technology TURBINE· COMBUSTOR DIFFUSER COMPRESSOF!

SEE FIGURE 23 SEE FIGURE 22 SEE FIGURE 24 SEE FIGURE 24 Figure 25. Gas Generator Configuration No.4 Candidate Technology ~ ......

-.I N rln r1

I11111

I H W

I H H

II111I

UuU

IURBINE COMBUSTOR DIFFUSER COMPRESSOR CENTRIFUGAL STAGE AXIAL STAGES SEE FIGURE 22 SEE FIGURE 22 SEE FIGURE 22 SEE FIGURE 22 • 1.2-1.4 P/P/STAGE • CAST INTEGRAL ROTORS AND ,- ::~ STATORS . ." w"- '~B Or.; • CAST INTEG RAL ,- ~.

.~ ;:..

COMPRESSOR 'r

~~

:;....;&,. • PM Ti ,,0 .... " t=1 • PM Ti-AI

:1-

.... en Figure 26. Gas Generator Configuration No. 5 Candidate Technology , ,

• •

•

-

nnn

I I I I I I

I I I I I I

I I I I I I

I I I I I I

LJLJU COMPRESSOR DIfFUSER COMBUSTOR TURBINE SEE FIGURE 26 SEE FIGURE 22 SEE FIGURE 22 SEE FIGURE 23 Figure 27. Gas Generator Configuration No. ~ Candidate Technology -.J W

"

..

•

•

-

o Per formance - Studies have shown that the powder metal approach would allow thinner blades and smaller radii and closer tolerances than casting approaches.

o Weight - The powder metal titanium approach would yield lower component weight as compared to cast steel designs.

Cast titanium configurations would be competitive with PM T i. Cast alumi num is not a cand idate for the single- stage pressure ratios of 6 to 10.

o Cost - The cost of the cast configur~tions should be less th~n the PM Ti approach.

o Risk - The PM Ti approach, par t icular ly for complex designs with a high number of blades and splitters, is considered high risk. The cast approach is lower risk but only if ~ lower performance level is accepted.

t-l.lchining the compressors is lower risk th.ln either the cast or PM 'l'i ~pproachcs but is very expensive (2 to 4 times) particularly when compound curvature is required.

The two-stage centr itu~al compressor was evaluated over a pressure-r~tio r.lnqc ot .lpproxim.ltely 8 to 16. Cast aluminum was considered tor the first stage in addition to the materials and processes considered for the sinqle-stage centrifugal compressor.

The ~x ial-ccntr i fugal .>mpressor was evaluated over a pressure-ratio r.lnge of 8 to 10. Candid.lte materials and manuf.lc- turiny approaches .lre: o C.)~t intcgral rotors .:md st.ltors (.lluminum, steel, and, tit .:10 i lim) o Cast integral compressor o Powder-meta~ titanium rotors o Powder-metal titanium aluminide rotors \ \ 4.1.2.2 Diffusers

\

, Four types of diffusers were considered for the GATE engines: , o Vane island

\

o Vane o Multi-vane o Pipe Trade-off studies included performance and cost. Materials and manufacturing processes included:.

o Cast steel and titanium o Powder metal (PM) titanium o Sheet metal construction o Sintered PM vanes brazed to cast or sheet metal side plates The fi rst-stage dlf fuser for the two-stage centr ifugal com- pressor is die-cast aluminum. The selection was based on extensive trade-off studies conducted for the TPE331 Engine series.

4.1.2.3 Combustors Annular, reverse-flow combustors operating at temperatures from 1255°K to l478°K (1800 F to 220QoF) were evaluated. Materials considered for the combustors included: '-..

'-'-.., o Hastelloy X o HSl88 Q ODS sheet alloys o Ceramics Thermal barrier coatings and photoetched/laminated construction techniques were investigated.

4.1.2.4 Turbines Gas generator turbines operating at rotor inlet temperatures o of l255°K to l478°K (1800 F to 220QoF) were evaluated. At 1478°K (2200 F), the rotor and vane are cooled. At l3110K (l900°F) the vane requires cooling and at l255°K (1800°F) the turbine is uncooled.

Candidate materials and fabrication processes are: o Integral castings using AF2-lDA and IN792 plus hafnium o Integral PM super alloy net shape o Laminated superalloy o Ceramics Turbi~e vane candidate materials and processes are: o Photoetched/laminated superalloy sheet o ODS extrusions o Cast and hot-isostatic-pressed superalloy o Ceramics Axial turbine rotor candidate materials and processes include: o Exothermic OS blades and powder metal super alloy hub.

o MAR-M 247 integral casting (OS blades and equiaxed hub).

o Hot-isostatic-pressed MAR-M 247 integral casting.

o Photoetched/laminated super alloy sheet.

o Ceramics.

4.1.2.5 Fans Low cost and satisfactoq' per formance in the fan component requires a low-cost manufacturing approach coupled with a mechan- ical design/materials approach that will meet bird ingestion requirements and allow the elimination of mid-span dampers. A pin- ned blade attachment appeared most promising as a mechanical design approach to satlsfy the bird ingestion requirements without mid- span dampers. Material and fabrication approaches considered for the fan blades includea: 0 PM titanium 0 PM steel 0 Composite 0 Cast steel 0 Forged aluminum 0 Forged steel 0 Forged titanium Material and fabrication approaches for the fan disk included: o PM titanium o PM steel 4.1.2.6 LON-Pressure Turbine The low-pressure (LP) turbine configuration selected for all Casting engines was an uncooled, shrouded, axial, cast design.

approaches considered were: o Investment o Rubber mold o AiRetrar.:* Other variations that were considered in the LP turbine design were: o Elimination of tip shrouds o One piece casting ot multi-stage turbine o Hot Isostatic Press (HIP) c~sti:1gs for improved proIJer- ties and higher yiel~.

4.1.2.7 ~earboxes In addition to conventional gears and housings, the following variations were evaluated: o Laser-hardened gears o Traction drives 4.2 Baseline Engine Design Engine trade-off studies, which will be discussed in more detail in d later section, were conducted on a sensitivity basis.

Changes in component performance, weight, and cost were related to *Pro?rietary Process, AiResearch Casting Co.

changes in engine performance, cost, and weight, which in turn were related to changes in airplane performance, cost, and weight.

The sensitivities were derived for baseline engines and for airplanes sized using the baseline engines. Three baseline engines were designed. They included a turboprop, turboshaft, and turbo- fan. The three engines had a common core, which was selected on the basis of prior studies.

4.2.1 T~rboprop B~~~!~n~ 4.2.1.1 Descrip~io~ A cross section of the turboprop baseline is shown in Figure 28. It is a two-spool, concentric-shaft, front-drive con- fiyuration comprised of a single-stage centrifugal compressor dr iven by a cooled single-stage radial turbine, a reverse-flow annular burner, a low-pressure t~.;o-stage axial uncooled turbine, and an offset two-stage reduction gearbox. The accessory gearbox is driven off the high-pressure spool and the engine is controlled by a low-cost, digital, electronic fuel control. In the component descriptions which follow, reference is made to current technology for comparison. Current technology is defined as that technology which could be committed to engineering development in 1978. As such, it is more advanced than technology in cur rent product ion engines.

Character ist ics of the single-stage centri fugal compressor are listed in Table 12. Three-dinensional blading is employed and the impeller is machined from a titanium forging.

The diffuser consists of 36 diffuser vanes followed by 58 deswirl vanes. Sheet metal construction is used.

t! "'0 aut 6U3 uu1 s .,. , a dOJdoqlnJ, au, lt se8 -.'9 pr;;:JPK: TABLE 12. BASELINE TURBOPROP - COMPRESSOR CHARACTERISTICS 6100 m (20,000 ft), 389 km/hr 210 Knots, Max. Power Type Centr ifugal Tip Speed 661 m/sec (2166 it/sec) Pressure Ratio 9.0 Relative* Efficiency +3.5 points Axial Clearance 0.013 cm (O.OOS in.)

Corrected Inlet Flow 1. 30 kg/sec (2.87 Ib/sec) Impeller Exit Mach No.

1.199 Diffuser Exit Mach No. 0.1S No. of Blades (full) 20 No. of Splitters 20 Compressor Diameter 27.196 cm (10.707 in.)

*Re1ative to current technology 9:1 pressure ratio, single-stage, centrifugal compressor -~ p 4l as x:xwec $ so QS Characteristics of the reverse-flow annular combustor are shown in Table 13. The combustor is rolled and welded from Inco 617. Ten airblast fuel nozzles investment cast in Hastelloy X are requirea. The operating temperature of this combustor is approxi- mately 311°K (100°F) higher than current technology.

The turboprop baseline high-pressure turbine design is summa- r ized in 'rable 14. 'i'he stator is an investment cast and brazed assembly of MAR-M 509. The rotor is machined from an AF2-1DA forg- ing. Cooli n9 passages are stem dr illed (electrostream). The exducer is investment cast from MAR-M 247.

Design cnaracteristics of the low-pressure turbine are listed in Table 15. The first-stage vane of the LP turbine is an integral investment casting in IN738 and the first-stage rotor is an inte- grul investment casting in IN792. The second-stage vane and rotcr are integrally cast from IN738. Both stages have-integral shrouds.

4.2.1.2 Baseline Turboprop Cycle ana Performance 'l'he basellne turboprop cycle was selected based on prior stuuies ana cycle selzction studies performeu in Task 1. Cycle cnaractcristics and perto:~ance at the engine design point (6100 m [20000 teet), 389 km/hr. [210 knots) true airspeed) and at sea- level static, standard day conditions, are shown in Table 16. A standard off-design thermodynamic model was used to predict engine performance throughout the flight envelope. This model includes

-

representatlons of component performance, thermodynamic routines, and matching procedures.

4.2.1.3 Baseline Turboprop Weight ~nd Cost r,etailcd estimates of turboprop baseline engine weight and cos~ were not available at the point in th2 program when baseline engine aata (size, weight, 2erformance, and cost) was required for WS¥PGiWCUC' ...... " a: l .. ; P:YZ;:UZt. .,..,.; ,€,w,wo __ ... .,.... ........... ,""".,. ......... , ... ...,.... ... i __ ... ~ .... _ .. "10_1'., ..... "" ... i'''''~ . .Y.,.---.-:""'7_~ TABLE 13. BASELINE TURBOPROP - COHBUSTOR CHARACTERISTICS Sea Level Static, Standard Day, Maximum Power Ti"pe Reverse-Flow Annular Inlet Pressure 78.12 N/cm (113.3 psia) Inlet Temperature 571.2°K (1028.l R) Inlet Flow 1.093 kg/sec (2.407 lb/sec) Combustor Exit Temperature l522°K (2739.7°R) Temperature Rise 950.9°K (1711. 6° R) Reference Velocity 6.85 m/sec (22.47 it/sec) Heat Release Rate 617 J/sec/m /Pa 3 6 (6.04 Btu/hr/atm/ft x 10 ) Pattern Factor 0.20 Liner Cooling, % Wa Pressure Drop, % ~P/P 3.0 Eff iciency 0.985

-~-. -"-. - ----------------

TABLE 14. TURBORP~OP BASELINE HIGH-PRESSURE TURBINE CHARACTERISTICS 6100 m (20, (l00 ft), 389 km/hr (210 Knots), Max Power Type Radial Specific Corrected Work, ~H/O 60,406 J/kg (25.97 Btu/lb)

Stage Work Coefficient, A = gJ~H

0.914 s u2.

tlP Pressure Ratio (tota17total) 2.492 Relative· Efficiency +5.5 Points 'l'ip Speed 583 m/sec (1910 ft/sec) Rotor Cooling Flow, • Wa 3.5 Bxit Mach No., v/a~r 0.33 Clearance 0.038 (0.015 in.)

No. i:.laues No. Vanes 17 Rotor Inlet Temperature 1477.6°K (2659.7°R) *Relative to a cu~rent technology ~oolea axial turbine at equal work.

;.

-~-~'.

~---.~- TABLE 15. BASELINE TURBOPROP LP TURBINE CHARACTERISTICS 6100 m (20,000 ft.), 389 km/~r (210 Knots), Maximum Power Type Axial No. Stages 2-1/2 Specific Corrected Weck, ~H/8 82,433 J/kg (35.44 Btu/lb)

Mean Work Coefficient h = gJ~H

2.3 , m 2 Urn Pressure Ratio 3.8 'rip Speed 320.5 m/sec (1051 ft/sec) Relative+. Efficiency +6 Points Ex it rtach No.

0.35 Clearance 0.038 cm (O.Ols in.)

No. Blades :12 No. Vanes Inlet Temperature l209"K (2176.4°R) Hub-to-Tip R~dius Ratio, Exit 0.698 *Relative to a current technology unc001eJ, axial, two-stage turbine at equal work coefficient.

TABLE 16. TURBOPROP BASELINE CYCLE AND PERFORMANCE CHARACTERISTICS, UNINSTALLED Altitude 6100 m Sea Level (20,0;)0 ft) Speed 389 km/hr.

Static (210 knots) Power Setting Maximum Power Haximum Power Standard Temperature Standard Shaft Power 239 kw 353 kw (320 hp) (473 hp) Shaft Specific Fuel Consumption 0.278 kg/hr/kll' 0.311 kg/hr/kw (0.455 lb/hr/hp) (0.511 lb/hr/hp) Corrected Airflow 1. 33 kg/sec 1.22 kg/sec (2.94 lb/sec) (2.693 lb/sec) Net Jet Thrust -15.13 N 87.67 N (-3.4 lb) (19.7 lb) Compressor Pressure Ratio 9.0 8.3 Turbine Inlet Temperature 1478°K 1478°K (2200°F) (2200°F) Nozzle Pressure Ratio 1.016 1.010 Gas Generator Speed, RPM 163,161 64,050 LP Spool Speed, RPM 128,000 28,000 Interturbine Pressure Drop % ~P/P 11.0 1.0 Overboard Leakage, % Wa 10.5 0.5 ., ~ ',- , airplane sizing and sensitivity studies. The goals established in Task I for cost an~ weight were therefore used. For the 353 kw (473 hp) baseline engine, this ODt cost goal was 60 dollars per kilowatt (45 dollars per horsepower). The weight goal for the baseline engine was 123 kg (270 lb). This goal, which translates to a relatively modest power-to-weight ratio, was set to allow meaningful trade- offs with respect to cost. Detailed estimates, performed later in the program, resulted in a significantly lower weight.

4.2.2 Turboshaft Baseline The turboshaft baseline had the same core and LP turbine design as the turboprop. The output gearbox was eliminated. It could be argued that the turboshaft cycle based on a single-stage centrifugal compressor would benefit from a slightly higher pres- sure rutio of approximately 10. This slight difference did not justify, however, the definition of a new baseline turboshaft. The turboshaft baseline engine is shown in Figure 29. Performance and component charilcteristics are identical to those previously listed for the bilseline turboprop.

4.2.3 Turbofan Baseline 4.2.3.1 Description A cross section of the turbofan baseline is shown in Figure 30. It is a two-spool, concentric-shaft, geared-fan, separately exhilusted configuration. The gas generator or high- pressure spool i~ compr ised of a single·-stage centr ifugal com- p.:~ssor dr iven b,! a cooled, single-stage radial turbine and a reverse-flow annular burner. The low-pressure spool is comprised of a single-stage axial fan driven by an uncooled, two-stage axial turbine through a simple, offset, reduction gearbox. The accessory gear bo); is dri ven off the high-pressure spool and the eng ine is controlled by a low-cost, digital electronic fuel control.

c------">")

--I --....

, / ~-.'

.....

" .

~. - ..

co co / The character istics of the fan are shown in Table 17. The single-stage fan is comprised of a forged titanium hub and pinned, forged, titanium blades and stators.

The compressor is a 1.3:1 scale of the turbopr.op baseline com- pressor. Its characteristics are shown in Table 18.

Th~ turbofan baseline combustor is a scale of the turboprop baseline combustor and its characteristics are listed in Table 19.

TIH~ high-pressure turbine is a scaled version of the turbopro:;.

high-pressure turolne and its characteristics are shown in Table 20.

The low-pressure (LP) turbine is a scaled version of the tur- boprop oasellne LP turbine ~nd its cllaracteristics are identical to those listed in T~ble 15.

4.2.3.2 Baseline Turbofan Cvcle and Performance The baseline turbofan cycle was selected on the basis of com- monality with tne turboprop baseline gas generator and on the basis of pr ior stud ies and cycle selection work per formed in TasK I.

Cycle characteristics and performance at the engine design point (6l00m [20,000 feet), 389 km/hr [210 knots) true airspeed) anci at sea-level static, standard day conditions, are shown in Table 21.

4.2.3.3 Baieline Turb~f~n Weight and Cost The OEN c':)st and weight targets established in Task I were used for the turbofan baseline. The data, which was revised later in Task II, was found to be conservative. The Task I targets for the tur- bofan O~\ cost and weight were $6. 7ol/N ($30/1b) of thrust and 134 kg {296 Ib), respectively.

TABLE ~7. TURBOFAN BASELINE FAN CHARACTERISTICS 6100 m (2C,000 Ft.), 389 km/hr (210 Knots), Maximum Power Inlet Corrected Flow 21. 08 kg/sec (46.43 Ib/sec) Bypass Ratio 8.0 Bypass Pressure Rat~o 1.5 Core Pressure Ratio 1.5 Corrected Tip Speed 381 m/sec (1250 ft/sec) Relative* Efficiency +1. 5 Points Hub-'l'lP Radius Ratio 0.452 Fan Speed, rpm 15,739 No. of Blades 17 No. of Stators 39 *Relative to a current technology 1.5 pressure ratio, single- st<lge fan.

TABLE 18. TURBOFAN BASELINE COMPRESSOR CHARACTERISTICS 6100 m (20,000 Ft), 389 km/hr (210 Knots), Maximum Power Type Centrifugal Tip Speed 648 m/sec (2124 ft/sec) Pressure Ratio 9.0 Relative* Efficiency +:.5 points Axial Clearance O. 013 cnl (0.005 in.)

Correctea Inlet Flcw 1. 69 kg/sec (3.73 lb/sec) Impeller Exit Mach No.

1.199 Diffuser Exit Mach No. 0.l5 No. of Eludes (Full) 20 No. of Splitters 20 Compressor Diameter 31.01 cm (12.21 in.)

*Relutive to current technology 9:1 pressure ratio, single- stage, ~e~trifugal compressor.

, TABLE 19. TURBOFAN BASELINE COHBU5TOR CHARACTERISTICS SEA LEVEL STATIC, STANDARD DAY, MAX. POWER Type Reverse-Flow Annular ?

Inlet Pressure 109.9 N/cm- (159.5 psia) Inlet Temperature 637.7°K (1147 .SOH) Inlet Flow 1.901 kg/sec (4.187 lb/sec) Combustor Exit Temperature 1522.1 K (2739.7°R) Temperature Rise 884.4°K (1591. 9° R) Reterence Velocity 8.42 rn/sec (27.6 ft/sec) Heat Release Rate 638.94 J/scc/m /pa , (6.25 btu/hr/atm/ft x lOu) Pattern Factor 0.20 Liner Cooling, % Wa 42 Pressure Drop, % ~P/P 3.0 Eft icicncy 0.985 ....

TABLE 20. TURBOFAN BASELINE HIGH-PRESSURE TURBINE 6100 m (20,000 Ft.), 389 km/hr (210 Knots), Maximu~ Power

Type Radial~

Specific Corrected Work, lH~

65, 221 ~/k~_. . I

(28.04 Btu/1b)

Stage Work Coefficient, A = gJlH

0.909 s 2 U TIP Pressure Ratio (total-to-tota1) 2.708 +5.5 Points Relative* Efficiency Tip Speed 607 m/sec (1990 ft/sec) Rotor Cocling Flow, % Wc 3.6 0.33 Exit Mach No., v/a~r C1earunce 0.038 cm (0.015 in.)

No. of Blades 14 No. of Vanes 17 Rotor Inlet Temperature 1477.6°K (2659.7°R) *Relative to a current technology, cooled, axial turbine at equal work.

~AB~E 21. TURBOFAN BASELINE CYCLE AND PERFORMANCE Alt .. tude 6100m (20,000 ft) Sea Level Sj?eeo 389 km/hr (210 kts) Static Standard Temperature Standard

.--

Net Thrust, lb 1740 N 4294 N (391 lb) (965 lb) Thr~st S~ecific Fuel O.OGl kg/N-hr 0.041 kg/N-hr Consumption (0.601 lb/hr/1b) (0.402 lb/hr/lb) F~~ Inlet Corrected Flow 21.1.0 kg/sec 18.75 kg/sec (46.48 lb/sec) (H.j lb/sec) ~ore Correctec Flow 1. 69 i<g/sec 1. 58 kg/sec (3.73 lb/sec) (3.49 lb/sec) Fan Pressure Ratio l.5 1.4 Corr.pressor Pressure Ratio ~.O 8.2 Turoine Inlet Temperature 1478°K 1478°K (2::00°F) (2200° F) B.O E?t:.:lsS Ratio 8.0 Compressor Speed, RPM 58,014 59,117 Fan Speed, RPM 15,736 15,075 Fan Duct .lP/P 0.025 0.025 Fan Nozzle Thrust Coefficient 0.985 0.985 Core Nozzle Thrust Coefficient 0.985 0.985 ~.3 Aircr3ft Sizina and SenGitivitv Studies The definition of fixed-wing aircraft characteri-tics was sub- contracted to Cessna Aircraft Company. Their task ~:as to define the general requirements and detailed characteristics of the air- planes selected for study in Task II, ~amely, the pressurized twin and the light twin. The characteristics of turbcprop- and turbof;m-powered pressurized twins were defined. T\,'o variations of the turbofan-powered ai rcraft were invest igated, namely, an aft- fuselage-mounted engine and a wing-mounted engine. The character- istics as defined by Cessna were based on their experier.c,;:: an0 engine data provided by AiResearch. Also, Cessna supplied weight and drag correlations, which allowed the weight and drag breakdownG to be adjusted as mission perforr.1ance and airplane synthesis was accolr.t'lished. Our ing Task II, the General Aviation Synthes is Program (GASP) was used by AiResearch to size the aircraft and establish th~ power requirements ~nd wing loading. The planform drag buildup and \oJeight breakdo\m wer~ not altered fro!!! those supplied by Cessna except as dictated by (1) the correlations for the effects of gross weight and wing loading, and (2) the modifi- cations necessary to allow modeling the airplanes in GASP. In the latt~r case, Cessna was consulted and recommended the required modifications. Advanced technology airplanes \oJere not defined.

The designs IJrovided by Cessna were slight extensions of current fixed-wing aircraft. Additional airframe advanced technology could ba postulated for 1988 but it would be more difficult to separate the i~provements due to the engine and those due to the advanced technology airframe.

The general characteristic~ and performance requirements of the designs supplied by Cessna are shown in Tables 22 and 21.

Design numbers were assigned for each of the airplanes, namely: • " , , TABLE 22. GENERAL AIRCRAFT CHARACTERIS'J'ICS - C-ATE STUDY Design t-~o. 1 1A 2 Pressurized PresslIrizc(l P:-cssllrizec1 Light Description Engine Type T\'1i n Tl'rbofan T\'li n Turhofan T\'d n Turboprop Twi n 'rurhoprop Estimated SHP/Thrust Class 6675N 373 hi 224 hi

-- (500 hp) (300 hp)

(1500 lbs) Estimated Weights 28110 kg

i Gross

..

(6300 lbsl 1544 kn 1317 kQ

I Empty

I

(2900 lb5) (3400 lbs)

-

I Approximate Wing Area 16.71"1

I

..

I (l80 ft2) I Seatinq (Illcluding Pilot)

I

I

..

Maximum

I I I

Norrr,al

--

3 3

I I

Cabin Volume 3.6m 4.6m 3 ..

(::'30 ft3) (165 ft ) Cabin Pressure Differential 3.24 N/cr.t ..

(4.7 ps i) 0

I I

\D ...J 7.\SLE: :;;3. PERFOR.'lANCE REQUIREl-IENTS - FIXED-WIlIG AIRCRAI'T - GATE PROGRAH

'-]

Design ~lo. 1 lA 2 4 Speed

I

482 km/hr 444 k::/hr "axi~um ..

I (260 ktsl (240 <';51

\

I I :-Iaximum Cruise 444 kr.l/hr 417 k:: /hr

i ~ I (240 ~tsl (225 ~';51 I , Range·

I At ~al(imum Cruise 1556 km

2037 "m (940 NMI () 1 00 ~~Il

-

I I

At Speed for Min COC·· 1945 km 2~08 "<:T1 • (l05a N'II (l~OO ~~I) : : !

Fay10lc (:ncludi~g Pilot) 518 kg 345 <a I • (1140 1b) (7~0 :!:-) I I I Ser'11c,? C~ilin~ I I I 9150 m Twin Encine 6: 00 '" • (30COO ft) (20000 ttl I

I

,

I 4575 m Sinqle Engine 2135 m

I (15000 ft) (7000 !t) !

i Rate of Cdmb

I Twin Engine (SL Std) 488 m/min

I ;

-

(160.,0 f~/.min) I

I !

I 9_ :::, inln Single E:''line (Sr. St=) (300 ft/minl I !

, Takeoff Oi~tance (Flaps i 0.~55 rad (15 deg). SL Std) I 458 m Ground Run 1:;6 ""' ..

I (l5GO ftl (l1CO !ti

I

!

I I

To I:; m (50 ft) Altitude 671 m 488 :': • (2200 ftl (1600 :t I

i

i Landino Oi"tance (,laps 0.51 rad (3~ degll I !

Ground Roll 259 m 229 ."

(850 ftl (750 !tl

-

i

I I 610 m

From 15 m (50 Ctl A1t~tude 1_ 4:8 :"1

I I (2000 ftl , (lSCO :tl

- I

L --- ·At 5490 m (18000 ftl Cor Nos. 1. 11\ and 2. and 3050 :!I (l0000 ftl for No.4.

··Direct Operating Cost ~d o Turbofan-Powered, Pressurized Twin (wing mounted) - uesign 1 Turb0fan-Powered, Pressurized Twin (aft fuselage o mounted)- Design lA o Turboprop-Powered, Pressurized Twin - Design 2 o Turboprop-Powered, Light Twi~ - Design 4 Detailed fixed-wing airplane characteristics as defined by Cessna are listed in Appendix II. The characteristics as supplied formed the basis for modeling the pressurized twin and light twin for the General Aviation Synthesis Program (GASP) used for airplane sizing, mission analysis, and sensitivity studies.

4.3.1 Fixed-Wing Aircraft Sizing and Mission Analysis Airplane sizing and mission analysis were performed assuming fixed mission performance requirements and varying airplane takeoff gross weight (TOGW), wing loading (W/S), and engine size to meet th~ mission requi rements. The character istics as supplied by Cessna were not varied except as required for changes in TOGW, W/S, and engine size. Specifically, wing and empennage g~omet~ic char- acteristics, fuselage dimensions, standard and optional equipment, and the high-lift system were unchanged. Wing area varied as TOGW and W/S were varied.

The weight breakdown as supplied by Cessna varied in the fol- lowing groups: , \.,ring 0 Vertical Tail 0 Horizont.Jl T.Jil 0 Hain Gear 0 Uose Gear 0 Controls 0 Retraction System The following weight groups were not allowed to vary: 0 Power Plant 0 Nacelle 0 Fuselage 0 Stand.Jrd Equipment 0 Furnishings Exterior Finish 0 Optional Equipment 'rhe tllselage weight remains constant since its size is fixed by cabin ~olume, which is a function of the number of passengers. The nacelle and power plant group would have beeo v~ried as engine size varied. However, GASP contilined routines for resizing the nacelle and associated equipment, which gave optimistic results. To avoid a I.-.ajor modification of ,GASP, engine weight was fixed and the results \vere .Jdjusted at a l.:lter point in the study, based on engine weigi1t sensltivities. 'fhe standard equipment group and optionill equipment does not vary with gross weight for a particul.Jr aircraft category. Furnishings and exterior finish were also ~ssumed to be fixed weights.

The drilg polar, as supplied by Cessna, varied as the .Jirplane was reslzea to account for change in wing area and .J change in ~;j~ relationship of aircr.Jft wetted area to wing area. The ch.Jnge is consistent witn the Cessna drag buildup.

Engine size varied as gross weight and wing loading varied.

The wing loading initially supplied by Cessna was an estimate and was iterated to find the wing loading that resulted in the lowest gross weight while meetir.~ all mission requirementn.

Aircraft and engine sizing was accomplished by the General Aviation Synthesis Program (GASP). Installed engine performance maps based on the baseline engine off:"design deck were utilized.

Assumed engine installation losses were as follows: Pressurized Twin Light Twin , TPE TFE TPE Bleed Air, kg/min/eng 2.0 2.0 0 (lb/mi n/eng) (4.5) (4.5) Power Extraction, kw/eng 3.7 3.7 3.7 (hp/eng) (5 ) (5 ) (5) Total Pressure Recovery 0.995 Ratio 1.0 1.0 The bleed air rate decreased linearly at the rate of 0.23 kg/min/eng/305rn (0.5 lb/min/eng/10,000 feet). The propeller effi- ciency, weight, and price were calculated by the propeller routine contained in GASP. These parameters were computed for a three- bladed propeller based on a fixed rotational speed and diameter of 2500 rpm and 1.9m (6.2 ft), respectively. The design character- istics of the propeller are: Activity factor/blade Design lift coefficient 0.5 Number of blades Efficiency (cruise) 0.87 '.

4.3.1.1 Pressurized Twin Mission requirements for the Prcssl.rized Twin (Designs 1, lA, and 2) are shown in Figure 31. Airplanes were sized by GASP at wing loadings of 137 to 205 kg/m (28 to 42 lb/ft ). At each wing load- ing evaluated, the aircraft were sized to m~et takeoff, cruise, and range requirements. Climb performance, landing distance, and ser- vice ceiling were evaluated- as a function of wing loading.

The results of the wing loading study for Design No. 2 (turbo- prop medium pressurized twin) are shown in Figures 32 thrcugh 35.

At each wing load ing shown in these figures, the requ i rements of takeoff distances are met or exceeded and all wing loadings meet the range requirement of 1556 km (840 nm) at 5490 m (18,000 feet) and 441 km/hr (240 knots). A , .. ing loading of 185 kg/m (38.0 Ib/ft2) was selected on the basis of meeting the single-engine ser- vice ceiling requirement of 4575m (15,000 ft), as shown in Figure 32. At 4575 m (15,00C' ft), a wing loading of 185 kg/m (38.0 lb/ft2) is the highest wi 0g loading that allows a 31 m/min (100 ft/min) rate of climb. This fi9ure also shows that the twin-engine rate of climb at 9150 m (30,000 ft) exceeds 31 m/min (100 ft/min) at all wing loadings. Figure 33 shows the variation of takeoff dis- loading. Below approximately 200 kg/m (41 tance with wing lb/ft ), takeoff requirements are exceeded and the engines are sized by the cruise requirement. Above 200 kg/m (41 lb/ft2) the engines are sized to provide sufficient power for takeoff. Figure 34 shows the variation of installed power at sea level, static, standard day, takeoff" power as a tunction of wing loading. At the selected wing loading, power is nea~ minimum. Figure 35 shows the variation of gross weight and fuel consumed versus wing loading.

Lower gross weights would result if a higher wing loading was selected but fuel consumption is close to minimum. At the selected wing landing, climb and landing requirements were exceeded.

, D E F . . ---- ..

A ., A B I 1556 km

•• •• (840N.M.1 SEGMENT DESCRIPTION A· B TAXI· 5 MINUTES AT IDLE TAKEOFF S·C CLIMB TO 5490m (18,000 FTi C·D D· E CRUISE AT 5490m. 444 km/hr (18,000 FT, 240 kts) E·F RESERVES· 45 MINUTES AT CRUISE CONDITIONS - --- --- - MISSION PERFORMANCE REQUIREMENTS (STD DAY) SPEED 482 km/hr (260 kts) MAXIMUM RATE-OF·CLI-1B SINGLE ENGINE, SL, MAX· 92m/min (300 FT/MIN) TWIN ENGINE. SL. MAX· 488m/min (1600 FT/MIN) FIELD PERFORMANCE (SL) GROUND RUN· 458m (1500 FT) TO 50 FT ALJITUDE • 671m (2200 FT) SERVICE CEILING SINGLE ENGINE· 4575m (15,000 FT) TWIN ENGINE· 9150m (30,000 FT) Mission Requirements - Pressurized Twin (Designs 1, r'igure 31.

lA, and 2).

, AIRCRAFT DESIGN NO.2 SERVICE CEILING REQUIREMENTS (MAXIMUM CLIMB POWER) • 1556 km (840 N.M.), 5490 m (18,000 FEET), 444 km/hr (240 KNOTS) • 671 m (2,200 FEET) T.O. DISTANCE TWIN ENGINE RIC 9150 m (30,000

:E 400[

FEET),ISA 0...

u..

c:

~ 30t E

-- E

u.. 200 w ~ I cd: 100 a:

I

170 180 190 200 kg/m2 34 36 38 40 42 WING LOADING· LB/FT2 Figure 32. Turboprop-Powered Pressurized Twin.

:'04 ~IRCRAFt DESIGN NO.2 • 1556 km {840 N.M.) 444 km/hr {240 KNOTS), 5490 m (18,000 FEET) 2,300 r- 700 I

t 2,200

I ~ E 660

~ I I 4 L.. TAKEOFF SIZED

u. E ENGINES [671m g 2,100 ~ 640;fr (2,200 FEET).

O. I T.O. DISTANCE) I- 0 w I- ~ w 620

« ~

I- 2,000 «

en l- e ~ 600

::: c I ...... CRUISE-SIZED ENGINES

o [444 km/hr (240 KNOTS), ~ 5486 m (18,000 FT,) ~ 1,900 580 MAX. CRUISE POWER) 170 180 190 ~OO 210 1,800 L kg/m 34 36 38 40 42 44 WING LOADING - LB/FT2 Figure 33. Turboprop-Powered Pressurized Twin-Engine Sizin~ AIRCRAFT DESIGN NO.2 • 1556 km (840 N.M.), 5490 m (18,000 FEET) 444 km/hr (240 KNOTS) • 671 m (2,200 FE:T), T.O. DISTANCE W CJ W a: W ~ W a: :::J O~ W:x: a: • ;:

-

..0lIl{'\.

..:J(.

~ a:<t 400

w(I) :: -.

"

0(1) ~...J ~- (I) l- LL <t :x: (I) w ...J ...J <t I- (I) 160 170 180 190 200 kg/m 2 36 38 40 32 34 WING LOADING - LB/FT2 Figure 34. Turboprop-Powered Pre$surized Twin-Engine Sizing

lao

AIRCRAFT DESIGN NO.2 • 1556 km (840 N.M.), 5490 m (18,000 FEET), 444 km/hr (240 KNOTS) • 671 m (2,200 FEET), T.O. DISTANCE TAKEOFF GROSS WEIGHT 5,600 [

'-

~ 5500 2500

- In W .:e.

G' l

i 5,400 I 245"

5,300 L

170 180 190 200 210 kg/m2 I I I I 34 36 38 40 42 WING LOADING - LB/FT2 FUEL USED co -J 400 I C In W .:.: en ::> -J W ::> u..

360_ 180 190 200 210 160 170 kg/m2 I I I I 34 36 38 40 42 WING LOADING - LB/FT2 Figure 35. Turboprop-Powered Medium Twin Sizing.

For the turbofan-powered pressurized twins (Design No.1), 2 2 all wing loadings above approximately 155 kg/m (32 Ib/ft ) allowed service ceiling and rate-of-climb requirements to be met. All wing loadings investigated resulted in acceptable landing per- formance. Figure 36 shows that acquisition cost is minimum at a 2 2 wing loading of approximately 150 kg/m (31 Ib/ft ), which is too low for per formance requirements. Operating cost is minimum at 2 2 approximately 165 kg/m (34 Ib/ft ), as shown on Figure 37. Fuel 2 2 consumption is minimum at approx im~tely 185 kg/m (38 Ib/f t ) , as shown in Figure 38. The best compromise did not appear to be sig- nificantly different from the wing loading originally chosen 2 2 by Cessna, namely 167 kg/m (34.23 Ib/ft ).

Characteristics and performance of the turbofan- and turboprop-powered pressurized twins are shown in Table 24. At the selected wing loading, both configurations meet or exceed the maxi- mum speed requirement of 482 km/hr (260 knots). There is a large difference between the turbofan- and turboprop-powered aircraft in gross weight, cruise fuel consumption, total mission fuel and engine core size required. For the speed and takeoff requirements of this application, the turboprop-powered confi~uration is clearly superior.

'rhe effects of relaxed field per formance and high-alt i tude cruise were investigated for the turbofan conf iguration. The results ar~ also shown in Table 24. Takeoff distance was increased to 862 m (2800 ft) and the airplane was allowed to cruise at 7625 m (25,000 ft). The difference between the turboprop and turbofan versions decreases, although the turboprop is still superior. The range requirement on the turbofan was increased to 1637 km [884 NM (+5 percent») to offset the increased altitude sinc~ the turboprop would also cruise more efficiently at 7625m (25,000 ft). Further improvements in the turbofan configuration may be P?ssible if

, -- ._-- -- - --_ ... ----- .. _. __ ., -- .. _" -_.---- -

AIRCRAFT DESIGN NO.1 PRESSURIZED TWIN (2) TFE MODEL 1054 TURBOFAN ENGINES 1.02 MINIMUM WING LOADING Q 1.01 I--- (CLIMB & SERVICE CEILING I- c:x: a: REQUIREMENTS)~ -l I- eI) o CJ ~ 1.00 ~ en ::J o CJ c:x: 0.99

~~

I

0.98 I~--------~----------~--------~--------~~--------~--------~ 130 140 150 160 170 180 190 kg/m 2

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

28 30 32 34 36 38 WING LOADING, LB/FT2 Figure 36. Relationship of Acquisition Cost to Wing Loading for Aircraft Design No. 1

'_.- ._-----

.---~ ._---- ------- .. ".-- -

AIRCRAFT DESIGN NO. 1 PRESSURIZED TWIN (2) TFE MODEL 1054 TURBOFAN ENGINES 1.10

,-- I

MINIMUM WING LOADING (CLIMB & SERVICE CEILING a: >- -.

en a: :I: o III N SELECTED a: WING w 0...

LOADING o

I ___

/ ~ <t 1.00 a: I- en o (J C!)

Z ~ a: w 0...

o 0.90 140 150 160 170 180 kg/m 2 '30 28 32 34 36 38 WING LOADING, LB/FT2 Figure 37. Relationship of Operating Cost to Wing Loading for Aircraft DeSign No. 1 .1.10 # ------ ----------- .... ----, AlB CRAFT DESIGN NO. 1 PRESSURIZED ... -JIN MINIMUM WING LOADING 1600, 720 (CLIMB & SERVICE CEILING REQUIREMENTS) (I) OJ ~ 1500, 680 a ~ w (I) :::> -I SELECTED WING LOADING W :::>

u. I

-I I 640 ~ 1400 t- - -- - - - --- ---- --- 130 140 150 160 170 180 kg/m2 I I I ~ ~ ~ ~ ~ WING LOADING - LB/FT2 • 671m (2200 FT) TAKEOFF DISTANCE • 1556km (840 NM), 5490m (18000 FT), 444km/hr (240 KTS) Figure 38. Mission Fuel Consumption

~

TABLE 24. AIRCRAFT SIZING SUMMAR¥ .....

I

Alternate Mission

~ '---------------=r-------'----r --- -----r------

- Aircraft Type P:eEsurized Twin Pressurized Twin Pressurized T~in Engine Type Turbofan Turboprop Turbofan t Takeoff Gross Weight 2825 kg 2470 kg 2706 kg (6223 1b) (5441 lb) (5960 lb) E.llpty Weight 1550 kg l4B5 kg 1524 kg (3413 Ib) (3271 Ib) (3357 lb)

I

2 2 2 Wing Loo-ding 167 kg/m lB5 kg/m 167 kg/m (34.2 1b/ft2) (3B.0 Ib/ft2) (34.2 lb/ft2) Maximum Speed/Altitude 50,/6100 km/hr/m 482/5490 km/hr/m Not (275/20,000 kts/ft) (:50/18,000 kts/ft) Available Range at Cruise Speed/Altitude 1556 km 1556 km 1637 km (840 nm) (840 nm) (884 nm*) Rate of Cli~b, 2 Engineti 547 m/min 607 m/min Not (1795 ft/min) (1991 ft/min) Available Takeoff to 15m (50 ft), Std Day 649 m 641 m B54 m (2128 ft) (2100 H) paoo H) Cruise Fuel Consumption 212 l/hr 132 l/hr 171 l/hr (56.0 gal/he) (34.7 gal/he) (45 gal/he) 4'12 liters Block Fllp.l ~ 782 liters 694 li ters (2C6.4 gal) (.27.2 gal) (lB3 gal) En'Jine 51.S Takeoff Power/Thrusl/Eng* 4695 H 336 kw 3627 N (l100 Ill; (4')0 hI') (615 ltl) Enyinc SLS Core Airflow I l.~l kg/sec 1.16 kg/sec 1.34 kg/sec

________________ 1 (3.9r, lb/sec_) I? "t; Ib/sec)

(2.95 lb/sl'c) 444 km/Ill (240 kts), 7,625 m (25,001) ft.)

**uninstdlled cruise speed was increased. Hcwever, the resulting airplane is out I of the category of the pressurized twin and cost could escalate I sharply.

' I (a .

. to~, J ;:I\.~1t

thlS .t. The comparison indicates that a competitive turbofan in I ~ J ~ • elf-f.

and smaller categor ies is unl ikely unless takeoff

distance is I~ -d''lP-' ~

increased, cruise speed and altitude are raised, and the cost of 'I ~ sophisticated high-lift systems is acceptable.

4.3.1.2 Light T~in Mission requirements for the light twin are shown in Figure 39. Airplanes were sized by GASP at wing loadings of 112 to 2 ' 2 146 kg/m (23 to 30 Ib/ft) to meet takeoff, cruise, and range requirements. Climb performance, landing distanc::!, and service ceiling were evaluated as a function of wing loading.

The results of this study for Design No.4, the light twin, are summarized in Figures 40 through 42. All performance require- ments were exct::eded over the range of wing load ings investigated 2 2 [112 to 146 kg/m (23-30 Ib/ft )]. The selection was therefore based on gross weight, fuel consumption, and engine size. On this basis, a wing loading of 139 kg/m (28.4 Ib/ft2) was 5elected.

This sel::ction re5ults in minimum gross weight, engine size, and fuel consumption. The variation of these parameters with wing loading is shown in Figures 40 and 41. Figure 42 shows the varia- tion of takeoff distance with wing loading. At wing loadings below 2 2 the selected value vf 139 kg/m (28 Ib/ft ), the engines are cruise sized. At higher wing loadings, the engines are taKeoff sized.

Characteristics and performance of the light twin a~e shown in Table 25. At the selected wing loading, the airplane meets the 445 k~/hr (240 knots) maximum speed requirement.

o E F ... . ---- ....

A B

Ioc 2037 km ~ I

(1100 N.M.)

SEGMENT DESCRIPTION A·B TAXI- 5 MINUTES AT IDLE B·C TAKEOFF C·D I CLIMB TO 3048m (10,000 FEET) D·E CRUISE AT 3048m (10,000 FEET)417 km/hr (225 KNOTS) E·F RESERVES - 45 MINUTES AT CRUISE CONDITIONS MISSION PERFORMANCE REQUIREMENTS (STD DAY) SPEED 444 km/hr (240 KNOTS) MAXIMUM RATE·OF·CLlMB SINGLE ENGINE. SL. MAX 92m/min (300 FT/MIN) lWlN ENGINE, SL. MAX 488m/min (1600 FT/MIN) FIELD PERFORMANCE (SL) GROUND RUN 336m (1100 FEET) TO 50 FT ALTITUDE 488m (1600 FT) SERVICE CEILING SINGLE ENGINE· 2135m (7000 FEET) lWlN ENGINE· 6100m (20,000 FEET) Figure 39. Mission RequireITlents - Light. Twin , 2650 00 1,300

~ 5.700

1 2550

I 5,600 1 1,200 ~

FUEL USED I I l- x C> w ~ 2500 l ~ ~ 5500 CI) w ' ::J :: ..J W ~ -11,100 ~ 2450r GROSS

ffi SAOOr

WEIGHT 6.300 6,200 1,000

l

~ ________ ~ __________ ~ ________ -L ________ ~1450 110 120 130 140 150 kg/m2 26 28 30 32 22 24 WING LOADING - LB/FT2 SIZED FOR. RANGE - 2037 KM (1,100 NM) AT 3048M (10,000 FT) ANO 417 KM/HR (225 KNOTS) t-' t-' UI

• to. DISTANCE OF 488M (1600 FT)

" figure 40. Tqrboprop Li9ht ~win SIZED FOR • RANGE - 2037 KM (1,100 NM; AT 3048M (10,000 FT) AND 417 KM/HR (225 KNOTS) Q.

J: • 1:0. DISTANCE OF 488M (1600 FTl W Z c:l 400 z I W a: W Q.

C W a:

~

3 350 0« W(I)

"

a:- ~

~

a: (I)" ~ 240 w..J I"'-- ..

~ ~cr.

\.I o Q. 300 I- u.

« J: I (I) - --- -_. - C UJ :::j 250

«

~ 120 130 140 150 z kg/m 2 24 26 28 30 WING LOADING - LB/FT2 Figure 41. Turboprop-Powered Light-Twin Eng:ne Sizing ", SIZED FOR • RANGE - 2037 KM (1,100 NM) AT 3048M (10,000 FT) AND 417 KM/HR (225 KNOTS) • T.O. DISTANCE OF 488M (1600 FT) 1,600 SELECTED WING LOADING -""PI!I.II.I~.---~ E 480 I- I u.

E 1.0 I- ~ 1,500 o TAKEOFF·SIZED ENGINES I- u.

[488 m (1600 FEET) T.O.

'w o U DISTANCE) 1.0 2 440 o ~ I- I- eI) w U C 2 1,400

~ CRUISE·SIZED ENGINES

~ u. 420 I ~ [417 km/hr (225 KNOTS), I- u.

eI) o MAX. CRUISE POWER) w C ~ u.

u. < I- 400 o w ~ 1,300 ~ I- 110 120 130 150 kgim2 , 1,200 22 24 26 28 30 32 WING LOADING - LB/FT2 Figure 42. Turboprop Light-Twin Engine Sizing " TABLE 25. LIGHT-TWIN SIZING SUMMARY Light Twin Aircraft Type Engine Type Turboprop Takeoff Gross Weight 2374 kg (5228 lb) Empty Weight 1352 kg (.2978 lb) Wing Loading 139 kg/m (28.4 lb/ft2) Maximum Speed/Altitude 444 km/hr - 3050m (240 ~ts - 10,000 ~t) Range at Cruise 5peed/Altitude 2037 km (1100 NM) Rate of Climb, 2 Engines 569 m/min (1864 ft/min) Takeoff to 15m (50 Ft), Std Day 486 m (1595 ft) Cruise Fuel Consumption 128 liters/hr (33.8 gal/hr) Block Fuel 637 li ters (168.2 gal) Engine SLS Takeoff Power* 251 kw (336 hp) Engine SLS Core Airflow 0.87 kg/sec (1.91 lb/sec) *Uninstalled liS , 4.3.2 Sensitivity Studies The effects of engine weight and specific fuel consumption on aircraft characteristics were evaluated by resizing the aircraft with the use of GASP for changes in these parameters. The baseline aircraft described in 4.3.1 were used. For each of the changes, the aircraft were resized to meet the takeoff, range, and cruise conditions. The results of th~se sensitivity studies are contained in Appendix III.

4.4 Engine Trade-Off Studies The majority of the engine trade-off studies were made using tt~ turboprop baseline engine described in 4.2.1.1. This engine, hereinafter referred to as Engine A, is " free-turbine engine, and is comprised of a single-stage centrifugal compressor driven by a cooled radial turbine, ~ reverse-flow annular combustor, and a two- stage axial uncooled power turbine driving a two-stage r~duction gearbox. Two groups of trade-off studies were conducted on this engine. The first group considered cycle and configuration and included the following items: o Cycle o Compressor type o High-pressure turbine type o Spool arrangement (single shaft versus free turbine) The second group cons isted of more detailed trade-ofis on a C~jn ponent level and included the following: o Single-stage centrifugdl compressor fabrication o Combustion system fabrication and fuel nozzles o High-pressure turbine fabrication and materials o Low-pressure turbine fabrication and materials o Single-stag~ versus two-stage power turbine , o Gearbox type and fabrication o Sheet metal versus cast construction As the trade-off studies were conducted, promising engine cycles and configurations were more fully defined and carried forward to an evaluation on a system or aircraft basis.

In addi t ion to the turboprop eng ines, two turbofans and one turboshaft eng ine were def ined. These three eng ines incorl?orated features identified in the turboprop studies.

4.4.1 Cycle and Configuration Trade-Off Studies 4.4.1._ Cycle The first cycle trade-off studies performed were accompli~hed for the baseline configuration, designated Engine A. The charac- teristics of Engine A are shown in Table 26. The maximum c~mpr~3- sor pressure ratio for the single-stage centrifugal compressor was determined to be 10 for the technolog'y level be ing investigated.

The rar.ge of compressor pressure ratios investigated was 6 to 10.

The variation in compressor efficiency assumed is shown in Figure 43. The efficiency shown is relative to the efficiency which could be achieved in a production compressor designed in 1977. Turbine rotor inlet temperature was also varied from l255°K to 1478°K (l800°F to 2200°F). At 1255°K (18DOOF), the turbine is uncooled, at l3110K (1900 F) the turbine nozzle is cooled and at 1478°K (220QoF) the nozzle and rotor are cooled. Turbine effi- ciency varies with the level of turbine inlet temperature. Levels of turhine efficiency assumed relative to a 1977 radial design are shown in Figure 44. The results of design-point calculations at cruise conditions are shown in Figure 45. Shaft power and SFC are shown as a function of turbine inlet temperature and pressure ratio. For all turbine inlet temperatures, shaft power is near optimum at a pressure ratio of 9, with specific fuel consumption near minimum.

, T/.IlLE 2(, TURIl(I!'i<OP CAtllJlPATE ENGUIE CIlAHACTt:HISTICS, SEA LEVEL, STATIC, STAllDAI<D 1)/,';, H.KEon' Po\-;EH UIHllSTALLED Free Turbine Engjn~s Ga~ Generator Tur~ln~ Type Hadial· AXial· 2 Sto'le Compressor Type I Stage Centrifugal Centr Ifugal 1 Stage Centrifugal

-

Lng in" A En'~ine B Eng i -,e C Engine 0 Engine E i OK Tur~lne Inlet Temp, 1476 125~ 1476 1478 12~~ (0 F) (2200) (l800) (22)0) (2200) (1800 ) Tur~lne Cooling Yes flo Yes No Yes (~0mpre~so, Pressure Rat10 8.3 8.~ 8.3 12.0 8.3 lrdet Corr~cted flow, Kg/sce 1.22 1. 22 1. 22 1. 22 1. 2:!

(I ~/sec) (2.69) (2.69) (2.69) (2.69) (2.69) ~thi(:r , KW 3~3 2~J 3j9 345 254 (t,P) (473) (347) (4~4 ) (463) (340) !ihaft 5~ecific fuel r:"~gn~7l~n 0.311 0.315 0.~95 0.321 0.325 (It,/tlr /tIP) (0.511) (O.:>ii , (0.533) (0. 41l4) (0.528) luqlnc Weiqtlt, I:g ~: 95 93 95 95 (lb) (210) (210 ) (219) (210) (210) lr.glne Ofl-l Cost, •••• $ (1'J77) 4~495 45891 39530 H908 37434 Sing.e Sh~ft Engines CC.Jnl",ct.:550[ Type 1 Stage Centrifugal Tuc~lOe Type Hadia1/Axial·· All Axial··· Engine f r- En'Jine G Engine II Engine 1 1 ur~lnc Inlet, T':mp, 0 Y. 1478 1255 1478 1255 (0 f) (2200) (1800) (2200) (1800, 'j'llr~lnc C(){Jlir.'J Yes ItO Yell Ito t:(JrT,£Jfl'uBOr l'rcG!;u(c katlf') 7.(, 7.6 7.t 7.6 Kr;/!;C(; lulct Corccclcd flow, 1.12 1.12 1.12 ,12 (ltJ/6cc) (2.47) (2.1,7) (2.47) (2.47) Power, kw 304 227 304 ~i~ (408 ) (401l) (294) (hi') (304 ) JilCJt t !lfJ":ci f 1'= Fd(:l cr'~~:J~:h~i~~n 0.324 O.3~£j ().333 0.319 I !t./tlr Ifl[" (0. 5lJ) (CJ.'J1U) (0.~2l) (0.~47) lnq I ne I,e .ghl, 8(,.3 85.8 K'l 86.3 85.8 (I ~()

(II.) I (J 90) (1891 (189)

38UB2 t..r<J I. OC 0111 CC~ t,"" * *: ... (1 fJ7 7 ) 173H 34426 46325

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.... ·Two-~t~gc dXI~l L~ lurLlne !;ur,~: , .. r.·'J(".i~t.t'ri s! 1 C£ h},r,·,./!1 in thi!i t .:JtJl L iJre lil~f(jrr_ .

t.

"U00 r~Oj~l ~t~~c ~nd one ~x~u, hta~~ j'P-'/'.f."· I~t;a !(.iL" .IJV J~ • ..:.:tJ lccJjrlr.."loyy 1,/.:('...: •• OTl.rel: .-oXloll Sl<l'Jc!; J r.::I"""'('i/)r,)t.~d.

··.·!s i.:J rj:. :r)Cf) J:'1i·.;:' J,t..(' ,"'Jr

COMPRESSOR EFFICIENCY· CORRELATION 1985 TECHNOLOGY en

t

~ 4 <1 ~ Z w :E w > 1985 TECHNOLOGY o 3 a: SINGLE-STAGE CENTRIFUGAL Q..

:E COMPRESSORS > c.J Z w ~ u.

u.

w c.J "RELATIVE TO CURRENT (1977) lECHNOLOGY ~ CD S ~ OLI ________ -L ________ ~ ________ ~ ________ ~ ______ ~ 2 4 6 8 10 12 COMPRESSOR PRESSURE RATIO Figure 43." Compressor Efficiency 1985 Technology.

TURBINE Er;FICIENCY·CORRELATION 1985 TECHNOLOGY RADIAL TURBINES

6r

en I- a..

I- ~ <1 w en 1. ZERO CLEARANCE <!

w 2. REREF = 4.62 )( 10 a: u 3. 14 BU~DES (FULL) >- 3 4. 0.5 .;;; W-JO .;;; 1.5 u w U u..

u..

w w ..

~ III a: /

:\S = gJ..lH/O

:J I-

/

(UT/,':iJ)2 ·RELATIVE TO CURRENT (1977) 11'- / RADIAL TURBINE TECHNOLOGY 0, , I I , I 0.8 1.0 1.2 1.4 1.0 1.8 STAGE WORK COEFFICIENT - AS ~i~urF 44. Turbine Efficiency, 1985 Technology.

ONE STAGE CENTRIFUGAL COMPRESSORS ___ ENGII NE A

~ l >r

~ (:300)·- 147soK ~ w ~(2200°F) TURBINE

---

;: INLET 0 (280, Q., TEMPERATURE I- ~200 u..

(260) <I: 1311°K :t ft E B

en - /-ENGIN

(1900 F) (240) I- :::> 1255·K Q., I- (lS00·F) (220) :::> ------ S 7 S 9 10 COMPRESSOR PRESSURE RATIO z o

Ii: lS)

:i!: I 0.35 147S·K :::> TURBINE en (2200·F) Z INLET b....

o TEMPERATJRE u- 1255°K

-l !E (.5 ~0.30

(l000°F) .= 1311°K w- ::Jet: en

- (1900~F)

..:.: u..:t

-I

u- _co 0.25 u..-l 5. S 7 S 9 10

--

~u COMPRESSOR PRESSURE RATIO Q.,u..

en en NOTES 1) CRUISE CONDITIONS 2) AIRFLOW IS CONSTANT Figur~ 45. Turboprop Cycle Selection - Free-Turbine Engines The selection of a pressure ratio of 9.0 can be made with con- fidence that factors such as cost and weight would not dictate a lower pressure rat.io. Reducing the pressure ratio to 6.0 would increase the SFC 10 percent and the cost and weight would not be , significantly different.

The selection of turbine inlet temperature is more complex, since differences in cost and weight were expected between the cooled and uncooled _ eng ines. To allow a complete evaluation of these differences, a more detailed definition of an uncooled ver- sion [1255°K (1800 F) T4 of the baseline engine was accomplished.

This uncooled version of the free turbine baseline was designated Engine B and its characteristics are shown in Table 26. Engine B has the same airflow as Engine A but produces less horsepower due to its lower temperature. Specific fuel consumption is only sl ightly higher. At equal airflow, the we ight dlt ference between Engines A and B was found to be insignificant but the cost of the uncooled engine at equal airflow was approximately 6 percent :ess.

Cycle analysis re3ults shown in Figure 45 show that a pressull:~ ratio of 9:1 is near optimum for the uncooled engine in,terms of specific power and specific fuel consumption.

Another cycle trade-off involved pressure ratios higher than could be obtained with a single-stage centr~fugal compressor. At 1478°K (220QoF), the pressure ratio range was increased to a maxi- mum pressure ratio of 16. Two-stage centrifugal and axial- centrifugal compressors were evaluated at cruise conditions. The adiabatic efficiencies of the two-stage centrifugal compressor and an axial-centrifugal compressor relative to 1977 designs are shown in Figure 46 for a corrected inlet flow of 5 pounds per second.

This efficiency correlation was corrected for size effects for the GATE study.

t-' N 0"\ COMPRESSOR EFFICIENCY* CORRELATION 1985 TECHNOLOGY W-.jO /0 ~ 5.0 LB/SEC 3"10 CLEARANCE I- Z *RELATIVE TO CURRENT (1977) TECHNOLOGY <0.2 EXIT MACH NO.

w

:e

~ 3 o a: Q.

:E 1985 TECHNOLOGY >~ _ _ AXIAL/CENTRIFUGAL, COMPRESSORS uQ. 2 z.

w u -~ [ ~.,

---- -

u.

u.

w

~........- -< - - ---

) --.

U

--

~ co <!

~ 1985 TECHNOLOGY 2-STAGE CENTRIFUGAL COMPRESSORS o I I I ...L I I I I 4 6 8 10 12 14 16 18 COMPRESSOR PRESSURE RATIO Figure 46. Compressor Efficiency, 1985 Technology Figute 47 shows the results in terms of shaft power and SFC as a tunction of comrressor pressure ratio for three compressor con- figurations. The single-stage data is shown for comparison, since this study was performed at a slightly different cruise condition than that used for Figure 45. The two-stage centrifugal compressor is clearly superior to the axial-centrifugal compressor, due to its higher eff iciency. The difference in SFC between the single- and two-stage centr ifugal compressors was signif icant (5 percent at 12.0 pressure ratio) and prompted the def ini t ion of Eng ine C for aircraft evaluation. Engine C has a two-stage centrifugal com- pressor in lieu of the single-stage centrifugal compressor and its characteristics are listed in Table 26. The pressure ratio selec- ted for Engine C is 12.0. This provides near minimum SFC without incurrlng a large penalty in specific power. The axial-centrifugal compressor was not given further considel3tion.

Cross sections of Engines A and C are shown in Figure 48. The upper cross section shows the single-stage centrifugal compressor and the lower shows the two-stage centrifugal compressor.

4.4.1.2 Configuration Trade-Offs Substitution of an axial high-pressure turbine for the radial high-pressure turbine was one of the conf igura tion trade-of fs.

Figure 49 compares Engine A to Engine 0, which is the cooled axial turbine version of Engine A. Characteristics of Engine Dare listed in Table 26. The performance differences are due to lower axial turbine efficiency. The cost difference is due to the inserted blade design chosen for the axial turbine. An uncooled axial version, designat<!d Engine E, was also defined to show the differences between cooled and uncooled cost when using axial tur- bines. Engine E is only 5 percent more expensive than Engine S, whereas Engine D is 15 percent more than Engine A.

, ,".e it ttl e r .,- 1 - .• d' • l} t" z« ._;,...er:!1!'f:.::.:n;"...........-

I

\ [ ~ t r-- N co Q.

I 0.34 220 'iI:O.55 AXI/CENTRIFUGAL 1 STG CENTRIFUGAL I I I I ell

--

2 STAGE CENTRIFUGAL -J Z a 0.32 L \\: I 210 I I\--J.oF= 4""-....... I - .. .2 ... - Cl ~ Q.

Q. 280 I ~ :::> a: en 0.5 :: w z a ~ 0 30 I h~~<+I---t--~ ~ ..

..c • u ~ ~ 200 1 I I ... 1'< Q. Jr '" CI

-J --

~ w I- u..

:::> u..

<!

0.281 .. m. .. ".,- - j--?>'" ~ 260

U "L

190 I I / \: I

~.45

w AXI/CENTRIFUGAL Q.

en 0.26 L' __ -'-_--:-'::---::::-_-:: 6 8 10 12 14 COMPRESSOR 180 I , 240 PRESSURE RATIO 4 6 8 10 12 14 16 COMPRESSOR PRESSURE RATIO NOTES: CRUISE CONDITION T4 = 1478°K (2200°F) CONSTANT AIRFLOW.

Figure 47. Compressor Pressure Ratio Trade-offs

"

/ v 3NmN3

I

j

I

.

I

I

a 3NH:>N3

v 3NH:>N3 The major configuration trade-off was a change in spool arrangement from a free turbine (two shaft) to a single shaft. Past experience indicated that the single-shaft engine is less expensive than the free-turbine engine. Four single-shaft configurations were defined. Engine F is a 1478°K (2200 F) turbine inlet tempera- ture engine comprised of a single-stage centrifugal compressor, a rever se-flow annular burner,. and a two-stage turbine composed of one radial stage and one axial stage. Engine G is a 1255°K (1800 P) -version ot Engine F. Engines Hand J arc cooled and un..:ooled versions of Engine F with two stages of axial turbines substituted for the single radial stage. A comparison of Engines F and H is shown in Figure 50 and the characteristics of all four single-shaft eng ines are shown in Table 26. Although the sea- level, static, shaft power of the single-shaft engines is less than comparable free-turbine engines, they produce equivalent power at cruise conditions and have essentially the same core flow at their design points. On the basis of €equal cruise power, the single- shaft engines are less expensive tna'l comparable free turbines, ~:though they have a slightly higqer SFC.

4.4.2 Detailed Component Trade-Offs 4.4.2.1 Single-Stage Centrifugal Compressor The single-stage centrifugal compressor incorporated in most of the engines defined earlier requires three-dimensional (3-D) blading to produce the high efficiency assumed. Presently, research compressors employing 3-D blading are machined and are very expensive. Blading formed from straight line segments can be m~chined less expensively on 5-axis machines but incur a perforre- ance penalty. The alternatives for low-cost manufacturing are power metal titanium (PM Ti) or casting (steel or titanium). Con- ventional castings result in large performance penalties.

The powder metal approach promises mechanical properties " / " "" " .. .!, ..

, .~ :.': ~- "-", :) H 3NmN3 approaching those of a forging and efficiency levels equlvalent to machined designs. A comparison of the three alternatives is shown in Table 27.

The major difference between the PM Ti compressor and the aavanced cast compressor is approximately 4 percent in efficiency.

If the efficiency of the cast compressor is improved, the cost of the cast part will increase due primarily to a lower yielr' of acc~ptable parts. The difference between the machined anD PM Ti c;:;r.\pressor is component cost. The PH Ti approach is higher risK since attainment or efficiencies equivalent to the machined design in P.-\ Ti requires extensive research and development. The PM Ti approach was selected for all the candidate engines employing single-stage ~entrifugal compressors.

4.4.2.2 Combustion System The comDustion system, including fuel nozzles, did not require nor allow extensive trade-offs except with respect to the fuel noz- zles. Ceramic combustors were eliminated because of their develop- ment status. Based on current ceram~c combustor development pro- grams, this approach will not be feasible for low-cost man-rated engines entering service in 1988. Ceramic materials could be used in non-man-rated engines by this time period. Good combustor dur- ability at the 1478°K (2200 F) temperature level of Engine A will require advanced cooling schemes. The baseline design assumed that the cooling passages would be photoetched before the combustor is rolled and welded. lnco 617 was selected as the combustor material. Another candidate is oxide dispersion strengthened (ODS) sheet "alloys. ODS sheet is more expensive than lnco 617 and would have to provide an increaze in durability to be a successful candidd~e. The only change to the baseline combustion system resulting in a cost decrease was fuel nozzles. A low-cost airblast nozzle was ,:onceived, which resulted in a one-percent reduction in engine cost.

, TABLE 27. SINGLE-STAGE CENTRIFUGAL COMPRESSOR COMPARISON Engine A 9:1 Pressure Ratio 1478°K (2200°F) Turbine Inlet Temperature I PM Advanced Machined Ti Cast

I

I

I L

I Relative Efficiency 1.0 1.0 0.96 - Relative SFC 1.0 1.0 1.04 Relative Specific Power 1.0 1.0 0.96 Relative Engine Cost* 1.0 0.96 1.03 .

Relative Engine Weight* 1.0 1.0 1.04

I

*For equal power The improvements relative to the combuBtor and fuel noz7.les are understated. Current production or development combustors in the 373 to 746 kw (~dJO to 1000 hp) class are operating at 1366°K (2000 F) or less rather than 1478°K (2200°F). Thus the baseline turboprop eng ine incorpora tes a combustor that is signif icantly ir.lproved relative to todays combustor. The 1478°K (2200°F) technology has been demonstrated as feasible in recent research programs. The transition from research to development or production status is a majer task. It is complicated by the need to utilize alternate fuels such as diesel, synjet and broad specification kerosene. The ability to utilize these fuels may result in further improveme~ts if the alternate fuel is less expensive, e.g. d:esel.

4.4.2.3 High-Pressure Turbine The high-pressure turbine ir. the baseline engir.e, Engine A, is a radial turbine comprised of Mar-M 509 cast nozzles, vanes brazed to Hastelloy "X" bands, and a forged and machi.ned AF2-10A ·whe""l jo ined to a cast Mar-r-i 247 exducer. Cool ing holes were stem or electrostream drilled. The advanced technology approach was lamin- ated construction using photoetcll~d O.040-inch Waspalloy or- Astroloy. The laminated approach results in a six percent savings in engine cost, a small increase in efficiency and a small decrease in cooling flow. Taole 28 shows the results of the comparison ~etween the baseline and the laminated approach.

The other advanced technology trade-off performed in the high- pressure turbine area focused on the axial high-pressure turbine selected for Engine o. The baseline configuration hdd segmented cast nozzle vanes and .a rotor comprised of a forged and machined hub arid inser ted cast blades. The advanced technology approach co~sisted of laminated vanes and an integral laminated wheel con- str~cted from photoetched 0.010- and 0.020-inch sheet. The sheet in the axial turbine is thinner gauge than in the radial 2.35 '- TABLE 2B. RADIAL HP TURBINE COMPARISON Engine A 9:1 Pressure Ratio 1478°K (2200°F) Turbin~ Inlet T~mperature I Baseline Laminated Relative Efficiency 1.0 1.014 Relative Cooling Flow l.0 0.8 Relative SFC l.0 0.99 Relative Specific Power l.0 1.0: (l.~4 Relative Engine Cost* 1.0 Relative Engine Weight* 1.0 1.0 - - - ---- ---- ---- - - -For equal power "0 wheel. to accomrnoaate the higher curvature requirt:u. The thinner material results in higher proportionate cost. The change to laminated construction resulted in an increase in efficiency as well as a decrease in cost. Table 29 is a comparison of the baseline axial turbine design and the laminated high-pressure axial turbine.

The benefits identifiad for the lamlnated cooled radial tur- bine are applicable to Engine C, the two-stage centrifugal compres- s~. design, as well as to Engine F, the cooled single-shaft engine employing a radial/axial turbine. The laminated cooled axial tur- bine is applicable to Engine H, the cooled single-shaft engine employing an all-axial turbine.

4.4.2.4 Other Trade-Offs In addition to the trade-off studies discussed above, a number of other trade-oft studies were conducted including the following: o Clearance control o Single-stage versus two-stage power turbine o Conventional versus laser-hardened gears o Sheet metal versus cast turbine plenum The clearance-control trade-off study showed that efficiency could be increased 1.0 perce~c in the HP turbine and the LP turbine by reducing the turbine cle~r~nce from 0.015 to 0.010 inches. The cost pen.3lty for achieving this reduction in clearance is very small if passive means such as abradables are workable.

A single-stage power turbine was investigated for the free- turbine engines but the reduction in efficiency offset the reduc- tion in cost based on the airplane sensitivities developed earlier.

'.

TABLE 29. AXIAL HP TURBINE COMPARISON Engine D 9:1 Pressure Ratio 1478°K (2200 F) Turbine Inlet Temperature Baseline** Laminated Relative Efficiency 1.0 1.01 Relative SFC 1.0 0.99 Relative Specific Power 1.0 1.01 Relative Engine Cost* 1.0 0.94 Relative Engine Weight* 1.0 1.0 *For equal power **Single stage axial, cooled, inserted blades '- Gearbox cost reduction scudies identified laser hardening as an al ternative to ~onventional hardening. Estimates show a 3- percent reduction in eng ine cost due to a reduction in machin ing required on the laser-hardened gears.

A reducticn in engine cost of two percent was identified for sheet metal fabr ication of the turbine plenum as opposed to a cast/forged/s e~ metal assembly. This item was not recommended as an advanced technology program since it should result from normal development.

In addition to the engine cost savings described above, fur- ther cost savings were assumed for items such as static structure, bearings, and shafting. These technologies are classified as low r i5k and should result' from on-going company- and Government- sponsored R&D. The magnitude of the low-risk technology category was assumed to be a function of the remaining engine cost, after the cost of items that were specifically investigated was removed.

Specifically, in the case of engine A, the components that were subjected to trade-off studies represented approximately one-third of the engine cost. For these components the application of advanced technology resulted in a 16-percent reduction in total engine cost. The application of advanced technology to the remain- ing components, which account for two-third of the engine cost, was assumed to result in additional cost savings of 8 percent.

4.4.2.5 Su~~ary - Detailed Component Trade-Off Studies The results of the detailed component trade-off studies ident- ified cost reductions of 19-25 percent for Engines A through I listed in Table 26. For Engine A, the C03t reduction is 24 percent and breaks down as follows: , Compressor 4 % Combustor 1% HP Turbine 6% Laser Hardened Gears 3% Sheet Metal Turbine Plenum 2% Low Risk Technology Category 8% TOTAL 24% Additional cost reductions due to advanced technology are implicit in the baseline engine. The candidate engines listed in Table 26 include advanced technology such as: o High efficiency, high-pressure-ratio compressor o High turbine inlet tmperature in the case of the cooled engines o Integrally cast shrouds on the low-pressure turbine o Low-cost digital electronic fuel control The cost reductions due to these i terns were l,ot evaluated in detail. An approximation of their contribution can be ~rrived at by compar ing the baseline engine OEM cost wi th current production engine cost. Table 26 lists the OEM cost of the baseline engines before the cost reductions due to the advanced technology discussed

in 4.4.2.1 through 4.4.Z.4. For example, the specific cost Of]

Engine A is approximately 84 dollars per horsepower. Current ?ro- .

duction turboprops at equivalent power and production volume would sell for 100 dollars per horsepower or more. Therefore~ it can be inferred that the advanced technology in the baseline engine results in a cost reduction of 16 percent. Therefore, the maximum cost reduction due to advanced technology is the sum of tne advanced technology benef.~s identified with respect to the base- line and the advanced tE'; '010gy included in the baseline. For Engine A, this is the su;. Ilf 24 and 16 percent for a total of 40 percent cost reduction due to advanced technology.

4.~.3 Turbofan and Turboshaft Detailed engine traile studies were not performed on the tu;- batan and turboshaft engines with the exceptiv(. of the fan compo- nent and cycle on the turbofan. The benefits identified in the turboprop engine trade studies were applied to the turbofan and t~rboshaft engines where appropriate.

Turbofan cycle optimization studies identified small improve- ments in performance. Figure 51 shows the results of fan pressure ratio and bypass ratio investigations. Fan pressure ratio should be reduced to 1.4 and bypass ratio increased to 10 for minimum thrust specific fuel consumption. Relative to the baseline cycle, this change \vould result in a 4-percent reduction in fuel consump- tion and no loss in crui3e thrust. This decrease, however, is not sufficient to offset the difference between the fuel consumption of the turboprop and turbofan. Addi tional cycle work would involve the optimlzation of core pressure ratio and additional configura- tion work could include booster stages d~iven by the LP turbine, 2- stage centr ifugal compressors, and axial/o:::entr ifugal compressors.

None of these approaches, however, could significantly diminish the 61 percent difference in fuel consumption identified by the initial sizing results. Significant changes in the characteristics of the aircraft (higher speed and altitude, longer takeoff distances/more sophisticated high-life systems) would be necessary before the tur- bofan could compete with the turboprop.

" Ok '("" ,-0"; .. ,, ,. ...... -- '::'''j " .... ;. - ,,;,.- ... .- .... ~~~~~ ____ _ ....

"'- tv • 6096 M, 389 KM/H (20000 FT, 210 KNOTS) ISA • CORE PRESSURE RATIO - 9:1 0 o • T4 = 1478 K (2200 F) z ~ o

t

:E => en 260 I

~ 1.1[°.11

z ..J

8 1.0 0.10

I 240 I-

..J5

tn

I:J..... 0.9 - 0.09 ~ 220 979 => =>a: a: a: u.. ~ 0.8 0.08 :I: :I: 200 uen l- I-

-..J

I- 180 I-

G 0.71- 0.07 801

w w Z z ~ 0.6 - 0.06 CI) .70 Ii; => 1.1 1.2 1.3 1.4 1.5 1.6 1.1 1.2 1.3 1.4 1.5 1.6 a: :I: FAN PRESSURE RATIO FAN PRESSURE RATIO I- Figure 51. Turbofan Cycle Selection.

"

Two tur.bofan engines were defined and are designated Engines J and K. Engine J is the turbofan baseline engine defined earlier and Engine K is an uncooled version of it.

One turboshaft engine was defined and is designated Engine L.

It is a turboshaft version of Engine ~.

4.4.4 Summar~- Engine Trade Studies The application of -,,,dvanced technology to the GATE engines identified pE:formance improve~ents such as efficiency incr~ases, reductions in c001ing flows, ..lnd cost reductions. In addition to the advanced tec~nolog~ investigations, a study was ..llso conducted to determine the effect of high volume-production. The OE}l engine cost listed in Table 26 assumed d production rate of 1000 units per year. P0tential production of the GATE engines is 10,000 units per ycar. The scope of the GATE st~dy did not allow a detailed study of the ~enefit::; of tligh volume production. Fortunately, data was available trom the AiResearch GT601 gas turbine truck program. As part of the GT6Ul program, detailed estimates ~ere made for cost reductions attrlbutable to high volume production at the rate of 10,000 uni ts per year. The GT601 gas turbine is a rf>cuperated sh.:lft enyine in the .:;.n kw (600 hp) cluss. The benefit of high volume productiun was escablished for the GT601 by compari:1g esti- mates of engine costs at 1000 and lO,OOO units per year. The major benefit iden;.itied is the rerluct:ion of set-up time through use of dcdicat"d or captured machines. Setup is labor intensive and ac~ounts tor a large portlon of the fat-rication cost. Based on the GT601 studies, the cost at the udvanced-techno~')gy engine can be reduced by 40 percent due to the decrease in fabric-ltiLn cost .)ssociated \-Jitn ilign-volume production.

Table 30 lists the cost reduction~ due to adv~nced technology anti high volume production for the 12 ~anjidate engines. Figure 52 su~narize5 the cost reJuctions with respect to curr2~t production , ·~~,/", ....... d"'i')·trrl ee= . >.... - ..• ,,-eM ... -M. <- I I ~

!

,- ..

..

\

I

ThDLE 30. O.E.H. COST OF THE TWELVE ChNDljA'fE EUGIrH:S (1~77 $) fcee-Tuc~lne TucbopC0P . Single-Shaft Turboprops Turbofans l'urboslla{ t

--------,E

Cooled

~'";1~;~00';;1~c"""" COO,", ~;"OOO'O,- COO," "".~,,' CuO'"' ""000'"

\

Had!al 1 Radial Cooled Radial AXlal Axial HiJd/Ax Rad/Ax Axial Axial Cooled Uncooled Radial ~ A IJ C 0 E f G II I J K L 0"--- -- .. -.. - - - - - -

t Ilaficllne.*$ 39,908 31,434 42,495

45,891 39,530 37,344 34,426 46,32~ 38, S8 2 47,878 45,610 35,937 \ Aov Tech':' $ 9575 7934 r,495 10,558 7530 9344 6926 10,658 7382 IO,~45 9118 0270 ~. t.olJ till: COl; t ',ov Tecl,* 30,333 29,500 34,000 35,333 32,000 28,000 27,500 35,667 ll,500 37,3)) 36,500 27,667 lIi'Jh Vol ~ Adv Tech

-----' ~n'-J i nc Cos t*· $

18,200 17,700 20,400 21,200 19,200 16,800 16,500 21,400 18,900 21,900 22 .. 400 16,600 ______ . - L--.

*IO()O units **10,000 units NOTES: 1. ENGINE A 2. CURRENT PRODUCTION ENGINE - Sl00/HP ()v{ )(,-, I/.)/I ~ CURRENT

PRODUCTION ·V,c,v ~

Ii;. 1.0

) ( 0 "v"J

a u

{ 1- '(0 ~

w Z ADVANCED , e" Z w w HIGH PRODUCTION > 0.5 ADVANCED I- TECHNOLOGY oCt ..J W a: 0' V///////J [//""/4 vceC//CCI ~igure 52. Engine Cost Reduction

"

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

11" engines. Approximately,}Q'"" percent is due to advanced technology .l}u and ~ percent is due to high-volume production. Note that Figure 54 shows engine cost reductions with respect to current pro- duction turboprops and not the GATE baseline engines used as the reference in Table 30. Table 31 lists the performance character- istics of the 12 engines after adjustments for advanced technology improvements in component per formance and reduced cooling flows.

A 1'1 the turboprops and the turboshaft listed in Table 30 and Table 31 have identical core flows at their design points. The core flow of the turbo.~an is 30 percent higher.

The 12 candidate engines, as d(fined in Table 31, _ere evalu- ated in the next element of Task II - aircraft/engine trade-off studies.

4.5 Engine/Aircraft Trade-off Studies The selection Ol the optimum engines for the medium pressu- rized twin and the light twin was made on the basis of the following criteria: - Total three-year cost - Operating cost - Acquisition cost - Fuel consumption 'rhe twelve candidate engines described earlier were evaluated in the two aircraft configurations where appropriate. In addition to the twelve. candidate engines, the performance, weight, an9 cost 0 0 characteristics of the 1311 K (1900 F) versions of five of the engines previously disc~ssed were defined. The chara~teristics of these engines are shown in Table 3~.

11\bl.1. 31. LI.GWl:. C1iI,I,;'t:'lr;.<I~TlC:' WITH l,!JVM.CLD 1LCH~1Ol:.()C;,{ lH!'P.OVLMLtlTS r'[e~-Tuc~ine Tu[~o~rG~s Slnglc-~h~tt 1uruoprnps TurLoof.J,ns

I Tu,!Jo~h"lt

2 StrJ Ccnte I ( Cooled Uncool<:d C(hleJ Cooled Uncoolcu Co()le(J Unc()oled Coolcd Cooled Uncnule:d H.:.d1,,1 11",01",1 H"d1,,1 /,x I 'll lI"d/Ax \("o/"x /'x 1.:.1 I .. ", .. al Cooled Uncooled Hadiill /'x I 'll [) A B C L F G II 1 I J K 1.

- kw 365 265 357 350 260 ;18 2Jt 4423N 338611 31.5 SII!'/FW 324 233 (II!') (470) () 17 • Sl.~, 'fO, 1469 ) (355 ) (479) (348 ) (427 ) (434) (312) (994 Ib) (761 Ib) (489) S~·C, :;(.~, TO O. )22 kg/he/Kw 0.300 0.)08 0.284 0.313 0.325 0.326 0.319 0.333 k~ /11- he kg/Il-hr 0.300 o 040 0.036 i It,/Ile /top, (0.493 ) (0. SOSj (0.467) (0. ~ 14) (0.529) (0.533 ) (0.536) (0.523) (0.547) Ib/hr /Ib lb/hr/lb (0.49] ) (0. )90) (0.)~7~ Lnrjinc Weight

I

kg 95 9~ 99 95 95 86.3 86.3 !l5.8 85.8 111 III 75 ( It.) (210) (210) (210) (190) (190) (189) (189 ) (245) (165) (210) (219) iH5) t:n<plOc Cos t 18,200 17,700 20,400 21,200 19,200 16,000 16,500 21,400 18,900 22,400 21,900 16,600 S/kw or II 6t.9 74.0 49.9 57.1 60.5 52.7 69.:1 66.1 81.2 5.1 6.5 4S.6 lb ]7.2 49.9 42.6 ~5.1 55.2 39.3 52.1 49.3 60.6 S/hl' or ".~. ~ l~r!J~

I~ ~

b:J iii

C"I'!:r .at.l}l I~~~

~ l"?~

~ H I~

~ ~ ~

~ n~

~~~

....

..

-a ,.

...

co "'~iJLE 32. i311°r. (l~OO°F) TUi<BII,t. U;GIrI£ CJiid{AC'I'LRISlIC:>, St.A LLVEL, :;1'AIlL.MUJ ur .. , , TAKEOFf POWER, UUIU:>1ALLLD.

Slogle-:;halt 1.nlJ 1 r,(: T'Ii~(: " we Lvi an frc~-1'ur~ine TurLGrlOr Tu[tJopro!, () £nYlne De&ign~tioo M II P Q G,,~ Gen~rator Tur~ine kadial k"dlLlI A'I ... I kau/Ax All Axial 'furlJlr.c ,~lct l~ffilJ~ratuce, OK 1311 1)11 I 1I1 1311 1311 (0 t') (1900) (I '.l00) (l ·,00) (1900) (190(1} Compressor presHure ratio 1l.0~ El.3 &.3 7.6 7.6 f~rl (Jrc~suce ratio 1. 3.1 ili'P~~5 ratio 7.8 Cornpr~G~or C0rrccted 1.12 1.12 t I"w, Kg/scc 1. 5t .- I.n 1.22 (1~/:;ec ) (3.47 ) (2.£.9) (~. 69) (2.47 ) (2.47) ~cl thruut/sh"ft horsepo~cr 3605 II 29!- kw 286 kw 260 kw 259 kw (347 stlp) (1l10 1~) (Jill :;hp) (14~ shp) ()!iL "til" ~lJt:<.: , l 'l: (U(: 1 (;ur,:'; Utl!(Jll (Iii O. )04 kg/IIC Ikw 0.037 1"lIN-I.[ O.2~4 k'.l/hr/kw 0.301 k'JII,r/kw 0.306 kg/hr/kw (0.31;3 Ib/t.r/lt.) (0.4"2 H,/hr/hp) (0.494 b/hr/hl') (0.502 Ib/hr/hp, (0.499 1~/hr/hp) ~ngine wClgnt kg III '.l5 9'; 116 86 (l~) (210) (210) (190) (189 ) (~45) Logln<: ()t;M C05t S(l~77) 22,200 18,000 20, :!OO 16,700 20,200 " 4.5.1 Definition of Evaluation Criteria 4.5.1.1 Total Three-Year Cost The market forecast established that owners of turboprop- powered aircraft and the larger twins trade-in their airplanes for newer or larger versions on the average of every 30 months. This was the basis for developing total cost on a three-year basis.

The real total cost for general aviation aircraft varies con- siderably depending on the type of owner - corporate, personal, etc., the tax situation of the owner, utilization, and other fac- tors. Therefore, any total cost model can only provide data for comparisons on a relative basis. The total cost for the GATE air- planes is defined as the acquisition cost, plus the loan interest, plus the three-year operating cost, minus the trade-in price. The loan interest is based on a six-year loan at 10 percent interest and assuming 20 percent down. The three-year operating cost was based on 506 hours/year utilization. The resale value of the air- plane was assumed to be 75 percent of the acquisition cost, which is approximately equivalent to the high wholesale "blue book" price at three years for an aircraft with a mid-time engine, i.e., half- way through the overhaul period. Since an engine overhaul reserve is maintained as part of the operating cost, the time on the accounted for and no adjustment of the resale price is engines is required. Tax advantages or the imputed interest (time value of money) on the down payment are not considered.

4.5.1.2 Operating Cost The operating cost is separated" into variable and fixed costs.

Variable costs include: '0 o Fuel and oil o Inspection and periodic maintenance o Engine overhaul reserve o Avionics reserve o Propeller overhaul reserve Fixed costs include: o Hull insurance o Liability insurance Tie-down and landing fees, local taxes, and other miscellaneous items, such as catering fees, were not included.

Fuel cost was based on the 1977 average jet fuel pr ice of 70.85 cents per gallon. The oil cost for turbine engines was negligible. Inspection and periodic maintenance was based on a survey of Phoenix fixed base operators. The data received from th is survey ind icated that per iod ic inspection and maintenance costs after 250 hours of utilization for aircraft in the ~ight-and medium-twin classes ar~ as follows: a Ten hours airframe labor o Fifteen hours engine labor a $300 to $500 for ·parts a Labor cost, $17.50 per hour These costs were for reciprocating-powered airplanes. On the same basis, turboprop-powered aircraft have reduced engine labor and parts cost. Typical periodic maintenance labor hours for current turboprops is 2.5 hours per 250 hours or 1 hour per 100 flight hours. The lower end of the parts cost range ($300) for the recip- rocating-powered aircraft was selected for the turbine-powered air- craft. Tne parts cost for the reciprocating engine= is believed to '.

cover failures such as magnetoes, oil pumps, etc. The parts cost for the turboprop does not cover failures such as the fuel control, thermocouple harnesses, speed pickups etc., since these are rela- tively high cost items on a turbine engine. Detailed estimates of these costs are related to mean times between failure for these components, which were not estimated in this program. Some allow- ance for component failures is included in the overhaul reserve.

An allowance of 50 percent of the OEM engine cost at 3500 hours is provided for overhaul. Data on overhaul cost is available for AiResearch engines, as well as other gas turbines. The avail- able data shows overhaul cost to vary between 15 and 60 perc~nt of the original engine cost. The variance is due to different philos- ophies regarding replacement versus repair, ~emanufactured versus overhauled, and overhaul specifications. The higher side of the range was chosen not only to allow for realistic overhaul but also to provide for random component failures and per iodic hot-end inspections. Hot-end inspections on current turboprops are required at between 1500 and 2000 hours. The effect on operating cost ranges between 4 and 18 cents per hour per engine:. In the future, this cost may be reduced further through higher durability, on-the-wing inspection, and modular constiuction.

The avionics reserve was based on a formula used by Cessna.

This formula computes the avionics overhaul reserve as 10 percent of the avionics options purchase price at 1000 hours. For example, if the avionics options are $30,000, the overhaul reserve is $3.00 per hour.

Available data suggests a propeller overhaul cost of 750 dol- lars for propellers used on curre~t light and me:dium twins. Ade- quate data on time between overhauls was not available. However, recent studies suggest that there is no reason for the propeller not to have a TBO equal to or better than the engine. The TBO interval for the propeller was therefore selected to be 3500 hours.

'.

Hull insurance yearly rates were obtained from Cessna and vary from 1.0 to 1.5 percent of the acqui~ition cost of the airplane on a sliding scale. A rate of 1.25 percent of the acquisition cost was asnumed for all airplanes.

Liability insurance rates are a function of the number of pas- sengers. According to this schedule, the annual rate for the mec- ium pressurized twin and the light twin is 550 dollars per year.

4.5.l.3 Acauisition Cost The acquisition cost of the airplanes is a function of the airframe weight and the maximum speed. The acquisition cost algo- rithm, shown below, was supplied by Cessna.

= 0.008031 (W )1.76063 x (V )0.486512

Airplane Retail Price E max + [Retail Cost of Engine(s) Propeller, Optional Equipment] Standard Empty,Weight minus the weight of the engine(s),

W =

E propeller and optional equipment

v = Maximum speed (kts)

max The retail cost of the engines and propeller is the OEM cost mUlti- plied by 1. 75.

4.5.2 Trade-off Studies 4.5.2.1 Turbine Inlet Temperature Trade-offs The medium pressurized twin was evaluate~ with 1255, 1311, and 1478°K (1800, 1900, and 2200°F) versions of the turbofan engine and the free-turbine turboprop engine equipped with a single-stage com- pressor ana a radial gas generator turbine. Tables 33 and 34 list TABLE 33, TURBOFAN-POWERED MEDIUM PRESSURIZED TWIN, TURBINE INLET TEMPERATURE TRA[E-OFFS J M K Engine Designation Turbine Inlet Temperature, oK 1478 1311 1255 (0 F) (2200) (1900 ) (1800 ) Engine Net Thrust, SLS, T.O., N 4459 4632 4859 (1002 ) (1041) (1092) (lb) 0.037 Engine TSFC, SLS, T.O., kg/N-hr 0.039 0.036 (lb/hr/lb) (0.390) (0.363) (0.357) Engine Weight, kg 121 140 160 (lb) (247) (309) (352 ) Engine Cost, $ (1977 ) 22,500 26,800 28,700 Airplane Gross Weight, kg 27:a.a 2793 2929 (lb) (5987) (6151) (645l~ 1485 1559 1624 Airplane Empty Weight, kg ( lb) (3270) (3433) (3578) Acquisition Cost, (1977 ) 250,497 266,847 2BO,672 $ r 59.00 Operating Cost, $/lIr (500 Hrs/Yr) 56.72 63.06 220,886 Total Cost, $ (1977 ) 197,806 209,049 Interest, $ 50,099 53,369 54,134 3 \r. Operating Cost, $ 85,083 94,584 88;503 'frock-In, $ 187,873 201,533 210,504 fuel Consumplion, li ter/hr 200.98 204.51 221. 49 t-' lJl (gal/hr) (53.03) (53.96) (58.44) VJ --

"

......

TABLE 34. TURBOPROP-POWERED MEDIUM PRESSURI ZED TWIN, 'l'URBINE lJ1 INLET TEMPERATURE TRADE-OFFS

"'"

Engine Designation A N B Turbine"Inlet Temperature, oK 1478 1311 (0 F) (2200) (1900) (1800 ) Shaft Power, SLS, T.O., kw 312 339 356 (hp) (419 ) (454 ) (478 ) 0.300 0.302 0.308 Engine SFC, kg/hr/kw (0.493) (0.496) (0.505) (lb/hr/hp) ~ 84 113 131 Engine Weight, kg (lb) (l86 ) (248) (298) Engine Cost, $ (1977) 16,200 19,900 22,100 Airplane Gross Weight, kg 2297 2459 (5672) (lb) (5060) (5416) Airplane Empty Weight, kg 1366 1497 1576 (3009) (3297) (3471 ) (lb) 202,854 229,553 245,741 Acquisition Cost, $ (1977) 39.38 42.88 45.57 Operating Cost, $/Hr (500 Hrs/Yr) Total CUbt, ~ (1~77) 150,357 167,616 178,944 40,571 45,911 49,148 Interest, $ 59,073 64,317 68,361 3 Yr. Operating Cost, $ 152,141 172,165 184,306 ~Tade-In, $ 122.06 131. 44 140.00 Fuel Consumption, li ter liar (gal/ht) (32.21) (34.68) (36.94)

"

- ... _-._ .. -.- ... -. -.--------~---------

the character istic data for these aircraft and Figures 53 and 54 show the relative values of the evaluation criteria and engine cost. The results show that the 1478°K (2200°F) turbine inlet temperature results in superior airplanes. The additional cost of the high-temperature components is offset by higher specific power or thrust, which results in smaller components and lighter weight.

The effect of the smaller engines on aircraft drag and nacelle weight was not accounted for and would result in additional, though small, improvement. It is of interest to note that the difference 0 0 between the 1478°K (2200 F) turbofan and the 1311 K (1900 F) turbo- fan is not as great as that between the l478°K and 1311 K (2200 F and 1900°F) turboprops. If a turbofan-powered aircraft was of interest, more detailed turbine inlet temperature comparisons would be desirable, as well as further optimization of fan pressure ratio and compressor pressure ratio as discussed earlier.

4.5.2.2 Engine Configuration Trade-off Studi~s The remalnlng engine configuration trade-off studies concen- trated on lll;? high-temperature configurations [1478°K (2200°F) 1.

The results are shown in Tables 35 and 36 for the pressurized and light twin and are summarized in Figures 55 and 56. In terms of the primary evaluation criterion, total cost, engine F (radial/axial single-shaft) is superior for both applications. The all-axial single-shaft engine, H, and the free-turbine engine, A, are within 5 percent of F. The two-stage centr ifugal compressor conf igur- 3tion, eng ine C, by virtue of its high-pressur'=! ratio and com- pressor efficiency, has the lowest fuel consumption in both appli- cations. Operating cost differences between engines A, C, F and H are very small. Aircraft acqulsi tion and engine cost show more pronounced differences.

If both applications and all evaluation criteria are consid- ered, engines A and F appear to be the best selections, with the " -.~-~--.......,.,.

,. .

....

• I

U.9' I - I ~. I _ I. I g

TOTAL COST FUEL OPERATING ACQUISITION ENGINE CONSUMPTION COST CCST COST , Figure 53. Turbofan PreSSUrized Twin Turbine Inlet Temperature Trade-Offs •

. '

0 0

D 1478 K (2200 F)

•

~ .

0 o ,

rn 1311 K (1900 F)

0 o FREE TURBINE ~ 1255 K 11800 F) SINGLE-STAGE CENTRIFUGAL COMPRESSOR RADIAL GAS GENERATOR TURBINE t 1.3 -1 w t :> ..J j ~ >

l

w > ~ ..J W a: r.

~

r//'~{- ;.',ylm v.t

:, "

l'0r-_ - r-~<m - - ~~ -- Y.xi.:--

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TAL COST FUEL _ ..... _a_ ... _ A __ III_nl,.' _JIIo.I_.&1

CONSUMPTION COST Figure 54. Turboprop Pressurized Twin Turbine Inlet Temperature ~rade-Offz • ......

\1l -~ ,.

1;,

.'

....

1> '" TABLE l<;. PHESSURIZED 'l"rlIN, ENGINE CONFIGURATION TRADE-OFF STUDIES 1478°K (220ll"f) ENGINES '" -, r-- -- J::ngine Type Turhofan TllC ~,,?P~op f---- -, r'- ::-Turbint.

Sin-'Lle-Shaft ...

A C· J J::ngine ueE~gnation D F II ~.

t::n" inc Ilet The ust/Powee, kw l12 318 314 274 256 4459N (hpi (4191 (4271 (421) (3601 (3431 (1002 1111 " Engine SFC, kg/hr/kw O.lOO 0.281 0.281 0.l25 0.319 0.040 kg/N-h (0.49l) (0.462) (0.514) (0.533) (0.523) (0.390) Ib/hr/lb (lb/hr/hp)

t

,~ Engine Weight, kg 84 92 89 75 73 121 (lb) (186) (202) (195) (165) (160) (247 ) Engine Cost, $(1977) 15,000 17,900 22,500 16,200 !8,700 19,500 Airplane Gross Weight, kg 2,297 2,285 2,339 2,245 2,223 2,728 (lb) (5,033) (5,15)) (4,946) (4,897) (5,987) (5,060) " Airplane t::mpty Weight, kg 1,366 1,384 l,nl 1,319 1,309 1,485 " (lb) (3,009) (3,049) (3,042) (2,906) (2,883) (3,270) 250,497 Acquisition Cost, $(1977) 202,854 212,233 215,636 194,109 203,120 Operating Cost, $/lir

---

41.89 56.72 (500 IIrs/'ir) 39.38 39.21 38.81 39.30 145,568 150,355 197,806 Total Cvst, $(1977) 150,357 154,313 159,865 Interest, $ 40,571 42,446 43,127 38,8n 40,625 50,099 3 'ir Operating Cost 59,073 58,809 62,829 58,218 58,950 50,099 Trade-In, $ 152,141 159,175 161,727 145,581 152,340 187,673 116.05 128.46 121.96 119.01 200.98 Yucl Consumption, li ter /hr 122.06 (gal/he) (32.21) (30.62) (33.90) (32.16) (31. 40) (53.03) GoJ!i (icn~rtJt()r

~ ~ HJI ~

H ~

*"wo-SlaCju f'unlrifug.,1 Compressor "

--

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224J 2L I~ 2285 2202 2179 AlrpJ~n~ Gr~Gu WCiqtlt, kq ( )t,) (4,872 ) (5,034) (4,799) (4,940) (4,851 ) 12)8 1230 126) 1267 1280 l'lCl'lanc Empty Weight, r.~ (2, ~27) (2,710) r (lb) (2,781 ) (2,790) (2,820) 18), ~ 17 174,'123 186,141 t.C'l.JI!;itlf,n C,,:ot, S(I'>77) 178,706 i90,767 (Jr".·rat Hl'l Cust, S/lfr J7.09 36.60 37.07 i~(J(I IIrr;IYrj 36.42 39.34 136,050 In,389 144,85', 133,622 139,606 1f,tal Co"t, $(1177) Interest, $ 35,741 36,783 38,154 34,985 37 228 54,627 59,009 54,906 55,605 1 Yr. Operating Co~t 55,632 14),075 Ill, 1 ~2 139,6011 'rrade-In, S 134,0)0 137,938 112.94 119.46 112.56 125.26 117. 07 Flu ... 'l Cl.Jntjuffilll Ion, 1 J t(.'! Itl( (lL 52) (29.70) (33. 05) (30.89) (29.&0) (gal/he) G~!,; (;, ... o'.'r (J t()r

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fixed-shaft engine (F) having a slight edge. In considering other applications, particularly rotary wing, the single-shaft engine is not as attractive as the free-turbine engine. The rotary-wing applications ,are important since they represent at least 10 percent of the market and would probably be the first applications (in a turboshaft version) of an engine of this type.

To allow a choice between the two, Task III was oriented toward investigating the possibility of a common core, which would allow the fixed-wing market to enjoy the benefits of the single- shaft approach and give the rotary-wing market the free-turbine engine it requires.

4.6 Benefit Analysis A benefit analysis was conducted to compare the GATE appli- cations studied to aircraft powered by current technology turbo- props, turboshafts, and reciprocating engines. Comparison to existing aircraft would be misleading, due to technology differ- ences between 'the GATE a ircraft and current aircraft. The GATE aircraft incorporate small improvements in aerodynamics and mate- r ials technology. The compar ison was also done using the same ground rules for computing aircraft acquisi tion, .operating, and total cost, and for identical missions.

4.6.1 Current Technology Turboprop There is no suitable current technology turboprop that can be compared to the GATE engines. Turboprops employing recent tech- nology are in the 447 to 746 kilowatt (600 to 1000 shaft horse- power) class and some of these are improved derivatives of engines designed 12 to 15 years ago. The Detroit Diesel Allison 250 Engine has been very successful and is at·the upper end of the GATE size but is used primarily in turboshaft applications. It is a compact, 162.

,

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durable, light-weight engine but only approaches the specific fuel consumption of the PT6, TPE331, and LTPIOI in recent growth ver- sions of approximately 447 kw (600 hp). A comparison of the 250 series with the GATE engines would show large benefits for the GATE designs but this approach would not yield a comparison to what could be achieved with readily available low-risk technology. To allow a fair comparison, a "current ~echnology" engine was synthe- sized.

The turboprop configuration selected was a two-spool eng~ne comprised of a two-stage centrifugal compressor, driven by a single-stage, axial, cooled turbine, a reverse-flow annular combus- tor, and a two-stage power turbine. Cycle characteristics and per- formance of the syn:hesized current-technology turboprop is com- pared to the GATE free-turbine engine (eng\ne A). The results are shown in Table 37. The synthesized current-technology turboprop has lower fuel consumption· than current gas turbines of comparable size, but is slightly heavier. The relatively low power-to-weight ratio is consistent with the philosophy followed on the GATE designs. \'leight was traded in the GATE engines for lower manu- factur ing cost. T,he cost of the current-technology turboprop was der ived by adding the increased costs of a t\oJo-stage centrifugal compressor, an inserted blade, cooled axial turbine, and a current electronic fuel control to the baseline cost of Engine A. This cost was then reduced by 40 percent for the effects of high-volume manufacturing.

Table 38 details the differences in cost, efficiency, specific fuel consumption, specifi~ power, and weight by component. Note that cost and weight differences are not specified at equal power.

The eng ines are compared as they are def ined in Table 37. For equal power, the cost and weight differences would be greater. The advanced, high-pressure, radial turbine is a major contributor to the gains indicated. Compared wi th the current-technology ax ial turbine, the laminated radial turbine results in a 16-percent savings TABLE 37. CURRENT TECHNOLOGY TURBOPROP CHARACTERISTICS Current Technology GATE oK Turbine Inlet Temperature, 1478 1478 (0 F) (2200) (2200) Cycle Pressure Ratio 8.3 8.3 Compressor Corrected Flow, kg/sec 1.22 1. 22 (lb/sec) (2.69) (2.69 ) kw Shaft Power, SLS, T.O., 315 365 (hp) (422) ( 489) Specific Fuel Consumption, kg/hr/kw 0.345 0.301 (lb/hr/hp) (0.567) (0.493) Engine Weight, kg 99 95 ( Ib) (219) ( 210) Engine Cost, S(1977) 27,671* 18,200* I No. of Compressor Stages 2 No. of HP. Turbine Stages 1 Axial

1 Radi:~

No. of LP Turbine Stages 2 *For Production Quantities of 10,000/year ..

TABLE 38. ADVANCED TECHNOLOGY BENEFITS* (DOES NOT INCLUDE SYSTEM EFFECTS) .:lWT <lCost <lSFC <lTJ I <lSHP % %

Technology % PTS % 1 WA'

PM T. Single-Stage Compressor -4 -1.0 +1.4 -1.56 -4.3 Low-Cost Fuel Nozzles and (l Combustor -1 0 0 0 I Laminated High-Pressure Turbine** +9.8 -7.4 +9.4 0 -16 I High-WorkfLow-Speed Power I Turbine 0 +6.0 -7.0 +8.11 I -2 Electronic Control 0 0 0 I Laser-Hardened Gears -3

- 0 0

Sheet Metal Turbine Plenum -2 C 0 0 Other -6 0 0 0 Total -34 -13.0 +15.9 -4.3 ~-- - I *Relative to current technology engine (Table 37) **Includes effects of turbine cooling and clearance control , a 9.8 percent increase in efficiency (ir.!proved in engine cost, initial value on page 84), and a 7.4 percent relative to the decrease in SFC. This table also does not include the airframe/ engine synergistic effects, i.e., the total system benefits from the lower weight, improved efficiency, and higher specific power of the GATE engine.

The turbine inlet temperature selected for the current- technology engine is 339 to 366°K (150 to 200°F) beyond the capa- bility of the latest TPE331 turbines and, as such, is somewhat beyond "readily available, low-risk technology".

The results of the comparison of the GATE free turbine engine (engine A) and the current-technology turboprop, as installed in the GATE airplanes, are shO\m in Table 39. The same propeller was used for the current-technology and GATE engines.

4.6.2 Reciprocating Enqines Currently, a number of reciprocating engine concepts are being investigated for aircraft applications. These include: o Rotary engines o Light-weight diesels o Advanced spark-ignition engines The rotary and diesel engines were considered only very briefly since current information en performance, durability, weight, size, and cost were not readily available. Both engine types could con- ceivably competp. with the conventional, reciprocating, spark igni- tion engine but durability, performance, and weight are problems that must be surmounted.

Quantitative data on the advanced, reCiprocating, spark igni- tion er.gines was also not readily available. Various projections have been made as to the level of fuel consumption improvement that will be possible. These projections range from 0 to 20 percent, , TABLE 39. COMPARISON OF GATE AND CURRENT TECHNOLOGY ENGINES Medium Pressurized Twin Light Twin GATE Current GATE Current Tech Engine Tech Engine TPE A TPE A Shaft Power, kw 327 312 253 244 (hp) (438 ) (419) (339 ) l327 ) Engine SFC, kg/hr/kw 0.345 0.300 0.345 0.300 (lb/hr /hp) (0.567) (0.493) (0.567) (0.493) Engine Weight, kg 103 80 64 (lb) (227 ) (186) (176 ) (140) , Engine Cost, $(1977) 28,454 16,200 .i 23,480 13,500 Airplane Gross Weight, kg 2501 2297 2393 2243 .. (5510) (lb) (5060) (5271) (4940) Airplane Empty Weight, kg 1465 1366 1357 (lb) (3226) (3003) (2990) (2781) Acquisition Cost, $(1977) 238,699 202,854 218,735 178,706 Operating Cost, $(1977)/lir 47.&9 39.38 43.97 37.09 $ (1977) 165,555 150,357 164,386 136,050 Total 3 Year Cost, 47,740 40,751 43,747 35,741 Interest 3 Year Operating Cost, $(1977) 71,835 59,073 65,955 55,632 'rrade-in, 179,024 152,141 164,051 134,030 $(1977) I Fuel Consumption, liter/hr 146.82 122.08 143.83 119.30 ~ 0\ (ga1/hr) (38.74) (37.95) (31.52) '32.21) -- - ----- -..J " but factors such as durability and cost were not always considered.

To compare reciprocating spark ignition engines to the GATE tur- bines, two levels of engine performance were assumed: o Current technology o Fuel consumption improvement of ten percect Reciprocatin~ ~ngines representative of both levels of performance were evaluated in the GATE airplanes. The ground rules and assump- tions followed in evaluating the reciprocating engines are listed in Table 40. The only change made in the basic empty weight break- down of the airplanes is the change in engine weight and its effect on the weight of the wing, empennage, and landing gear. Nacelle and other engine-related weights were not changed and the propeller weight and cost were identical to those used for the gas turbines.

Specific fuel consumption for current-technology reciprocat- ing engines at cruise conditions varies from 0.262 to 0.305 kg/~r/kw (0.43 to 0.50 lb/hr/hp). The 0.268 (0.44) level is typical of moderate sized engines.

I Maximum cruise power for the reciprocating engines was limited to 75 percent of maximum power. Some current applications allow 79 percent of maximum but the majority recommend 75 percent for acceptable life.

The acquisition cost of the reciprocating-engine-powered air- craft was developed using the same equation supplied by Cessna fo~ the GATE aircraft. The optional equipment and propeller cost \oJas identical to that used for the gas-turbine-powered aircraft. The reciprocating engine cost was estimated using data developed in Task I. The 20-percent increase in reciprocating engine pr ice assessed in Task I was not applied in Task II.

The major differences in operating costs were the fuel price, oil, the cost of engine overhaul, and inspection and routine main- tenance costs. The price of· aviation gasoline is the national average pr ice for 1977. The cost of oil and the eng in'=! overha.,;l , TABLE 40. GROUND RULFS AND ASSUMP'rIONS Evaluation of Reciprocating Engines Medium Pressurized Twin Light 'fwin Current Advanced Current Advanced 'l'echnology Technology Technology Technol~9Y Engine Type Turbocharged Turbocharged Naturally Naturally Aspirated Aspirated Power-to-Weight Racio 0.7 0.7 0.7 0.7 SFC, kg/hr/kw 0.268 0.241 0.268 0.241 (lb/hr/hp) (0.44) (0.396) (0.44) (0.396) Max Cruise Power 75% of 75% of 75% of 75% of Max Power Max Power Max Power Max Power Acguisition Cost Basis Cessna Cessna Cessna Cessna Equation Equation Equation Equation Propeller * * * * ..

Optional Equipment * * * Engine Cost, 44 44 39 39 $/kw (29) (29) ($/hp) (33) (33 ) I OEerating Cost Fuel Price,¢/gal 77 77 77 77 Oil, 0.45 0.45 0.45 0.45 $/Hr TBO, hr 1500 1500 1500 72% of 72% of Engine Overhaul Cost 87% of 87% of O.E.M. Cost O.E.M. Cost O.E.M. Cost O.E.M. Cost -~ ----- -- .-.

\0 *Same As Gas Turbine , cost was obtained from data made available through Cessna dealers.

They provide a service to potential customers called the Transpor- tation Analysis Plan, which analyzes the operating cost of Cessna aircraft.

The results of the comparisons of the optimum GATE engine to the reciprocating-powered airplanes are shown in Table 41. The primary difference between the gas turbine and the reciprocating engine aircraft is engine weight. The difference in engine weight - approximately 454 kg (1000 pounds) - results in an empty weight increase of over 772 kg (1700 pounds). This increase in empty weight increases fuel required, the size of the engines, and the acquisition cost. The difference in operating cost is primarily due to the difference in eng ine overhaul reserve per hour and secondarily to higher fuel prices, lower volumetric energy content of aviation gasoline, and the higher inspection and routine main- tena:1ce cost.s of the reciprocating engine. The engine overhaul rate per hour of utilization is higher for the re~iprocating engine due to: o An overhaul period of less than half that of th: gas tur- bine o Higher percentage of the original engine price for over- haul o An original engine cost equal to that of the gas turbine, due to higher power requirements For the medium pressurized twin, the higher per hour overhaul rate accounts for 54 percent of the operating cost difference between the gas turbine and reciprocating engines. Fuel and oil cost, inspection, and insurance account for 29, 11, and 6 percent of the operating cost difference, respectively.

, .... t. ..... - t ::;;.>~~ .... - J- . J:3 ... -~,', ___ _ t / TABLE 41. COMPARISON TO REA:IPROCATIIlG POWERED AIRCRAFT Light Twin Pressurized Twin GATE Current Advanced GATE Current Advanced Engine TechnolO<JY Technology Engine Technology Technology "A" Recip. Recip. "A" Reeip. Recip.

~nglne Shaft Power, kw 312 346 338 244 283 277 ,371) (419) (464 ) (453) (327) (380) (t·P) Er.11ne SfC. kg/hr/kw 0.1 0.2611 0.241 0.300 0.268 0.241 (0.493) (O.H) (0.396) (0.493) (0.44) (lb/hr/t.p) (0.396) 24:) Engine Weight, kg 84 301 292 64 250 (le) (l86) (664) 1643) (140) (550) (536) Engine COBt, S(1~77) 15,)12 14.949 11,020 10,759 16,200 13.500 Airplane Gross Weight. kg 2297 3097 2997 2243 2815 2715 (lb) (5060) (6821 ) (6601) (4940) (6200) (5980) Airplane E=pty Weight. kg 1366 2117 2068 1263 1846 1800 (lb) (3009) _ (4664) (4065) (4555) (2781) (3965) Alrplane Acquisiton Cost. $(1977) 252,;972 202. BS4 245.534 178,706 206,694 199,781 Oporating Cost, (500 Hrs/Yr) )9.38 6), 63 59.88 37.09 51. 32 48.26 S/Ilr Total 3 Ye~r COBt $(1~17) 1~0.351 20'l.2k9 200,308 131;,050 169,989 162.291 Interest 40,751 50.595 49,101 35.741 41,339 39,956 3 Yr Operdting Cost 59,073 95,451 89,817 55,632 76,977 72,396 Trade-in 152,141 la9,729 184.151 134.030 155,021 149,836 His~ion Fuel Consu=ption. liter/hr 122.0e 145.5 129.6 119.46 129.9 117.1 (3,.21 ) (38.4) (34.2) (gal/hr) (ll.52) (34.3) (30.9) - -- --- --- ....

....

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4.6.3 Current Technology Turboshaft Analysis. of the single-engine utility helicopter was a com- bined trade-off study and benefit analysis. The analysis was per- formed by Bell Helicopter. The turboshaft engine used in the analysis was engine L described in 4.5. The characteristics of the light helicopter and the results of sizing and mission analysis are listed in Table 42.

Bell compared the above results with a helicopter that used a current-production turboshaft engine. These results are shown in Figure 57. The engine characteristics of the advanced-technology engine (GATE) and the current production engine are shown at the bottom of the figure. At a constant payload of 377 kg (830 pounds), the advanced technology engine results in a helicopter that is 20-percent lighter. This reduction in gross weight trans- . lates to lower a~quisition and operating costs, as well as markedly lower fuel consumption.

, TABLE 42. SINGLE-ENGINE UTILITY HELICOPTER SIZING AND MISSION ANALYSIS RESULTS* Parameter Value Gross Weight, KG (T.B) 1109 (2442) Empty Weight, KG (LB) 549 (1210) .............

I I Payload, KG (LB) 377 (830) Max Speed KM/HR (KTS) 215 (116) r Cruise Speed, KM/HR (KTS) 189 (102) Range, KM (N.M.) 611 (330) Critical Altitude, M (FT) 1967 (6450) \ \ SHP, SLS, T.O. Max Power, KW (HP) 204 (274 )

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./ I I I 200_ 1200 1300 1400 1500 1600 900 1000 1100 GROSS WEIGHT, KG , I I I 2000 2400 2800 3200 GROSS WEIGHT, LBS ENGINE CHARACTERISTICS ADV CURRENT TECH PRODUCTION SFC, SLS, T.O. 029 KG/HR/KW 0.48 KG/HR/KW (0.47 LB/HR/HP) (0.78 LB/HR/HP) , POWER/WEIGHT 4.57 KW/KG 3.51 KW/KG (2.17 HP/LB) (2.13 HP/LB) Figure 57. Comparison of a Turboshaft Version of the GATE Free-Turbine Turboprop with a Current-Production Turboshaft in a Light, Utility, Single-Engine Helicopter.

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SECTION V 5.0 COMr-tON-CORE CONCEPT The common-core concept study was envisioned as an effort to compromise engine requirements, as def ined in the broad scope trade-off studies, and define a single gas generator, which would satisfy all requirements to a degree and achieve lower cost through parts commonali ty. It was anticipated that different types of engines would be optimum for the various applications, i.e.

o Turbofan o Turboprop o Turboshaft The Task II results showed that all fixed-winq applications require a turboprop. Turbcfans would not be competitive unless cruise speeds were increased and takeoff performance was relaxed. The most signif-icant difference between optimum engines is believed to lie in the difference between turboprop and turboshaft configura- tions. The optimum engine for fixed-wing applicati~ns, by a small margin, is a single-shaft configuration. Although a single-shaft turboshaft is workable for single-engine rotary-wing applications, it introduces large compromises and may be unsatisfactory in twin- engine installations. The common-core concept' study was therefore oriented toward determining if there was a common core that would app'roach the characteristics of the single-shaft engine in the fixed-wing applications, and was suitable for use as the core for a free-turbine turboshaft for rotary-wing applications.

The common core that resulted is shown in Figure 58 as are the turboprop and turboshaft engines which result from this common core. The common core is comprised of a single-stage centrifugal compressor, a reverse-flow annular combustor, and a radial inflow turbine.

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Figure 58. Common-Core Concept

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The single-shaft engine shown on the lower half of the figure requires one more axial turbine stage than the optimum single-shaft turboprop with the radial/axial turbine shown in Figure 50. The additional turbine stage is the result of unloading the radial turbine in the optimum single-shaft engine. The radial turbine must be unloaded so that it is compatible wi th the turboshaft engine shown in the upper half of Figure 58. This engine is a rear- drive, free-turbine turboshaft. The radial turbine in the optimum engine drives .the compressor and ~lso supplies part of the output power. As a part of a common core for the free-turbine turboshaft, it only needs to drive the compressor.

The benefits of the common core are: o Increased parts commonality o Increased turbine efficiency due to more lightly loaded stages.

The disadvantages are: o Additional turbine stage in single-shaft version (cost and we ight) o High-temperature power-turbine bearing compartment in free-turuine turboshaft o A common-core turbofan is not easily derived.

The common-core approach would provide a single-shaft engine for the fixed-wing aircraft and a free-turoine turboshaft for the helicopter. However, a preliminary study indicates that the superiority of the common-core, single-shaft engine over the non- common-core, free turbine engine (Figure 48) would be diminished ....c..... ___ """-----~- 'ref ............. .;;.;.;.", ____________ _ ....

-----.-------" ' .. - ---- and the relative difference would be very small, due primarily to the cost of the additional turbine stage.

On the basis of this analysis, Engine A, which is a free- turbine engine comprised of a single-stage, centrifugal compressor and radial turbine, is recommended as the preferred engine config- uration. It is close to optimum, offers turbofan derivatives, and is compatible with rotary-wing applications.

, - ~--'---'- .-~- .. -- .. _-- .- .. -. ....... _-------.

SECTION VI TECHNOLOGY PROGRAM PLAN 6.0 Program Sco~e An integrated program approach is recommended to establish technology readiness for general aviation turboprop and turboshaft engines in the 298 kw. (400 shp) class. The scope of the recommended experimental program is shown in Table 43. The first task consists of a preliminary design of an experimental engine incorpurating the advanced technology components to be demonstrated. The advanced comIJonent test hardware will be designed in paralleL Each of these components will be extensively evaluated in full-scale com- ponent test rigs. The high-pressure spool components will then be further evaluated in a gas generator core. The core performance is critical in establishing a successful technology demonstration.

The highest pressures and temperatures are encountered in the core and significant performance improvements can be made as a result of optimization of the gas generator component system. After separate component testing, the low-pressure turbine system and output reduction gearbox will be cOlilbined with the high-pressure core' to form the complete experimental engine. Additional evaluation tests will be conducted to demonstrate the technology readiness for full- scale development. System analys is and eng ine def ini tion \-li 11 be performed throughout the program to insure that engine design trade-offs do not result in undue compromises in aircraft cos~ or capability.

6.1 Preliminary Design The preliminary design effort will establish the configuration of the experimental engine. The engine cycle will be defined and the components sized in order to establish the design requirements for each of the GATE advanced technology components. The experi- -_01. ~ ••• ~". ... -- ..... - -_."

j , TABLE 43. GATE EXPERIMENTAL PROGRAM PLAN - PROGRAM SCOPE.

o Preliminary Design o Coreponent Technology Development o High-Pressure Turbine - Rotor - Nozzle o Compressor o Clearance Control o Combustion System o Low-Cost Digital Electronic Control o Hign-Work/Low-Speed Power Turbine o La~er Hardened Gears o Gas Generator Technology Development o Experimental Engine Technology Development o Engine System Analysis and Definition , mental engine design will be based on a front-drive, concentr ic shaft, free-turbine, turboprop engine configurati(\n, as shown in Figure 59. The nominal ~akeoff power rating for the en~ine will be approximately 313 kw (420 hp).

Design objectives will be established for each of the compo- nents. These objectives will be compatible with the overall engine technology requir~d. The experimental engine will be designed as a demonstrator only and would not necessarily, in all areas, have flight-weight or production-type components. In areas where new technology is not being developed, the components will be designed with an objective of best program economy, while ensuring that the experimental engine will provide a representative demonstrator for both steady-state and dynamic operation.

6.2 Component Technolog~ 6.2.1 High-Pressure Turbine The high-pressure turbine is an integrally cooled radial tur- bine designed for a rotor inlet temperature of 1478°K (2200°F).

Both the nozzle and rotor are of low-cost laminated construction.

The objective is to provide the technology for a small, ceoled, radial turbine with a 9.8 percent efficiency improvement over cur- rent, 3mall, cooled, axial turbines, while reducing engine cost by 17 percent. The critical elements of technology to be addressed are shown in Table 44. Figure 60 shows the program plan and sched- ule for the component technology development.

6.2.1.1 Rotor Task Advanced process res~arch will address the need for a low-ccst sheet alloy wi th high stress-rupture strength'. Candidates are Astroloy and AF2-lDA. As shown in Figure 61, the conventional rol- ling process has a yield ot only 35 percent when making photoetch --_._--'.

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TABLE 44. LAMINATED RADIAL GAS GENERATOR TURBINE MATERIALS AND PROCESSING o Low-Cost AF2-1DA Sheet (Rotor) o Low-Cost AF2-1DA or ODS Material (nozzle) o Low-Cost/High Strength Bonding o Non-destructive Evaluation Techniques o Photoetching Process for ODS Materials o Material Characterization o 3-D ECM AERODYNAMICS o Minimize Clearance Losses Shroud Treat~ent Decreased Clearance o Minimize Cooling Penalty (Tip Discharge) o 3-D Blading Decreased Incidence Loss Increased Blade Loading o Reduced End Wall Losses 3-D Velocity Diagram 3-D End Walls C-'?

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CAP ELECTROSTATIC HIP REMELTING CONSOLIDATION CONSOLIDATION

-

I

FORGE TO SLABS

I

HOT & COLD HOT & COLD SLICE CIRCULAR ROLL ROLL LAMINATES - I-- __ L ____ ..... _ -- __ L _________ L ___ SHEET YIELD 90% 50% 80%

-----1- -- -r-----T--------i---

PHOTO ETCH PHOTO ETCH PHOTO ETCH SQUARE TO ROUND SQUARE TO ROUND ROUND TO ROUND 1- ___ .1-- _____ - - -- -.1 ________ .L. ___ 35% LAMINATE YielD 63% 72%

'-- - - - - - - - - - - - - - - - - - -- - - - - --

Figure 61. Sheet Alloy Processes '.

quality sheet stock. The powder rolling process and powder billet sli~ing method have projected yields of 63 percent and 72 percent, respectively. Development of either of these processes would remove the cost barrier and allow use of the high-strength Astroloy or AF2-1DA material.

Both of the advanced sheet alloy formi ng processes will be evaluated by small-scale pilot processing. The most promising approach will be selected and an adequate quantity of material fab- ricated for the final gas generator and experimental engine test components.

New methods of bonding the sheet alloys will be evaluated to obtain high bonding strength using a low-cost method. Two methods will be evaluated: (a) sputtering and (b) electroless nickel plat- ing. After initial evaluati~n using small bonded test stacks, the superior method will be selected and further optimized by bonding larger test stacks.

Nondestructive test methods using computer enhancement tech- niques will be optimized with the use of known defect test samples.

Additionally, the electrochemical method (ECM) of final machining the airfoil surface will be optimized to. permit fully three- dimensional reproduction of the desired aerodynamic shape. All of these processing advancements will be used to fabricate a turbine rotor to be used in the second series of gas generator tests.

The objectives of the aerodynamic technology program may be summarized as: o Ref ine per formance effects of increased rotor blockage resulting from internal cooling.

, o Determine the performance effects of coolant airflow dis- charge into the main gas stream.

o Determine clearance effects on performance of a cooled radial turbine.

o Define shroud treatment to reduce clearance losses.

o Evalua te approaches to increase blade loading so that fewer blades and proportionally less cooling air will be required.

o Evaluate methods tv decrease incidence losses by 3-D design methods.

The baseline tests ;lill assess rotor blockage effects, performance effects of coolant airflow discharge, and clearance criteria.

Data from the baseline tests will be combined with advanced analytical methods ~nd used to make a final 3-D design. The base- line test series will be repeated on the 3-D design and additional tests will be made to evaluate blade loading and incidence losses.

Flow predictions will be confirmed by use of Laser Doppler Veloci- meter (LDV) test methods. Three test series are planned with modi- fications to optimize performance.

6.2.1.2 Nozzle Task The initial program will test nozzles in conjunction with the baseline rotor. Both integral and segmented, cooled nozzle designs will be tested. Predictions of performance, cost, and life will be made for both designs. Laminated construction will be used for both designs. Performance will be evaluated in a test rig based on sta tor ex i t pressure surveys and torque measurements on a down stream rotor.

'.

After testing the integral stator, the segmented stator will be tested to assess the leakage effects. Cooling flow shall be var ied in each test to determine the effects of the quantity of cooling flow on nozzle per formance. A final test shall be con- ducted on either the integral or segmented nozzle to assess the performance effect of smoothing the laminated vane surface.

Trade-offs must be accomplished in order to select the optimum nozzle laminated sheet material. Candidates are Astroloy, AF2-lDA and oxide-dispers ion-strengthened (ODS) mater ial such as MA956E.

The effort would involve preliminary sheet fabrication, photoetch, and bonding method development to assess the relative merits of the candidate materlals. The selected material would be used for final hardware, which will be tested in conjunction with the 3-D rotor.

The orthotropic material properties of the selected laminated material will be e~tablished, including strength, elastic modulus, Poisson's ratio, low- and high-cycle fatigue, and creep /creep rup- ture.

Trade-off studies will be made of the 3-D design approach to obtain an optimum balance of performance and cost. A cost versus performance design trade-off will also be accomplished f0r 3-D end walls versus 2-D end walls.

6.2.2 Compressor The advanced technology compressor component is a 9:1 pressure ratio, single-stage, centr ifugal compressor fabr icated to essen- tially net shape from powder metal titanium. The objective is t~ provide the technology for improving the efficiency of the compres- sor by 3.5 points relative to a current technology 9:1 single-stage machined centrifugal compressor, while reducing cost to the point where it is competitive with cast designs.

, The critical elements of the compressor technology are shown in Table 45 Figure 62 shows the program plan and schedule for the component design and evaluation. In the first task, two or more candidate PM titanium processes will be evaluated to determine the one most sui table for meeting the G.n.TE compressor requirements.

Impellers of an existing configur3tio~ will be made from each pro- cess and a comparison made of properties, shrinkage, blade-shape reproduction, surface finish, and potential production cost.

Concurrently, the baseline GATE compressor ~ill be designed.

The design will be compatible with an existing test rig. The com- pressor shall be fabricated by machining and tested for performance in the component rig. It is anticipated that five tests will be conducted including impeller modifications and diffuser redesigns.

Prototype fabrication of the GATE impeller design will estab- lish the process limitations and provide information for a redesign of the impeller. The redesigned impeiler will be fabricated both by machining and the PM titanium process.

Tests shall be run to compare the performance of the machined and PM tit.:mium impellers. A final test is anticipated with a redesigned diffuser.

Using all the available data, cost and performance trade-offs will be made to determine if any design changes are required for the compressor to be carried into gas generator testing.

6.2.3 Clearance Control Clearance control on small engines is exceedingly important because the clearance per formance loss penal ties ar e relat ively greater than on large engines. However, s:nall engine clearance control is difficult because of various factors: , TABLE 45. REQUIRED COMPRESSOR TECHNOLOGY - 9:1 SINGLE STAGE.

3-D Blading for High Efficiency - High Tip Speeds - High Inducer Mach Number 0 3-D Diffuser Research 0 Improved Surge Margin 0 PM Ti Technology , YEAR MAJOR TASKS 1 2 3 PM TECHNOLOGY EVALUATION DESIGN BASELINE CENTRIFUG.AL COMPRESSOR FAB & RIG TEST BASELINE COMPRESSOR I

FINALIZE PREFERRED PM METHOD A

DETERMINE GEOMETRIC LIMITS 0;: PM DESIGN IMPELLER FOR PM FABRICATION FAB PM & 3-D MACHINED IMPELLERS I I RIG rEST PM & 3-D MACHINED IMPELLERS EVALUATE COST & PERFORMANCE TRADE-OFFS

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

Figure 62. PM Titanium Centrifugal Compressor I-' ID I-' o Tolerances do not scale o Small engines ~ave larger thermal gradients o High-speed rotor systems have more critical speed prob- le~s o Rotor excursions are relatively larger o Conventional abradable systems cause rel=~ively high tip wear o Clearance control means must be simple and low cost The major technology tasks that must be addressed for substantially improved clearance control are shown in Table 46. The tech~ology progra;.i plan is summar ized in Figure 63. Design concepts will be evaluated to establish viable candidate approaches to clearance control. Dynamic thermal analysis methods will be used to evaluate candidate clearance control concepts dUling transient operation.

In order to minimize operating clearances, controlled growth structure approaches shall be analyzed considering selec.tion of optiMum material expansion rates. Materials with thermal expansion coefficien~s that vary with temperature will be considered. Effec- tive use of cooling air will be analyzed as an additional method of controlling differential expansion rates of structure and rotating components. Although emphasis will be on non-active clearance con- trol because of the GATE Engine low-cost emphasis, simple active clearance control methods will also be considered and trade-offs made on cost versus performance. S~lected controlled-growth approaches will be designed, fabricated, and tested in an existing engine, using dynamic clearance meaaurement instrumentation.

CLEARANCE C~NTROL TECHNOLOGY.

TABLE 46.

o Controlled Growth Structur~s o Variable Expansion Rate Materials/Designs o Effective Use of Cooling Air o Dynamic Simulation of Small Engines o Rotor Damping o Shroud ~reatment o Grooves o Cooling Flow Discharge o Abradable Materials o Tip Wear o Life YEAR MAJOR TASKS 1 3 DESIGN CONCEPTS/ ANALYSIS COMPONENT RESEARCH CONTROLLED GROWTH STRUCTURES ROTOR DAMPING ADVANCED SHROUD TREATMENT ABRADABLE MATERIALS PRELIMINARY ENGINE TESTS Figure 63. Clearance Control )_ 94 Reduced rotor excursion is desirable to allow both clo£er rotor-face operating clearances and smaller seal clearances in order to improve performance. Rctor system damping method~ will be evaluated to select concepts that offer the pontential of reducing rotor excursions when operating through critical speeds. This will i~clude evalua~;on of shafting designs, bearing support structure, and hydraulic al.J/or mechanical bearing damping sys~ecs. Candidate systems will be designed, fabricated, and tested on a dynamics test rig to determine the optimum system approach.

Inv':stigations shall be made into reducing effective clear- ance. Two approac~es will be considered. In the first, the effect of grooves or labyrinths in compressor and turbine shrouds will be analyzed and tested in component test rigs to establish their effect on performarce. The affect of boundary layer bleed in the compressor shroud/diffuser interface will be analyzed and tested to determine whether any performance benefit exists.

Alternate methods of high-pressure turbine cooling flow dis- charge shall be studied and any promising methods will be tested in a component test rig.

Abradable materials will be evaluated including investigation of rotor tip confiquration alternatives and rotor tip hardening methods. Promising com0inations of abradable shroud materials and rotor tip confi9urations and materials will oe evaluated and tested in existing engines.

Additional testing using existing engines shall be accom- plished combining the most promising apprcaches derived from the controlled-growth structures, rotor damping, shroud treatment, and rotor-tip/abradable-shroud efforts. This testing will provide for the determination of performance effects of zero engine time and after limited running.

19J 6.2.4 Combustion System The GATE combustion system consists of a reverse-flow annular combustor combined with a minimum number (8-10) of low-co!jt air- blast fuel nozzles. The combustor outlet temperature required to provide a turbine rotor inlet temperature of 1478°K (2200 P) is 1522°K (2280 P). The system will require low-cost combustor con- struction and low combustor wall temperatures with minimum gradi- ents for long life. Technology for good starting, operating, and relight characteristics, with either Jet A (current jet fuel), broad speci f ication, synthetic or diesel fuels is required. The cr i tical elements of combustor system technology are shown in Figure 64. Figure 65 shows the program plan and schedule for com- bustor component development.

A full-scale, baseline, combustion system will be designed using advanced empirical/analytical design methodology. The design "equirements will be consistent with the GATE Engine and will ~mphasize minimum fuel impingement on walls to reduce the carbon forming tendency. An objective is a 30-percent improvement in exit pattern factor from currer.t technology. The improved pattern fac- tor can be achieved with proper matching of fuel nozzle character- istics and combustor flow field with the use of advanced analytical model ing. Two alternate advanced wall cooling schemes will be designed for comparison with the baseline design. Low-cost photo- etch fabr ication methods wi 11 be used in these 'ldvanced cooling schemes.

Approximately ten tests on the baseline combustor will be accomplished to optimize its per formance character istics. Six additional tes~s are planned on the advanced wall cooling config- urations. B,1sed on test evaluations and life-cycle cost predic- tions, the most promising configuration will be selected for further evaluation over the entire operating envelope.

S5 x tests , OBJECTIVE TECH~OLOGY ADVANCEMENT PRIMARY LINER ZONE CARBON INJECTOR COOLING DESIGN FORMATION STABILITY DESIGN EFFICIENCY REDUCED CHANNEL HEIGHT X X X X X X INCREASED DURABILITY X X X X LOWER PATTERN FAC10R X X AL TERNATE FUEL CAPABILITY X X X ~ J - ----- -- - - -- -------- • EMPIRICAL/ANALYTICAL COMBUSTOR DESIGN METHODOLOGY • IMPROVED FILM EFFECTIVENESS - COOLING SCHEMES - LOW-COST FABRICATION • PRIMARY ZONE DESIGN Figure 64. Combustor Technology t-' \D --.J ," I-' \.0 co YEAR MAJOR TASKS 1 2 DESIGN BASELINE COMBUSTOR ..

DESIGN TWO ADVANCED COOLING SCHEMES

-

FABRICATE TEST COMBUSTORS I RIG TEST BASELINE COMBUSTOR RIG TEST ADVANCED COOLING SCHEMES VERIFY OPf.RATION ON BROAD-SPEC.

~ & ALTERNATE FUELS Figure 65. Low-Cost ~ombusto~ 1478°K (2200 F) " are expected for this final optimization. Finally, combustor per- formance will be demonstrated with Jet A, broad specification kero- sene fuel, diesel fuel, and an additional fuel to be specified by NASA.

Cr i tical elements of the low-cost airblast nozzle component technology are shown in Table 47. Figure 66 shows the program plan and schedule. Conceptual design will be done on several candidate airblast atomize:s. Three designs will be selected bazed on cost and performance projections. These designs will be evaluated to determine spray characteristics. One or two of the best configura- tions will be selected for further evaluation in combustor rig testing on the baseline combustor. One configuration will be selected for further evaluation along with the baseline combustor in the baseline gas generator. A comparison of gas generator test and rig test r~sults will be made and used to accomplish the final optimi~ation of the combustion system on the component test rig.

Starting characteristics and limited endurance evaluation will be conducted on the baseline gas generator and any required improve- ments would be incorpoLated into the final component rig evaludtion tests.

6.2.5 Low-Cost, Digital, Electronic Control Control systems for small general aviation gas turbine engine applications must be low cost and reliable. The least expensive, most reliable control is a simple, hydromechanical type provided there are few sensed parameters, outputs, or automatic features, and a relatively high pilot workload is acceptable. The cost and weight penalty oE hydromechanical mechanization oE features such as torque limiting, automatic starting and sequencing, automatic transEer and protection, and provisions Eor optihlum engine perform- ance, noise abatement, and emission reductions is very high.

" TABLE 47. LOW-COST AIRBLAST NOZZLE TECHNOLOGY.

o Eliminate Air Assist o Design for High Production Quantity o Simplify Piloting Requirements o Improve Spray Quality o Improve Functional Reliability o Alternate Fuel Capabilitj , YEAR MAJOR TASKS 1 L 3 -

-

DESIGN CANDIDATE AIRBLAST ATOMIZERS

- FABRICATE SELECTED DESIGNS

- FLOW TEST SPRAY CHARACTERISTICS

-

FABRICATE ENGINE SETS OF. BEST CONFIG:S

-

COMBUSTOR RIG TEST , ENGINE PERFORMANCE

-

ENGINE START & ENDURANCE TEST SELECTED CONFIGURA"fION Figure 66. Low-Cost Airblast Fuel Atomizers tv o I-' The only feaslble approach to a control that provides these features and retains low cost an:1 high reliability is a digital electronic control. The current ph ilosophy of a full author i ty electronic control with hydromechanical backup will be retained.

Since a gas turbine requires no electrical power to sustain oper- ation, the backup control should not require electrical power to function.

A low-cost, high-reliability fuel control offering automatic sequencing and protection will require new approaches to closed- loop control, advanced microprocessors, and resolution of the tem- perature and vibration environment problems of the electronic hard- ware. Cost reduction will result from advances in microprocessor design and reducing the number of ser.30rs and output devices.

The critical elements of the technology are shown in Table 48.

Figure 67 shows the program plan and schedule for the component technology development.

The objectives of this program are to: (a) Continue the control philosophy trade-off study - The additional cost and weight of a backup hy'romechanical control must be continually substantiated versus reli- ability.

(b) Finalize the selected approach and mechanization study.

Determine the electronic/hydromechanical/fluidic split.

Select the optimum closed-loop control.

(c) Sensor and output devices definition - Characteris~ics, life, cost, and physical size as related to gas path blockage.

TABLE 48. LOW-COST DIGITAL EI.ECTRONIC CONTROL o Low-Cost Electronic Control Required To Meet GATE Fuel Control Requirements.

o Low-Cost Approach to Prime Control and Hydromechanical or Fluidic Backup R~quired o Clo~ed-Loop Control Philosophy o Microprocessor Design o Senser and Output Device Definition Compatible with Engine Size and Cost Objectives Iv o .;" • I I > YEAR i \ I MAJOR TASKS 1 2 3 CONTROL PHILOSOPHY TRADE-OFF STUDY FINALIZE CONTROL APPROACH I I I COMPLETE MECHANIZATION STUDY

-

- SENSOR & OUTPUT DEVICE DEFINITION· BREADBOARD SELECTED ELEMENTS BENCH TEST BREADBOARD ELEMENTS

•

- ENGINE TEST BREADBOARD ELEMENTS

I Figure 67. Low-Cost ~igital Electronic Control Cd) Breadboard selected elements for bench and engine tests/ designs will be compatible with the existing TPE33l turboprop control system, thus obtaining actual relative performance type data.

The reprogrammable feature of the electronic control will allow the electronic breadboard to be tested on the TPE331 Engine and on the baseline gas generator as well. This will permi t an early determination o-f the control characteristics under act~al engine transients as well as on the GATE gas generator. As a result of testing, modifications will be made to optimize the control per- formance f0~ t~e GATE requirements.

6.2.6 High-Work/Low-Speed Power Turbine The GATE power turbine, as is typical in small eng i nes, is required to run at lower than optimum a~rodynamic sp~ed because of critical speed limitations. Experience has shown that a wide mar- gin must be held between th~ operating range and critical speeds in order to achieve high bearing system reliability and avoid exces- sive seal and tip clearances. A high turbine work coefficient is then required to minimize the number of stages and their cost. The power turbine technology will include an objective of a 5 to 6 point improvement in turbine efficiency, to be accomplished utiliz- ing low-cost Cilst rotors wi th integral tip shrouds to minimize clearance J~sses. Currently, low-cost cast designs are unshrouded.

Shrouded designs require inserted blades which is an expensive design.

Problems tc be addressed in improving efficlency in the high- work design include: o High plade-row turning o Low stator and rotor reaction due to high inlet velocity o High exit swirl requiring downstream turning vanes o High Mach numbers The program is based on an adv~nced two-stage,high-work, low-blade- speed design with exit guide vanes. The critical technoloqv tasks that will be accomplished are shown in Table 49. Figure 68 shows the program plan and schedule for the component technology effort.

The analysis task will establish the design method and conduct the trade-offs required to optimize the design. Using these results as a baseline, a design will be made for rig test evaluation. The baseline design will be made compatible with machined components to allow early initiation of testing and facilitate rapid modifica- ticns. Three tests are planned to optimize the stator and deswirl vane settings.

Concu.'" rently ~ an integral, shrouded, cast, turbine assembly design shall be made. Processing technology iterations will be conducted to optimize the method of casting the integral shrouded rotors. The final task include~ three tests of the cast version of the t~rbine to assess any differences in performance from the machined version and to evaluate clearance effects.

Based on all of the available data, cost and p~rformance trade-offs will be conducted to evaluate the need for any design changes necessary prior to experimental engine testing.

6.2.7 Laser-Hardened Gears - . -------- The long life and high reliability requirements of propulsion engines requires hardening of gear teeth with the use of methods such as carburizing. Quenchi~g after carburizing results in dis- tortion. This distortion is corrected by final grinding operations that amount to nearly 37 percent of the gear cost. The technology advancement for the GATE Engine gears consists of replacing car bur- izing with laser contour hardening. Additional cost savings, not 2()6

--------.------.----------

TABLE 49. HIGH-WORK/LOH-SPEED POWER I'URBINE.

Aerodynamic Technology J Define Los~ Correlations for High Turning/Low Reaction (2-D Analysis and Available Data) o Optimize 3-D Velocity Diagram o 3-D Blade Design Solidity Blade Loading Stack and Contour o 3-D Vane Design Lean End Wall Contour Exit Guide Vane Optimizatior.

o Clearance Effects Tighter Clearances Shroud Treatment _~ __ ...."..,-~-".~. 'is UCW4 _ (%$ --~.-~..,..."..--, ~ YEAR MAJOR TASKS 2 3 ANALYSIS BASELINE DESIGN BASELINE FABRICATION

-

BASELINE TESTING CAST DESIGN CAST FABRICATION CAST TESTING Figure 68. High-Work/Low-Speed Power Turbine .... _, ........, ¢ e. 3. OS ... ""', .. £( .. [&>3_<"''''' ,",.,~J(MMi1Iaugc;;"::::W"ZA;....wwRA.,_S;';_J1.MW~..-..:4c;ae - quantified, will accrue due to eli~ination of copper plating and stripping operations, deleting the requirement for use of natural gas, and reduced material requirements. The laser hardening process is highly compatible with automation and promises extended gear life due to improved surface hardness through closer control over case depth and the ability to have ductile mate=ial layers between hardened zones.

The initial program task consists of an expp.rimental effort to optimiz~ the laser hardening technique. An investigation would be made of the desired gear material cr.aract~ristics followed by selection of candid~te material(s). Material coatings to enha~ce the laser hardening will be evaluated and the best coating selec- ted. Laser hardening experiments will be conducted on sample gears for sequential tooth hardening, hardening teeth sequentially 0Ppo- si te each other, and simultaneous scanning of the entire gear.

Resul ti:1g mater ial property character istics and gear distortion will be evaluated for each of these techniq~es, and the best method selected for further evaluation. Gears made using the best harden- ing method will be designed, fabricated, and tested in a gear ~est rig for approximately 100 hours Results \':ill be evaluated fol- lowed by fabrication of gears for endurance testing. These gears will be tested on a piggy-back basis on either an APU or propulsion engine test wherein. a substantial number of hours may be accumu- lated. Following endurance testing, the gears will be compara- tively evaluated with respect to conventional gears.

6.2.8 Gas Generator The gas generator effort will ensure early discovery of criti- cal component integration requirements. Figure 69 snows the pro- gram plan for the gas generator. After completion of the experi- mental engine preliminary design and the initial components deSign, the design of the baseline gas generator will be initiated. It YEAR MAJOR TASKS 1 2 3 4 BASELINE DESIGN FABRICATION

-

TEST SERIES 1 REDESIGN

-

FABRICATION TEST SERIES 2

-

Figure 69.

Gas Generator Program Plan ' .

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

will include the baseline machined compressor, the baseline combus- tor, and the baseline high-pressure turbine, and will provide an engine environment test bed for the fuel atomizers, electronic con- trol, and gas generator clearance control features.

The baseline gas gen~rator will be tested prior to the com- pletion of the component test efforts. This will substantially reduce program risk and provide early data to substantiate the com- ponent test data in-an actual engi·ne environment. The integrity of the gas generator design will be proven to ensure successful evalu- ation of the final components in the second gas generator test series.

Extensive performance, mechanical, and thermal instrumenta- tion will provide data for comparison with desigll predictions dur- ing the IJaseU.ne test ser ies. The test ser ies will include the follm'ling: o Mechar.ical checkout o Starting o Combustor performance o Clearance-control evaluation o Transient control operation o Structure temperature survey o Performance evaluation A total of 75 hours of testing is planned for the baseline gas generator test series. The baseline gas generator design will be modified to incorporate any desirable changes indicated by the ini- tial testing, and will incorporate the final component configur- ation established in the component test effort.

This modified gas generator test series Hill include the fol- lowing: " o Controls evaluation o Turbine cooling evaluation o Clearance control sy~tem o Performance testing o Alternate fuel tests o Transient thermal cycles o Limited durability testing A total of 125 hours is expected to be accumulated dur ing this second test series.

6.2.9 Experimental Engine The experimental engine effort will demonstrate the technology readiness of the GATE components and provide the final data needed to assess the GATE eng ine per formance and production potentia:.

Tha experimental engine will consist of an integration of the GATE gas generator and the low-spool components. The exper imental engine prcgram schedule is shown in Figure 70.

The design of the low-spool components shall include the low- pressure turbine and exhaust system, the low-pressure turbine shaft and bearing system, and an output power gear system. The experi- mental engine will not represent a final production engine design but will be a test bed to integrate components to the extent neces- sary to assess overall performance, component interactions, and mechanical system technologies.

The experimental engine design will begin' near the completion of the final gas generator design effort. Two experimental engines will be fabricated. Engine Serial No. 1 will emphasize perform- ance, com~ustion, and controls testing, and engine Serial No. 2 will emphasize mechanical and durability testing.

\ YEAR MAJOR TASKS 3 4 5 DESIGN FABRICATION

- •

SERIAL NO.1 TESTS SERIES 1

II ·1

SERIAL NO.2 TESTS SERIES 1

l

r •

Figure'70. Experimental Engine Program Schedule " " The test plan for the two experimental engines is shown in Figure 7~. Four test series are ~lanned on each engine with neces- sary modifications incorporated as testing progresses. A total of 400 hours testing is planned to be accumulated utilizing the two engines.

6.2.10 E!!9 ine Sy~tellL~!!aly~is a!!~ __ Qef ini!:ion Throughout the GATE experimental program, analysis will con- tinue to refine and update the previous engine definition. The results of component, gas generator, and experimental engine test ing of the GATE design will be evaluated; and engine cost, life, weight, performance characteristics, and trade-offs will be updated. Technology from other sources such as company efforts will be evaluated for applicability to the GATE Engine. Using the updated engine character istics, the GATE engine performance and economic benefits 1n an aircraft system will be updated.

f.2.ll S~hedule The schedule of ea~h of the program elements has been des- cribed. Figure 72 shows the overall program schedule and relation- ship of the program elements. The program schedule and task inter- relationships are based on minimizing program risk with an economi- cal program approach. An engine preliminary design is accomplished early to ensure compatibility of the components in the gas gener- ator and experimental engine. Gas generator testing is started as soon as initial component readiness is established. Early gas gen- erator results will insure that final component testing is properly directed. Final g~s generator testing is completed prior to exper- imental engine testing and will minimize experimental engine test problems. System analysis and definition continues throughout the program to ensure proper assessment of available data and help direct the design and test efforts. Program milestones and reviews , " STZ m m Z Z ~ -I -I -I -I -I -I -I G) G) m m m m m m m m en en en en en en en en Z Z -I -I -I -I -I -I -I -I m m en en en en en en en en en en m m m m m m m m m rn :c ::c ::c ::c ::c ::c ::c ::c ::c ::c m m m m m m m m l> l> en en en en en en en en r- r- A to) I\) ... to) I\) ...

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DETAILED DESIGN TRADE-OFFS PRELIM. DESIGN AND BASELINE SYSTEM DESIGN DEMONSTRATE BASIC TECHNOL· OGY FOR LOW-COST. HIGH- PERFORMANCE COMPONENTS COMPONENTS EARLY DISCOVERY OF CRITICAL GAS COMPONENT INTEGRATION GENERATOR REQUIREMENTS DEMONSTRATE GATE TECHNOL· OGY READINESS AND PER· EXPERIMENTAL FORMANCE AND DRODUCTION ENGINE POTENTIAL CONTINUING COST/PERFORM· SYSTEM ANCE TRADE-OFFS AND UPDATE ANAL YSIS AND OF PREDICTED ENGINE AND .

DEFINITION SYSTEM PERFORMANCE AND ECONOMICS MILESTONES ~ PRELiMH~ARY • DETAIL DESIGN • FIRST • TECHNOLOGY DESlm~ REVIEW REVIEW TEST READINES3 R~V!EW Figure 72. Recommended GATE Experimeqta1 Program , with NASA to obtain approval of the approach are indicated in the overall program plan.

6.2.12 Technology Development - Benefit Analysis The critical components identified In the previous paragraphs are high-risk development items but have significant payoff with respect to the GATE engine. To quantify this payoff each component was evaluated with respect to the cost of demonstrating technology readiness and the payoff of the particular component to the engine and aircraft application. A summary of the evaluation is shown in Table 50. The improvements in engine weight and cost are on a sys- tem basis, i.e., GATE engines are compared to current-technology engiLes at, the power level required to meet perf~rmance requi~e ments. This comparison, therefore, includes synergistic airplanp/ engine effects, which were not included in Table 38. The :mptove- ments in component efficiency and specific fuel consumpti~n are independent of applications. The benefit/cost ratio is the 20-year fleet total cost savings for the pressurized twin divided by the cost to demonstrate technology readiness. A benefit analysis for clearance control and the combustor was not conducted. Clearance control was assumed in component design in order to achieve the tight-clearances desired.

, N ......

CD TABLE 50. PAYOFF RELATIVE TO CURRENT TECHNOLOGY TURBINE ENGINE.

A A A Engine Engine, Engine Benefit/Cost ~rJ PTS % Cost, Weight, % SFC, % Ratio liP Laminated Turbine +9.8 -7.4 -21 -7 561 PM Titanium Single-Stage +1. 4 Compressor -1.0 -6.0 232 -4 Low-Cost Fuel Nozzles -1 0 0

- 144

Electronic Control - -2 0 0 132

High-Work/Low-Speed LP Turbin'? +6.0 -5 -7.0 -7.0 498 Laser-Hardened Gears - -3 0 0 226 Total -36 -20 -13.0 402 (Avg)

I -"

NOTES: 1) Changes are relative to a hypothetical current-technology turbine engine (Table 37) 2) Clearance control benefits are included in the above.

SECTION VII 7 • 0 CONCLU S IONS This report summarizes the results of the General Aviation Turbine Engine Study. Small gas turbine engines in the 336 k~ (450 hp) class were defined and evaluated in appropriate aircraft. The performance and economics resulting from the use of these engines were eVuluated, and comparisons were made between aircraft powered by reciprocating and turbine engines. Identical aircraft tech- nology levels were assumed in all aircraft comparisonb. Overall ~onclusions that were drawn as a result of the study program are: o The general aviation market was predicted to continue to grow at current rates.

o Compared to current-technology reciprocating engines and current-technology turboprops, significant reductions in aircraft fuel consumption and weight were projected with the 1988 GATE technology engines. Reduced aircraft I lnici~l ~ost and operating cost were also estimated, based on proj€~tions of new technology and high manufac- turing quantities. The barrier technology which mUDt be overcome through development of new technologies is the achievement of this low manufacturing cost without major sacrifice in performance.

o A turboprop engine is the most suitable propulsion system for the medium- and light-twin aircraft investigated.

'furbofans .at the flight speeds, altitudes, and takeoff C:!istances stipulated have higher: fuel consumption and require larger engines than do turboprops und therefore are more costly.

o A single-shaft turboprop is slightly superior to a free- turbine turboprop for the aircraft studied but the dif- ference is slight and the free-turbine engine is the most likely choice if the needs of the rotary-wing market are considered.

o High-temperature engines [1478°K (2200 F)] are superior o to lower temperature engines [1255 to l3ll K (1800 to 1900°F)-] • o Study results indicated that a GATE turboshaft would allow a reduction in helicopter gross weight of 20-percent when compared to a helicopter designed with a current-production turboshaft.

o Component research and development integrated wi th an experimental engine program is required to realize the benefits of the GATE engines.

APPENDIX I

GENERAL AVIATION MARKET DATA

During Task I, data was compiled on each of the ten recipro-

ca ting-engine-powered fixed-wing categor ies, the turboprop cate-

gory, and the three rotary-wing categories. This data includes,

for most models, the engine model_and rated power, the 1977 average

equipped price, number of seats, cruise speed, engine time between

overhaul, and service ceiling. This data is displayed in Tables 51

through 64.

>J IJ IJ TADLE 51. TWl" PLACE LIGIIT SINGLE ENGINE.

Cruise Speed Aircraft At Ceiling Service I Aircraft /,\"g.

Manulacturer Engine Equip. Price flo. of Seats (Recommended) Engine TDO Ce i ling and Model Engine Type Standard km/hr. (mph) Hours (tt. )

kw (hp) I 1977 m

Deech Sport 19 LYC 112 (150) 29,376 2 19B (123) 1200/2000 3,553 (11,650) 0-320-E30 Del1anca Citabria LYC ECA 0-235-Cl B5B (115) 19,010 . 2 19B (123) 2000 3,660 (12,000) GCM 0-320-A20 112 (150 ) 22,575 2 20B (129) 1200/2000 5,165 (17,000) KCAD 10-320-E2A 112 (150 ) 23,460 2 20B (129) 1200/2000 5,lBs (17,000) Ucllancil Lye 10- Decathlon 320-EIA 112 (150) 26,705 2 219 (136) 1200/2000 4,BBO (16,000) CaNT (100) Cessra 150/152 75 0-200-/\ 75 (1(,0) IB,255 2-1/2 195 (121 ) 1800 4,667 (15,300) Grumman- LYC 0- American Trainer 2J5-L2C 858 (115) 19,853 2 200 (124) ;WOO 3,889 (12,750) Cherukee LYC 0- 112 (150) 24,b15 2 20) (126 ) 1200/2000 ),338 (l0,950) Piper Cruiser 320-E30 .

I'A20-140 ----- ----- ~---- ---- ---- ------ -- -- ----- ---- - TAIlL!; ~2. UTlI,ITY HIGH Pl::RFORMANCI:: SIIIGLE ENGINE.

-

Cruise Speed Aircraft Aircraft Avg. At Ceiling Service Manufacturer Engine Equip. Price No. of Seats (Recommended) Engine TaO Ceiling I::ngine Type kw and Model (hp) 197; Standard km/hr. (mph) lIours m (f t. )

-

tiL'11 ijnc:a Scou l !,YC 134 (180) 26.600 196 (122) 2000 (14.500) 2 4.423 0-)60-C2A Skywa9gon Ceasnd COtlT In (230) 43.552 6 253 (157 ) 1500 5.399 (17.700) 160 0-470-U All Carryall COtn 224 (lOO) 49.252 6 227 (141 ) 4.087 (13.400) Cosana 10J 10-520'F) i Cessna 207 CONT 224 (300) 64.610 7 264 (l6~) 1200/1S'JO 4,OS7 (13.300) I IO-S20F

Cessna Turbo 207 CONT 231 (310) 70.455 7 298 (185) 1400 7,930 (26.000) I

TSI0-520M I

I·

Maule Hackel CON'r 157 (210) 21.245 4 390 (242) 1200/1S00 S,490 (18,000

i

10-360

!

I Pipor Super Cuu LYC 112 (150) 24.140 2 (ll5) 185 1200/2000 5.795 (19.000) I 0-)20-A2A 1'1118

I

"

'J W l' t

!

'"

...

'"

i.

TABLE 53. FIXED GEAR IIIGH PEIlFORMANCE SINGLE ENGINE Cruise Speed I Aircraft Aircrar.t Avg. At Cciling Scrvice (Rccommcnded) M.Hlulilcturcr eng inc Equip. Price No. of 5eats Enginc T80 Cci ling and Modcl Engine Typ" kw (hp) 19'7 Standard km/hr. (mph) Hours m (ft. ) I ':,,:;:;nil Sky 1<,n" CONT 172 (230 44,/45 4 1/2 267 (166 ) 1500 5,033 (16,500) Ill<! 0-470-U I I

~

Cesona 206 CONT 224 (300) 57,665 6 272 (169) 4,514 (14,600) 1O-520-F

~

i CeSGnii Turbo CONT 63,135 (192) 1400 8,235 (27,000) 231 (310) 6 309 206 TSIO-520M Cessna Hcim5 CONT 157 (210) No Price 1200/1500 Ho<.:k et t'H-1 7 2 10-3600 Givcn

r

~ (1],550)

C,"" uk"" PA28-235 LYC 175 (235) 47,3Z5 4 233 (145) 1200/2000 4,133

pipcr pathfindcr 0-540-13485 (158) 1200/2000 3,904 (12,800) I Pipor Cherokee LYC 194 (260) 54,235 7 254 PA32-260 0-540-E485 (16,250) Piper Cherokee LYC 224 511,005 7 282 (175) 2000 4,956

(300) I

IO-540--K1G5 PA32-300

.'

--- -- -- ---

"

.AbU. ~... H/JI< !-LACE LIGIlT SIHG:'E 1:.~.r.I!II::.

C--' Cruise Sf'eE:d i Aircraft Aircraft A-/g. a~ CE:iling Service , , 110. of Seats (Reco;r.;nendE:d) Engine T80 Ceiling Man.Jf~cturE:r EnginE: Equq:.. !-rlce (a.ph) and Mud'.:l Erl'j In', T"lpe kw 1~77 Standard krr./r,r /lours In (ft. I (r'f'1

'"'

IH:(:cr, Sur,d0W'(,eC (1 fI 0) (141) (12.600) :''iC 134 37.373 4 227 1200/2000 3.843 rJ- 16O-,;4Y (l40) (14.200) Ct::o!.\nii SJ(,r.ililllf.. L'iC 11 !I (160 ) 3u.0~Ci 4 1/2 n5 2000 4.3H 0-3~0-1l2A(j (180) (150 ) Cesl.r,a Caraina1 !.'iC 134 39,1~5 4 In 242 1200/2000 4.453 (14.600) u-360-Alf6D Cessna F r enci, COllT 108 (145) 26.850 0-300-0 S~yhawk

,\

aUllt t.y RCIlIIS 1'-172 Rolls Royce (17,000) Cessn" /law~ XP COliT 145 (1 ~5: 3a.680 4 1/2 241 (151) 1200/1500 S.laS 10-360)( L'IC (150) 27,285 4 203 (126) 1200/2000 3,874 (12,7()0) Ct.cro,;e<: 112 0-)20-030 I-Iper Warrior 151 2)5 (146) Pif'er Warrior 161 119 (160) 28.700 4 3.965 (13,0001 "'1C (WarClOr II) Q-32:l-03G (lSI) Ct,erol(£:e L'lC 134 (180) 33.930 4 243 1200/2000 4,163 (ll,650) Plf'cr Arcr,er 0-360-h4li PA-28-181 I GI ua.T,~n Ct,E:E:tat, L'iC 112 (150 ) 31. 2'i4 . 4 237 '147) 1200/2000 3.858 (12.650) i M~A 0-32u-E2G Gr ufr.irI..;,an r 1(j(:( L'lC (1&0) 25& (160) 4,209 114 ~~. 7&0 4 1200/2000 (13.eOO) I , .v,5a 0- 360-A4Y.

I - -- --- -- to ..

. "

"

IJ '''' a .\' 1h8L~ 55. LIGHT RETRhCTABLES.

Cruise 5pecd Aucrdtt Aircraft Avg. at Ceiling Service Manufacturer Engine Equip. Price /lo. of Seats (Recommended) Engine TBO Cei li n'l <lnu Hodel Engine Type kw (hpj 1977 Standard km/hr (mph) Hours m (f t I ~

=

ileech Sierra LIC 149 (200) 53,5~4 6 ~54 (158) 1200/1600 4,697 (15,400) C-24-11 IO-360-AlI16 (,,:.:.n .. C"ruin"l 1.'1C 149 (~OO) 50,O~5 4 274 (170) 1200/1600 S,216 (17,100) l77-IIG IO-)60-AlI160 Mooney Ranger L'IC 134 (lElO) 44,185 264 (16; ) 1200/2000 5,033 (16,000) M~OC 0-360-/.1 0 M00nci EX~cutl~e L'Ie 14~ (~OO) 48,~60 4 288 (179 ) 1200rOOO 5,734 (18,800) M20F 10- 360-AIA M(,one)l 201 L'iC (200) 149 55,310 314 (195) 1200/1600 5,734 (16,&00) M~O,l IO-360-AI860 Lye (200) "II"H Ar r ow II 14~ 47,&50 4 266 (165) 1200/1600 4,575 (15,000) PA 2SI! 200 10-360-CIC Piper Arrow III L'iC 149 (200) 50,320 264, (164 ) 1200/16UO 4,941 (16.200) PA 2BR 201 10-360-CIC6 Piper Turt.o CO/IT 149 (200) 54,975 4 319 (198) 1400 6,100 (20,000) Arrow III TSIO-360F PI, 28H 201T Hockwe11 112 LYC 149 (200; 61.295 4 262 (163) 1200/1600 4,590 (15,050) (1128) 10-360-CI06 I!ockwell 112 TCA L'iC 157 (210) 65,295 4 301 (187 ) 1200 6,100 (20,000) TO- 360-CIMO i -- - -- - --- '------ / I

r

t

TABLE 56. IIEAVo, RETAACTABLES.

Cruise Speed Aircraft Aircraft Avg. at Ceiling Service Manufacturer Engine Equip. Price No. of Seats (Recomznendedl Engine TBO Ceiling Eng i ne To,pe kw (hp) and Model 1917 Standard klll/hr (mphl lIours m (f t I Beech Bonanza h36 CONT (285 ) 96.545 6 (193) 213 311 1200/1500 5.063 (16.600) , 10- 520-Bh

t

lIeech Bonanza V35 corn 213 (285 ) 89.355 5 319 (1981 1200/1500 5,444 (11.850) 1O-520-UA

~

I Ueech Bondnza F33 corIT (285 ) i 213 84.224 5 319 1198 I 1200/1500 5,447 (17.858) 10-520-IJA I

I

I lJellanca Viking Lye 224 (300) 68.259 4 306 (190) 2000 5.551 (18,2001 17-31A 10-540-l<lE5 Bellanca Turbo LYC 224 (300) 79.090 4 357 (222) 2000 7.320 (24.0001 Viking 17-31 ATC 10-540-l<lES Cessna Centurion CONT 224 (300) 11.335 6 317 (1971 1200/1500 5,277 (17,300) 10-~lC.-L Cessna Turbo COtlT (310) 231 17.455 6 367 (228) 1400 8,693 (28.500) Centur ion TSI0-520R Piper Lance LYC 224 (300) 72,120 6 293 (182) 4,453 (14,600) I PA 32R-300 10- 540-l<lG5D ROCKwell 114 LYC 194 (260) 70.800 4 291 (181 ) 2000 5,)07 (17,400) I 10-540-T4B5D I - --------~ ....

'"

....

"

1 > I> a; 'fABl.E 57. AGRICULTURAL.

Cruise Speed Aircraft AlCcraft Avg. at Ceiling Service Manufacturer Engine Equip. Price No. of Seats (Recommended) Engine TBO Ceiling and Model Engine Type kw (hp) 1977 Standard km/hr (mph) flours m (et) Ces~na AG Carryall corn 224 (300) 55,205 6 227 (141 ) 1200/1500 4,0117 (13,4CJO) 10- 520-0 Cessna AG Wagon cOIn 224 (300) 51,485 1 182 (113) 1200/1500 3,386 (11,100) 10- 520-0 IU8 Cessna AG Truck CONT (300\ 54,310 20g 224 1 (130) 1200/1500 3,386 (11,100) 10-520-0 Grumman AG Cat P & W 336 (450) 69,005 190 (118) 4,270 (14,000) I R985ANl Piper Pawnee 235 LYC 175 (235) 39,880 1 183 (114) 1200/1500 3,965 (13,000) , I'A-25-2350 0-540-B2C5 Piper Pawnee 260 LYC (260) 194 42,350 1 187 (116) 1200/2000 4,819 (15,800) PA-25-2600 0-540-GIAS Piper Brave 285 CONT 213 (285) 54,305 1 237 (147 ) 3,965 (13,000) PA-36-285 TIARA 6-28'; Piper Urave laO Lye 224 (300) 55,605 1 . 227 (141) 2000 3,660 (12,000) IO-!.40-KIG5 PA-l6-300 Rockwell Thrush P .. W 447 (600) 78,500 1 200 (124) 900 4,575 (15,000) S2R600 R1340ANl TAB/.I:; ~8. LIGHT 'l'WUlS • ..

Cruille Speed Service Aircraft Avg. at Ceiling Aircraft No, of Sea~s (Recommended) Engine TBO Ceiling Manufacturer I:;ngine Equip. Price km/hr (mph) Hours (tt) Engine Type kw (hpj 1977 Standard m and Model 1200/1500 5,673 (l8,600) COf;T 2lJ (28S) 187,115 6 370 (230) Beech Baron ~8 IO-S20C (19,100) (285) 167,3Si 6 370 (230) 1200/1500 5,826 Beech Daron £~S CONT 213 IO-S20C 1200/1500 5,887 (19,300) (260) 142,844 6 348 (216) Beech Bacon B5S Cm:T 194 IO-470L (310) 214,&66 6 446 (277 ) 1400 7,625 (25,000) Beech Baron 58TC CONT 231 TSIO-520L (192) 1200/1500 5,490 (l8,OOO) CONT 157 (210) 102,15S 6 309 Cessna Skymastec 337 1O-360G 359 (223) 1500 6,024 (19,750) 213 (28S) 152,440 6 Cessna 31 0 CO NT 10-520:-1 & (27,400) 170,880 412 (256) 1400 8,357 Cessna T31il CONT 213 (285 ) : 'TSIO-520B (17,600) 6 325 (20~ ) 1200/2000 5,368 Piper Aztec LYC 186 (250) 137,835 IO-S40-C4[lS PA23 , ] 400 7,625 (25,000) 102,180 6 353 ,219) Piper Seneca cOIn 149 (200) TSIO-)60£ PA34R 327 (203) 1400 5,917 (19,400) LYC 216 (290) 242,700 8 kockwcll Shrike IO-S40-1::1DS !

(274) 2000 6,466 (21,200) (290) 171,170 6 441 Aerostar 600 I.YC 216 10- 540-1<1t'S (290) 1800 9,181 (30,100) 216 (290 J 189,170 6 467 Aerostar 601 LYC I IO-S40-S1AS - -- .~

'J

. .,

"

':.

( " \ \ u ...

o TABLE ~9. CAHIll CLASS UIlPRESSURIZED TWINS.

--

Cruise Speed Aircraft Aircraft Av'j. at Ceiling Service !

Manufacturer Engine Equip. Price No. of Seats (Recommended) Engine TBO Ceiling !

and Model I::ngine Type kw (hp) 1977 Standard km/hr (mph) lIoucs m (f t) Cl:ssna 402 CONT 224 (300) 225,675 10 386 (240) 1400 7,985 (26,180) I T510-520-E I I CONT (375) 3n,665 399 (248) 7,930 (26,000) Cl:ssna Titan 404 280 10 1200 I GT510-520-M I Piper Navajo LYC 231 (310) 232,490 8 398 (247) 1500/1800 8,022 (26,300) PAll-310 TI0-540-A2C Piper Navajo " LYC 261 (350) 263,485 10 409 (254) 1600 8,052 (26,400) Chieft~in PA31-350 TIO-540-J2BD .'

1hbLE 60. PRESSURIZED TdIUS.

Cr u 1I1e Sl'c~d I Aircraft Aircralt Av'l. . at Cciling Scrvice Manuf dcturer Englnc Equip. Price 110. of Seats (Recommended) Engine TOO Ceiling j and Model En'1lne Type kw (hp) 1977 Standard km/he (mph) /lours m (f t) i !

Oeech Duke 60 LYC 2&3 (380) 3)0,090 6 443 (275) 9,150 (30,000) TlO-~41-E!C4 i (25,000) I Ueech Daron S&TC CONT 231 (310) 265,908 6 452 (281) 1400 7,625 'I'S I 0-::, 2 0 L ~ Ce~3nil Pressurized cOIn 168 (225) 146,155 5 380 (236) 1400 6,100 (20,000) I S,-ymastcr TS 10- 360C Ce:;~n~ 340 COUT 231 (31 0) n5,24~ 6 430 (167 ) 1400 9,089 (29,800) TSIO-5~Orl " Ccssna 414 cOIn 231 (310) 271,870 412 (256) 1400 9,562 (31,350) TSIO-520N Chancellor Cessna 421 COUT 280 (375) 381,000 8 448 (278) 1200 9,211 (30,200) Golden Ea']le GTSIO-~20/l Pil'er /lavajo 425 LYC 317 (425) 390,255 407 (253) 1200 6,845 (29,000) TIGO-541-EIA Aerout"r 601 P LYC 216 (290) 247,940 6 467 (290) 1800 6,037 (26,350) 1O-540-;aM '----- IJ ....

.....

TABLE 61. TURBOPROPS. *

Average

Equipped

Aircraft Engine kw hp Price

Beech King Air Super 200 PT6A-41 634 (850) 1,128,200

Beech King Air B100 TPE331-6-252B 533 (7 Pi) 956,000

Beech King Air A100 PT6A-28 507 (680) 926,100

Beech King Air E90 PT6A-28 410

(550) I 807,500

Beech King Air C90 PT6A-21 410 (550 )

614,900

Piper Cheyenne P'!'5A-28 462 (620)

665,000

Rockwell 690 A/B TPE331-5-251K

535 (718) 781,190

Merlin III A TPE331-303 626 (840 )

1,078,070

Merlin IV A TPE331-303 626

(840) 1,175,970

Metro II TPE331-303

701 (940 ) 1,055,900

Cessna Conquest TPE331-8

41;6 (625 ) 850,000

I

- - ----- *Manufactured by Gamma member.

, TABLE 62.

HELICOPTERS - SINGLE ENGINE PISTON

, , ~

Aircraft

Aircraft 1977

Manufacturer Engine Engine Avg. Equipment

Type hp

I and Model kw Price

•

(1 ) Robinson R22 LYC-O-320 92 (124 )

•

(2 ) Brantly B2B IVO-360-A1A Lye (180 ) $48,950

(3 ) Enstrom F-28A LYC HIO-360-C1A (205 ) 64,500

Hughes 300 C HIO-360-D1A LYC (190 ) 65,450

(4 ) 142

(5 ) Enstrom 280 Shark HIO-360-C1A LYC (205) 71,000

(6 ) Enstrom F-28C LYC HIO-360-E1AD (205) 71,000

(7) Brantly 305 IVO-540-B1A LYC 227 (305) 79,950

(8 ) Enstrom F280C HIO-360-E1AD LYC 153 (205) 76,000

VO-540-C2A LYC .

(9 ) Hiller UH-12E 227 (305 ) 78,000

------ ----- --- I ---_.- '-------

N W W / t,J w ~

TABLE 63. HELICOPTERS - SINGr.E ENGINE 1URBINE

Aircraft 1977

Aircraft

Manufacturer Engine Engine Avg. Equipment

Type kw hp Price

and Model

400 $209,000

(10) Hughes 5000 Alli~on T-63 298

Model 500 250-C20B

(369)

Allison 250-C20 400 212,500

(11 ) Bell 206B 29a

592-650 235,000

(12 ) Aerospatia1e Asta~ 441-485

!

350 (1 ) Ar riel or

SA350 Ecureu 11 (1) LTS 101

!

(13 ) Aerospatia1e Gaze11~ 440 590 300,000

I

SA341

(14 ) Allison 250-C20B 313 420 309,500

Bell 206L

- -- - -- -- --- - --- -- -- -- - --- --- ----- -- --- -- -- --- - --- -----~--- I , ----------- -----.

TABLE 64. HELICOPTERS - TWIN ENGINE TURBINE

Aircraft Aircraft 1977

Manufacturer

Eng ine Engine I Avg. Equipment

Type

and Model kw hp Price

(15 ) Allison 250

MBB BO-105C B2 313 420 $ 385,000

C20/20B

(16 ) Agusta A-109 (2) Allison 336 450 700,000

A-109A 250-C30

(17) Bell 222 LTS-10l· 447 600 750,000

(18 ) Aerospatia1e Dauphin 2 (1 ) Au ie1/SM365 317

425 620-865,000

SA365 LTS 100/5A366

(2)

(19 ) Sikorsky S-76 All ison· 250-C30 522

L-..- ________ t..J VJ Ul "

APPENDIX II

AIRCRAFT DESIGN CHARACTERISTICS

1.0 FIXED-WING AIRCRAFT

During :ask I and early in Task II, the Ces.sna Aircraft Com-

pany, Pawnee Division, defined the characteristics of the aircraft

to be used in the GATE study. These characteristics were used to

-

model the airplanes for the General Aviation Synthesis Program

(GASP). GASP resized the airplanes as required for wing loading

changes and changes in takeoff gross weight required to meet the

mission requirements. Checks were performed during the GASP analy-

sis to ensure that f ideli ty to the or iginal character istics, as

supplied by Cessna, were maintained.

Table 65 shows the weight brea~down of the four designs

studied. Designations are as follows:

Design No. Description

1 Turbofan-Powered (wing mounted)

Medium Pressurized Twin

Turbofan-Powered (fuselage mounted)

lA

Medium Pressurized Twin

2 Turboprop-Powered Medium Pressurized

Twin

4 Light Twin

Cessna's weight breakdown philosophy is ~xplained in NASA

CR-15l973, "Conceptual Design of Single, Turbofan-Engine-Powered

Light Aircraft", Section 3.2~4, pages 42-46. The methodology has

been modified for the GATE study, based on larger Cessna models, in

order to handle the medium-twin configurations. The powerplant

installation weight was based on engine data supplied to Cessna

~- :iEC£:>."NG PAGE BlANK NOT FILMED

TABLE 65. WEIGHT BREAKOOWN, KG (LB) Configurations 1 lA 2 4 Components 283.9 283.~ 283.9 271.9 Wing (625.4) (625.4) (625.4) (598.9) Includes control surfaces, attachment hdrdware, fairing, carry-thru in fuselage 426.4 332.5 Power Plant .nst3llation (See Table 66)1 322.9 313.5 (711.4) "(690.6) (939.2) (732.4) Includes everything supported by engine mount, intake ard exhaust systems, filters, pumps, controls 63.0 145.3 9().1 145.3 ~ (320.0) (138.8) (320.0) (198.4) Includes cowling, attacnment, engine mount 25.0 20.8 16.6 20.8 '.tcrtic.:ll T.:s.il (45.8) (55.1) (45.8) (36. ';) 34.6 30.9 26.1 30.9 Horizontal Tail (68.0) (76.2) (68.0) !:;7.5) I 71.6 71.6 71.6 69.0 Main Gear Assemblv (157.7) 1157.7) (157.7) (15:'.0) Includes tires, wheels, brakes, gear legs, shocks 25.8 25.8 25.8 H.9 No~e Gear Assemblv (56.8) (56.8) (56.8) (55.0) 47.1 47.1 47.1 44. S Retraction System (103.7) (103.7) (103.7) (98.0) Includes actuators, valves, lines, pumps, selectors, reservoirs, fluids 292.7 292.7 292.7 234.3 Fuselage (644.8) (64L8) (644.8) (516.0) Includes structure~ doors, hatches, windows, attachment fittings, brackets, floors 52.5 52.5 52.5 50.8 Controls (liS .6) (115.6) (115.6) (112.0) Flight and englne 93.3 93.3 93.3 91.7 Equipment (205.6) (205.6) (205.6) (202.0) Electrlcal, battery, box, regulator, basic lnstruments 147.8

Furnishings I 147.8 147.8 134 . .&

(325.5) (325.5) D25.5) (290.0) Includes seats, restraint systems, ventl13tion system, soundproofing 15.9 15.1 ll.9 Extenor Finish 15.2 (33.5) (35.2) (33. :) (28.5) :38 w~:,-'lT BREAKDOWN, KG (LB) (Contd) :"BLE ">5.

, Configurat ions Components 2 1 lA

_L I

~

"

Drl! EmEt~' Weight IDEWl 1549.9 1466.9 1653.2 1399.8 (3231.0) (3413.8) (3641.3) (3083.2) Basic Emotl! Weight IB~~! 1660.0 1557.0 1763.3 1517.5 (365(;.4) (3473.6) (3883.9) (3342.6) Assumed Gross Wei9ht IGWl 2860.2 2860.2 2860.2 2724.0 (630:1.0) (6300.0) (6300.0) (6000.0) Constants, 549.2 458.5 753.3 607.5 a , kg l (1209.7) (1010.0) (1659.2) (1338.2) (lb) a , kg/m 2.859 3.266 2.859 3.304 2 (lb/ft 2) (0.58603) (0.66916) (0.58603) (0.67695) a (dimensionless) 0.13518 0.13616 0.13518 ().12616 (0.29775) (0".29991) (0.29775) (0.27789) a ,11m 0.0000193 0.0000183 0.0000193 0.0000185 4 (11ft 2) (0.0002078) (0.0001974) (0.0002078) «().0001991) '- , .. ,. .,.. _or' .. , .. " ~.:_~ ...... '- ...... 4 .• "· ... "'---·-~-·..l-.-::.-..-.....c"""""'5"-"'''.~_ r~.-- ___ :'--: «"''''r'Itiz!'il1a~_,

e.uly in Task I .1nd 1".15 revised az the detailed engine weight

bec.lme .1vai labll'. The difference between dry empty weight and

basic empty \ ... ei'.Jht includes option.11 equipment and unusable fuel

and oil. 'rhe a~,sumed 9ro::>;. I"eiqht was b.1sed on the basic empty

weight plus payload (passengers plus bagg.1ge minus option.11 equip-

ment) and Cessna's estimate of the fuel required. The constants .11

tllrough .1.\ I.;ere ~'Hlppl ied by Cessn.) to .111ow Garrett to check the

\.:ing weight calculatcd by G,\SP as TOG\'1 .1nd \'1/S varied. Use of these

cnnst,lnts i::; ('Xpl.1 ined in the previously ci ted reference. The

propulsion ~ystem weight breakdown iz detailed in Table 66.

\'1i nq .md empenn.1ge qeomet ric char.1cter i st ics arc shO\.;n in

T.1b le 67. CC;'Gna recommended a b.1sic wing having an aspect ratio

l)t 7 and .1 t.lper r.lt il) of ll. 7, wi th a thickness-to-chord rat io

v;nyinq line,lrly trom 0.17 .1t the root to 0.13 .1t the tip. The

b,l~.ic I"ing de~.iqn W.l:; .1d.lpted to each application by adding wing

[,)ot pluqs t,) .1chieve the desir(·d Idnq area. The addition o[ the

winq root pluus increases ~~rect ratio and decrea:.es taper ratio.

Optil)n.l1 l'quipml'nt li~.t:. fl~r ,:'ach of the confiqur:.tion! •

. He :,hown in T.lb 11':. hB ;lnd 6'). 1'he i terns :,(~lected are thone

i:1clllded in l.'I':,:,n.1':, poplll.lr "-II" t.lctory im,talled accessory

rack~ges. The rrice~ are liGted [or each itcw installed separately

.1nd mu:,t be .1d-ju!,tt'd (or f.lctory-in:.talled packaqes. 1\ package

instal1.1tion rl'dllce~; the tot.l1 cost by 17 percent. The I.;eights of

tile option.l1 equipment ,1re cllarqed against the payloads stipulated

ill the desiqn [,'qui rt'ml'nts.

Dr.1l1 pol.ns tor .111 confiqurations are shown in Table 70.

TIH'Y are (or tilL' I.;inq areas selected by Cessna and with the gear

retracted. Tile dra~ polar:> supplied by Cessna were used to cali-

brate tile G~Sr dr~q subroutines. The equivalent flat plate area of

the Lmdinq IJI'.H is 3.5 sq. [t. b.1sed on a nose gear tire si7.e of

6.00-6 ~nd a main qear tire size of 6.50-10. The flap system lift/

2·H)

, ..... _~_.~,~~,,_,.~ ,. • ...... . _'r'?t ....... '{'"'O~ y .:J-,·II.'~'-' ,;;,~~.."..,.:~.

TABLE 66. POWER PLANT INSTALLATION DETAILS WEIGHT, KG (LS) Configur at ions 1 lA 2 COIIlponents 134.4 134.4 123.0 Engine 90.S (296.0) (296.0) (271.0) (200.0) Propeller 53.6 40.4 (118.0) _ (89.0) Spinner 3.6 2.3 (7.9) (5.0) 12.3 12.3 12.3 Starter Generator 12.3 (27.0) (27.0) (27.0) (:7.0) Propeller Pitch Control 1.4 1.4 (3.0) (),O) Hydraulic Pump and Ptopeller Governor 3.5 3.5 (7.6) (7.6) Prcssurc Switch and Voltage Rcgulator 0.73 0.73 0.73 O. i3 (1.6) (1.6) (1.6) (1.6) Oil Pressure Transducer 1.1 1.1 1.1 1.1 (2.4) (2.4) (~. 4) (2.4) Drain Tubes 1.1 1.1 1.1 1.1 (2. 4) (2.4) (2.41 (2.4 ) 3.ti Electric Boost Pum~s (l) 3.6 3.b ),6 (8.0) (8.0) (8.0) (8.0) Unfeathering Pump 1.1 1.1 (2.4) (2.4 ) Oil Coolcr and Mount 3.4 3.4 3.4 3.4 (7.5) (7.5) (7.5) (7.5) Control Linkage on Engine 0.18 0.18 0.18 0.18 (0.4) (0.4) (0~4) (0.4) Tailpipe 4.7 0.0 4.7 4.5 ~ ~ ~ ~ TOTAL (Per Engine) 161.5 156.8 213.2 Ibt.3 (355.7) (345.3) (469.6) (366.2) ':~l .~ ".Ir -,. .,', .0' ~'t"'--'"-.··d--.:;e,,..-.c"',-,,,,,,,,,,,-,.c:'·-"""'''''-..r ,-,l:~"""" • .;..-:T}-,-"'·- .• _ .. ~·"- .!.ii..-...«':~~~~.

r

..

..

..

lAfl!.£ (,7. WWG AIILJ DlH.lHIAGE GEOME7kIC CHARACTEklSTICS WH.q Horizontal Tail Vertical Tail _Area 1a11 LJilit ~Area2 Sweef' Area Tail Dist Sweep 2 2 2 ).

(;r ... n11 gur ilt lCJO m (t t I AI< SW(:(:f' m" (ft ) err, ! in. J A m (f t ) em (in.)

Ai< rad (d<:'l) Ak A r"d (deg) 4.3;' 1 17.17 7.71 0.~7 0 ~Ga ).~d 0.60 0.14~3 2.60 ~23 1. 45& 0.338 0.7~j (1&4.06) (47.0~) (20C) (&.53) (28.85) (206) (n) 11. 17.17 7.71 0.67 0 4.90 459.7 5.35 1. 00 0.00 1.22 445 1. 221 0.369 0.815 (184.06) (52.74) (181) (34.6)) (115) (SO) .--- ..

~ 17.11 1.71 0.67 0 4.30 ~ua 3.9a 0.60 0.14!1) 523 1.458 2.60 0.ll8 0.153 (ld4.06) (41.00) (200) (9.53) (29.95) (206) (43) 4 16.75 1.60 0.68 0 3.13 406 4.19 0.67 0.111) 2.16 432 1.495 0.348 0.75) (180.09) (40.16 ) (160) (6. )6) (23.26) (110) (431 ------- - - -- ------- ---- ---- --"

"

TABLE 68. CONFIGURATIONS 1 AND 2 - OPTIONAL EQUIPMENT LIS1.

(Taken trom the Pressurized Model J40II) Price Wei'lht Item (1977 ) kg ( 1bs) ~OOB Nav-O-Mat1c (AF-550A) S 8,595.00 14.98 (3).0) Basic Avionics Kit 1,135.00 2.72 (6.0) 300 $C[lCS Avionics $ystem-TSO'd B,1l5.00 ~7.69 (61.0 ) ~OJ Transponder (RT-452A) - lIigh Altit;Jde 795.00 3.1t!

(7.0) 400 DME (RTA-476A) Distancc ~casuring Equipment 3,495.00 6.81 (15. 0) Indicator, Economy Mixture 610.00 1. 04 (2. j) -:ontrols, Dual 680.00 3.45 (7.6) Cabin Pressure Control System, Variable (Exchange) 1,895.00 O.~l (::.0) Fuel $ystem, Auxiliary-Wing 2)9 litcrs (63 gallons) 4,680.00 30.55 (67.3) Ground Service Plug Rcceptaclc 295.00 2.50 (5.5) Light, Landing, RH ~)O.OO ~.72 (6.0) L1ght, Taxi 80.110 0.68 U.S) Li'lhts, Strobc (Thrce) 1,295.00 4.99 (11.11 ) ,~ Locator Beacon, Economy :50.~0 l.

(2. ~ I ~osc .... hee 1 Ft.. .lde [ 75.110 0.45 i 1.0) $tat1c Dischargers (sct ot fivc) 135.110 0.09 Ill.:: ) Indicator, Outs1de Air Temperat~[e (Elcctric) 150.00 11.';5

_'!.:.i!t

S32,710 • .1O l.1L'> (~30. ~) ~4j .-t ... ___ ..... '_.. + -. rt . rl -:Awe .. ,;.,.s".? ~ t.

.~ TABLE 69. ~ONFIGUru\TION 4 - OPTIONAL EQUIPMENT LIST.

(Taken trom the Moael )lOll and Turbo 310 II) Price Wel.,ht Item (1977) <Q ("051 4003 l1;JV-0-;'\;JtlC (AF-550A) S 8.595.00 H.99 (33.0) B~SlC Avionlc! 'it 1.010.00 :.:!7 (5.0) lOa SerIe~ AVIonics System-TSO'd 8.115.00 :7.69 (6~.Q) 4~~ Tr.ln~punder (RT-459A) - lIiqh ,\ltitude 795.00 3.18 (7.0) Ilh:::ic..lt0r, uut:":;l~e Air Tempc[.ltu(C (Electric) ISO.OO J.45 (:.0) IndlC;Jlur. Ecunumy MIxture 610.00 :.04 1:.3) ~untrol'" Ll".l1 540.00 :.'jll (~.5) Lio ... ' [ , u.Jyq.lqe- L • .l[";L' $1 ze t EX~:·.lnqL') 660.00 ~.81 ll:.e I Fuel ~)stem. Auxill;JrY-Wlnq 1':39 liter~ IliJ ~allons)1 4.365.00 :~ • .3 3 (f:.':l ~roun~ SerVLce Plu~ Rec~pt~cle :!:!S.OO :.3::!

,S.l) . -, Ll.;ht, ~..l:1alr.~, HH He.CO ... '''' (';.0 ) l.l-,jht. 'f.lX i 80.00 C.68 (!.5 ) l.l~hts. ROL~tinq de;Jcun (on rudder) )10.00 :.tid (1.5 ) L0C..ltor ilc.lcon, Economy ::50.00 :.15 1:.6) Nose "tle~l Fender 75.00 : .45 (:.J) Sl.ltlC 'bL~ch.)r~;ers ~set ot tlVC) 1::0.00 ': .. J:i (~.::) SC.lt:Lnq '\[r.Jnqe:!1t!nt - IJptlon ::.::85.00 19 .. 75 (~3. 5) S28.595.00 1 :'':.1 ..

(';47.0) .:4J -.-- -_... ---.--.

TABLE 70. ESTUIA'fED DRAC PGL.~:·: FOR Tl!E GATE STUDY CONFIGURA'rIONS

S b .. f

dC

S'ET

',lns D

2 REF2 " w 2

CD

(ft2) f.".l (ft )

CC:ltiguration In (L t ) I.i (ft) i.K dC 2

m e 0

L

(.37.67)

1 .I. 7.11 7.71 83.';2 0.39 0.765 0.0229

0.05"0

(184 ) (897 ) (4.22)

1A 17.11 (37.67) 7.71 88.12 0.41 0.758 0.0242 0.0544

(184 )

(947.5) (4.45 )

(37.67)

2 17.11 7.71 82.49 0.45 0.750 0.0260 0.0551

(184 )

(887 ) (4.79)

4 16.74 (36.983) 7.60 69.75 0.38 0.0225

0.769 0.0545

(180 ) (750)

(4.05)

tv

""" 1J1

"

drag character istics supplied by Cessna would have requi red re-

progral':ming GASP. A comparison of the six options contained in

GASP indicated that GASP option No. 3 (Split Flap) approximated the

Cessna data sntisfactorily.

Three-view drawings of thp fo~r aircraft used in the

GATE study are shown in Figures 73 through 76.

" 1-'-.,;;;;-- '21~8-:.:---t-~ .. - ~~ .... = • ..:..- ...=-==_=...:=t:"'-=_=_=-:_==-=-=_=:x:r.::= ____ J- _______ -L __

=::..,~:.~ ~,~ -~~ 2]E?' {t1

.............. " •• , .. !..'.!'" ,.. 0 -~-

1------ C'u1' S'l6() w:> L'L66---~f------('u1' O'Z!;V) Ill:> 8vtT--------t

L------,

t ----.----- '-_. _____ -.:1 _____ -.t... ___ . ___ _

,

o

o

o o

o

'31

rf:=J1 '0 0 0 t -, -----(Ou"J: BOZ6£) ill=:> L"L66---"- _,,~ ___ (Ou"[ O·ZSt) ill=:> Btll-------1 co "1' N . VIa . (TOHrT

o 'vL) w::> 88T--

,-0-,/0

W::> -----(·UT 8'~6(} w::> L'L66--->-----(·U1: O'Z5v) w::> Bvl1------tl .. ~1

. - - - - - - - - - - ~

tv

-, -------r _H -- ) ---

111 _ J==:-'

_ _~. ________ ~2~======~====~;=~~===

o

FTf ---:.:: -:- ,-p

D

---"t

c c

I:

• •

• •

~---------1127.3 em (443.8 in.)--------~ 840.2 em (330.8 in.) ~

_--B.

'A, ~

em (74 in.) .~ .. : PROP DIA.

"

Figu~o 76. GATE Dosign No.4

APPENDIX III SENSITIVITY DATA As described in Section 4.0, sensitivity data was developed for each of the three baseline aircraft. Specifically, the rela- tionship of empty weight, gross weight, fuel weight and power or thrust requi:ed to engine weight and specific fuel consu~ption was defined. Base values for the sensitivities were: Pressurized Twin Light Twin Turboprop Turbofan Turboprop Design 1 4 GrL ,,~. vieight 2470 kg 2825 kg 2374 kg (5441 Ib) (6223 Ib) (5228 Ib) Empty Weight 1485 kg 1550 kg 1352 kg (3271 Ib) (3413 Ib) (2978 Ib) 468 kg 758 kg Fuel Weight 590 kg ::'030 Ib) (1670 Ib) (1300 Ib) Th!"IJst or Power 336 kw 4579 N 251 kw (450 hp) SLS,TO (1029 lb) (336 hp) Engine Weight 123 kg 134 kg 91 kg I (271 Ib) (296 lb) (200 lb) !

Speci f ic Fuel 0.31 kg/hr/kw 0.065 kg/N.h

0.31 kg/r.l,'kw I

Consumption* (0.51 Ib/hr/hp) (0.64 lb/hr/lb) (0.51 Ib/hr/hp) -- -- -- *Cruise conditions, installed shaft or thrust SFC as appropriate.

Figures 77 through 82 show sensitivity data for the three baseline aircraft.

Engine sensitivity data was also developed during the program and is included in Tables 71 and 72. These data were generated for the baseline turboprop and turbofan engine.

'.

0.10 0.08 0.06 WEIGHT 0.041

L1L~lL'

l- I (!)

0.02 I-W :I: ~ (!)

w w Z ~ -oJ <l w -0.02 en <t: III / / I W/S = CONSTANT

RANGE = 1556 KM (840 NM)

5490 M (18,000 FT) 444 KM/HR (240 KNOTS) 671 M (2200 FT) TAKEOFF DISTANCE 0.06

/

., 0.04

Iffi

~

V

Il..

0.02 W en I

/

1l..1~

:I::I: , <lw

Z /

-0.02 -oJ W en

/

<t: -0.04 III

V

-0.06_

- -

- 0.6 0.8 1.0 12 1.4 1.8 1.6 ENGINE WEIGHT BASELINE ENGINE WEIGHT Engine ~"leight Sensitivities, Turboprop Figure 77.

Pressurized Twin, Design No.2.

... -"'- ' 0.25

I I

L FUEL WEIGHT -

0.20 0.15

I- -""V I

0.10 :I: CJ I-w GROSS WEIGHT L <Il!!!!.

0.05 :I::;:

1.-

CJ w

~ -

w

z ~

EMPTY WEIGHT _ :: ...J w <l tI)

P7

<t -0.05 al ~ .

---

I

-0.10 kS = CONSTANT

RANGE = 1556 KM (840 NM)

5490 M (18,000 FT)

/

-0.15 444 KM/HR (240 KNOTS) 671 M (2200 FT) TAKECFF DISTANCE I

-0.20 L

0.0 ~

/'

0.0 !

/

./

0.0 1 Q..

:I:

L

)

Q..I~

:I:-

//

<l ~ -0.0

<t

/'

al

/"

) -0.0

~/

-0.0 -0.041 0.8 0.9 1.0 1.1 12 1.4 1.6 CRUISE SFC BASELINE CRUISE SFC Figure 78. Fuel Consumption Sensitivities, Turboprop Pressurized Twin, Design No.2.

:>0.. ___ '. -C3_~

•

0.04 W/S = CONSTANT RANGE .. 1556 KM (840 NM) 0.03 5490 M (18.000 FT) 444 KM/HR (240 KNOTS) 671 M (2200 FT) TAKEOFF 0.02 DISTANCE 0.01 t- :I: (!)

t- W :I: :: (!)

FUEL WEIGHT -G.Ol W W

z

3: <1Ld -0.02 en ct co -0.03 -0.04 EMPTY WEIGHT -0.05 -0.06 I . 1 0.04 I 1 1 0.03 :- t- en ::J

J

0.02 t- t-a: en:I: ::Jt- a: I 1/

J

W ~ :I: 0.01 t-~ I

-r- IZI

<1 .J I W If-- ~ o en ct co

IL ~

-0.01

Z_ I

,1

ZI I I I

-0.02

-

0.7 0.6 0.8 0.9 1.0 1.1 12 ENGINE WEIGHT BASELINE ENGINE WEIGHT Figure 79 • Engine Weight Sensitivity Studies, Turbofan Pressurized Twin, Design No.1.

25-!

.' ....... , ....

, 0.30

t-----t----+---+---Yf.:A FUEL WEIGHT

0.25 0.20 0.15 I- :I: (!)

0.10 I------+-----If---~j"..£-+I----+I GROSS WEIGHT

I- -

w :I: (!)

I

== 0.05 w w z EMPTY WEIGHT ==

~I~

-0.05 W IS -= CONSTANT

r ' I RANGE = 1556 KM (840 NM)

r 5490 M (18,000 FT) 444 KM/HR (240 KNOTS) -0.15 671 M (2200 FT} TAKEOFF DISTANCE -0.20 0.06

/

0.04 I- cn ::J

/

I-a: 0.02 cn:I: ::J I- a:

Z

w :I: Z I-

<1~ /

-0.02 cn q: co

/

-0.04

L

-0.06 0.8 0.9 1.0 1.1 1.2 1.3 1.4 CRUISE SFC BASELINE SFC CRUISE Figure 80.

Fuel Consumption Sensitivity Studies, Turbofan Pressurized Twin, Design No.1.

, ...-tL-_ •. ~'*' •• ~.:~~~.,. '-4 ,;.,.-~_', .... S, ......... ·_·

\

0.10 0.08 EMPTY WEIGHT 0.06

~

0.04 I-

~ROSS WEIGHT

J:

..........t_

~

C,!)

0.02 ~ I-w FUEL WEIGHT

.£ t::::----

J:=: C,!)

~ w ~ ..u Z =: ...I

l---::::= ~

<1 w -0.02 til ~ en

~

-0.04

W/S = CONSTANT

-0.06

-

RANGE = 2037 KM (1100 NM) 3048 M (10,000 FT) -0.08 417 KM/HR (225 KNOTS)

-

448 M (1600 FT) TAKEOFF DISTANCE L...

-1.00 0.08 0.06 r- 0...

0.04 J:

o...lW

J:z 0.02 <]...1

w ----

~

til ~ en

---- ~

~ --- - -

-0.02 -0.04 0.8 0.9 1.0 1.1 12 1.4 1.3 ENGINE WEIGHT BASELINE ENGINE WEIGHT Figure 81. Engine Weight Sensitivities, Turboprop Light Twin, Design No.4.

.. .,: I ~4 ....... _ ..... -', ....... -~."'" .......... ..

-- .~ .

..- "I .. :~ •• , .......... ~ .->ti'v ' ..... c.,IIr..4-..""'. ~~-.,.

",i "'" ,"'-' 0.30 I 0.26 FUEL 0.2C If-----t----. WEIGHT ""717-+----+----1 0.15 0.10 0.05 / yL" l:: 7 EMPTY WEIGHT

. £ WIS = CONSTANT

,I ::;;>,,4 j RANGE = 2037 KM (1100 NM)

3048 M (10,000 FT) 417 KM/HR (225 KNOTS) 448 M (1600 FT) TAKEOFF DISTANCE

-0.151 / I

-0.201 / 0.06 ./ 0.04

Q. V

J: 0.02

V

I

3:/ ~ 0

<lw en

/

~ -0.02 al

/

-0.0

·V

I 0.0 - - 0.8 0.9 1.0 1.1 12 1.3 4 CRUISE SFC BASELINE CRUISE SFC Fuel Consumption Sensitivities, Turboprop Light Twin, Figure 82 .

Design No.4.

"

'.r~"-:".""""""".--"'.·'-~ .. OW ... , ..... "".,-.5- ..... ; ;..c·.-:..:-tt)..;=:;~ ~#r'_"".·- d~rw==- J TABLE 71. l478°K (2200 F) BASELINE TURBOPROP SENSITIVITY OF PERFORHANCE TO COHPONENT PARAl-1ETERS 6100 H (20,100 FT), 389 KH/Hn (210 KTAS), STD DAY (ENGI~E A) .

Base .1 .1% .1% Parameter Value Value Power SFC Ram recovery 1.0 -0.02 -3.32 1.36 Compressor efficiency Base (.1T/T=C) -0.02 -2.63 2.72 Co~pressor_efficiency .P/P=C) Base -0.02 -2.18 1. 52 Pressure ratio 9.0 -0.8 -1.21 2.74 Compressor bleed 0.043 +0.02 -4.01 1. 90 Turbine cooling flow Base +0.02 -3.04 0.90 Burner .1P/P Base +0.02 -1. 31 1. 34 Burner leakage !lase -0.02 -3.70 3.82 HP turbine efficiency Base -0.02 -1. 28 1. 29 HP-LP turbine .1P/P Base +0.02 -1.28 1. 29 Horsepower extraction (GG) 5 +5 -1. 56 1. 58 HP turbine leakage Base +0.02 -1.97 2.00 Base -0.02 -2.28 2.34 jPower turbine efficiency Horsepower extraction ,IP.T.) 0 +5 -1.70 1. 73 p Turbin diffuser .1P/P Base 0.02 -1. 28 1. 30

I

" , .... '''' ..... ., .......... _ ..... ,_ .... ";...--.,.., ...... · ....... -·i?trtl.:.....--·' ............... -~·· .. ;;,· ....... it_~- TABLE 72. 1478°K (220GoF) BASELINE TURnOFAN SENSITIVITY OF PERFORMANCE TO COHPONENT PARAl>lETERS 6100 H (20, 000 FT), 389 KH/HR (210 KTAS), STD DAY (ENGINE J') Base .1% .1% Value Parameter Value Thrust TSFC Ram recovery 0.995 -0.02 -5.78 +3.85 Fan eff iciency (.1T/T=C) Base -0.02 -2.51 +1.67 Fan efficiency (P/P=C) Base -0.02 -1. 76 +1. 34 Fan pressure ratio 1.5 -0.05 -3.81 +1.67 Fan duct .1P/P Base +0.02 -3.02 +3.01 Fan-Comp .1P /P Base +0.02 -2.91 +10.84 Comp efficiency (.1T/T=C) Base -0.02 1.84 +1.84 Comp efficiency (P/P=C) Base -0.02 -1.89 +11.00 Comp pressure ratio 9.0 -0.8 -0.05 +1. 84 Fan duct leakage Base -+0.02 -2.79 +2.84 Compressor leakage Base +0.02 -3.58 +1.51 Turbine cooling flow Base +0.02 -2.43 +0.17 Burner .lP/p Base +0.02 -0.89 +0.84 HP turbine efficiency Base -0.02 -1.23 +1.17 I HP turbine leakage Base +0.02 -1.84 +1. 84 Horsepower extraction 5 +5 -0.64 +0.50 HPT-LPT .1P/P Base 0.02 -0.87 +0.84 LP turbine efficiency Base -0.02 -1.48 +1. 34 LP turbine leakage Base +0.02 -0.47 +0.33 Turbine diffuser, JP/p Base +0.02 -0.89 +0.84 Core thrust coefficient 0.985 -0.02 -0.49 +0.50 Fan thrust coefficient 0.985 -0.02 -2.84 +2.84 Bypass ratio I 8.0 '-2.0 -8.85 +9.7 I - , izX to" t~ it ,3"'-". n ... ,.--Aa:tm .. • .... csT? "jet: .;=;;·c .. m,·,iff:e:'ei , ..... --·"".:··l .... ,...·r .. '· .,.....",.. .....-;;or ..... .,.., , _ "'C~~ ___ -' "~ REFERENCES 1. Merrill, G. L., G. A. Burnett, C. C. Alsworth, et al., A Study of Small Turbofan Engines Applicable to General Aviation Aircraft Final Report, NASA CR 114630, September 1973.

2. Hildenbrand, R. W., G. L. Merrill, and G. A. Burnett, Study of Small Civil Turbofan Engines Applicable to Military Trainer Airplanes Final Report, NASA CR 137575, April, 1975 3. Merrill, G. L., Study of Small Turbofan Engines Applicable to Single-Engine Light Airplanes, NASA CR 137944, September 1976 <t-~ # ...

. rl y" - ... ,-- eGCt .a_ --¥ .. ---- .... -.-- --"-

End of Document